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An efficient numerical method is applied to investigate the influence of different tip gap shapes on the performance of an axial-flow compressor stator under the condition of rotating hub or stationary hub. Tip gap shapes considered in this paper include zero gap, uniform gap, and linearly varying gap. The analysis of the results shows, at least for these test cases, tip leakage vortex can be found to generate earlier for stationary hub cases than for rotating hub cases. On the same working conditions, the blade row with zero gap has not the highest efficiency, while the blade row with smaller tip gap can attain higher efficiency due to the role of tip leakage flow the direction of which is opposite to the secondary flow in the passage near the hub. For the same circumferential leakage area in the tip region, the blade row with linearly expanding gap can obtain higher efficiency. The work of this paper can provide a useful reference for optimum design of compressor stator.Copyright © 2001 by ASME

This paper is Part III of a comprehensive study on the blade leading edge sweep/bend of supersonic and transonic axial compressors. For the curved leading edge shock surfaces generated by high loading swept axial flow rotors, using an analytical structure model of 3-D shock surface, this paper investigated the characteristics of the shock surface structure, then calculated the distributions of the flow variables in front of and behind the shock. The research focused on the influences of the relatively forward- or backward-swept rotors on the flow field behind but near the leading edge shock surface. By the analytical model, this paper answered the one hand of the question of the swept rotors’ functions. For the first time it quantitatively showed the existence of the phenomenon of 3-D shock induced span-wise and circumferential mixing in the primary flow.Copyright © 2001 by ASME

In this second part of a two-part paper on the application of the NASA multistage aerodynamic simulation CFD tool, the APNASA code, work is presented on how the code was used successfully in the design of a brand new, four-stage axial compressor aimed for a modern turbofan engine. In particular, the code was used to guide the blade geometry changes such that the right stage matching throughout the compressor was achieved. The overall performance is shown to be in good agreement with the map generated prior to the detailed design from a one-dimensional model, including the surge line. This success is attributed to the fact that the compressor has been successfully matched at its design intention. More detailed comparisons of the measurements with the prediction demonstrate capability of APNASA in capturing details of flow, such as the rotor exit pressure and temperature profiles and the stator exit flow angle. Largely as a result of the application of APNASA, the compressor described in this paper has been designed “right the first time”, resulting in significant cost savings in a new engine development program for Honeywell Engines and Systems.Copyright © 2001 by ASME

Part I of this paper described how the HP turbine model rig of the Rolls-Royce Trent 500 was redesigned by applying non-axisymmetric end walls to both the vane and blade passages, whilst leaving the turbine operating point and overall flow conditions unaltered. This paper describes the results obtained from testing of the model rig and compares them with those obtained for the datum design (with conventional axisymmetric end walls). Measured improvements in the turbine efficiency are shown to be in line with those expected from the previous linear cascade research at Durham University, see Harvey et al. [1] and Hartland et al. [2]. These improvements are observed at both design and off-design conditions. Hot wire traverses taken at the exit of the rotor show, unexpectedly, that the end wall profiling has caused changes across the whole of the turbine flow field. This result is discussed making reference to a preliminary 3-D CFD analysis. It is concluded that the design methodology described in part I of this paper has been validated, and that non-axisymmetric end wall profiling is now a major new tool for the reduction of secondary loss in turbines (and potentially all axial flow turbomachinery). Further work, though, is needed to fully understand the stage (and multistage) effects of end wall profiling.Copyright © 2001 by ASME

This paper presents a computational investigation of flows in a single axial flow fan with and without casing treatment. It analyses the effect of the recess casing treatment on stall margin improvement as well as its influence on global performance parameters. The paper seeks to offer a contribution to the understanding of the physical processes occurring when approaching stall and the working mechanism by which casing treatments improve stall margin. A Reynolds-Averaged Navier-Stokes CFD code was used for the analysis and the numerical investigation of the overall performance, efficiency and work-input characteristics of the fan were found to agree very well with previously reported experimental results. The effect of casing treatment was investigated using two types of configurations, vaneless and vaned casing. The vaneless casing treatment produced a sizeable stall margin improvement with negligible loss of efficiency. The recess was fitted later with vanes and was shown to offer both a further stall margin improvement and an increase in the pressure rise coefficient without any significant drop in efficiency at design conditions.

