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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

Aerodynamic investigations in a cold flow annular sector high-pressure turbine cascade test facility and an annular sector cascade facility operating at near-engine conditions are presented. The test section of both facilities is a 36° sector cascade of a modern turbine stator consisting of 6 vanes. The two facilities have been designed in order to gain detailed information concerning film cooled gas turbine vanes. Due to the operation conditions of the hot annular sector cascade it takes over the part of detailed investigations of the influence of film cooling on the heat transfer. In the cold annular sector cascade facility investigations on the aerodynamic behavior of the cascade are performed. Both facilities together will lead to a better understanding of the complicate three-dimensional flow in modern gas turbines. A detailed description of both facilities is given in this paper. Aerodynamic investigations in both facilities were performed. The in- and outlet Mach number and profile Mach number distribution is in good agreement in both of them and shows a periodic flow filed. Aerodynamic performance measurements in the cold flow facility have been conducted by means of a five-hole pneumatic pressure probe traverses 106% of cax downstream of the cascade to gain information about the quality of the flow field across flow passages “+1” and “–1” in terms of yaw angle, pitch angle and primary loss distribution. Comparison with a three dimensional Navier Stokes solvers show a very good agreement with the measurements. In order to deduce the external heat transfer coefficient on the vane a transient test procedure was adopted in the high-pressure hot facility. The dependency of the heat transfer coefficients on the Reynolds number is presented in the paper. The experimental results show reasonable agreement with calculations using a two dimensional boundary layer code.Copyright © 2001 by ASME

The paper presents a compilation of experimental data on the effects of wake-induced transition on a highly loaded LP turbine cascade intended to be used for further numerical work. Although the underlying physics is not yet completely understood, the benefits of wake passing are already known and employed in the design process of modern gas turbines. For further optimizations, the next step seems to be now to enable numerical simulations detailed enough to capture the major effects while being as uncomplicated as possible at the same time to be cost-effective. The experimental results constituted in this systematic investigation are available for download and should serve as a basic data set for future calculations with different turbulence and transition models, thereby shedding some light on the complexity and modeling required for a suitable numerical treatment of the wake-induced transition process. The data introduced in this test case was acquired using a turbine cascade called T106D-EIZ with increased blade pitch compared to design point conditions in order to achieve a higher loading. A large separation bubble forms on the suction side and allows to study boundary layer development in great detail. The upstream blade row was simulated by a moving bar type wake generator. The measurements comprise hot wire data of the bar wake characteristics in the cascade inlet plane (velocity deficit and turbulence level), boundary layer surveys with surface-mounted hot films sensors and a hot wire probe at various locations and measurements of the total pressure loss coefficient. Unsteady pressure transducers are embedded into the suction side of a cascade blade and in a wake rake to resolve the local pressure distributions over time. They yield quantitative values easily comparable to the output of numerical simulations. The objective of this paper is to enable and to invite interested researchers to validate their code on the data set. From the extensive test program, a very limited number of operating points have been selected to focus the work. The standardized data files include a “reference” case with an exit Reynolds number of 200.000 and an exit Mach number of 0.4 as well as two points with higher Mach or lower Reynolds number for constant wake passing frequencies and background turbulence levels.

An analysis of laser Doppler velocimeter (LDV) data for the axial velocity flowfield in the rotor of the Mach 2 inlet flow supersonic throughflow fan (SSTF) is presented in this paper. The paper starts with a short description of the SSTF test package to highlight the specifics of the SSTF operation. It is followed by a detailed description of a dedicated LDV system for measurement in a supersonic throughflow fan and the experience gained. Most of the experimental data presented were acquired in a low supersonic throughflow regime (inlet Mach number of 1.4). The results and conclusions presented are based mainly on the experimental data only. A limited amount of computational fluid dynamics (CFD) predictions were used for comparison with the experimental results. The CFD methods, however, are not discussed in this paper. As shown in this paper, a reasonably good agreement between the LDV data and the CFD predictions was found for the low supersonic throughflow regime. The design point data (inlet Mach number of 2.0) exhibited an unexpectedly high noise in the velocity data in comparison with the data for low supersonic throughflow operating conditions. For the off-design supersonic regime (shock in rotor), substantial differences exist between the experimental and computational data.

