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IN THIS VOLUME


Turbomachinery

1996;():V001T01A001. doi:10.1115/96-GT-028.
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Mass transfer analogies have long been used to experimentally determine the distribution of cooling fluid through pointwise sampling for turbine blade, nozzle or combustor film cooling. The behavior of a turbulent jet or plume flowing into its surroundings cannot be fully understood from point measurements alone, however. Full-field measurement of the instantaneous distribution of cooling fluid can reveal the structure and mechanisms governing cooling performance. This paper describes an improved dual light sheet PLIF (Planar Laser Induced Fluorescence) technique developed for full field concentration measurements. An analytical model of laser light sheet / fluorescent dye interaction was formulated and used to evaluate the light sheet attenuation corrections. With the more common single light sheet technique, these corrections lead to substantial concentration uncertainty which can be substantially reduced by using a dual light sheet. The dual light sheet technique was used to study the time-varying position and area of concentration isopleths for a round jet issuing into quiescent surroundings. Results show that, although concentrations at any point vary widely with time, the area within a given concentration isopleth remains virtually constant.

Commentary by Dr. Valentin Fuster
1996;():V001T01A002. doi:10.1115/96-GT-029.
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Over the years it has been speculated that the performance of multi-stage axial flow compressors is enhanced by the passage of a wake through a blade row prior to being mixed-out by viscous diffusion. The link between wake mixing and performance depends on the ability to recover the total pressure deficit of a wake by a reversible flow process. This paper shows that such a process exists, it is unsteady, and is associated with the kinematics of the wake vorticity field. The analysis shows that the benefits of wake total pressure recovery can be estimated from linear theory and quantified in terms of a volume integral involving the deterministic stress and the mean strain rate. In the limit of large reduced frequency the recovery process is shown to be a direct function of blade circulation. Results are presented which show that the recovery process can reduce the wake mixing loss by as much as seventy percent. Under certain circumstances this can lead to nearly a point improvement in stage efficiency, a nontrivial amount.

Commentary by Dr. Valentin Fuster
1996;():V001T01A003. doi:10.1115/96-GT-030.
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We present two recent CFD techniques developed for numerical flux computation and grid generation. The numerical flux scheme, AUSM+, yields significant improvements over current popular schemes in terms of accuracy, simplicity of implementation, and generality to include other equations of conservation laws. The scheme has been shown to be free from inaccuracies/anomalies such as the “carbuncle phenomenon” associated with some celebrated upwind schemes. A new hybrid grid technique, termed DRAGON grid, has been developed to maximize the strengths of the Chimera grid and unstructured grid. Validation tests of these techniques and applications to rotating flows were conducted and their typical results are included in the paper.

Commentary by Dr. Valentin Fuster
1996;():V001T01A004. doi:10.1115/96-GT-031.
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The differences of two distinct numerical schemes implemented in one code called ITSM3D are presented for a turbine stage test case. Thus both schemes are used with exactly the same computational infrastructure, e. g, same grids, boundary conditions, acceleration strategies, time-stepping, turbulence model etc. The two methods are based on an explicit Runge-Kutta-type finite volume scheme expressed in cylindrical coordinates and have been developed at the Institut für Thermische Strömungsmaschinen und Maschinenlaboratorium of the University of Stuttgart. One scheme is a node centered 3rd order TVD scheme according to Osher and the other belongs to the cell vertex central difference type with the concept of artificial viscosity. The model of Baldwin-Lomax is used in order to simulate turbulent effects. Non-reflective boundary conditions are taken at stator inlet and rotor outlet to avoid non-physical reflections. A multigrid technique in combination with implicit residual smoothing and local time-stepping is employed to accelerate the computation.

The test case for this comparison is the last stage of a low-pressure turbine. The computational results obtained are discussed and compared to each other as well as to experimental data. They are presented as pressure and Mach number isoline contours and diagrams of circumferential averaged quantities at inlet and outlet planes of stator and rotor versus radial position.

Commentary by Dr. Valentin Fuster
1996;():V001T01A005. doi:10.1115/96-GT-037.
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A three-dimensional space marching code is used for the numerical modelling of the flow in an isolated axial flow compressor rotor. The rotor is analyzed at four operating points, up to near stall conditions. Numerical results are first validated versus available experimental data and then further exploited in order to illuminate flow patterns in the inter-blade region. The tip leakage impact on the main passage flow and losses level as well as the effect of blade loading on the hub corner stall extent and the radial displacement of the flow are fully detailed. In order to account for the rotor geometry, the modifications performed in an existing software are mainly concerned with the accurate modelling of the clearance which is formed above the curved blade tip; for this purpose, a local H-type mesh is embedded to the main passage grid.

