Engineering › Aerospace Engineering

Turbomachinery Performance and Optimization

Works Count: 42346

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Description

This cluster of papers focuses on the aerodynamics and heat transfer phenomena in turbomachinery, with specific emphasis on topics such as turbine performance, film cooling techniques, boundary layer transition modeling, stall and surge control, and separation control. It also includes research on blade tip heat transfer, axial compressor behavior, and the application of large-eddy simulations in understanding complex flow phenomena.

Keywords

Turbine; Film Cooling; Transition Model; Axial Compressor; Heat Transfer; Stall and Surge; Boundary Layer; Separation Control; Blade Tip; Large-Eddy Simulation

Axial Flow Compressor Noise Studies

1962-01-01

John M. Tyler , T. G. Sofrin

Design of radial turbomachines

1990-01-01

A. Whitfield , Nicholas C. Baines

Part 1: Outline of the book history and applications the centrifugal compressor the radial turbine non-dimensional parameters for performance assessment and design selection. Part 2 Fundamental fluid mechanics and thermodynamics: … Part 1: Outline of the book history and applications the centrifugal compressor the radial turbine non-dimensional parameters for performance assessment and design selection. Part 2 Fundamental fluid mechanics and thermodynamics: the basic equations the dynamics of compressible, perfect gas flow definitions of efficiency loss coefficient definitions stator performance parameters. Part 3 Preliminary design and analysis of centrifugal compressors: non-dimensional parameters for design selection application of basic thermo-fluid dynamics to compressors impeller design diffuser performance, design and analysis. Part 4 Preliminary design and analysis of radial turbines: basic design concepts preliminary design of the rotor performance correlations rotor passage preliminary design of the stator exhaust diffuser. Part 5 Generalised computer based one-dimensional performance prediction: generalised gas dynamic analysis empirical flow models and loss correlations summary of prediction procedure. Part 6 Modelling of compressor impellers with separated flow: performance prediction models application to experimentation. Part 7 Geometric description of turbomachine blades and passages: surface definition by two-dimensional projection functional representation of two-dimensional curves geometry definition by three-dimensional surfaces functional representation of three-dimensional surfaces. Part 8 Computation of internal flows: categorisation of methods calculation domains inviscid solution methods viscous solution methods coupled viscous-inviscid methods modelling of stator flows. Part 9 Special problems associated with turbomachine design and operation: compressor surge and techniques to suppress it high pressure ratio compressors pulse operation of radial turbines variable geometry stator for radial turbines special features of radial turbine rotors turbine rotor cooling size effects and scaling. Appendices: preliminary design of centrifugal compressor impellers - computer program listing, description of the computer program preliminary design of radial turbine rotors - computer program listing, description of the program a summary of the governing equations for three-dimensional fluid flow - the conservation of mass, or continuity, equation, the momentum equation, the energy equation discretisation methods - finite difference, finite element, finite volume.

Fluid mechanics, thermodynamics of turbomachinery

1977-02-01

Simon Dixon

Gas Turbine Engineering Handbook

2006-01-01

Meherwan P. Boyce

Gas Turbine Heat Transfer and Cooling Technology

2012-11-27

Je-Chin Han , Sandip Dutta , Srinath V. Ekkad

Fundamentals Need for Turbine Blade Cooling Turbine-Cooling Technology Turbine Heat Transfer and Cooling Issues Structure of the Book Review Articles and Book Chapters on Turbine Cooling and Heat Transfer New … Fundamentals Need for Turbine Blade Cooling Turbine-Cooling Technology Turbine Heat Transfer and Cooling Issues Structure of the Book Review Articles and Book Chapters on Turbine Cooling and Heat Transfer New Information from 2000 to 2010 References Turbine Heat Transfer Introduction Turbine-Stage Heat Transfer Cascade Vane Heat-Transfer Experiments Cascade Blade Heat Transfer Airfoil Endwall Heat Transfer Turbine Rotor Blade Tip Heat Transfer Leading-Edge Region Heat Transfer Flat-Surface Heat Transfer New Information from 2000 to 2010 2.10 Closure References Turbine Film Cooling Introduction Film Cooling on Rotating Turbine Blades Film Cooling on Cascade Vane Simulations Film Cooling on Cascade Blade Simulations Film Cooling on Airfoil Endwalls Turbine Blade Tip Film Cooling Leading-Edge Region Film Cooling Flat-Surface Film Cooling Discharge Coefficients of Turbine Cooling Holes 3.10 Film-Cooling Effects on Aerodynamic Losses 3.11 New Information from 2000 to 2010 3.12 Closure References Turbine Internal Cooling Jet Impingement Cooling Rib-Turbulated Cooling Pin-Fin Cooling Compound and New Cooling Techniques New Information from 2000 to 2010 References Turbine Internal Cooling with Rotation Rotational Effects on Cooling Smooth-Wall Coolant Passage Heat Transfer in a Rib-Turbulated Rotating CoolantPassage Effect of Channel Orientation with Respect to the RotationDirection on Both Smooth and Ribbed Channels Effect of Channel Cross Section on Rotating Heat Transfer Different Proposed Correlation to Relate the Heat Transferwith Rotational Effects Heat-Mass-Transfer Analogy and Detail Measurements Rotation Effects on Smooth-Wall Impingement Cooling Rotational Effects on Rib-Turbulated Wall ImpingementCooling New Information from 2000 to 2010 References Experimental Methods Introduction Heat-Transfer Measurement Techniques Mass-Transfer Analogy Techniques Liquid Crystal Thermography Flow and Thermal Field Measurement Techniques New Information from 2000 to 2010 Closure References Numerical Modeling Governing Equations and Turbulence Models Numerical Prediction of Turbine Heat Transfer Numerical Prediction of Turbine Film Cooling Numerical Prediction of Turbine Internal Cooling New Information from 2000 to 2010 References Final Remarks Turbine Heat Transfer and Film Cooling Turbine Internal Cooling with Rotation Turbine Edge Heat Transfer and Cooling New Information from 2000 to 2010 Closure Index

Stall, Surge, and 75 Years of Research

2015-08-31

I. J. Day

Work on rotating stall and its related disturbances have been in progress since the Second World War. During this period, certain “hot topics” have come to the fore—mostly in response … Work on rotating stall and its related disturbances have been in progress since the Second World War. During this period, certain “hot topics” have come to the fore—mostly in response to pressing problems associated with new engine designs. This paper will take a semihistorical look at some of these fields of study (stall, surge, active control, rotating instabilities, etc.) and will examine the ideas which underpin each topic. Good progress can be reported, but the paper will not be an unrestricted celebration of our successes because, after 75 years of research, we are still unable to predict the stalling behavior of a new compressor or to contribute much to the design of a more stall-resistant machine. Looking forward from where we are today, it is clear that future developments will come from CFD in the form of better performance predictions, better flow modeling, and improved interpretation of experimental results. It is also clear that future experimental work will be most effective when focussed on real compressors with real problems—such as stage matching, large tip clearances, eccentricity, and service life degradation. Today’s topics of interest are mostly associated with compressible effects and so further research will require more high-speed testing.

Recent Progress in the Understanding of Basic Aspects of Secondary Flows in Turbine Blade Passages

1985-04-01

C. H. Sieverding

The present paper is an attempt to summarize the results of experimental secondary flow research over the past decade in order to give a full picture of our present knowledge … The present paper is an attempt to summarize the results of experimental secondary flow research over the past decade in order to give a full picture of our present knowledge and uncertainties of basic secondary flow aspects. The paper gives a detailed description of secondary flow vortex structures and their effect on end wall boundary layer characteristics and loss growth through straight turbine blade passages.

A Theory of Post-Stall Transients in Axial Compression Systems: Part I—Development of Equations

1986-01-01

F. K. Moore , E. M. Greitzer

An approximate theory is presented for post-stall transients in multistage axial compression systems. The theory leads to a set of three simultaneous nonlinear third-order partial differential equations for pressure rise, … An approximate theory is presented for post-stall transients in multistage axial compression systems. The theory leads to a set of three simultaneous nonlinear third-order partial differential equations for pressure rise, and average and disturbed values of flow coefficient, as functions of time and angle around the compressor. By a Galerkin procedure, angular dependence is averaged, and the equations become first order in time. These final equations are capable of describing the growth and possible decay of a rotating-stall cell during a compressor mass-flow transient. It is shown how rotating-stall-like and surgelike motions are coupled through these equations, and also how the instantaneous compressor pumping characteristic changes during the transient stall process.

Gas Turbine Film Cooling

2006-03-01

David G. Bogard , Karen A. Thole

The durability of gas turbine engines is strongly dependent on the component temperatures. For the combustor and turbine airfoils and endwalls, film cooling is used extensively to reduce component temperatures. … The durability of gas turbine engines is strongly dependent on the component temperatures. For the combustor and turbine airfoils and endwalls, film cooling is used extensively to reduce component temperatures. Film cooling is a cooling method used in virtually all of today's aircraft turbine engines and in many power-generation turbine engines and yet has very difficult phenomena to predict. The interaction of jets-in-crossflow, which is representative of film cooling, results in a shear layer that leads to mixing and a decay in the cooling performance along a surface. This interaction is highly dependent on the jet-to-crossflow mass and momentum flux ratios. Film-cooling performance is difficult to predict because of the inherent complex flowfields along the airfoil component surfaces in turbine engines. Film cooling is applied to nearly all of the external surfaces associated with the airfoils that are exposed to the hot combustion gasses such as the leading edges, main bodies, blade tips, and endwalls. In a review of the literature, it was found that there are strong effects of freestream turbulence, surface curvature, and hole shape on the performance of film cooling. Film cooling is reviewed through a discussion of the analyses methodologies, a physical description, and the various influences on film-cooling performance.