This paper is part II of a comprehensive study on the blade leading edge sweep/bend of supersonic and transonic axial compressors. The paper explores and analyzes the kinematic characteristic variables of three-dimensional (3-D) swept shock surfaces. In the research field studying the sweep aerodynamics of axial flow compressors and fans, many types of high loading swept blades are under intensive study. So, in both direct and inverse design methods and experimental validations, an accurate grasp of the sweep characteristic of the blade’s 3-D swept shock surface becomes of more concern than before. Associated with relevant blading variables, this paper studies the forward and zero and backward sweeps of shock surfaces, defines and resolves every kind of useful sweep angle, obtains dimensionless sweep similarity factors, suggests a kind of method for the quantitative classification of 3-D shock structures, and proposes the principle of 3-D shock structure measurements. Two rotor blade leading edge shock surfaces from two high loading single stage fans are analyzed and contrasted. This study is the foundation of the kinematic design of swept shock surfaces.

Magnetic bearings are widely used as active suspension devices in rotating machinery, mainly for active vibration control purposes. The concept of active tip clearance control suggests a new application of magnetic bearings as servo-actuators to stabilize rotating stall in axial compressors. This paper presents a first-of-a-kind feasibility study of an active stall control experiment with a magnetic bearing servo-actuator in the NASA Glenn high-speed single-stage compressor test facility. Together with CFD and experimental data a two-dimensional, incompressible compressor stability model was used in a stochastic estimation and control analysis to determine the required magnetic bearing performance for compressor stall control. The resulting requirements introduced new challenges to the magnetic bearing actuator design. A magnetic bearing servo-actuator was designed which fulfilled the performance specifications. Control laws were then developed to stabilize the compressor shaft. In a second control loop, a constant gain controller was implemented to stabilize rotating stall. A detailed closed loop simulation at 100% corrected design speed resulted in a 2.3% reduction of stalling mass flow which is comparable to results obtained in the same compressor by Weigl et al. (1998) using unsteady air injection. The design and simulation results presented here establish the viability of magnetic bearings for stall control in aero-engine high-speed compressors. Furthermore the paper outlines a general design procedure to develop magnetic bearing servo-actuators for high-speed turbomachinery.

This paper is Part I of a comprehensive study on the blade leading edge sweep/bend of supersonic and transonic axial compressors. The objective of the paper is to develop an analytical shock surface model which links a blade leading edge with the 3-D shock structure, and can be eventually explored as a means to estimate the direct and indirect shock losses. The model surface is composed of helixes on a set of cylinders with aerodynamically determined ascendent for each helix. One kind of the relation between the ascendent and 2-D aerodynamics on S1 stream surface has been established for a leading edge attached shock surface. The succeeding research has revealed that the concepts of a forward/backward swept shock surface, and their qualitative and quantitative natures, can be analysed by the model in uniformity. The model is constructed as such that it can also be used in the inverse method of design process. Two leading edge shock surfaces from two high loading transonic fans are simulated by the model.Copyright © 1999 by ASME

Throughflow theory, as a design problem, is extensively used in the design of transonic axial flow compressors and fans. The losses caused by a simplified passage shock named shock loss model were taken into account in the design of transonic stages. However, the influences of shock waves on flow fields are still under development for transonic design. This problem is studied in the present paper. The basic equations of the throughflow method are re-derived under the concept of discontinuous passage averaging, and then, the distribution shock forces acied on the averaged throughflow fields are obtained. To evaluate these effects on averaged throughflow fields, a passage shock structure model is developed based on the previous models. Finally, the shock influences are assessed and compared in the present paper. It is demonstrated that the distribution shock forces have dramatically effects on the results of design.

This paper describes the development of a computational model for the moderate to long lengthscale flowfield behaviour of high-speed multi-stage compressors. The simulation is a time accurate solution of the fully compressible non-linear equations for fluid flow through an entire compressor. Flow in the non-bladed regions is calculated using a two-dimensional Euler solver whilst the individual blade rows are modelled using multiple one-dimensional flowfields with body forces. The body forces are chosen to give the required steady state blade row performance and a physically based circulation model is used to give the correct non-steady flow dynamics. The capabilities of the simulation are demonstrated in this paper by calculating how the flowfield breakdown of a four stage compressor is affected by operating speed and a range of different inlet flow distortions.Copyright © 1997 by ASME

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.Copyright © 1997 by ASME

This paper deals with the numerical simulation of flow through a 1.5 stage axial flow turbine. The 3-row configuration has been experimentally investigated at the University of Aachen where measurements behind the first vane, the first stage and the full configuration were taken. These measurements allow single blade row computations, to the measured boundary conditions taken from complete engine experiments, or full multistage simulations. The results are openly available inside the framework of ERCOFTAC 1996. There are two separate but interrelated parts to the paper. Firstly, two significantly different Navier-Stokes codes are used to predict the flow around the first vane and the first rotor, both running in isolation. This is used to engender confidence in the code that is subsequently used to model the multiple bladerow tests, the other code is currently only suitable for a single blade row. Secondly, the 1.5 stage results are compared to the experimental data and promote discussion of surrounding blade row effects on multistage solutions.