Commentary by Dr. Valentin Fuster
1996;():V001T01A006. doi:10.1115/96-GT-038.
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The complex three-dimensional flow in the wicket gate and the runner of a Francis turbine is investigated by applying both a quasi-three-dimensional and a three-dimensional computational method. The computations were conducted on a double grid containing the stationary wicket gate and the rotating runner. The equations for inviscid and incompressible flow are solved, assuming that the relative flow field in the runner is stationary. In the quasi-three-dimensional method the governing equations are solved on stream surfaces using a Finite-Element-Method. In the three-dimensional method, the equations of continuity and motion are solved by a Finite-Volume technique using Denton’s code for incompressible flow. Both methods are used in order to compute the flow in a Francis-runner of high specific speed at the operating point of optimum efficiency. The results of the calculations are compared with measurements taken at the draft-tube inlet. Differences between results of computations and measurements are presented.

Commentary by Dr. Valentin Fuster
1996;():V001T01A007. doi:10.1115/96-GT-039.
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An iterative procedure for 3D blade design is presented. The three-dimensional blade shape is modified using a physical algorithm, based on the transpiration model. The transpiration flux is computed by means of a modified Euler solver, in which the target pressure distribution is imposed along the blade surfaces. Only a small number of modifications is needed to obtain the final geometry.

The method is based on a high resolution three-dimensional Euler solver. An upwind biased evaluation of the advective fluxes allows for a very low numerical entropy generation, and sharp shock capturing.

The method is first validated, by redesigning an existing geometry, starting from a different one. It is further used to redesign a transonic compressor blade, to achieve, for the same mass flow and outlet flow angle, a shock free deceleration along the suction side. The last example concerns the design of a low aspect ratio turbine blade, with a positive compound lean to reduce the intensity of the passage vortices. The final blade is designed for an optimized pressure distribution, taking into account the forces resulting from the blade lean angle.

Commentary by Dr. Valentin Fuster
1996;():V001T01A008. doi:10.1115/96-GT-040.
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A quasi three-dimensional method has been developed to calculate the transonic flow in a compressor rotor. The method allows changes in the stream tube thickness and mean radius evolution through the rotor. This blade-to-blade method uses original concepts in order to be rapid (≈ 3–4 sec CPU on a HP 715 workstation), to find the precise location of the shocks (and thus the flow values on them) and the flow distribution around the profile, and finally to cover a large field of “Off-Design” operating points. Thus, it can be used in the process of optimization of a transonic compressor. The supersonic flow at the inlet as well as the oblique shock configuration inside the cascade are calculated by methods based on the characteristics theory. Conditions with attached shock wave (and thus unique incidence angle) or with detached shock wave can be calculated. Quasi-3D equations were developed (Bölcs and Tsamourtzis, 1991). The subsonic flow is calculated by a streamline curvature method, with some new concepts, and finally the shock in the inter-blade canal is found by a combination of the supersonic and subsonic flow values. This method was combined, in a S1 - S2 calculation, with a throughflow method (Sayari and Bölcs, 1995) in order to be validated by comparing the results with the measurements provided by NASA - Lewis Research Center, on a transonic compressor rotor (Strazisar T.; 1994).

Topics: Compressors , Rotors
Commentary by Dr. Valentin Fuster
1996;():V001T01A009. doi:10.1115/96-GT-041.
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A wave rotor topped gas turbine engine has been identified which incorporates five basic requirements of a successful demonstrator engine. Predicted performance maps of the wave rotor cycle have been used along with maps of existing gas turbine hardware in a design point study. The effects of wave rotor topping on the engine cycle and the subsequent need to rematch compressor and turbine sections in the topped engine are addressed. Comparison of performance of the resulting engine is made on the basis of wave rotor topped engine versus an appropriate baseline engine using common shaft compressor hardware. The topped engine design clearly demonstrates an improvement in shaft horsepower and SFC. Predicted off design part power engine performance for the wave rotor topped engine is presented including that at engine idle conditions. Operation of the engine at off design is closely examined with wave rotor operation at less than design burner outlet temperatures and rotor speeds. Challenges remaining in the development of a demonstrator engine are addressed.