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.

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.

The present study is part of an on-going project concerned with the effects of off-design incidence on losses. Previous work has examined the profile losses (Benner et al., 1995). This study focusses on the secondary flow, the endwall flow in the absence of tip clearance. This paper describes an experimental study in a large-scale, low-speed, linear turbine cascade for which the incidence was varied. Three different values of incidence were investigated: 0, +10 and +20 degrees. Detailed flow field measurements were made upstream and downstream of the cascade and static pressure distributions were measured on the blade surfaces. The data were supplemented by extensive surface oil flow visualization. This paper concentrates on the physical insights into the relationship between the behaviour of secondary flows and incidence which have been gained from the measurements and the flow visualization. A preliminary discussion of the effect of incidence on the losses is also included.Copyright © 1997 by ASME

This paper analyses the flow in a transonic turbine stage, using numerical results from a full 3D Navier-Stokes computation over the whole stage. This analysis, based on the combination of time and space Fourier transforms, points out how a spatio-temporal periodicity phenomenon is involved in the so-called stator-rotor interaction. A first comparison is made between the computational results using geometrical approximation and the results achieved by the complete simulation. This analysis points out how the signal is enforced by the numerical conditions at the circumferential boundaries. The stator-rotor interaction produces low spatial harmonics which are identified as spinning modes by means of Fourier series. The analytical theory from Tyler and Sofrin (1962) is then used to define a general model of stator-rotor interaction. This model allows the determination of the angular speed of rotation of the spinning modes, and results are compared with the global numerical simulation. Finally, a first analysis of the axial propagation of the spinning modes is proposed, and attention is focused on the fact, that only few parameters are necessary to reproduce the unsteady signal.

In a twin-spool jet engine the HP-compressor experiences a rotating inlet distortion if rotating stall in the upstream LP-compressor occurs. This destabilizing effect on the compressor operating behavior may lead to HP-compressor instabilities like rotating stall or surge. In order to investigate this specific type of unsteady inlet distortion a 5-stage HP-compressor is operated with a newly developed high rotation speed distortion generator in the inlet duct. It produces a rotating inlet distortion with up to 65% of the compressor design speed rotating clockwise or counterclockwise with the compressor rotor. Kulite pressure sensors as well as hot-wire probes are installed to resolve the unsteady character of the inlet distortion, to define the compressor map and to analyse the stall inception process. Describing the rotating inlet distortion and its influence on the compressor operating behavior will be the major objective of this paper. In the experiment the distortion speed of an inlet distortion rotating clockwise with the compressor rotor is varied at different compressor speeds and the effects on the surge margin are analysed.Copyright © 2001 by ASME

The paper presents the detail investigation of temperature field evolution through multistage cooled turbines. An investigation bases on simple enough numerical simulation and allows for transient, heat transfer, viscous and some other important effects on temperature field transformation. Herewith the special test data for a number of cooled turbines are used. The developed numerical code has the following peculiarities: - a time-marching method for the unsteady Euler equation system; - a special algorithm of flow parameters averaging in mixing planes in the middle of axial gaps; - a monotone implicit scheme of second or third order accuracy in space and time. The code has been used for a numerical study of the flow pattern in a number of multistage aviation and industrial turbines. The described simulation demonstrates satisfactory correlation between the numerical and experimental data for temperature gradient attenuation in the flowpath of investigated cooled turbines.