By far the greater part of our understanding about stall and surge in axial compressors comes from work on low-speed laboratory machines. As a general rule, these machines do not model the compressibility effects present in high-speed compressors and therefore doubt has always existed about the application of low-speed results to high-speed machines. In recent years interest in active control has led to a number of studies of compressor stability in engine type compressors. The instrumentation used in these experiments has been sufficiently detailed that, for the first time, adequate data is available to make direct comparisons between high-speed and low-speed compressors. This paper presents new data from an eight-stage fixed geometry engine compressor and compares this with low-speed laboratory data. The results show remarkable similarities in both the stalling and surging behaviour of the two machines, particularly when the engine compressor is run at intermediate speeds. The engine results also show that, as in the laboratory tests, surge is precipitated by the onset of rotating stall. This is true even at very high speeds where it had previously been thought that surge might be the result of a blast wave moving through the compressor. This paper therefore contains new information about high-speed compressors and confirms that low speed testing is an effective means of obtaining insight into the behaviour of high-speed machines.Copyright © 1993 by ASME

This paper deals with a theoretical study of the active control of purely axial 1D instabilities in multi-stage axial-flow compressors. The paper briefly considers the development of a suitable surge model and continues with the derivation of a controller using linear optimal control theory. The controller is specifically designed to suppress the instabilities predicted by the linearised form of the surge model. The control technique involves a bleed being dynamically varied in response to fluctuations of variables. It is found that a stabilising optimal controller can always be obtained. The second part of the paper presents a non-linear simulation of the surge prediction model with the linear controller, using a 4th order Runge-Kutta scheme. This demonstrates how some non-linearities in the flow model can affect the controller operation. It is also shown that, the amplitudes and the frequency response required of the actuator to retain stability can exceed the practical limitations.Copyright © 1993 by ASME

Several methods are available for the optimization of basic design parameters and the preliminary efficiency prediction of axial flow turbine stages. However, their application is often questionable for stages having low specific speed and/or small volume flow rates. In particular, the question may arise whether a better performance is achieved by a partial admission, impulse stage or by a full admission reaction stage having lower blade height. The paper firstly reviews the available loss correlation methods applicable to partial admission turbines, then a comparison is performed between the efficiency achievable by partial and full admission stages designed for the same operating conditions. The turbine design procedure for both options is fully automatized by an efficiency optimization method similar to the one described in previous authors’ papers. The results of calculations are presented in the paper as a function of similarity parameters (specific speed, size parameter, expansion ratio). It is found that the results obtained with different correlations are relatively similar for “conventional” turbine stages (low expansion ratio, moderate size parameters), while important differences take place for very small sizes and/or in presence of important compressibility effects. The presented results can be useful: 1) to decide whether selecting full or partial admission solutions; 2) to optimize the degree of admission and the other basic design parameters, and 3) to predict with reasonable accuracy the stage efficiency.

Theory researches, numerical analyses and experiment studies of sweep aerodynamics in turbo-machinery had rapidly developed during the 1990s. Taking a low-aspect-ratio, high-loaded axial flow transonic rotor as the baseline, this paper attempts to probe the connotation of sweep aerodynamics of transonic fans and compressors. The definition of sweep angle was discussed. Flow field, key factors that affect structures of flow fields, linkage between leading edges as well as flow field structures in swept rotors were presented. In order to find out the optimum sweep leading edge curve under prescribed conditions, a numerical optimization program that can be added into design processes was developed.Copyright © 2001 by ASME