A Mean Line Prediction Method for Axial Flow Turbine Efficiency

1982-01-01

S. C. Kacker , U. Okapuu

A mean line loss system is described, capable of predicting the design point efficiencies of current axial turbines of gas turbine engines. This loss system is a development of the … A mean line loss system is described, capable of predicting the design point efficiencies of current axial turbines of gas turbine engines. This loss system is a development of the Ainley/Mathieson technique of 1951. The prediction method is tested against the “Smith’s chart” and against the known efficiencies of 33 turbines of recent design. It is shown to be able to predict the efficiencies of a wide range of axial turbines of conventional stage loadings to within ± 1 1/2 percent.

Effect of Mistuning on the Vibration of Turbo-Machine Blades Induced by Wakes

1966-03-01

D. S. Whitehead

The report gives a theoretical investigation of the effect of blade mistuning on the vibration of the blades induced by wakes in the incident gas flow which rotate relative to … The report gives a theoretical investigation of the effect of blade mistuning on the vibration of the blades induced by wakes in the incident gas flow which rotate relative to the blades in question. An upper limit to the effect of mistuning is found, which shows that the amplitude on one blade may increase by a factor of roughly ½(1 + √ N), where N is the number of blades in the row. This limit is only approached in special and unusual circumstances. In normal circumstances the amplitude will not increase by more than about 20 per cent. It is concluded that blades expected to work under forced vibration conditions should be made as nearly identical as possible. This is in contrast to the situation when flutter may occur, where the effect of mistuning is always favourable.

Crossflows in a Turbine Cascade Passage

1980-10-01

Lee S. Langston

Measurements of the subsonic flow in a large scale plane turbine cascade, that were given in an earlier paper, are examined in more detail from the standpoint of the endwall … Measurements of the subsonic flow in a large scale plane turbine cascade, that were given in an earlier paper, are examined in more detail from the standpoint of the endwall boundary layer. Representative data are presented in terms of normal and streamwise velocities, flow angle deviations, and polar plots, that can be used to substantiate analytical models of the endwall flow. The qualitative behavior of the endwall crossflow was found to be correlated by a relatively simple expression, based on the flow angle deviation.

Three-Dimensional Flow Within a Turbine Cascade Passage

1977-01-01

Lee S. Langston , Minor Nice , Richard Hooper

Detailed measurements of the subsonic flow in a large-scale, plane turbine cascade were made to evaluate the three-dimensional nature of the flow field. Tests were conducted at a passage aspect … Detailed measurements of the subsonic flow in a large-scale, plane turbine cascade were made to evaluate the three-dimensional nature of the flow field. Tests were conducted at a passage aspect ratio of 1.0 with a collateral inlet boundary layer. Flow visualization was done on airfoil and endwall surfaces. Velocity and pressure measurements were taken before and behind the cascade and in six axial planes within a cascade passage, using a five-hole probe. Hot wire measurements were taken in the endwall boundary layer within the cascade passage. The characteristics of the endwall boundary layer are presented, showing that three-dimensional separation is an important feature of end-wall flow. A large part of the endwall boundary layer was found to be very thin when compared to the cascade inlet boundary layer. Data showing the growth of aerodynamic loss through the passage are discussed.

Surge and Rotating Stall in Axial Flow Compressors—Part I: Theoretical Compression System Model

1976-04-01

E. M. Greitzer

This paper reports a theoretical study of axial compressor surge. A nonlinear model is developed to predict the transient response of a compression system subsequent to a perturbation from steady … This paper reports a theoretical study of axial compressor surge. A nonlinear model is developed to predict the transient response of a compression system subsequent to a perturbation from steady operating conditions. It is found that for the system investigated there is an important nondimensional parameter on which this response depends. Whether this parameter is above or below a critical value determines which mode of compressor instability, rotating stall or surge, will be encountered at the stall line. For values above the critical, the system will exhibit the large amplitude oscillatory behavior characteristic of surge; while for values below the critical it will move toward operation in rotating stall, at a substantially reduced flow rate and pressure ratio. Numerical results are presented to show the motion of the compression system operating point during these two basic modes of instability, and a physical explanation is given for the mechanism associated with the generation of surge cycle oscillations.

Discrete-Jet Film Cooling: A Comparison of Computational Results With Experiments

1994-07-01

James H. Leylek , R. D. Zerkle

Large-scale computational analyses have been conducted and results compared with experiments to understand coolant jet and crossflow interaction in discrete-jet film cooling. Detailed three-dimensional elliptic Navier–Stokes solutions, with high-order turbuence … Large-scale computational analyses have been conducted and results compared with experiments to understand coolant jet and crossflow interaction in discrete-jet film cooling. Detailed three-dimensional elliptic Navier–Stokes solutions, with high-order turbuence modeling, are presented for film cooling using a new model enabling simultaneous solution of fully coupled flow in plenum, film-hole, and cross-stream regions. Computations are carried out for the following range of film cooling parameters typically found in gas turbine airfoil applications: single row of jets with a film-hole length-to-diameter ratio of 1.75 and 3.5; blowing ratio from 0.5 up to 2; coolant-to-crossflow density ratio of 2; streamwise injection angle of 35 deg; and pitch-to-diameter ratio of 3. Comparison of computational solutions with experimental data give good agreement. Moreover, the current results complement experiments and support previous interpretations of measured data and flow visualization. The results also explain important aspects of film cooling, such as the development of complex flow within the film-hole in addition to the well-known counterrotating vortex structure in the cross-stream.

A Review of Surface Roughness Effects in Gas Turbines

2010-01-11

Jeffrey P. Bons

The effects of surface roughness on gas turbine performance are reviewed based on publications in the open literature over the past 60 years. Empirical roughness correlations routinely employed for drag … The effects of surface roughness on gas turbine performance are reviewed based on publications in the open literature over the past 60 years. Empirical roughness correlations routinely employed for drag and heat transfer estimates are summarized and found wanting. No single correlation appears to capture all of the relevant physics for both engineered and service-related (e.g., wear or environmentally induced) roughness. Roughness influences engine performance by causing earlier boundary layer transition, increased boundary layer momentum loss (i.e., thickness), and/or flow separation. Roughness effects in the compressor and turbine are dependent on Reynolds number, roughness size, and to a lesser extent Mach number. At low Re, roughness can eliminate laminar separation bubbles (thus reducing loss) while at high Re (when the boundary layer is already turbulent), roughness can thicken the boundary layer to the point of separation (thus increasing loss). In the turbine, roughness has the added effect of augmenting convective heat transfer. While this is desirable in an internal turbine coolant channel, it is clearly undesirable on the external turbine surface. Recent advances in roughness modeling for computational fluid dynamics are also reviewed. The conclusion remains that considerable research is yet necessary to fully understand the role of roughness in gas turbines.

Surge and Rotating Stall in Axial Flow Compressors—Part II: Experimental Results and Comparison With Theory

1976-04-01

E. M. Greitzer

This paper reports an experimental study of axial compressor surge and rotating stall. The experiments were carried out using a three stage axial flow compressor. With the experimental facility the … This paper reports an experimental study of axial compressor surge and rotating stall. The experiments were carried out using a three stage axial flow compressor. With the experimental facility the physical parameters of the compression system could be independently varied so that their influence on the transient system behavior can be clearly seen. In addition, a new data analysis procedure has been developed, using a plenum mass balance, which enables the instantaneous compressor mass flow to be accurately calculated. This information is coupled to the unsteady pressure measurements to provide the first detailed quantitative picture of instantaneous compressor operation during both surge and rotating stall transients. The experimental results are compared to a theoretical model of the transient system response. The theoretical criterion for predicting which mode of compression system instability, rotating stall or surge, will occur is in good accord with the data. The basic scaling concepts that have been developed for relating transient data at different corrected speeds and geometrical parameters are also verified. Finally, the model is shown to provide an adequate quantitative description of the motion of the compression system operating point during the transients that occur subsequent to the onset of axial compressor stall.

Turbine Rotor-Stator Interaction

1982-10-01

R. P. Dring , H. D. Joslyn , Larry W. Hardin +1 more

The aerodynamic interaction between the rotor and stator airfoils of a large scale axial turbine stage have been studied experimentally. The data included measurements of the time averaged and instantaneous … The aerodynamic interaction between the rotor and stator airfoils of a large scale axial turbine stage have been studied experimentally. The data included measurements of the time averaged and instantaneous surface pressures and surface thin film gage output on both the rotor and stator at midspan. The data also included measurement of the stator suction and pressure surface time averaged heat transfer at midspan. The data was acquired with rotor-stator axial gaps of 15 and 65 percent of axial chord. The upstream potential flow influence of the rotor on the stator was seen as well as the downstream potential flow and wake influences of the stator on the rotor. It was also seen that at the 15 percent axial gap, the stator heat-transfer coefficient was typically 25 percent higher than that at the 65 percent gap.

Performance Deterioration in Industrial Gas Turbines

1992-04-01

Ihor S. Diakunchak

This paper describes the most important factors affecting the industrial gas turbine engine performance deterioration with service time and provides some approximate data on the prediction of the rate of … This paper describes the most important factors affecting the industrial gas turbine engine performance deterioration with service time and provides some approximate data on the prediction of the rate of deterioration. Recommendations are made on how to detect and monitor the performance deterioration. Preventative measures, which can be taken to avoid or retard the performance deterioration, are described in some detail.