Topics: Waves , Gas turbines , Rotors
Commentary by Dr. Valentin Fuster
1996;():V001T01A010. doi:10.1115/96-GT-042.
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Extensive measurements of velocity and turbulence have been performed by means of a two-component fibre-optic laser Doppler velocimeter, to investigate the profile boundary layer development on a large scale turbine cascade.

Flow field investigation has been integrated with data obtained by surface-mounted hot-film gauges in order to get direct information on the boundary layer nature and on its time varying characteristics.

Measurements were detailed enough to allow constructing mean velocity and Reynolds stress boundary layer profiles giving an in-depth description of the boundary layer development along both suction and pressure surfaces through laminar, transitional and turbulent regimes.

Commentary by Dr. Valentin Fuster
1996;():V001T01A011. doi:10.1115/96-GT-055.
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This paper presents a procedure which cao generate highly stretched, unstructured, viscous meshes for essentially arbitrary three dimensional configurations. The procedure is based on a combination of the ideas behind Delaunay and moving front methodologies. Examples are presented for a variety of turbomachinery types of flow to demonstrate the potency of the new approach. Recent enhancements to the author’s flow solver are also briefly described.

Commentary by Dr. Valentin Fuster
1996;():V001T01A012. doi:10.1115/96-GT-056.
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Engine controller data have been interrogated for indications of incipient surge for three turbofan engines; a large Pratt and Whitney, a large General Electric, and a small Williams International. Versions of these engines are currently operating in the field and all have compression ratios of 18 or greater. The Pratt and Whitney engine was surged only at full power while the other two were surged at partial power and at full power. The interest in this work was in detecting the presence of warning signatures for a current inventory of engines. A constraint was imposed on the experiments to use only existing engine instrumentation. The frequency response of the controller and the engine instrumentation limited the high frequency detection capability to about 100 Hz for the large engines and about 200 Hz for the small engine, For the large engines, it was not possible to detect a surge warning but for the small engine a sufficient warning of incipient surge was detected.

Commentary by Dr. Valentin Fuster
1996;():V001T01A013. doi:10.1115/96-GT-057.
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In this paper, an application of Wavelet Transform, which is a newly developed time-frequency technique of signal processing, is demonstrated in analyzing compressor rotating stall signals. In contrast to conventional signal processing methods, e.g. Fourier Transform, Wavelet Transform is very suitable for analyzing transient processes as rotating stall inception in compressors. In this study, some typical rotating stall signals are processed via Morlet’s wavelet. It is concluded that Wavelet Transform has a great advantage in detecting rotating stall inceptions, which are usually very weak and embedded in relatively stronger noises. In the diagrams resulted from the transform, every emergence of precursor as well as full stall signals of a certain frequency is illustrated versus time.

Commentary by Dr. Valentin Fuster
1996;():V001T01A014. doi:10.1115/96-GT-058.
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A new methodology for interactive design of turbomachinery blades is presented. Software implementation of the methods provides a user interface that is intuitive to aero-designers while operating with standardized geometric forms. The primary contribution is that blade sections may be defined with respect to general surfaces of revolution which may be defined to represent the path of fluid flow through the turbomachine. The completed blade design is represented as a non-uniform rational B-spline (NURBS) surface and is written to a standard IGES file which is portable to most design, analysis, and manufacturing applications.

Commentary by Dr. Valentin Fuster
1996;():V001T01A015. doi:10.1115/96-GT-059.
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A highly loaded two stage transonic axial flow compressor, which forms a front stages of a multi stage compressor for industrial gas turbines, has been designed and tested. Overall pressure ratio is 2.25 and the first stage rotor tip Mach number is 1.15. Two airfoil types, Double Circular Arc airfoil and Multi Circular Arc airfoil, were designed for a transonic rotor blade under the same condition. MCA blade design method was devised and introduced. The blade design relied heavily on CFD techniques using a Euler code and a Navier Stokes code to cope with a precise treatment. The rig test was conducted by our compressor test facility to verify a validity of the transonic compressor design method and to compare the performance of the DCA and the MCA airfoils. This report describes the aerodynamic design and the test results as well as the test facility and instrumentation.