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 analysis and design methods used for turbomachinery components are mostly based on steady aerodynamics, neglecting the important unsteady nature of the flow field. An improvement in performance can however be achieved with a prior understanding, evaluation and modeling of the main unsteady loss sources generated in rotor/stator interactions, through new advanced experimental data coupled to systematic and controlled numerical simulations performed at the full unsteady level of approximation. But such calculations are even nowadays challenges to the CFD community, due to their high requirement in computer resources. To investigate the importance of unsteady loss mechanisms, a 1-1/2 axial turbine stage has been resolved at both quasi-steady and fully unsteady levels of approximation. In order to reduce the demand on computer resources, a scaling procedure can be applied to retrieve equal pitch distance on both sides of each rotor/stator interface. The space and time flow periodicity are then uncoupled and the unsteady flowfield may be resolved on a reduced number of blade passages per row without having to consider any time periodicity in the boundary treatment. The grid scaling however affects the turbine total pressure ratio and the position and strength of secondary flows, as the pitch-to-chord ratio is not kept constant. This effect is analyzed in the paper, with the objective to assess the associated approximation errors. Steady and unsteady numerical simulations are compared with the experimental data along three measuring stations placed downstream of each blade row. Even if steady results are in good agreement and allow capturing the main flow structures of the turbine stage, only the fully unsteady calculation resolves the complex loss mechanisms encountered mainly in the rotor and downstream stator components. These unsteady interactions are observed through time variations of the entropy, absolute flow angle and static pressure.

The paper describes the flow conditions found inside an optimized transonic, high pressure ratio centrifugal compressor rotor operated with a rotor tip speed of 586 m/s. The rotor was coupled with a vaned diffuser of rather small exit radius ratio. Measurements and calculations were carried out on this device. A newly developed 3-component hybrid Laser velocimeter was applied for the optical measurements inside the rotating system. For the first time the new measurement technique delivered three velocity components instead of only two for such a compact machine with very narrow flow channels. An available 3D calculation system with an embedded Navier-Stokes solver was used for the theoretical approach. The flow field development is described with the help of optical measurement results taken at measurement planes located at the rotor inlet and exit as well as at 40% and 60% flow path length. The measured 3D effects were highest in the middle of the flow channel, where the flow is turned from axial to radial. Here the curvature effects in the meridional plane and the enhancement of secondary flow are significant. The agreement between the theoretically and experimentally obtained third velocity component is quite good, especially in the area with strong secondary flows.

In this paper, a computational analysis of a high-speed centrifugal compressor stage for turbocharger applications is presented. Emphasis is focused on the effect of different number of diffuser vanes, and for this reason four different designs of the vaned diffuser are analysed. The first three of the diffusers consist of 11, 22 and 33 vanes, respectively, with their leading edge at a radius of 1.075 times the radius of the impeller tip. The fourth one consists of 22 vanes with its leading edge at 1.15 times the radius of the impeller tip. All the above vane designs are of double circular arc shape. A steady CFD analysis is carried out using the Reynolds-Averaged Navier-Stokes solver TASCflow at design and off-design operating conditions. An averaging approach is used at the interface between the impeller and the diffuser. A detailed comparison between the predicted and the available experimental data is performed in terms of pressure rise and efficiency characteristics, and very good agreement is accomplished. In addition, detailed flow distributions are compared and critically analysed. One of the most important conclusions is that while maintaining the overall throat area and the location of the leading and trailing edges of the diffuser, as the number of diffuser vanes increases, the pressure recovery coefficient at the semi-vaneless space at surge condition was found to reduce, the wake pattern becomes more pronounced and the velocity distribution at vaneless and semi-vaneless space becomes more distorted when passing the same mass flow rate and therefore the diffuser has a narrower flow range. On the other hand, it was found that the diffuser outlet to throat area ratio is not the dominant factor to influence the flow range when the number of vanes changes.

Poor pressure recovery and flow control in gas turbine exhaust ducts can significantly reduce overall engine performance. For aircraft installations, a configuration that presents a challenging design is the single port tractor exhaust that is used where the turbine is facing the direction of travel. This duct consists of an annular to rectangular transition and a turn of approximately 160°. The combination of a diffusing duct with strong curvature produces a flow that has a non-uniform velocity profile throughout the duct, large areas of separation and strong vortices. These flow conditions lead to insufficient diffusion and large pressure losses, making the design of a low loss duct difficult. The problem of design is further complicated by the need for the duct to operate over a range of flow Mach number and swirl angles. An experimental study has been undertaken to identify the main flow structures that lead to poor performance in tractor exhausts and how the key geometric and flow parameters affect the flow. A half scale model of the exhaust duct has been tested using cold flow over a range of swirl conditions. Flow measurements include static wall pressures, and inlet and outlet velocity profiles. This paper presents experimental results showing the effects of swirl on a gas turbine single port tractor exhaust duct.