The Calculation of Three-Dimensional Viscous Flow Through Multistage Turbomachines

1992-01-01

J. D. Denton

The extension of a well-established three-dimensional flow calculation method to calculate the flow through multiple turbomachinery blade rows is described in this paper. To avoid calculating the unsteady flow, which … The extension of a well-established three-dimensional flow calculation method to calculate the flow through multiple turbomachinery blade rows is described in this paper. To avoid calculating the unsteady flow, which is inherent in any machine containing both rotating and stationary blade rows, a mixing process is modeled at a calculating station between adjacent blade rows. The effects of this mixing on the flow within the blade rows may be minimized by using extrapolated boundary conditions at the mixing plane. Inviscid calculations are not realistic for multistage machines and so the method includes a range of options for the inclusion of viscous effects. At the simplest level such effects may be included by prescribing the spanwise variation of polytropic efficiency for each blade row. At the most sophisticated level viscous effects and machine performance can be predicted by using a thin shear layer approximation to the Navier–Stokes equations and an eddy viscosity turbulence model. For high-pressure-ratio compressors there is a strong tendency for the calculation to surge during the transient part of the flow. This is overcome by the use of a new technique, which enables the calculation to be run to a prescribed mass flow. Use of the method is illustrated by applying it to a multistage turbine of simple geometry, a two-stage low-speed experimental turbine, and two multistage axial compressors.

1997 Best Paper Award—Turbomachinery Committee: A Study of Spike and Modal Stall Phenomena in a Low-Speed Axial Compressor

1998-07-01

Thomas R. Camp , I. J. Day

This paper presents a study of stall inception mechanisms in a low-speed axial compressor. Previous work has identified two common flow breakdown sequences, the first associated with a short length-scale … This paper presents a study of stall inception mechanisms in a low-speed axial compressor. Previous work has identified two common flow breakdown sequences, the first associated with a short length-scale disturbance known as a “spike,” and the second with a longer length-scale disturbance known as a “modal oscillation.” In this paper the physical differences between these two mechanisms are illustrated with detailed measurements. Experimental results are also presented that relate the occurrence of the two stalling mechanisms to the operating conditions of the compressor. It is shown that the stability criteria for the two disturbances are different: Long length-scale disturbances are related to a two-dimensional instability of the whole compression system, while short length-scale disturbances indicate a three-dimensional breakdown of the flow-field associated with high rotor incidence angles. Based on the experimental measurements, a simple model is proposed that explains the type of stall inception pattern observed in a particular compressor. Measurements from a single-stage low-speed compressor and from a multistage high-speed compressor are presented in support of the model.

Turbulent Transport on the Endwall in the Region Between Adjacent Turbine Blades

1988-11-01

R. J. Goldstein , Ronald Spores

The complex three-dimensional flow in the endwall region near the base of a turbine blade has an important impact on the local heat transfer. The initial horseshoe vortex, the passage … The complex three-dimensional flow in the endwall region near the base of a turbine blade has an important impact on the local heat transfer. The initial horseshoe vortex, the passage vortex, and resulting corner vortices cause large variations in heat transfer over the entire endwall region. Due to these large surface gradients in heat transfer, conventional measurement techniques generally do not provide an accurate determination of the local heat transfer coefficients. In the present study, the heat/mass transfer analogy is used to examine the local transport coefficients for two different endwall boundary layer thicknesses and two free-stream Reynolds numbers. A linear turbine blade cascade is used in conjunction with a removable endwall plate. Naphthalene (C10H8) is cast into a mold on the plate and the rate of naphthalene sublimation is determined at 6000 + locations on the simulated endwall by employing a computer-aided data acquisition system. This technique allows one to obtain detailed contour plots of the local convection coefficient over the entire endwall. By examining the mass transfer contours, it is possible to infer information on the three-dimensional flow in the passage between the blades. Extremely high transport coefficients on the endwall indicate locations of potential overheating and failure in an actual turbine.

Surge Dynamics in a Free-Spool Centrifugal Compressor System

1992-04-01

David Allan Fink , N. A. Cumpsty , E. M. Greitzer

Turbocharger surge has been investigated in a radial impeller-vaneless diffuser free-spool system. Several different aspects are addressed. First, two very different compression systems, one with a large downstream volume and … Turbocharger surge has been investigated in a radial impeller-vaneless diffuser free-spool system. Several different aspects are addressed. First, two very different compression systems, one with a large downstream volume and one with the smallest possible downstream volume, are employed to examine stall initiation phenomena as well as the behavior of the compressor characteristics when operating in surge. The measurements show impeller stall at the inducer tips to be a key phenomena in initiating surge. The inducer stall is stationary and asymmetric, due to the presence of the volute, and is most severe near the volute tongue angular position. The compressor characteristic in the large volume system (which gave surge) is observed to be flatter and to lag that in the stabilized small volume system. The difference arises because of the slow development time and differing circumferential extent of the inducer stall present at a given mass flow. A nonlinear simulation of the system is also presented. The model deviates from previous treatments of unsteady flow in compressor systems in that the assumption of constant rotor speed is relaxed. Including a time lag on the order of the compressor throughflow time, together with proper treatment of speed variations, is shown to improve agreement with the observed surge behavior dramatically.

A Detailed Analysis of Film-Cooling Physics: Part I—Streamwise Injection With Cylindrical Holes

1997-02-01

D. Keith Walters , James H. Leylek

A previously documented systematic computational methodology is implemented and applied to a jet-in-crossflow problem in order to document all of the pertinent flow physics associated with a film-cooling flowfield. Numerical … A previously documented systematic computational methodology is implemented and applied to a jet-in-crossflow problem in order to document all of the pertinent flow physics associated with a film-cooling flowfield. Numerical results are compared to experimental data for the case of a row of three-dimensional, inclined jets with length-to-diameter ratios similar to a realistic film-cooling application. A novel vorticity-based approach is included in the analysis of the flow physics. Particular attention has been paid to the downstream coolant structures and to the source and influence of counterrotating vortices in the crossflow region. It is shown that the vorticity in the boundary layers within the film hole is primarily responsible for this secondary motion. Important aspects of the study include: (1) a systematic treatment of the key numerical issues, including accurate computational modeling of the physical problem, exact geometry and high-quality grid generation techniques, higher-order numerical discretization, and accurate evaluation of turbulence model performance; (2) vorticity-based analysis and documentation of the physical mechanisms of jet–crossflow interaction and their influence on film-cooling performance; (3) a comparison of computational results to experimental data; and (4) comparison of results using a two-layer model near-wall treatment versus generalized wall functions. Solution of the steady, time-averaged Navier–Stokes equations were obtained for all cases using an unstructured/adaptive grid, fully explicit, time-marching code with multigrid, local time stepping, and residual smoothing acceleration techniques. For the case using the two-layer model, the solution was obtained with an implicit, pressure-correction solver with multigrid. The three-dimensional test case was examined for two different film-hole length-to-diameter ratios of 1.75 and 3.5, and three different blowing ratios, from 0.5 to 2.0. All of the simulations had a density ratio of 2.0, and an injection angle of 35 deg. An improved understanding of the flow physics has provided insight into future advances to film-cooling configuration design. In addition, the advantages and disadvantages of the two-layer turbulence model are highlighted for this class of problems. [S0889-504X(00)01201-0]

Adiabatic Wall Effectiveness Measurements of Film-Cooling Holes With Expanded Exits

1997-06-02

Michael Gritsch , A. Schulz , S. Wittig

This paper presents detailed measurements of the film-cooling effectiveness for three single, scaled-up film-cooling hole geometries. The hole geometries investigated include a cylindrical hole and two holes with a diffuser … This paper presents detailed measurements of the film-cooling effectiveness for three single, scaled-up film-cooling hole geometries. The hole geometries investigated include a cylindrical hole and two holes with a diffuser shaped exit portion (i.e. a fanshaped and a laidback fanshaped hole). The flow conditions considered are the crossflow Mach number at the hole entrance side (up to 0.6), the crossflow Mach number at the hole exit side (up to 1.2), and the blowing ratio (up to 2). The coolant-to-mainflow temperature ratio is kept constant at 0.54. The measurements are performed by means of an infrared camera system which provides a two-dimensional distribution of the film-cooling effectiveness in the nearfield of the cooling hole down to x/D = 10. As compared to the cylindrical hole, both expanded holes show significantly improved thermal protection of the surface downstream of the ejection location, particularly at high blowing ratios. The laidback fanshaped hole provides a better lateral spreading of the ejected coolant than the fanshaped hole which leads to higher laterally averaged film-cooling effectiveness. Coolant passage crossflow Mach number and orientation strongly affect the flowfield of the jet being ejected from the hole and, therefore, have an important impact on film-cooling performance.

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The 1993 IGTI Scholar Lecture: Loss Mechanisms in Turbomachines

1993-10-01

J. D. Denton

The origins and effects of loss in turbomachines are discussed with the emphasis on trying to understand the physical origins of loss rather than on reviewing the available prediction methods. … The origins and effects of loss in turbomachines are discussed with the emphasis on trying to understand the physical origins of loss rather than on reviewing the available prediction methods. Loss is defined in terms of entropy increase and the relationship of this to the more familiar loss coefficients is derived and discussed. The sources of entropy are, in general: viscous effects in boundary layers, viscous effects in mixing processes, shock waves, and heat transfer across temperature differences. These are first discussed in general and then the results are applied to turbomachinery flows. Understanding of the loss due to heat transfer requires some discussion of cycle thermodynamics. Sections are devoted to discussing blade boundary layer and trailing edge loss, tip leakage loss, endwall loss, effects of heat transfer, and miscellaneous losses. The loss arising from boundary layer separation is particularly difficult to quantify. Most of the discussion is based on axial flow machines, but a separate section is devoted to the special problems of radial flow machines. In some cases, e.g., attached blade boundary layers, the loss mechanisms are well understood, but even so the loss can seldom be predicted with great accuracy. In many other cases, e.g., endwall loss, the loss mechanisms are still not clearly understood and prediction methods remain very dependent on correlations. The paper emphasizes that the use of correlations should not be a substitute for trying to understand the origins of loss, and suggests that a good physical understanding of the latter may be more valuable than a quantitative prediction.