The subject of the interaction between the mainstream annulus flow and seal flows is now receiving increased attention as a result of concerted attempts to further improve compressor performance. In an axial compressor the use of a shroud on a stator blade row necessitates a trench and results in two such regions of interaction at the upstream and downstream stator well cavities with flow limited by a labyrinth seal. The upstream stator well cavity, with its superposed radial outflow, has received some attention and possess similarities with the cavity downstream of nozzle guide vanes in axial turbines, which have received considerable interest. The downstream compressor stator well cavity has, however, been less extensively studied. Here the flow is dominated by inflow into the cavity. This inflow is driven by the pressure gradient across the blade row and hence the pressure gradient between the upstream and downstream cavities. This paper presents preliminary pressure measurements from a high-speed two-stage research compressor. The experimental results are compared with two and three-dimensional CFD models. In order to gain confidence with CFD for compressor stator wells, the case of inflow into a rotor-stator disc wheelspace has been revisited; validating a commercial code, Fluent, for a wide range of flow conditions. This code has then been used to predict the detailed flow structure for the conditions in the experimental research compressor.Copyright © 2000 by ASME

During the last few decades extremely powerful Quasi-3D codes and fully 3D Navier-Stokes solvers have been developed and successfully utilized in the design process and optimization of multistage axial-flow compressors. However, most of these methods proved to be difficult in handling and extremely time consuming. Due to these disadvantages, the primary stage design and stage matching as well as the off-design analysis is nowadays still based on fast 2D methods incorporating loss-, deviation- and end wall modeling. Only the detailed 3D optimization is normally performed by means of advanced 3D methods. In this paper a fast and efficient 2D calculation method is presented, which already in the initial design phase of multistage axial flow compressors considers the influence of hub leakage flows, tip clearance effects and other end wall flow phenomena. The method is generally based on the fundamental approach by Howard and Gallimore (1992). In order to allow a more accurate prediction of skewed and non-developed boundary layers in turbomachines an improved theoretical approach was implemented. Particularly the splitting of the boundary layers into an axial and tangential component proved to be necessary in order to account for the change between rotating and stationary end walls. Additionally, a new approach is used for the prediction of the viscous end wall zones including hub leakage effects and strongly skewed boundary layers. As a result, empirical correlations for secondary flow effects are no longer required. The results of the improved method are compared with conventional 2D-results including 3D loss- and deviation-models, with, experimental data of a 3-stage research compressor of the Institute for Jet Propulsion and Turbomachinery of the Technical University of Aachen and with 3D Navier-Stokes solutions of the V84.3A compressor and of a multi-stage Siemens research compressor. The results obtained using the new method show a remarkable improvement in comparison with conventional 2D-methods. Due to the high quality and the extremely short computation time the new method allows an overall viscous design of multistage compressors for heavy duty gas turbines and aeroengine applications.

This paper is concerned with the experimental characterisation of the unstable behaviour exhibited by a two stage axial flow compressor with variable geometry inlet guide vanes. The characterisation includes performance assessment stall transient and determination of three dimensional flow during one cell operation. The three flow components have been measured with a single slanted hot wire ensemble phase averaging technique based on several cell passages. This technique, already successfully applied for stall in an isolated rotor, is now tried on the more complex two stage compressor.Copyright © 2000 by ASME

The paper describes an experimental investigation of the stator hub and blade flow in two different stators of a highly loaded single-stage axial-flow low-speed compressor. The first stator (A) is a conventional design with blades of rectangular planform. The second stator (K) is an unconventional, more advanced design with blades of a special planform, characterized by an aft-swept leading edge with increasing sweep angle towards hub and casing. The experimental results show that stator K exhibits a much better hub performance than stator A, finally leading to a better overall performance of stage K compared to stage A. The better hub performance of stator K is, primarily, the result of a planform effect of the newly introduced blades with an aft-swept leading edge and the aerodynamics of an aft-swept wing.Copyright © 2000 by ASME

The history of Sir Frank Whittle’s invention of the jet engine is well known. Somewhat less well known is that the Royal Aircraft Establishment embarked in 1926 on developing the gas turbine as a way of driving a propeller. In 1938, A.R.Howell joined the team as a new graduate, and by 1944 he had played a major role in evolving successful axial compressor design methods, which were used in the first two generations of UK gas turbine engines. He was appointed Head of Aerodynamics Department in the National Gas Turbine Establishment when it was created in 1946, and led that team for twenty years. For many years he was a key figure in compressor design in the UK. He returned to personal research before retiring in 1980, and he died in 1988. This paper summarises his personal research contributions and some of the pioneering research he led in NGTE.Copyright © 2000 by ASME