The experimental program reported here was executed using full-scale vaneless counter-rotating engine hardware operating at nondimensionally scaled aerodynamic design point conditions. Measurements were obtained for three different pressure ratio values: design point, low pressure ratio, and high pressure ratio. For brevity, only the design point data will be presented in this paper. Time-averaged and time-resolved surface pressures on the high pressure turbine (HPT) vane, HPT blade, and low pressure turbine (LPT) blades are presented. Additionally, three-dimensional (3D) Navier-Stokes computational fluid dynamics (CFD) predictions are presented for comparison with experimental data. The results presented show that the predictions qualitatively capture the flowfield physics, but require some additional calibration to fully match experimental data quantitatively.Copyright © 2000 by ASME

The design of an aft-swept transonic compressor rotor and the investigation of its flow field compared to the radial baseline rotor are reported. Both rotors were designed using a 3D Euler solver with body forces. Through the choice of composite material using carbon fibres, high amounts of sweep and lean could be imparted to the blades without adding mass to the rotor assembly. No compromise between aerodynamic and structural constraints had to be made. The swept design as well as the radial base-line rotor were investigated using the commercial RANS solver CFX-TASCflow to validate the design. Complete speed lines at design speed were computed for both rotors. The analytical results showed similar peak efficiency and indicate that the swept rotor’s stall margin is reduced as compared to the base-line rotor. The resulting flow field of the aft-swept rotor is described in this paper and an attempt is made to discuss the fidelity of the design tools used.

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

This paper presents the results from three-dimensional (3D) steady and unsteady Navier-Stokes computations, performed on the transonic VKI BRITE EURAM test turbine stage. The work aimed at giving deeper insight in the aerodynamics of the turbine stage. The analysis has been carried out with the nominal stator trailing edge ejection slot geometry and cooling flow ejection. Additionally a simplified rounded stator trailing edge was employed. The results from the unsteady computations were compared with measured pressure perturbation traces at 22 locations around the rotor blade at midspan. Computations with both the ejection slot and the rounded stator trailing edge geometry were in good agreement with the measurements on the pressure side and half chord of the rotor blade’s suction side. Measurements and computations showed less good agreement downstream a weak shock on the suction side of the rotor blade. The measured pressure double peak in the rotor blade leading edge region is only observed in the computations with the ejection slot geometry.Copyright © 2000 by ASME

Inward flow radial and mixed flow turbines are effectively utilized in both turbochargers and small gas turbine auxiliary power units (APU’s), where moderately high levels of efficiency can be readily attained with simple cast components, less sensitive to blade endgap clearances than axial turbines. This paper first provides an overview of radial turbine performance characteristics for turbocharger and gas turbine applications as basically influenced by specific speed, reaction, velocity ratio, and rotor tip to exducer root mean square (RMS) diameter ratio. Since turbine rotor mass and inertia play important roles in structural integrity and engine acceleration characteristics, the effects of rotor blade solidity on turbine efficiency are then discussed based upon past research, and supported by analysis of a one dimensional friction model of the rotor passage. Finally the effects of rotor reaction on radial turbine flow versus pressure characteristics are examined pertinent to engine matching requirements.Copyright © 2000 by ASME

The paper presents detailed measurements of the midspan aerodynamic performance of a transonic turbine cascade at off-design conditions. The measurements were conducted for exit Mach numbers ranging from 0.5 to 1.2, and for Reynolds numbers from 4×105 to 106. The profile losses were measured for incidence values of +14.5 deg, +10 deg, +4.5 deg, 0 deg, and −10 deg relative to design. To aid in understanding the loss behavior and to provide other insights into the flow physics, measurements of blade loading, exit flow angles, trailing-edge base pressures, and the axial velocity density ratio (AVDR) were also made. It was found that the profile losses at transonic Mach numbers can be closely related to the base pressure behavior. The losses were also affected by the AVDR.