Tip Leakage Flow in Axial Compressors

1991-04-01

J. A. Storer , N. A. Cumpsty

Experimental measurements in a linear cascade with tip clearance are complemented by numerical solutions of the three-dimensional Navier–Stokes equations in an investigation of tip leakage flow. Measurements reveal that the … Experimental measurements in a linear cascade with tip clearance are complemented by numerical solutions of the three-dimensional Navier–Stokes equations in an investigation of tip leakage flow. Measurements reveal that the clearance flow, which separates near the entry of the tip gap, remains unattached for the majority of the blade chord when the tip clearance is similar to that typical of a machine. The numerical predictions of leakage flow rate agree very well with measurements, and detailed comparisons show that the mechanism of tip leakage is primarily inviscid. It is demonstrated by simple calculation that it is the static pressure field near the end of the blade that controls chordwise distribution of the flow across the tip. Although the presence of a vortex caused by the roll-up of the leakage flow may affect the local pressure field, the overall magnitude of the tip leakage flow remains strongly related to the aerodynamic loading of the blades.

A Review of Shaped Hole Turbine Film-Cooling Technology

2005-03-30

Ronald S. Bunker

Film cooling represents one of the few game-changing technologies that has allowed the achievement of today’s high firing temperature, high-efficiency gas turbine engines. Over the last 30 years, only one … Film cooling represents one of the few game-changing technologies that has allowed the achievement of today’s high firing temperature, high-efficiency gas turbine engines. Over the last 30 years, only one major advancement has been realized in this technology, that being the incorporation of exit shaping to the film holes to result in lower momentum coolant injection jets with greater surface coverage. This review examines the origins of shaped film cooling and summarizes the extant literature knowledge concerning the performance of such film holes. A catalog of the current literature data is presented, showing the basic shaping geometries, parameter ranges, and types of data obtained. Specific discussions are provided for the flow field and aerodynamic losses of shaped film hole coolant injection. The major fundamental effects due to coolant-to-gas blowing ratio, compound angle injection, cooling hole entry flow character, and mainstream turbulence intensity are each reviewed with respect to the resulting adiabatic film effectiveness and heat transfer coefficients for shaped holes. A specific example of shaped film effectiveness is provided for a production turbine inlet vane with comparison to other data. Several recent unconventional forms of film hole shaping are also presented as a look to future potential improvements.

Loss Sources and Magnitudes in Axial-Flow Compressors

1976-07-01

C. C. Koch , L. H. Smith

A method is presented for calculating the design point efficiency potential of a multistage compressor. Design parameters that affect the efficiency are vector diagram shape, aerodynamic loading level, aspect ratio, … A method is presented for calculating the design point efficiency potential of a multistage compressor. Design parameters that affect the efficiency are vector diagram shape, aerodynamic loading level, aspect ratio, solidity, clearances, airfoil maximum and edge thicknesses, annulus area contraction, Mach number, Reynolds number, airfoil surface finish, and part-span shroud placement. Losses associated with off-design operation, blading unsuited to the aerodynamic environment, or poor hardware quality are not considered. The loss model is constructed using rational fluid-dynamic elements, such as boundary layer theory, whenever feasible in an attempt to minimize empirical influences, although some empiricism inevitably enters. The resulting formulation is found to be in satisfactory agreement with multistage compressor experience that covers a wide range of the design parameters.

Stall Inception in Axial Compressors

1990-01-01

N. M. McDougall , N. A. Cumpsty , T. P. Hynes

Detailed measurements have been made of the transient stalling process in an axial compressor stage. The stage is of high hub-casing ratio and stall is initiated in the rotor. If … Detailed measurements have been made of the transient stalling process in an axial compressor stage. The stage is of high hub-casing ratio and stall is initiated in the rotor. If the rotor tip clearance is small stall inception occurs at the hub, but at clearances typical for a multistage compressor the inception is at the tip. The crucial quantity in both cases is the blockage caused by the endwall boundary layer. Prior to stall, disturbances rotate around the inlet flow in sympathy with rotating variations in the endwall blockage; these can persist for some time prior to stall, rising and falling in amplitude before the final increase, which occurs as the compressor stalls.

The Measurement and Formation of Tip Clearance Loss

1989-07-01

J. P. Bindon

The detailed development of tip clearance loss from the leading to trailing edge of a linear turbine cascade was measured and the contributions made by mixing, internal gap shear flow, … The detailed development of tip clearance loss from the leading to trailing edge of a linear turbine cascade was measured and the contributions made by mixing, internal gap shear flow, and endwall/ secondary flow were identified, separated, and quantified for the first time. Only 13 percent of the overall loss arises from endwall/secondary flow and of the remaining 87 percent, 48 percent is due to mixing and 39 percent is due to internal gap shear. All loss formation appears to be dominated by phenomena connected with the gap separation bubble. Flow established within the bubble by the pressure gradient separates as the gradient disappears and most of the internal loss is created by the entrainment of this separated fluid. When this high-loss leakage wake enters the mainstream, it separates due to the suction corner pressure gradient to create virtually all the measured mixing loss. It is suggested that the control of tip clearance loss by discharge coefficient reduction actually introduces loss. Performance improvements may result from streamlined tip geometries that optimize the tradeoff between entropy production and flow deflection.

Rotating Instabilities in an Axial Compressor Originating From the Fluctuating Blade Tip Vortex

2000-02-01

Ronald Mailach , Ingolf Lehmann , Konrad Vogeler

Rotating instabilities (RIs) have been observed in axial flow fans and centrifugal compressors as well as in low-speed and high-speed axial compressors. They are responsible for the excitation of high … Rotating instabilities (RIs) have been observed in axial flow fans and centrifugal compressors as well as in low-speed and high-speed axial compressors. They are responsible for the excitation of high amplitude rotor blade vibrations and noise generation. This flow phenomenon moves relative to the rotor blades and causes periodic vortex separations at the blade tips and an axial reversed flow through the tip clearance of the rotor blades. The paper describes experimental investigations of RIs in the Dresden Low-Speed Research Compressor (LSRC). The objective is to show that the fluctuation of the blade tip vortex is responsible for the origination of this flow phenomenon. RIs have been found at operating points near the stability limit of the compressor with relatively large tip clearance of the rotor blades. The application of time-resolving sensors in both fixed and rotating frame of reference enables a detailed description of the circumferential structure and the spatial development of this unsteady flow phenomenon, which is limited to the blade tip region. Laser-Doppler-anemometry (LDA) within the rotor blade passages and within the tip clearance as well as unsteady pressure measurements on the rotor blades show the structure of the blade tip vortex. It will be shown that the periodical interaction of the blade tip vortex of one blade with the flow at the adjacent blade is responsible for the generation of a rotating structure with high mode orders, termed a rotating instability.

Predictions of Endwall Losses and Secondary Flows in Axial Flow Turbine Cascades

1987-04-01

Om P. Sharma , T. L. Butler

This paper describes the development of a semi-empirical model for estimating end-wall losses. The model has been developed from improved understanding of complex endwall secondary flows, acquired through review of … This paper describes the development of a semi-empirical model for estimating end-wall losses. The model has been developed from improved understanding of complex endwall secondary flows, acquired through review of flow visualization and pressure loss data for axial flow turbomachine cascades. The flow visualization data together with detailed measurements of viscous flow development through cascades have permitted more realistic interpretation of the classical secondary flow theories for axial turbomachine cascades. The re-interpreted secondary flow theories together with integral boundary layer concepts are used to formulate a calculation procedure for predicting losses due to the endwall secondary flows. The proposed model is evaluated against data from published literature and improved agreement between the data and predictions is demonstrated.

Stall Inception in Axial Flow Compressors

1993-01-01

I. J. Day

Studies have been conducted on two laboratory test compressors to investigate the process leading to the formation of finite amplitude rotating stall cells. The measurements were obtained from circumferential arrays … Studies have been conducted on two laboratory test compressors to investigate the process leading to the formation of finite amplitude rotating stall cells. The measurements were obtained from circumferential arrays of hot wires and were spatially and temporarily analyzed to show that modal perturbations are not always present prior to stall, and when present, sometimes have little direct effect on the formation of the stall cells. The measurements lead to the conclusion that the occurrence of modal perturbations, and the formation of finite amplitude stall cells, are two separate phenomena, both occurring under roughly the same conditions at the peak of the pressure rise characteristic. The measurements also underline the hitherto unsuspected importance of short length scale disturbances in the process of stall inception. Examples are given of different ways in which stall cells can develop and the conclusions are backed up with a summary of current test data from various machines around the world.

Mini- and micro-gas turbines for combined heat and power

2002-12-01

P.A. Pilavachi

Criteria for Spike Initiated Rotating Stall

2008-01-01

Huu Duc Vo , C. S. Tan , E. M. Greitzer

A computational study to define the phenomena that lead to the onset of short length-scale (spike) rotating stall disturbances has been carried out. Based on unsteady simulations, we hypothesize there … A computational study to define the phenomena that lead to the onset of short length-scale (spike) rotating stall disturbances has been carried out. Based on unsteady simulations, we hypothesize there are two conditions necessary for the formation of spike disturbances, both of which are linked to the tip clearance flow. One is that the interface between the tip clearance and oncoming flows becomes parallel to the leading-edge plane. The second is the initiation of backflow, stemming from the fluid in adjacent passages, at the trailing-edge plane. The two criteria also imply a circumferential length scale for spike disturbances. The hypothesis and scenario developed are consistent with numerical simulations and experimental observations of axial compressor stall inception. A comparison of calculations for multiple blades with those for single passages also allows statements to be made about the utility of single passage computations as a descriptor of compressor stall.