Topics: Design , Axial flow
Commentary by Dr. Valentin Fuster
1996;():V001T01A016. doi:10.1115/96-GT-060.
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An experimental investigation has been carried out to determine the effects of inlet swirl on the flow field that develops within an annular S-shaped duct. The duct is representative of that used to connect the compressor spools on multi-spool gas turbine engines. By removing the outlet guide vanes from an upstream single stage compressor swirl angles in excess of 30° were generated. Results show that within the S-shaped duct tangential momentum (Wr) is conserved, leading to increasing swirl velocities through the duct as the radius decreases. Furthermore, this component influences the streamwise velocity as pressure gradients are established to ensure the mean flow follows the duct curvature. Consequently in the critical region adjacent to the inner casing, where separation is likely to occur, higher streamwise velocities are observed. Within the duct substantial changes also occur to the turbulence field which results in an increased stagnation pressure loss between duct inlet and exit. Data is also presented showing the increasing swirl angles through the duct which has consequences both for the design of the downstream compressor spool and of any radial struts which may be located within the duct.

Commentary by Dr. Valentin Fuster
1996;():V001T01A017. doi:10.1115/96-GT-061.
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Experiments were conducted on a diagonal-flow machine to study the behaviour of flow. Measurements showed that at reduced flow rates, reversal of flow occurs near the tip upstream of the rotor and near the hub downstream. At high flow rates, the flow reverses near tip at downstream only. In fact, there is only a limited regime of operation where the flow is not reversed before or after the impeller. The best fluid-dynamic efficiency was observed to be midway of this non-reversed flow regime.

Through-flow solutions of the mean hub-to-tip streamsurface were carried out by streamline curvature computation and compared with experimental results. The comparison showed good agreement of the predicted values with the experimental data. However, attempts to compare theoretical estimates of rotor losses with experimental measurements showed that the existing loss models are inadequate for loss prediction and further work is required in this direction.

The head-flow characteristic of the machine showed a droop at reduced flow rates, typical of what one usually notices in an axial-flow machine with the onset of blade stall. Study of the time history of velocity downstream of rotor illustrated that unlike rotating ‘stall-cells’ in axial-flow machines, the blade stall in the present case did not possess any regular pattern nor any unique speed of propagation. Near the hub at downstream of rotor, where the flow finally reverses upon reduction of flow rate, the stall appeared as patches of ‘blockage’ type disturbance over an otherwise systematic train of blade wakes when the flow coefficient reaches a value where the droop in the characteristic curve starts.

Commentary by Dr. Valentin Fuster
1996;():V001T01A018. doi:10.1115/96-GT-062.
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A throughflow method for designing and analysing compressors has to be supplied with loss, deviation, and blockage estimates for every blade row. The earliest methods used empirical correlations for profile loss and deviation, together with an empirical blockage or “work done” factor, and empirical estimates of additional losses near the endwalls. Previous papers by the author have described how to replace the empirical blockage factor and endwall corrections by explicit calculations using a new mathematical model of the endwall phenomena. Those papers illustrated the application of the method near design conditions, using either design profile loss and deviation figures or computations by a viscous-inviscid interaction blade-to-blade method.

In order to estimate off-design performance rapidly over the whole operating range, some way of estimating off-design profile loss and deviation must be chosen. In this paper, the previously-derived design point loss and deviation figures are retained, and an empirical correlation due to Miller, Wasdell, and Wright is used to predict the changes in loss and deviation off-design. It is shown by means of sample three-dimensional Navier-Stokes computations that the endwall model remains applicable off-design.

The method has been tested against two low speed and two high speed compressors, one of each example having controlled-diffusion blading. The low speed compressor characteristic maps are predicted only approximately, but the predicted high speed compressor maps are good. It is widely believed that endwall flow separation can initiate stall or surge. As stall or surge was approached the shape factor of the annulus wall boundary layer at one location rose sharply, but no single stall-predicting value could be found.

Topics: Compressors , Design
Commentary by Dr. Valentin Fuster
1996;():V001T01A019. doi:10.1115/96-GT-063.
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In many turbomachinery applications a compressor is directly driven by a turbine; for turbocharger applications a centrifugal compressor is usually adopted which is generally driven by a radial flow turbine, although mixed flow or axial flow turbines are occasionally required. A non-dimensional design procedure is developed to provide the basic dimensions and blade angles of centrifugal compressor impellers, whilst accounting for the turbine conditions as assessed through the matching requirements. The design of the turbine is then considered further in Part B.