In Part I of this paper a family of numerically optimized subsonic compressor airfoils for heavy-duty gas turbines, covering a wide range of flow properties, is presented. The objective of the optimization was 10 create profiles with a wide low loss incidence range. Therefore, design point and off-design performance had to be considered in an objective function. The special flow conditions in large scale gas turbines have been taken into account by performing the numerical optimization procedure at high Reynolds numbers and high turbulence levels. The objective of Part II is to examine some of the characteristics describing the new airfoils, as well as to prove the reliability of the design process and the flow solver applied. Therefore, some characteristic members of the new airfoil series have been extensively investigated in the cascade windtunnel of DLR Cologne. Experimental and numerical results show profile Mach number distributions, total pressure losses, flow turning and static pressure rise for the entire incidence range. The design goal with low losses and especially a wide operating range could be confirmed, as well as a mild stall behavior. Boundary layer development, particularly near stall condition, is discussed using surface flow visualization and the results of boundary layer calculations. An additional experimental study, using liquid crystal coating, provides necessary information on suction surface boundary-layer transition at high Reynolds numbers. Finally, results of Navier-Stokes simulations are presented which enlighten the total pressure loss development and flow turning behavior especially at high incidence in relation to the results of the design tool.

In this paper, a method for the performance prediction of multistage axial flow compressors through a steady, three-dimensional, multi-block Navier-Stokes solver is presented. A repeating stage model has been developed aiming at the simplification of the required global aerodynamic boundary conditions for the simulation of the rear stages of multistage axial compressors where only mass flow rate and exit average static pressure are required. The stage inlet velocity distribution is fixed to be equal to the one calculated at the stage exit and the exit static pressure distribution is fixed to have the same shape to that at inlet but maintain its own average value. A mixing plane approach is used to exchange information between neighbouring blade rows which allows both radial and circumferential variations at both sides of the interface. A pressure correction method with the standard k–ε turbulence model is used in combination with Stone’s two step procedure for the solution of the algebraic system of the discretised equations. A global iteration is carried out in order to establish the physical consistency between the blade rows. A combination of two structured grid blocks for the rotor blade row, one for the main passage and a second for the modelling of the tip clearance, is used for a detailed representation of the leakage flows. Computational results from two methods, the first by using the repeating stage model and the second by setting stage inlet velocity profile, are presented from the analysis of the third stage of the four-stage Cranfield Low Speed Research Compressor (LSRC). Good agreements with the experimental data are obtained in terms of total pressure, static pressure and velocity distributions at the inlet, exit and interface planes proving that the repeating stage model is a very economical and accurate alternative to the very expensive complete multistage simulations.

Compressors used for industrial applications are sometimes fitted with some form of variable geometry. Centrifugal compressors may have adjustable inlet guide vanes on one or multiple stages, or diffuser guide vanes at the discharge of the compressor. Axial compressors often have variable inlet guide vanes or adjustable stators. Such variable geometry compressors are becoming more common in the petroleum and chemical industries. For many variable geometry compressors (but not all), the boundary between the stable operating region and surge is a surface, not just a curve. In this paper, we discuss which forms of variable geometry result in a surge limit curve and which forms spread out into a surface. We describe some methods for determining the location of the operating point relative to the surge limit. An example is included showing actual surge points as found in the field. Surge testing to determine the complete surge limit surface is discussed.

A numerical 2D study of the unsteady flow in the first stage of an industrial axial flow gas turbine is presented. The study focuses on the influence of the axial spacing between stator and rotor on the unsteady flow field and relevant design parameters. For six different axial spacings between 18.5 % and 60 % of rotor chord length, time-averaged and time-resolved results are presented, focusing mainly on the pressure distribution on the rotor, stage efficiency and flow angle variation. The instantaneous pressure and entropy contours at midspan are presented for different rotor positions. The underlying physical principles are described making a clear distinction between the impact of the potential flow interaction and the wake-blade interaction. The pressure fluctuations over the stator and rotor blades and the seal air injection gap as well as the effects of axial spacing are investigated. The major focus of this paper is to examine the variation of stage efficiency and rotor incidence angle with a changing axial gap. A simple design rule is derived in order to optimize the vane-blade spacing and it is shown that the gap chosen for the real life machine offers a good compromise between the opposing design requirements.