Unsteady stator-rotor interaction in gas turbines has been investigated experimentally and numerically for some years now. Most investigations determine the pressure fluctuations in the flow field as well as on the blades. So far, little attention has been paid to a detailed analysis of the blade pressure fluctuations. For further progress in turbine design, however, it is mandatory to better understand the underlying mechanisms. Therefore, computed space–time maps of static pressure are presented on both the stator vanes and the rotor blades for two test cases, viz the first and the last turbine stage of a modern heavy duty gas turbine. These pressure fluctuation charts are used to explain the interaction of potential interaction, wake-blade interaction, deterministic pressure fluctuations, and acoustic waveswith the instantaneous surface pressure on vanes and blades. Part I of this two-part paper refers to the same computations, focusing on the unsteady secondary now field in these stages. The investigations have been performed with the flow solver ITSM3D which allows for efficient simulations that simulate the real blade count ratio. Accounting for the true blade count ratio is essential to obtain the correct frequencies and amplitudes of the fluctuations.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.

It is obvious that the Reynolds normal stresses uu¯ should always be positive in all directions, i.e. the computed turbulence stresses should be realizable. However, the commonly used two-equation turbulence models do not incorporate realizability. They take the turbulent viscosity as cμk2/ε with cμ a constant, and frequently generate negative normal stresses far from walls in the nominally inviscid sections of turbomachinery flows. Pressure gradients due to leading edge stagnation and blade turning create an inviscid strain field. These strains cause the calculation of negative normal stresses over significant portions of the flow field. The result can be erroneous increases in turbulence kinetic energy upstream of the leading edge by a factor of ten or more. This erroneous turbulence is then convected around the blade and through the blade row, significantly affecting the computed boundary layer development and profile losses. Frequently the problem of overproduction is avoided by using artificially high values of the dissipation, ε, at the inlet. But this incorrect procedure is not needed when realizability is incorporated in the turbulence model. The paper reviews some methods and models which ensure realizability in two-equation turbulence models. The extent of the problem and its solution are illustrated with examples from compressor and turbine cascades.

This paper, Part II of the study on wake-passing effect upon the aerodynamic performance of the turbine cascade, demonstrates the detailed measurements of the time-varying flow field downstream of the turbine cascade as well as of the moving bars. The experiment employs a single hot-wire probe to measure pitchwise distributions of the ensemble-averaged velocity at the blade midspan. The resultant data consequently provide clear images of the incident bar wakes that are bowed and directed to the suction side of the blade wake. A custom-made total pressure probe, instrumented with a miniature fast-response pressure transducer, are also adopted to understand time-resolved feature of the wake-affected stagnation pressure fields downstream of the cascade. Furthermore, a decay process of the bar wake through the test cascade is examined in detail, which serves for the discussion related to wake recovery and its impact on the stage loss.Copyright © 1999 by ASME

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.

The first part of this paper analysed the time-averaged flow in a transonic single stage turbine. In this second part, we analyse the three-dimensional unsteady flow and temporal correlations in the rotor frame of reference. Two main conclusions are deduced from the comparison of each correlation. First, all correlations have the same order of magnitude. This is an important conclusion for models that handle the interaction between blade rows, because the temporal correlations are usually neglected. Second, these correlations occur in complementary localizations. In particular, the largest values of spatial deterministic correlations is located between the nozzle trailing edge and the rotor leading edge, as a consequence of nozzle wakes. Conversely, temporal deterministic correlations are located within the rotor blade row. We also establish that entropy accumulations in the stator wake near the end-wall regions generate high fluctuations of the rotor passage vortices. Such a phenomenon may have great repercussions on the radial loss distribution in the stage.

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.

Laminar separation bubbles are commonly observed on turbomachinery blades and therefore require effective methods for their prediction. The location and size of the bubbles is critically dependent on the laminar-to-turbulent transition process. Therefore, in this paper the transition models of Solomon et al., Abu-Ghannam & Shaw, Mayle, Calvert, and Choi & Kang are incorporated into an upwind-biased Navier-Stokes solver and the computed results are compared with the measurements of Elazar & Shreeve in a cascade with controlled-diffusion blades. It is found that none of the models predicts the measured bubbles very well, although most of them give reasonable results as long as transition is predicted to occur within the bubble.