The procedure can be applied for any desired compressor pressure ratio and target efficiency to develop an initial non-dimensional skeleton design. No other parameters are required from the initial specification and the design is developed non-dimensionally without recourse to empirical loss models and the associated uncertainties as the target efficiency must be specified. The procedure provides graphical information with respect to the impeller discharge conditions and inlet conditions from which the designer must select the most appropriate design. The screen graphics interface enables the designer to search across the design options; as this search is carried out numerical data are displayed and continuously up-dated to provide immediate information on which an infnrmed assessment can be based.

In addition to the compressor design options which are provided the matching conditions for the drive turbine provides information, such as specific speed, non-dimensional mass flow rate and pressure ratio, relevant to the turbine design. Judgements with respect to the design options for the compressor can then be made with the consequences for the associated turbine design clearly in view.

The non-dimensional design can be translated into an absolute design through the specification of the required mass flow rate and the inlet stagnation pressure and temperature.

Commentary by Dr. Valentin Fuster
1996;():V001T01A020. doi:10.1115/96-GT-064.
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A procedure similar to that presented in Part A is described for the non-dimensional design of radial inflow turbines. The technique is developed non-dimensionally to provide the overall dimensions, including the leading and trailing edge blade angles, of radial flow turbine rotors. Consideration of the adoption of non-radial blades at rotor inlet, leading to mixed flow rotors, and discharge swirl is included.

The procedure is developed from a knowledge of the non-dimensional power requirement which is derived through the turbocharger matching conditions. To satisfy the matching requirement the mechanical efficiency of the turbocharger, the air fuel ratio, the inlet stagnation temperature relative to that at compressor inlet, and the gas constants for the hot exhaust gas must be specified. In addition a target stage efficiency, on a total to static basis, must be specified.

In essence the procedure provides graphically all the possible velocity vectors at rotor inlet and exhaust. The graphical presentation allows the designer to survey all the options available; as this is done relevant data is presented and continuously up-dated to aid in the selection of the preferred design conditions.

The non-dimensional geometry of the rotor is then developed from which the absolute dimensions can be derived through the specification of the gas mass flow rate and the inlet stagnation conditions.

Commentary by Dr. Valentin Fuster
1996;():V001T01A021. doi:10.1115/96-GT-065.
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This paper surveys the development of the primary and secondary flows in the rotors of radial-inflow turbines. Information previously scattered throughout the literature has been brought together, and it has been possible to create a coherent picture and a good understanding of the complex flow processes which occur. The secondary flow is generated by cross-passage forces due to the turning of the blades, and Coriolis forces. Near the leading edge these give rise to a strong vortex adjacent to the pressure surface, moving low momentum fluid from hub to tip. This feature helps to explain why best efficiency occurs typically at 20°–30° negative incidence. Attempts to correlate the optimum incidence angle using traditional slip factor expressions can give quite misleading results, but a new approach based on the blade loading shows considerable promise. Nearer the exit there is motion of fluid from hub to tip near the suction surface and a vortex in the suction surface-shroud corner, and this is linked to the highly non-uniform flow at exit. The latter effect makes the prediction and correlation of rotor deviation information very difficult, despite the development of a rational exit averaging procedure. The present deviation data are sparse and not easy to correlate.

Commentary by Dr. Valentin Fuster
1996;():V001T01A022. doi:10.1115/96-GT-066.
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The turbine volute is a complex flow device, about which a few papers on both measurements and CFD predictions have appeared. The main reasons for the difficulties being the complicated geometry which hinders measurements to be taken by both intrusive and non-intrusive techniques, and makes the numerical predictions difficult.

In this paper, the complex three-dimensional flow through a turbine volute with non-symmetric circular cross-section is studied by using a 3-D Navier-Stokes solver which has been developed by the authors. In this solver, the fully 3-D Reynolds averaged N-S equations coupled with high Reynolds number k-ε turbulence model together with the wall function under arbitrary curvilinear coordinate system are solved. The Semi-Implicit Method for Pressure-Linked Equations (SIMPLEC algorithm) with the non-staggered grid arrangement is used. In order to eliminate the decoupling between the velocity and pressure under non-staggered grid system, the physical covariant velocity component is selected as dependent variable in momentum equations and a momentum interpolation approach is employed.