A new family of subsonic compressor airfoils, which are characterized by low losses and wide operating ranges, has been designed for use in heavy-duty gas turbines. In particular the influence of the higher airfoil Reynolds numbers compared to aeroengine compressors and the impact of these differences on the location of transition are taken into account. The design process itself is carried out by the combination of a geometric code for the airfoil description, with a blade-to-blade solver and a numerical optimization algorithm. The optimization process includes the design-point losses for a specified Q3D flow problem and the off-design performance for the entire operating range. The family covers a wide range of inlet flow angle, Mach number, flow turning, blade thickness, solidity and AVDR in order to consider the entire range of flow conditions that occur in practical compressor design. The superior performance of the new airfoil family is demonstrated by a comparison with conventional controlled diffusion airfoils (CDA). The advantage in performance has been confirmed by detailed experimental investigations, which will be presented in Part II of the paper. This leads to the conclusion that CDA airfoils that have been primarily developed for aeroengine applications are not the optimum solution, if directly transferred to heavy-duty gas turbines. A significant improvement in compressor efficiency is possible, if the new profiles are used instead of conventional airfoils. @S0889-504X~00!02102-4#

The LM2500+ gas turbine, rated between 39,000 to 40,200 shaft horsepower (shp), was introduced for field service in 1998. This growth aero-derivative gas turbine is suitable for a variety of power generation applications, such as co-generation and combined cycle, as well as mechanical drive applications. At the heart of the LM2500+ 25% power increase is an up-rated derivative 17-stage axial compressor. This paper describes the aerodynamic design and development of this high pressure ratio single spool compressor for the LM2500+ gas turbine. The compressor is derived by zero-staging the highly efficient and reliable LM2500 compressor to increase the flow by 23% at a pressure ratio of 23.3:1. The aerodynamic efficiency of the compressor is further improved by using three-dimensional, custom-tailored airfoil designs similar to those used in the CF6-80C2 high pressure compressor. The compressor achieved a peak polytropic efficiency above 91 percent, meeting all its operability objectives. The technical requirements and overall aerodynamic design features of the compressor are presented first. Next, the zero stage match point selection is described and the procedure used to set up the vector diagrams using a through-flow code with secondary flow and mixing is outlined. Detailed design results for the new transonic airfoils in the compressor using three-dimensional viscous analysis are presented. The compressor instrumentation and performance test results are discussed. The performance of the zero stage is separated from that of the baseline compressor with the CF6-80C2 airfoils to show the improvement in efficiency with the new airfoils.Copyright © 1999 by ASME

The structure and propagation of rotating stall cells in a single- and a two-stage subsonic axial compressor is addressed in this paper using computational and experimental analysis. Unsteady solutions of the 2-D inviscid compressible (Euler) equations of motion are presented for one operating point in the fully-developed rotating stall regime for both a single- and a two-stage compressor. The inviscid assumption is verified by comparing the single-stage 2-D in viscid/compressible solution with an equivalent 2-D viscous (Navier-Stokes) result for incompressible flow. The structure of the rotating stall cell is analyzed and compared for the single- and two-stage cases. The numerical solutions are validated against experimental data consisting of flow visualization and unsteady row-by-row static pressure measurements obtained in a four-stage water model of a subsonic compressor. The CFD solutions supply a link between the observed experimental features and provide additional information on the structure of the stall flow. Based on this study. supporting assumptions regarding the driving mechanisms for the propagation of fully-developed rotating stall cells and their structure are postulated. In methodical respect the results suggest that the inviscid model is able to reproduce the essentials of the flow physics associated with the propagation of fully-developed, full-span rotating stall in a subsonic axial compressor.

This paper describes the modelling of axial compressor blade rows in an axisymmetric viscous throughflow method. The basic method, which has been reported before, includes the effects of spanwise mixing, using a turbulent diffusion model, and endwall shear within the throughflow calculation. The blades are modelled using a combination of existing two-dimensional blade performance predictions for loss and deviation away from the annulus walls and a novel approach using tangential blade forces in the endwall regions. Relatively simple assumptions about the behaviour of the tangential static pressure force imposed by the blades allow the secondary deviations produced by tip clearance flows and the boundary layer flows at fixed blade ends to be calculated in the axisymmetric model. Additional losses are assigned in these regions based on the calculated deviations. The resulting method gives realistic radial distributions of loss and deviation across the whole span at both design and off-design operating conditions providing a quick method of estimating the magnitudes of these effects in the preliminary design process. Results from the method are compared to measured data in low and high speed compressors and multistage three-dimensional viscous CFD predictions.