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

This paper discusses measurements of the discharge coefficients of gas turbine nozzle guide vane film cooling holes under fully engine representative conditions. These unique experiments were carried out in a large scale annular blowdown cascade which models the three-dimensional external, flow patterns found in modern aero-engines, including all secondary flow phenomena. Furthermore, the coolant system design allows the coolant-to-mainstream density ratio and blowing parameter to be matched to engine values, although they can be independently varied. The results confirm that the discharge coefficients of film cooling holes are significantly altered by external crossflow. The discharge coefficient is usually reduced by external crossflow, but under certain external flow conditions it can be increased over the non-crossflow case. This previously unhightighted phenomenon has been termed ‘the crossover effect’, and, although an initially surprising result, is of importance to aero-engine designers as taking account of it should lead to improved predictions of coolant consumption. As a consequence, more uniform blade cooling should be achieved and, in turn, the attainment of greater component durability will be possible.

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

A theoretical and computational study into the aerodynamics of trailing-edge-cooled transonic turbine blades is described in this part of the paper. The theoretical study shows that, for unstaggered blades with coolant ejection, the base pressure and overall loss can be determined exactly by a simple control volume analysis. This theory suggests that a thick, cooled trailing edge with a wide slot can be more efficient than a thin, solid trailing edge. An existing time-marching finite volume method is adapted to calculate the transonic flow with trailing edge coolant ejection on a structured, quasi-orthogonal mesh. Good overall agreement between the present method, inviscid and viscous, and experimental evidence is obtained.Copyright © 1997 by ASME

This paper presents both experimental and numerical results demonstrating the dramatic effect of a particular geometric discontinuity on the flow within a highly-loaded single-stage transonic compressor. This compressor features cantilevered variable-stagger stator vanes (VSV) with buttons at the outer radius partially standing out in the annulus, by a value of about 0.3 mm which amounts to 0.5% of vane height. The first tests of this compressor revealed a very large corner flow separation at the stator outer radius. This feature could not be reproduced by steady 3D Navier-Stokes computations of the stator flow field with a smooth flowpath. However, using a simple numerical approach to account for the actual geometric discontinuity of the annulus, the computation then showed flow trends similar to that of the experiment. Also, a detailed analysis of the computed flow field indicated that the flow turning (both radial and tangential) due to the button blockage contributed toward strengthening the effect of secondary flows. Subsequently, after the button geometry was corrected to match the flowpath, further tests confirmed a much more satisfactory flow behaviour within the stator vanes leading to a one point improvement in compressor isentropic efficiency. This study leads to the conclusion that ensuring a smooth flowpath can be critically important for the high stage leadings characteristic of advanced military applications. Moreover, the predictive value of 3D Navier-Stokes computations for the simulation of technological effects, such as VSV buttons / flowpath mismatch, is clearly demonstrated.

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 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

This paper presents the results of an investigation to determine the effects of variational particle rebounding models on surface impacts and blade erosion patterns in gas turbines. The variance in the particle velocities after the surface impacts are modeled based on the experimental measurements using Laser Doppler Velocimetry (LDV) under varying flow conditions. The probabilistic particle trajectory computations simulate the experimental variance in the particle restitution characteristics using cumulative distribution functions and random sampling techniques. The results are presented for the particle dynamics through a gas turbine flow field and are compared to those obtained with deterministic rebound models based on experimental mean values.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.

The behaviour of the relative flow in centrifugal turbomachines is extremely complex due to the existence of various fluid dynamic phenomena and their interaction. At design and off-design operating conditions, the relative flow is subject to stationary unsteadiness which includes the flow separation and wakes associated with passage pressure gradients, secondary flows, and boundary layer stability. It is also subject to periodic unsteadiness from the rotating stall and the cyclic flow phenomena induced by the casing. This paper describes the mechanical and optical design of a rotating laser-Doppler anemometry system which allows direct measurement of the relative flow by means of an optical de-rotator. By isolating the impeller rotational frequency from the sampling frequency, it allows direct time-average measurements of the stationary behaviour of the relative flow along with the ensemble (phase)-average measurements of its periodic behaviour. Its success is demonstrated with measurements conducted in a low specific speed centrifugal impeller fitted with a single volute. Sample results of the time-averaged blade-to-blade variation of total relative velocities along with their associated turbulence intensities are reported. The (periodic) cyclic variations of the impeller exit flow, induced by the volute at low flow rates, are also presented for the suction and pressure sides.