Flow Visualization in a Linear Turbine Cascade of High Performance Turbine Blades

1997-01-01

H. P. Wang , S. J. Olson , R. J. Goldstein +1 more

Multiple smoke wires are used to investigate the secondary flow near the endwall of a plane cascade with blade shapes used in high-performance turbine stages. The wires are positioned parallel … Multiple smoke wires are used to investigate the secondary flow near the endwall of a plane cascade with blade shapes used in high-performance turbine stages. The wires are positioned parallel to the endwall and ahead of the cascade, within and outside the endwall boundary layer. The traces of the smoke generated by the wires are visualized with a laser light sheet illuminating various cross sections around the cascade. During the experiment, a periodically fluctuating horseshoe vortex system of varying number of vortices is observed near the leading edge of the cascade. A series of photographs and video tapes was taken in the cascade to trace these vortices. The development and evolution of the horseshoe vortex and the passage vortex are clearly resolved in the photographs. The interation between the suction side leg of the horseshoe vortex and the passage vortex is also observed in the experiment. A vortex induced by the passage vortex, starting about one-fourth of the curvilinear distance along the blade on the suction surface, is also found. This vortex stays close to the suction surface and above the passage vortex in the laminar flow region on the blade. From this flow visualization, a model describing the secondary flows in a cascade is proposed and compared with previous published models. Some naphthalene mass transfer results from a blade near an endwall are cited and compared with the current model. The flows inferred from the two techniques are in good agreement.

A Correlation-Based Transition Model Using Local Variables—Part II: Test Cases and Industrial Applications

2004-03-01

Robin Langtry , Florian Menter , S. R. Likki +3 more

A new correlation-based transition model has been developed, which is built strictly on local variables. As a result, the transition model is compatible with modern computational fluid dynamics (CFD) methods … A new correlation-based transition model has been developed, which is built strictly on local variables. As a result, the transition model is compatible with modern computational fluid dynamics (CFD) methods using unstructured grids and massive parallel execution. The model is based on two transport equations, one for the intermittency and one for the transition onset criteria in terms of momentum thickness Reynolds number. The proposed transport equations do not attempt to model the physics of the transition process (unlike, e.g., turbulence models), but form a framework for the implementation of correlation-based models into general-purpose CFD methods. Part I of this paper (Menter, F. R., Langtry, R. B., Likki, S. R., Suzen, Y. B., Huang, P. G., and Völker, S., 2006, ASME J. Turbomach., 128(3), pp. 413–422) gives a detailed description of the mathematical formulation of the model and some of the basic test cases used for model validation. Part II (this part) details a significant number of test cases that have been used to validate the transition model for turbomachinery and aerodynamic applications, including the drag crisis of a cylinder, separation-induced transition on a circular leading edge, and natural transition on a wind turbine airfoil. Turbomachinery test cases include a highly loaded compressor cascade, a low-pressure turbine blade, a transonic turbine guide vane, a 3D annular compressor cascade, and unsteady transition due to wake impingement. In addition, predictions are shown for an actual industrial application, namely, a GE low-pressure turbine vane. In all cases, good agreement with the experiments could be achieved and the authors believe that the current model is a significant step forward in engineering transition modeling.

A Correlation-Based Transition Model Using Local Variables—Part I: Model Formulation

2006-01-01

Florian Menter , Robin Langtry , S. R. Likki +3 more

A new correlation-based transition model has been developed, which is based strictly on local variables. As a result, the transition model is compatible with modern computational fluid dynamics (CFD) approaches, … A new correlation-based transition model has been developed, which is based strictly on local variables. As a result, the transition model is compatible with modern computational fluid dynamics (CFD) approaches, such as unstructured grids and massive parallel execution. The model is based on two transport equations, one for intermittency and one for the transition onset criteria in terms of momentum thickness Reynolds number. The proposed transport equations do not attempt to model the physics of the transition process (unlike, e.g., turbulence models) but form a framework for the implementation of correlation-based models into general-purpose CFD methods. Part I (this part) of this paper gives a detailed description of the mathematical formulation of the model and some of the basic test cases used for model validation, including a two-dimensional turbine blade. Part II (Langtry, R. B., Menter, F. R., Likki, S. R., Suzen, Y. B., Huang, P. G., and Völker, S., 2006, ASME J. Turbomach., 128(3), pp. 423–434) of the paper details a significant number of test cases that have been used to validate the transition model for turbomachinery and aerodynamic applications. The authors believe that the current formulation is a significant step forward in engineering transition modeling, as it allows the combination of correlation-based transition models with general purpose CFD codes.

Secondary Flows in Axial Turbines—A Review

2001-05-01

Lee S. Langston

A bstract : An important problem that arises in the design and the performance of axial flow turbines is the understanding, analysis, prediction and control of secondary flows. Sieverding 1 … A bstract : An important problem that arises in the design and the performance of axial flow turbines is the understanding, analysis, prediction and control of secondary flows. Sieverding 1 has given a review of secondary flow literature, covering up to 1985. In this paper a brief review of pre‐1985 work is given, and then a survey of open literature secondary flow investigations since the Sieverding review is presented. Most of the studies reviewed deal with plane or annular cascade flows. Tip clearance effects are not covered. The basic secondary flow picture for a turbine cascade, as measured and verified by a number of investigators is described. Recent work that shows refined secondary flow vortex structures is examined. A flow parameter based on inlet boundary layer properties used to predict horseshoe vortex swirl is presented. Work on secondary flow loss reduction, involving airfoil geometry, endwall fences and endwall contouring is briefly reviewed. A new leading edge bulb geometry that has demonstrated impressive loss reduction is considered. It is concluded that accurate routine prediction of secondary flow losses has not yet been achieved, and must await either a better turbulence model or more experiments to reveal new endwall loss production mechanisms. Lastly, loss is examined from the standpoint of entropy generation.

Numerical Solution of Periodic Transonic Flow through a Fan Stage

1977-11-01

John Erdos , E. Alzner , W. D. Mcnally

A numerical method of solution of the inviscid, compressible, two-dimensional, unsteady flow on a blade-toblade stream surface through a stage (rotor and stator), or a single blade row, of an … A numerical method of solution of the inviscid, compressible, two-dimensional, unsteady flow on a blade-toblade stream surface through a stage (rotor and stator), or a single blade row, of an axial flow compressor or fan is described. A cyclic procedure has been developed for representation of adjacent blade-to-blade passages, which asymptotically achieves the correct phase between all passages of a stage. A shock-capturin g finitedifference method is employed in the interior of the passage, and a method-of-characteristiics technique is used at the boundaries. The blade slipstreams form two of the passage boundaries, and are treated as moving contact surfaces capable of supporting jumps in entropy and tangential velocity. The Kutta condition is imposed by requiring the slipstreams to originate at the trailing edges which are assumed to be sharp. Results are presented for a transonic fan stage. The rotor tip section solution is compared with an experimental pressure contour map. The subcritical stator solution is compared with results from a stream function method. Finally, the periodic solution for the stage, which has 44 rotor blades and 46 stator blades, is discussed.

The 1991 IGTI Scholar Lecture: The Role of Laminar-Turbulent Transition in Gas Turbine Engines

1991-10-01

R. E. Mayle

A critical study of laminar-turbulent transition phenomena and their role in aerodynamics and heat transfer in modern and future gas turbine engines is presented. In order to develop a coherent … A critical study of laminar-turbulent transition phenomena and their role in aerodynamics and heat transfer in modern and future gas turbine engines is presented. In order to develop a coherent view of the subject, a current look at transition phenomena from both a theoretical and experimental standpoint are provided and a comprehensive state-of-the-art account of transitional phenomena in the engine’s throughflow components given. The impact of transitional flow on engine design is discussed and suggestions for future research and developmental work provided.

Computation of Unsteady Nonlinear Flows in Cascades Using a Harmonic Balance Technique

2002-05-01

Kenneth C. Hall , Jeffrey Thomas , W. Scott Clark

A harmonic balance technique for modeling unsteady nonlinear e ows in turbomachinery is presented. The analysis exploits the fact that many unsteady e ows of interest in turbomachinery are periodic … A harmonic balance technique for modeling unsteady nonlinear e ows in turbomachinery is presented. The analysis exploits the fact that many unsteady e ows of interest in turbomachinery are periodic in time. Thus, the unsteady e ow conservation variables may be represented by a Fourier series in time with spatially varying coefe cients. This assumption leads to a harmonic balance form of the Euler or Navier ‐Stokes equations, which, in turn, can be solved efe ciently as a steady problem using conventional computational e uid dynamic (CFD) methods, including pseudotime time marching with local time stepping and multigrid acceleration. Thus, the method is computationally efe cient, at least one to two orders of magnitude faster than conventional nonlinear time-domain CFD simulations. Computational results for unsteady, transonic, viscous e ow in the front stage rotor of a high-pressure compressor demonstrate that even strongly nonlinear e ows can be modeled to engineering accuracy with a small number of terms retained in the Fourier series representation of the e ow. Furthermore, in some cases, e uid nonlinearities are found to be important for surprisingly small blade vibrations.

A Study of Spike and Modal Stall Phenomena in a Low-Speed Axial Compressor

1997-06-02

Thomas R. Camp , I. J. Day

This paper presents a study of stall inception mechanisms a in low-speed axial compressor. Previous work has identified two common flow breakdown sequences, the first associated with a short lengthscale … This paper presents a study of stall inception mechanisms a in low-speed axial compressor. Previous work has identified two common flow breakdown sequences, the first associated with a short lengthscale disturbance known as a ‘spike’, and the second with a longer lengthscale disturbance known as a ‘modal oscillation’. In this paper the physical differences between these two mechanisms are illustrated with detailed measurements. Experimental results are also presented which relate the occurrence of the two stalling mechanisms to the operating conditions of the compressor. It is shown that the stability criteria for the two disturbances are different: long lengthscale disturbances are related to a two-dimensional instability of the whole compression system, while short lengthscale disturbances indicate a three-dimensional breakdown of the flow-field associated with high rotor incidence angles. Based on the experimental measurements, a simple model is proposed which explains the type of stall inception pattern observed in a particular compressor. Measurements from a single stage low-speed compressor and from a multistage high-speed compressor are presented in support of the model.