This paper presents the results of three-dimensional unsteady Navier-Stokes simulations of the flow in a transonic compressor stage with inlet guide vanes and an axial flow air turbine stage with two identical stators for which detailed unsteady experimental data are available. Various unsteady flow phenomena are shown. The focus in the computations are stator/rotor interaction effects. Of special interest are their influences on the flow field downstream of the interface regions. The secondary flow effects are visualized via vector plots. The numerical results are compared with the experimental results. Although there is a good agreement in the major flow phenomena local deviations can be observed.Copyright © 1997 by ASME

Rotating stall phenomenon is usually characterized by 3D aerodynamic stability behavior. The earlier models mainly considered the flow effects in terms of 1D and 2D spatial variables. In order to involve the characteristics of the 3D flow of the compressor, it is necessary to improve the existing rotating stall stability models and further develop the models to consider the effects of the 3D disturbance. In this paper, a new aerodynamic stability model concerning the effects of a radial disturbance produced by the compressor, and explaining more mechanisms about the aerodynamic stability of compressor is presented. Using the developed rotating stall stability model, the stall margins are calculated and compared to experimental data for two axial flow compressors. The calculated results show that the developed 3D rotating stall stability model gives better stall margin prediction than that by the 2D model.Copyright © 1997 by ASME

The primary purpose of this study was to investigate improved numerical techniques for predicting flows through multistage compressors. The vehicle chosen for this study was the Pennsylvania State University Research Compressor (PSRC). The PSRC facility consists of a 3-1/2 stage axial flow compressor which shares design features which are consistent with embedded stages of modern gas turbine engine axial flow compressors. In Part 1 of this two part paper, several Computational Fluid Dynamics (CFD) techniques were applied to predict both steady and unsteady flows through the PSRC facility. Inter-blade row coupling via a circumferentially-averaged mixing plane approach was employed for steady flow analysis. A mesh density sensitivity study was performed to define the minimum mesh requirements necessary to achieve reasonable agreement with the experimental data. Time dependent flow predictions were performed using a time dependent interblade row coupling technique. These calculations evaluated the aerodynamic interactions occurring between the second rotor, second stator, and third rotor for the PSRC rig.Copyright © 1997 by ASME

Goto has presented data in the literature for the Deverson axial flow compressor rotor in a single stage environment. The flow features tip leakage and hub corner stall. In this paper, 3-d turbulent steady flow calculations for the rotor are used to determine if, and how, this rotor can be a useful CFD test case. The sensitivity of the CFD results to unknowns about the test case environment is computed. In particular the influences of blade twist, hub gap leakage flow between rotating and stationary components, and a departure from a radial equilibrium pressure distribution at the rotor exit are investigated. Evidence in support of these phenomena is presented and discussed, their likely magnitudes are suggested, and their influence on the rotor performance is quantified.Copyright © 1997 by ASME

Research and development of ceramic gas turbines (CGT), which is promoted by the Japanese Ministry of International Trade and Industry (MITI), was started in 1988. The target of the CGT project is development of a 300kW-class ceramic gas turbine with a 42 % thermal efficiency and a turbine inlet temperature (TIT) of 1350°C. Two types of CGT engines are developed in this project. One of the CGT engines, which is called CGT302, is a recuperated two-shaft gas turbine with a compressor, a gas-generator turbine, and a power turbine for cogeneration. In this paper, we describe the research and development of a compressor for the CGT302. Specification of this compressor is 0.89 kg/sec air flow rate and 8:1 pressure ratio. The intermediary target efficiency is 78% and the final target efficiency is 82%, which is the highest level in email centrifugal compressors like this one. We measured impeller inlet and exit flow distribution using three-hole yaw probes which were traversed from the shroud to the hub. Based on the measurement of the impeller exit flow, diffusers with a leading edge angle distribution adjusted to the inflow angle were designed and manufactured. Using this diffuser, we were able to attain a high efficiency (8:1 pressure ratio and 78% adiabatic efficiency).

The recent trend in using aerodynamic sweep to improve the performance of transonic blading has been one of the more significant technological evolutions for compression components in turbomachinery. This paper reports on the experimental and analytical assessment of the pay-off derived from both aft and forward sweep technology with respect to aerodynamic performance and stability. The single stage experimental investigation includes two aft-swept rotors with varying degree and type of aerodynamic sweep and one swept forward rotor. On a back-to-back test basis, the results are compared with an unswept rotor with excellent performance and adequate stall margin. Although designed to satisfy identical design speed requirements as the unswept rotor, the experimental results reveal significant variations in efficiency and stall margin with the swept rotors. At design speed, all the swept rotors demonstrated a peak stage efficiency level that was equal to that of the unswept rotor. However, the forward-swept rotor achieved the highest rotor-alone peak efficiency. At the same time, the forward-swept rotor demonstrated a significant improvement in stall margin relative to the already satisfactory level achieved by the unswept rotor. Increasing the level of aft sweep adversely affected the stall margin. A three-dimensional viscous flow analysis was used to assist in the interpretation of the data. The reduced shock/boundary layer interaction, resulting from reduced axial flow diffusion and less accumulation of centrifuged blade surface boundary layer at the up, was identified as the prime contributor to the enhanced performance with forward sweep. The impact of tip clearance on the performance and stability for one of the aft-swept rotors was also assessed.