For structured grid codes, the H type grid has become a standard, with deviations being used for exceptional circumstances. Among these circumstances are boundary layer or wake capturing, leading and trailing edge resolution and high grid shear. The latter of these is especially troublesome as some codes may converge with highly sheared grids to solutions which are inaccurate. In the case explored here, the trailing edge shock system of a high pressure nozzle guide vane is smeared over many grid cells using a conventional H grid, although better resolution can be obtained by careful attention to the grid. The paper demonstrates this on a very simple case and proposes the use of a grid that is nearly orthogonal to the flow to overcome the problem. The grid generation and treatment of periodic boundary conditions including a novel flux balance boundary condition across the mis-matched grids are discussed. Both two-dimensional and three-dimensional results are presented against test data to show the efficacy of this technique.Copyright © 1993 by ASME

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

This paper describes how allowance for the thermal effects of non-adiabatic flow, altered boundary layer development, changes in tip clearances and changes in seal clearances have been incorporated into a general gas turbine transient program. This program has been applied to a two-spool bypass engine. Revised predictions of surge margins in three common transients have been obtained. When the engine undergoes a “cold” acceleration, the thermal effects on the trajectory and on the surge line give a much increased proportion of unused surge margin in the H.P. Compressor, as compared to adiabatic predictions. In a “hot” acceleration this improvement is considerably reduced.Copyright © 1985 by ASME

The use of non-conventional diffusing systems for centrifugal compressors has been mandated by the smaller diameter of many advanced turbofan engines. Optimizing the design of these geometrically complex diffusers by rotating component tests is both expensive and time-consuming. Methods of designing and static blow testing low-diameter, controlled-contour passage diffusers have been previously reported in which scaled-up models of six diffuser passages were constructed and tested at a Reynolds number comparable to that encountered in rotating component tests. One design was fabricated in full size and run in both a static blow test and a rotating component test. This paper reports on the performance of this design at several Reynolds numbers up to those encountered in engine application for several diffuser throat blockages. It is shown that as with geometrically simpler passages such as conical diffusers, static blow tests can also be used to optimize the design of geometrically more complex diffusing passages.Copyright © 1981 by ASME

It is important that turbocharger compressors have a broad operating range between surge and choke. The surge limit is controlled by flow stability in both the compressor impeller and vaneless diffuser. Theoretical studies, and to a limited extent, experimental studies, have shown that gradients of total pressure and angular momentum between the hub and shroud surfaces at diffuser inlet, as produced by the impeller, cause separation of the flow in the vaneless diffuser. This paper presents the overall performance of a turbocharger compressor employing three mixed flow impellers with the discharge blade tip designed to reduce the enthalpy gradients at diffuser inlet. It is shown that overall stability is improved by using a mixed flow impeller with a horizontal tip cut off.Copyright © 1981 by ASME

Blade tip losses represent a major efficiency penalty in a turbine rotor. These losses are presently controlled by maintaining close tolerances on tip clearances. This two-part paper outlines a new methodology for predicting and minimizing tip flows, and focuses on the control of tip leakage through minimization of the discharge coefficient to control the normal leakage flow component. Minimization of the discharge coefficient was achieved through viscous analysis and was supported by discharge-rig testing. The analysis for the discharge cross-flow used a stream function-vorticity formulation. Support testing was conducted with a newly developed water table discharge rig in which tip-coolant discharge could also be simulated. Experimental and numerical tip-leakage results are presented on a discharge coefficient parameter for five different tip configurations. In addition, numerical studies were conducted for stationary and rotating blades with and without tip coolant injection.Copyright © 1981 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