The validity of the free-vortex assumption is reviewed. The computation results are compared with a set of experiments performed previously by one of the authors. The flow features in the volute are discussed.

Topics: Turbines
Commentary by Dr. Valentin Fuster
1996;():V001T01A023. doi:10.1115/96-GT-067.
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A numerical study of the flow and heat transfer in secondary flow elements of the entire inner portion of the turbine section of the Allison T-56/501D engine is presented. The flow simulation included the interstage cavities, rim seals and associated main path flows, while the energy equation also included the solid parts of the turbine disc, rotor supports, and stator supports. Solutions of the energy equations in these problems usually face the difficulty in specifications of wall thermal boundary conditions. By solving the entire turbine section this difficulty is thus removed, and realistic thermal conditions are realized on all internal walls. The simulation was performed using SCISEAL, an advanced 2D/3D CFD code for predictions of fluid flows and forces in turbomachinery seals and secondary flow elements. The mass flow rates and gas temperatures at various seal locations were compared with the design data from Allison. Computed gas flow rates and temperatures in the rim and labyrinth seal show a fair 10 good comparison with the design calculations. The conjugate heat transfer analysis indicates temperature gradients in the stationary intercavity walls, as well as the rotating turbine discs. The thermal strains in the stationary wall may lead to altered interstage labyrinth seal clearances and affect the disc cavity flows. The temperature, fields in the turbine discs also may lead to distortions that can alter the rim seal clearances. Such details of the flow and temperature fields are important in designs of the turbine sections to account for possible thermal distortions and their effects on the performance. The simulation shows that the present day CFD codes can provide the means to understand the complex flow field and thereby aid the design process.

Commentary by Dr. Valentin Fuster
1996;():V001T01A024. doi:10.1115/96-GT-068.
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A simple time-accurate algorithm is presented for the computation of the unsteady stator-rotor interaction. The algorithm is based on the scalar approximate factorization method originally developed for the computation of complex three-dimensional steady flows. The method introduces a physical time step, used to march in time, and a numerical time step to iterate in between physical time steps. The method is formulated so as to take full advantage of the implicit formulation and provide an implicit treatment of the unsteady terms. A set of preliminary tests on a turbine stage, still in the experimental testing phase, proved the speed and accuracy of the method which was able to capture the essential features of a transonic stage.

Topics: Algorithms , Rotors , Stators
Commentary by Dr. Valentin Fuster
1996;():V001T01A025. doi:10.1115/96-GT-069.
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This paper presents a computational method for the calculation of unsteady three-dimensional viscous flow in turbo-machinery stages. The method is based on a Finite-Volume Navier-Stokes solver for structured grids in a multiblock topology. The meshes at the stator/rotor interface are overlapped by two grid cells. An implicit residual smoothing method applicable to global time-stepping is used to accelerate the solution process.

The problem of periodic boundary treatment for unequal pitches is handled using a method of time-inclined computational domains for three dimensions. The method applies a time transformation to the stator domain and to the rotor domain and uses different time-steps in the two domains.

The results of a numerical simulation of the flow in a transonic turbine stage with a pitch ratio of 1.364 are presented. The time-averaged solution is compared to experimental data and satisfactory agreement is stated. Complex 3D-unsteady flow phenomena (shock motion, vortex shedding) are observed. Unsteady blade pressure fluctuations at various positions in spanwise direction are shown and the fluctuations are found to vary considerably along span. Instantaneous distributions of static pressure, Mach number, and entropy are presented.

Commentary by Dr. Valentin Fuster
1996;():V001T01A026. doi:10.1115/96-GT-070.
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A steady/unsteady, three-dimensional Navier-Stokes solver that utilizes a semi-implicit, pressure-based solution procedure is developed to simulate the three-dimensional, incompressible flow through a single stage compressor. The present numerical scheme features the implementation of a second-order plus fourth-order artificial dissipation formulation to prevent the numerical oscillation due to central differencing schemes. A low-Reynolds-number form of the two-equation turbulence model is used to account for the turbulence effects. For unsteady flow computations, the coupling between the mean flow properties and the turbulence is enhanced by an inneriteration procedure during each time step. The steady flow field in the rotor passage is computed first. This is used as input for the computation of the unsteady flow in the subsequent stator. The predicted unsteady pressure on the stator blades and unsteady Velocities at several locations inside the passage are compared with the experimental data. The unsteady pressures on the stator blade surfaces are in good agreement with the experimental data. The predicted unsteady velocity components at various locations inside the stator blade rows are generally smaller than the measured values in the endwall regions. The phase angle variations of the unsteady velocity are in good agreement with the measured values. The effects of the rotor wake, secondary and tip clearance flows on the unsteady flow through the subsequent stator are studied. An attempt is also made to quantify the contributions of incoming tip leakage flows and the endwall boundary layers on the unsteady flow through the downstream stator. It was found that the endwall boundary layers and tip leakage flows have a much stronger influence on the unsteady flow development than the wake.