In this paper, a 1D unsteady compressible viscous flow model of a generic compression system previously developed by the authors is applied to a multi-stage axial compressor experimental rig configured for single–stage operation. The required model parameters and maps are identified from experimental data. The resulting model is an explicit system of 9 first order ODE’s. The model inputs are compressor speed, nozzle area, compressor discharge bleed area, plenum bleed area, inlet total pressure and entropy, and nozzle and bleed exit static pressures. The model and experimental data are compared with respect to both open–loop uncontrolled and closed–loop controlled behaviors. These comparisons focus on i) forced transients and ii) global nonlinear dynamics and bifurcations. In all cases the comparison between the model and experimental data is excellent. Of particular interest is the ability of the model, which does not include any hysteretic maps, to predict experimentally observed hysteresis with respect to the onset and cessation of surge. This predictive capability of the model manifests itself as the coexistence of a stable equilibrium (rotating stall) and a stable periodic solution (surge) in the model at a single fixed set of system input values. Also of interest is the fact that the controllers used for closed–loop comparisons were designed directly from the model with no a posteriori tuning of controller parameters. Thus, the excellent closed–loop comparisons between the model and experimental data provide strong evidence in support of the validity of the model for use in direct model based controller design. The excellent agreement between the model and experimental data summarized above is attributed in large part to the use of effective lengths within the model, as functions of axial Mach number and nondimensional compressor rotational speed, as prescribed by the modeling technique. The use of these effective lengths proved to be far superior to the use of physical lengths. The use of these effective lengths also provided substantial improvement over the use of physical lengths coupled with fixed first order empirical lags, as proposed by other authors for the modeling of observed compressor dynamic lag. The overall success of this model is believed to represent a positive first step toward a complete experimental validation of the approach to control–oriented high–frequency turbomachinery modeling being developed by the authors.Copyright © 1993 by ASME

In this paper, a three-dimensional computational model for the solution of the time-averaged Navier-Stokes equations, based on a pressure correction method and the k-e turbulence model, is presented and implemented for the viscous flow modelling through a series of centrifugal compressors. Theoretical calculations with the current fully elliptic method are carried out and the results are compared critically with available experimental data and with results from other computational models. A radial and two backswept high-speed subsonic compressors with different geometrical and operating characteristics are analysed at design and off-design conditions. In all cases, a wake flow pattern is evident and strong secondary flows are discerned. The tip clearance effects on the relative flow pattern are found to be important and are appropriately simulated. The predictive capability of the current flow model is judged to be encouraging taking into consideration the limitations of the physical models and the numerical schemes involved in the computations.Copyright © 1993 by ASME

The objective of this paper is to apply a three-dimensional Navier-Stokes calculation to the second stage rotor and stator of a large scale compressor and to compare the computed results with measured data. The comparisons include: (1) airfoil and endwall surface streamlines, (2) radial-circumferential distributions of exit total pressure, (3) spanwise distributions of circumferentially averaged inlet and exit flow quantities, and (4) fullspan airfoil pressure distributions. This assessment demonstrates that the computational state-of-the-art has advanced to a point where calculations such as this can be applied with reasonable confidence in both design and analysis situations.

This paper presents the results of the experimental investigations on the effects of leaned and curved blade of plane axial compressor cascade. The ink trace method has been used for the surface flow visualization. The five-hole probe measurements are performed at the cascade exit 12.5 percent chord downstream, and the static pressure distributions are measured on the endwalls and blade surfaces.Copyright © 1993 by ASME

In this paper we present the results of a detailed experimental study of the development of small rotating stall, as it appears in a one stage axial compressor. Stationary hot-wire probes are used to measure the variation of amplitude and propagation speed of the disturbances caused by small stall. Measurements near the rotor blade surface with rotating probes provide additional information on the nature of the phenomenon. The development of the cell pattern for different operating conditions is studied. The different character from what is known as “big stall” is demonstrated.Copyright © 1985 by ASME