Loss Mechanisms in Turbomachines

1993-05-24

J. D. Denton

The origins and effects of loss in turbomachines are discussed with the emphasis on trying to understand the physical origins of loss rather than on reviewing the available prediction methods. … The origins and effects of loss in turbomachines are discussed with the emphasis on trying to understand the physical origins of loss rather than on reviewing the available prediction methods. Loss is defined in terms of entropy increase and the relationship of this to the more familiar loss coefficients is derived and discussed. The sources of entropy are in general: Viscous effects in boundary layers, viscous effects in mixing processes, shock waves and heat transfer across temperature differences. These are first discussed in general and then the results are applied to turbomachinery flows. Understanding of the loss due to heat transfer requires some discussion of cycle thermodynamics. Sections are devoted to discussing: Blade boundary layer and trailing edge loss, tip leakage loss, endwall loss, effects of heat transfer and miscellaneous losses. The loss arising from boundary layer separation is particularly difficult to quantify. Most of the discussion is based on axial flow machines but a separate section is devoted to the special problems of radial flow machines. In some cases, eg attached blade boundary layers, the loss mechanisms are well understood, but even so the loss can seldom be predicted with great accuracy. In many other cases, eg endwall loss, the loss mechanisms are still not clearly understood and prediction methods remain very dependent on correlations. The paper emphasises that the use of correlations should not be a substitute for trying to understand the origins of loss and suggests that a good physical understanding of the latter may be more valuable than a quantitative prediction.

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Centrifugal Compressors: A Strategy for Aerodynamic Design and Analysis

2000-01-01

Ronald H. Aungier

While several books are available that provide a general overview of centrifugal compressor aerodynamic technology, this book is unique in that it fully describes a working design and analysis system … While several books are available that provide a general overview of centrifugal compressor aerodynamic technology, this book is unique in that it fully describes a working design and analysis system with all of the interacting procedures, design guidelines, and decision processes required. This book describes the author’s own centrifugal compressor aerodynamic design and analysis system, and the strategy he uses while applying it. He provides a description sufficiently complete that both new and experienced compressor aerodynamicists will fully understand the methods used. This includes the basic thermodynamic and fluid dynamic principles, empirical models, and key numerical methods, which form the basis of these design and analysis methods. This book provides a comprehensive aerodynamic design and analysis system for centrifugal compressors that has produced significant performance improvements in recent years. It uses practical and efficient methodology and requires minimal resources for its implementation. A personal computer of modest capability is adequate for implementing and using all of the procedures described in this book.

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Compressor Surge and Stall Propagation

1955-05-01

Howard W. Emmons , Carl E. Pearson , H. P. Grant

Abstract Compressor surge is shown by the application of several types of instrumentation (notably a hot-wire anemometer) to consist of two distinct types of phenomena. The whole compressor flow system … Abstract Compressor surge is shown by the application of several types of instrumentation (notably a hot-wire anemometer) to consist of two distinct types of phenomena. The whole compressor flow system may be unstable in the manner of a self-excited Helmholtz resonator. The theory of this instability is presented and is shown to explain some of the observed pulsation symptoms. The stalling of the flow through the blade rows, which usually is assumed to be the origin of pulsation, is shown to occur in propagating groups of 1 to 5 regions involving from 2 to 20 blades each. The theory of this “stall propagation” shows the propagation velocity relative to the wheel to be dependent upon boundary-layer growth parameters and hence the frequency (relative to a stationary probe) to be proportional to the wheel speed. Another part of observed compressor pulsation thus is explained. These two phenomena frequently interact to produce complex performance characteristics. The theories presented are essentially correct as shown by experimental verification, but much remains to be done to make quantitative compressor-performance prediction practical.

Film Cooling With Compound Angle Holes: Adiabatic Effectiveness

1996-10-01

Donald L. Schmidt , Basav Sen , David G. Bogard

Film cooling effectiveness was studied experimentally in a flat plate test facility with zero pressure gradient using a single row of inclined holes, which injected high-density, cryogenically cooled air. Round … Film cooling effectiveness was studied experimentally in a flat plate test facility with zero pressure gradient using a single row of inclined holes, which injected high-density, cryogenically cooled air. Round holes and holes with a diffusing expanded exit were directed laterally away from the free-stream direction with a compound angle of 60 deg. Comparisons were made with a baseline case of round holes aligned with the free stream. The effects of doubling the hole spacing to six hole diameters for each geometry were also examined. Experiments were performed at a density ratio of 1.6 with a range of blowing ratios from 0.5 to 2.5 and momentum flux ratios from 0.16 to 3.9. Lateral distributions of adiabatic effectiveness results were determined at streamwise distances from 3 D to 15 D downstream of the injection holes. All hole geometries had similar maximum spatially averaged effectiveness at a low momentum flux ratio of I = 0.25, but the round and expanded exit holes with compound angle had significantly greater effectiveness at larger momentum flux ratios. The compound angle holes with expanded exits had a much improved lateral distribution of coolant near the hole for all momentum flux ratios.

Gas Turbine Engineering Handbook

2012-01-01

JOHN W. SAWYER

A One-Equation Local Correlation-Based Transition Model

2015-07-04

Florian Menter , Pavel E. Smirnov , Tao Liu +1 more

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Experimental Study on Radial Compressor Instability Caused by Inlet Airflow Restriction

2025-06-24

Marián Hocko , Maroš Divok , M. P. Petrenko | Acta Avionica Journal

This work examines the issue of unstable performance in aircraft turbine engine compressors, focusing on radial compressors. It investigates the causes, mechanisms, and effects of compressor instability, such as airflow … This work examines the issue of unstable performance in aircraft turbine engine compressors, focusing on radial compressors. It investigates the causes, mechanisms, and effects of compressor instability, such as airflow pulsations, pressure variations, engine vibrations, and the risk of compressor blade damage. The theoretical section explains how deviations in airflow rate from design specifications can result in flow separation and a transition from laminar to turbulent flow, which may lead to surge and possible reverse airflow. The analysis of real MPM-20 engine measurements and results for key parameters such as pressures, temperatures, and airflow rates showed significant agreement with models, confirming the theoretical assumptions. These findings underline the crucial role of stable airflow management and effective compressor design in maintaining the safe operation of aircraft engines, especially under varying and extreme flight conditions.This work examines the issue of unstable performance in aircraft turbine engine compressors, focusing on radial compressors. It investigates the causes, mechanisms, and effects of compressor instability, such as airflow pulsations, pressure variations, engine vibrations, and the risk of compressor blade damage. The theoretical section explains how deviations in airflow rate from design specifications can result in flow separation and a transition from laminar to turbulent flow, which may lead to surge and possible reverse airflow. The analysis of real MPM-20 engine measurements and results for key parameters such as pressures, temperatures, and airflow rates showed significant agreement with models, confirming the theoretical assumptions. These findings underline the crucial role of stable airflow management and effective compressor design in maintaining the safe operation of aircraft engines, especially under varying and extreme flight conditions.

Geometric Optimization of Coanda Jet Chamber Fins via Response Surface Methodology

2025-06-23

Hui Zhang , Kai Yue , Yiming Zhang | Aerospace

A highly loaded axial flow compressor often leads to significant flow separation, resulting in increased pressure loss and deterioration of the pressure increase ability. Improving flow separation within a compressor … A highly loaded axial flow compressor often leads to significant flow separation, resulting in increased pressure loss and deterioration of the pressure increase ability. Improving flow separation within a compressor is crucial for enhancing aeroengine performance. This study proposes adding a fin structure to the jet cavity of the Coanda jet cascade to improve flow separation at the trailing edge and corner area. The fin structure is optimized using response surface technique and a multi-objective genetic algorithm based on numerical simulation, enabling more effective control of the simultaneous separation of the boundary corner and trailing edge of the layer. The response surface model developed in this study is accurately validated. The numerical results demonstrate a 2.13% reduction in the optimized blade total pressure loss coefficient and a 12.74% reduction in the endwall loss coefficient compared to those of the original unfinned construction under the same air injection conditions. The optimization procedure markedly improves flow separation in the compressor, leading to a considerable decrease in the volume of low-energy fluid on the blade’s suction surface, particularly in the corner area. The aerodynamic performance of the high-load cascade is enhanced.

Mechanism of improvement in stability margin and choking flow capacity of a centrifugal compressor with variable geometry casing treatment

2025-06-21

Zhiyuan Liu , Ben Zhao , Wentao Shang | Proceedings of the Institution of Mechanical Engineers Part G Journal of Aerospace Engineering

Casing treatments in centrifugal compressors conventionally employ a single port positioned above the impeller tip, which effectively increases the surge margin but provides limited enhancement of choke flow capacity due … Casing treatments in centrifugal compressors conventionally employ a single port positioned above the impeller tip, which effectively increases the surge margin but provides limited enhancement of choke flow capacity due to a mismatch between port location and the shroud surface pressure gradient. In this study, a dual-ported shroud casing treatment with two ports arranged in meridional direction over the impeller tips is proposed to resolve this mismatch. Compressor aerodynamic performance was quantified through complementary experimental tests and three-dimensional numerical simulations. Results show that the dual-ported configuration not only yields a more pronounced improvement in surge margin than the conventional single-port design, but also increases choke flow capacity by up to 14.15% in mass flow rate at the maximum rotational speed. This enhancement arises because the dual-ported shroud provides an auxiliary flow path into the tip region—bypassing the impeller throat and mitigating local flow separation. Finally, an active control strategy is introduced, whereby the shroud ports are dynamically switched to optimize isentropic efficiency and extend the compressor’s stable operating envelope.

Investigations on the Effect of Inclination Angle on the Aerodynamic Performance of a Two-Stage Centrifugal Compressor of a Proton Exchange Membrane Fuel Cell System

2025-06-18

Wenke Wang , Dengfeng Yang , Guo Li +5 more | Energies

This study examines how leading-edge inclination angles affect a two-stage centrifugal compressor’s aerodynamic performance using numerical and experimental methods. Five impellers with varied inclination configurations were designed for both stages. … This study examines how leading-edge inclination angles affect a two-stage centrifugal compressor’s aerodynamic performance using numerical and experimental methods. Five impellers with varied inclination configurations were designed for both stages. The results show that negative inclination improves the pressure ratio and efficiency under near-choke conditions, with greater enhancements in the low-pressure stage. Positive inclination significantly boosts the pressure ratio and efficiency under near-stall conditions, particularly in the low-pressure stage. Negative inclinations optimize blade loading and choke flow capacity, while effectively reducing incidence angle deviations induced by interstage pipeline distortion and decreasing outlet pressure fluctuation amplitude in the high-pressure stage. Positive inclinations delay flow separation, suppress tip leakage vortices, and extend the stall margin.

Verification and validation of blade-resolved viscous-flow tidal turbine simulations

2025-06-16

Manuel Rentschler , Tiago Gomes , Guilherme Vaz +2 more | International Marine Energy Journal

Tidal turbines are a renewable energy source on the rise. The exceptional predictability of tidal currents contributes to a high reliability of this technology, which represents a key advantage in … Tidal turbines are a renewable energy source on the rise. The exceptional predictability of tidal currents contributes to a high reliability of this technology, which represents a key advantage in the endeavor to become a major contributor to the energy mix. To foster the development and to support the design process of tidal turbines, reliable numerical modeling techniques are required. This paper presents verification and validation work performed within the framework of the Supergen ORE Tidal Turbine Benchmarking Study. Viscous-flow CFD code ReFRESCO is used to conduct blade-resolved simulations of the towing tank experiments. In a first approximation, a steady-state frozen-rotor approach is chosen. A transition model, Gamma-ReTheta, is employed to predict the flow state transition on the turbine blades. In the process, the sensitivity to input turbulence quantities is highlighted. The numerical uncertainty is estimated based on mesh refinements. Finally, a conclusion is drawn to which accuracy the presented numerical models can predict the outcome of the experiments.

Vibration and Flutter Analysis of Fan Blades at Design Speed in an Aircraft Turbofan Engine

2025-06-16

Dong Hyun Kim , Seoung-Jin Go , Dong-Il Kim +1 more | The KSFM Journal of Fluid Machinery

Performance Sensitivity Analysis of a Power Generation Gas Turbine Subject to a Control Curve

2025-06-16

Young-Kwang Park , Tong-Seop Kim | The KSFM Journal of Fluid Machinery

Experimental and CFD-Based Response Surface Methodology Optimization of Rotor–Stator Blade Configuration in Micro-Kaplan Turbines

2025-06-16

Kada Kada , Areej Khalil , Mohamed Abousabae +3 more | Journal of energy resources technology.

Abstract Kaplan turbines are widely recognized for their efficiency and adaptability in low-head hydropower systems, making them critical components in renewable energy infrastructure. This study combines experimental testing, high-fidelity computational … Abstract Kaplan turbines are widely recognized for their efficiency and adaptability in low-head hydropower systems, making them critical components in renewable energy infrastructure. This study combines experimental testing, high-fidelity computational fluid dynamics (CFD), and Response Surface Methodology (RSM)-based Design of Experiments (DOE) to evaluate the influence of rotor blade count, stator blade count, and inlet velocity on turbine performance. A three-level factorial DOE with four center points was used to efficiently capture nonlinearities and interactions among the design factors. The CFD mode was validated by comparing its power predictions with experimental data from a closed-loop hydraulic test rig under controlled flow conditions. The model showed good agreement with experimental results, with deviations of less than ±5% across the tested range of velocities, confirming its accuracy. This validated CFD model was then used to construct second-order regression surfaces for key performance metrics: power, efficiency, and pressure drop. Results indicated that while an increase in blade count improves power output, it also leads to higher pressure losses and reduced efficiency beyond an optimal configuration, reflecting the inherent trade-offs in Kaplan turbine design. Significant interactions between blade geometry and flow conditions highlighted the need for coordinated rotor and stator tuning. This integrated approach offers a predictive, resource-efficient framework for optimizing micro-Kaplan turbines, aiding the development of compact, high-performance hydropower systems that contribute to global sustainability goals.

Objective Functions for the Blade Shape Optimisation of a Cross-Flow Tidal Turbine under Constraints

2025-06-16

Karla Ruiz-Hussmann , Pierre-Luc Delafin , Cyrille Bonamy +3 more | International Marine Energy Journal

Hydro-kinetic cross-flow tidal turbines (CFTT) are omni-directional and offer higher area-based power density compared to horizontal-axis tidal turbines, making them very attractive for tidal energy exploitation. However, the rotating motion … Hydro-kinetic cross-flow tidal turbines (CFTT) are omni-directional and offer higher area-based power density compared to horizontal-axis tidal turbines, making them very attractive for tidal energy exploitation. However, the rotating motion around the vertical axis results in continuously varying angles of attack, causing alternating loads, which may lead to fatigue failure and structural damage. The OPTIDE Project addresses these challenges by implementing intracycle blade pitching to individually control the angle of attack, increasing the power coefficient CP and reducing structural loads. For this purpose a Darrieus turbine is designed with embedded actuators in each blade. Firstly a blade shape optimization will be conducted to fit the actuator at the quarter-chord position while ensuring sufficient thickness. The optimization procedure couples Computational Fluid Dynamics (CFD) with a Genetic Algorithm. The employed optimizer OPAL++ sets ten variables for each individual, which describe the hydrofoil shape, length and tip speed ratio (TSR). A smooth hydrofoil shape is generated from the variables, followed by an automatic mesh generation. Subsequently, numerical simulations of each individual at the desired TSR are conducted, while keeping the blade pitch angle constant. Simulation results provide the CP and stress acting on the turbine blades, which are the two optimization objectives (maximize CP while minimizing stress). This process is repeated during the optimization, aiming to determine the most suitable blade shape, that fits the actuator, and operating point (TSR) in a trade-off between CP and structural loads. This will lead to the increase of efficiency and a longer turbine lifetime.

experimental optimization environment for developing an intracycle pitch control in cross flow turbines

2025-06-16

Stefan Hoerner , Roberto Leidhold , Shokoofeh Abbaszadeh +7 more | International Marine Energy Journal

Cross-flow tidal turbines have not reached the efficiency of horizontal-axis turbines. Among various improvement approaches found in the literature, the intracycle pitch control is one of the most promising ones. … Cross-flow tidal turbines have not reached the efficiency of horizontal-axis turbines. Among various improvement approaches found in the literature, the intracycle pitch control is one of the most promising ones. It consists in actively pitching each blade as function of the azimuth angle. This paper presents the development of an experimental environment to optimize the intracycle control for cross-flow tidal turbines, the aim of the project OPTIDE, a collaboration of the Otto-von-Guericke-University Magdeburg, Germany, the University Grenoble-Alpes in France and the University of Applied Sciences Magdeburg-Stendal. The experimental system consists of a turbine flume model equipped with a controlled generator. The three bladed turbine will have independently controlled pitch actuators embedded in each blade with a force sensing system. Local high dynamic systems control the speed of the generator and the pitch position as a function of the azimuth angle. A superimposed governing system performs an experimental optimization process. Evolutionary algorithms are to be employed to solve two objectives: (1) maximizing the efficiency and (2) minimizing the structural loads. These kind of problems are usually solved with numerical simulations, which commonly comes with very high computational efforts and uncertainty. For this reason simulations are replaced by an experimental optimization technique. It is expected that this method could quickly and reliably find optimal pitch trajectories once the setup is fully installed. In the course of the project selected cases will be analysed using particle-image velocimetry (PIV) with synchronized force measurements to research the hydrodynamic mechanisms and the effects of the flow control performed by the pitching motion.In this paper we provide an overview over the decision and design process of the experimental flume setup and the project goals.

Comparative Study of Core Driven Fan Stage Aerodynamic Layout

2025-06-16

Jian Xiao , Guoping Huang , Chen Xia +1 more | Advances in transdisciplinary engineering

The core driven fan stage (CDFS) serves as a key component for realizing wide bypass ratio adjustment in variable cycle engines (VCE). Significant variations in rotor outlet flow angle across … The core driven fan stage (CDFS) serves as a key component for realizing wide bypass ratio adjustment in variable cycle engines (VCE). Significant variations in rotor outlet flow angle across different operating modes pose challenges for matching CDFS with downstream components. To address the dual requirements of blade performance and component matching under various operating conditions, three CDFS type types were compared through analysis. The non-stator type demonstrates structural compactness and high efficiency, yet suffers from limited stability margin and poor downstream compatibility. While the classical stator type achieves optimal downstream matching, it exhibits lower efficiency. The small angle stator type effectively balances efficiency, stability margin, and downstream compatibility. Compared to the non-stator type, the small angle stator type enhances CDFS stability margin by 4.0% in single bypass mode and 4.4% in double bypass mode. These findings establish the small angle stator type as the aerodynamically optimal CDFS layout solution.

Effect of pressure-side squealer shape and coolant supply on tip aerothermal characteristics in a turbine blade

2025-06-14

Hui Li , Han Yan , Songtao Wang +1 more | International Communications in Heat and Mass Transfer

Film cooling and its combination with thermal barrier coating for hypersonic engine

2025-06-13

Huikun Cai , Haorui Liao , Haiping Li +2 more | International Journal of Thermal Sciences

Film Cooling Performance and Superposition Method of an Actual Turbine Vane at High Freestream Turbulence

2025-06-12

Peng Chu , Yongfeng Sui , Bin Dai +5 more | Aerospace

This study aims to enhance the understanding of film cooling performance in an actual turbine vane by investigating influencing factors and developing more precise numerical prediction methods. Pressure sensitive paint … This study aims to enhance the understanding of film cooling performance in an actual turbine vane by investigating influencing factors and developing more precise numerical prediction methods. Pressure sensitive paint (PSP) testing and Reynolds-Averaged Navier–Stokes (RANS) simulations were conducted. The findings indicate that the current design blowing ratio of S1 holes (0.89) is too high, resulting in poor film cooling effectiveness. However, the blowing ratios of P3 (0.78) and P4 (0.69) holes are relatively low, suggesting that increasing the coolant flow could improve the film cooling effectiveness. It is not recommended to design an excessively low blowing ratio on the suction surface, as this can lead to poor wall adherence downstream of the film holes. A slight increase in turbulence intensity enhances the film covering effect, particularly on the suction surface. Additionally, a novel superposition method for multirow fan-shaped film cooling holes on an actual turbine vane is proposed, exhibiting better agreement with experimental data. Compared with experimental results, the numerical predictions tend to underestimate the film cooling effectiveness with the examined k-ε-based viscosity turbulence models and Reynolds stress turbulence models, while the SST demonstrates relatively higher accuracy owing to its hybrid k-ω/k-ε formulation that better resolves near-wall physics and separation flows characteristic of turbine cooling configurations. This study contributes to the advancement of turbine vane thermal analysis and design in engineering applications.

Understanding savonius turbine wake for 2-D and 3-D simulations: proper orthogonal decomposition of velocity components

2025-06-10

Shivam Singh Tomar , Anupam Dewan | Sustainable Energy Technologies and Assessments

Optimal Management of Hybrid Thermal-Wind Sources and FACTS Devices Using Improved Quadratic Interpolation Optimization

2025-06-10

Belkacem Mahdad | Engineering Research Express

Abstract The growing integration of various types of renewable energy sources (RES) in coordination with Flexible AC transmission systems (FACTS) presents new challenges for modern power system management. This paper … Abstract The growing integration of various types of renewable energy sources (RES) in coordination with Flexible AC transmission systems (FACTS) presents new challenges for modern power system management. This paper introduces a novel stochastic planning strategy based Improved Quadratic Interpolation Optimization (IQIO). The proposed IQIO is characterized by adopting a novel adaptive search mechanism to efficiently explore the search space during exploitation stage to handle the complexities arising from the inherent stochastic nature of various types of RES generation in particular wind and photovoltaic sources. The security energy management is optimally analyzed and evaluated by minimizing individually and simultaneously the total cost of thermal units and wind sources, the total power loss, and the total voltage fluctuations. The stochastic aspect associated to wind farms is considered and modeled using Weibull distribution. In the other hand, to ensure flexible and sustainable reactive power support, multi shunt compensators, and multiple series controllers based FACTS devices are modeled and optimally integrated at specified buses. The effectiveness of IQIO is evaluated on the modal and multi modal benchmark functions and to the practical power system IEEE 30-Bus under normal and abnormal scenarios such as critical load growth. For single objective function, the proposed method effectively integrates nine Static VAR compensators (SVC) and two Thyristor Controlled Series Capacitor (TCSC) to achieve a competitive total cost (thermal and wind) of 806.9972 $/h. On the other hand, the innovative IQIO approach prioritizes high loading margin stability (1.676 p.u) at lower penetration level of wind farms (28.4223 %), and a significant reduction in TR.P.L (14.8375 MW) while strictly adhering to security constraints. Comparisons with other recent competitive metaheuristic approaches demonstrate its superior performance in terms of solution quality and convergence accuracy. This research suggests that IQIO is a promising tool for optimizing large scale modern electric power systems considering the growing integration of both renewable sources and various types of FACTS technology.

Review of Film Cooling Techniques for Aerospace Vehicles

2025-06-10

Edidiong Michael Umana , Xiufeng Yang | Energies

Film cooling, a vital method for controlling surface temperatures in components subjected to intense heat, strives to enhance efficiency through innovative technological advancements. Over the last several decades, considerable advancements … Film cooling, a vital method for controlling surface temperatures in components subjected to intense heat, strives to enhance efficiency through innovative technological advancements. Over the last several decades, considerable advancements have been made in film cooling technologies for applications such as liquid rocket engines, combustion chambers, nozzle sections, gas turbine components, and hypersonic vehicles, all of which operate under extreme temperatures. This review presents an in-depth investigation of film cooling, its applications, and its key mechanisms and performance characteristics. The review also explores design optimization for combustion chamber components and examines the role of gaseous film cooling in nozzle systems, supported by experimental and numerical validation. Gas turbine cooling relies on integrated methods, including internal and external cooling, material selection, and coolant treatment to prevent overheating. Notably, the cross-flow jet in blade cooling improves heat transfer and reduces thermal fatigue. Film cooling is an indispensable technique for addressing the challenges of high-speed and hypersonic flight, aided by cutting-edge injection methods and advanced transpiration coolants. Special attention is given to factors influencing film cooling performance, as well as state-of-the-art developments in the field. The challenges related to film cooling are reviewed and presented, along with the difficulties in resolving them. Suggestions for addressing these problems in future research are also provided.

Impact of mass flow ratio on the aero-thermal performance of a film-cooled nozzle guide vane

2025-06-10

Babji Gudla , Arun Kumar Pujari | International Journal of Turbo and Jet Engines

Abstract Film cooling is critical in protecting gas turbine nozzle guide vanes from extreme thermal loads, enhancing their durability and efficiency. Understanding the interaction between aerodynamic performances and cooling effectiveness … Abstract Film cooling is critical in protecting gas turbine nozzle guide vanes from extreme thermal loads, enhancing their durability and efficiency. Understanding the interaction between aerodynamic performances and cooling effectiveness is essential for optimizing vane design. This study investigates the aero-thermal characteristics of a film-cooled gas turbine nozzle guide vane under subsonic conditions. ANSYS simulation software was used for the computational study, and the results were experimentally validated using a low-speed cascade wind tunnel setup with a coolant flow unit. The study explored how film cooling in different regions of the vane (Suction surface, pressure surface, leading edge) influences total pressure losses. Findings indicated that both effectiveness and loss coefficient increased with rising MFR. Specifically, an increase in MFR from 0.004 to 0.005 led to a 1.3 % improvement in cooling effectiveness and a 2.57 % rise in pressure loss coefficient. Similarly, when the MFR increased from 0.005 to 0.006, effectiveness improved by 10.86 %, and the total pressure loss coefficient increased by 3.25 %.

Secondary Flows Control in a Modern Aero-Engine Fan using Vortex Generators

2025-06-01

Roberto Putzu , M. Carta , Tiziano Ghisu +2 more | Aerospace Science and Technology

Investigation on the Optimization Design of Non-Axisymmetric Endwall of a Contra-Rotating Turbine

2025-06-01

Zhiyuan Cao , Qiang Gu , Xinyu Hao +2 more | Aerospace Science and Technology

Characteristics and flow mechanism of a transonic variable pitch fan

2025-06-01

W. Su , Shengfeng Zhao , Jingquan Zhao +3 more | Physics of Fluids

Variable pitch fan (VPF) can effectively expand the operating range of the fan stage and broaden the flight envelope by adjusting the rotor stagger angle. Leveraging experimentally validated numerical simulations, … Variable pitch fan (VPF) can effectively expand the operating range of the fan stage and broaden the flight envelope by adjusting the rotor stagger angle. Leveraging experimentally validated numerical simulations, this study investigates the impact of stagger angle variations on the aerodynamic performance and flow field characteristics of transonic variable pitch fans. A dimensionless enthalpy loss coefficient analysis method is employed to quantitatively evaluate the internal flow losses of the VPF. The results indicate that within the studied stagger angle variation range (Δθ = ±8°), the design-point mass flow range varies by 33.3% compared to Δθ = 0°. Under opening stagger angle conditions, the rotor tip flow field structure improves, but suction surface flow separation intensifies in both the rotor and stator, leading to VPF efficiency reduction primarily attributed to increased losses in the rotor and stator. Conversely, under a closed stagger angle, leading-edge spillage occurs in the tip region, causing extensive passage blockage; the decline in VPF efficiency is dominated by rotor losses due to enhanced viscous dissipation and tip leakage vortex interaction.

Integration design for inverted inward-turning inlet and waverider airframe with boundary-layer removal effect

2025-06-01

Xiaogang Zheng , D. N. C. Lin , Zhancang Hu +2 more | Aerospace Science and Technology

Discovery, Interpretation, and Model-based Judgment of a Hidden Operating Condition during the Process from Compressor Stall to Surge Induced by Inlet Distortion

2025-06-01

Yafei Qiao , Wuli Chu , Jiafu Chen +3 more | Aerospace Science and Technology

Loss analysis and evaluation of axial compressor with casing treatment based on local entropy generation method

2025-06-01

Zhidong Chi , Zhiguo Wang , S. R. Hu +1 more | Physics of Fluids

Casing treatments are widely used to enhance compressor stability, yet their impact on compressor efficiency penalty mechanisms remains incompletely quantified. This study employs control volume entropy generation analysis to evaluate … Casing treatments are widely used to enhance compressor stability, yet their impact on compressor efficiency penalty mechanisms remains incompletely quantified. This study employs control volume entropy generation analysis to evaluate the internal flow loss characteristics induced by slot-type casing treatments in an axial compressor. The axial distributions of entropy generation reveal that casing treatments introduce two distinct peaks: a primary enhancement at the rotor leading edge and a secondary peak near the slot trailing edges. The spanwise distributions demonstrate increased losses above 80% span but modest reductions between 65% and 80% span. Flow field visualization identifies that suction effects deflect the tip leakage vortex toward the suction surface, creating localized high-entropy regions, while injection effects promote radial entropy redistribution. Parametric analysis based on the response surface method establishes slot width and axial overlap as the dominant geometric parameters governing loss generation in the upper end wall region, with slot width exhibiting linear correlations and axial overlap showing nonlinear behavior. Self-organizing maps analysis further confirms a strong correlation between efficiency penalty and upper end wall losses. These findings provide quantitative guidelines for optimizing casing treatment geometries to balance stability enhancement and efficiency preservation in axial compressors.