Commentary by Dr. Valentin Fuster
1996;():V001T01A027. doi:10.1115/96-GT-071.
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The three dimensional flow around an extensively investigated slot film cooled turbine blade is numerically investigated using a multi block finite volume Navier-Stokes solver.

Three blowing rates are simulated including the whole geometry of the interior blade cooling system and slots. Due to the ejection at the blade leading edge and the geometry of the cooling slots a very complex turbulent three dimensional flow field is generated.

The size and shape of the flow separation zones depending on the film cooling ejection is systematically investigated using several two-equation models, e.g. the standard and low Reynolds k–ε-Model of Lam and Bremhorst (1981) r[4], the extension of Kato/Launder (1993) [3] and the k–ω-Model of Wilcox (1991) [10], whereas the results of the standard k–ε-Model are presented. Experimental data obtained by Laser velocimetry, oil-flow pictures and pressure probes are used to understand the complex flow field and to validate the Navier-Stokes solver.

The multi-block code applies a traditional Jameson type solver and an implicit solver using several spatial discretization schemes for the convective fluxes. The two-equation models are solved using an RED-BLACK implicit technique with first order spatial upwind discretization to guarantee stability.

Commentary by Dr. Valentin Fuster
1996;():V001T01A028. doi:10.1115/96-GT-072.
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The possibility of predicting the total pressure loss radial distribution, due to the tip clearance presence, is examined in this paper. Models advanced for the diffusion of a line vortex are used for the simulation of the leakage vortex induced velocity and pressure fields, with sufficient success. The leakage vortex strength seems to control directly only a small part of the total pressure loss distribution, the one connected with the pressure deficit and the rotating flow. The remaining profiles result as functions of a free parameter — the constant of integration — and an assumption is needed to close the problem. The widely proposed observation for lost secondary jet kinetic energy is considered as a method of predicting the total amount of tip clearance loss in successive planes inside and downstream the blade passage.

A calculation procedure for predicting the tip clearance effects in the flow field inside and downstream the tip clearance, has been developed. The method, being compatible with a meridional flow calculation procedure, accounts for the calculation of the peripherally mean deficit profiles of the various flow quantities. The predictive capability of the calculation procedure is established in a wide range of test cases, including axial flow compressor cascades, isolated rotors and multi-row machines. The radial variation of tip clearance pressure loss is calculated with sufficient accuracy for engineering purposes.

Commentary by Dr. Valentin Fuster
1996;():V001T01A029. doi:10.1115/96-GT-073.
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A model for the prediction of the leakage vortex circulation was developed, based on the assumption that the leakage jet flow enters as a whole the vortex core, increasing its radius and its moment of momentum in the direction of the vortex axis. Using the assumption that the leakage vortex has a solid body rotation, an expression was derived for the vortex circulation, which demonstrates that this circulation is proportional to the square root of the corresponding tip clearance height. This theoretical result is supported by the available experimental data for both compressors and turbines. A simple model was developed, which demonstrates the ability of the proposed theory to calculate the leakage vortex circulation, provided that the vortex trace is known.

A method for predicting the tip clearance effects in the flow field inside and downstream the blade passage, compatible with a meridional flow calculation procedure, has been developed by the authors. The method uses incompressible considerations and accounts for the calculation of the circumferentially mean deficit radial profiles of the various flow quantities. In the calculation procedure the tip clearance flow effects are considered as a modification to the basic flow, existing in the absence of tip clearance. The complete calculation procedure was used in order to calculate the leakage vortex circulation and the induced velocity field in various axial flow cases, with satisfactory results.

Commentary by Dr. Valentin Fuster
1996;():V001T01A030. doi:10.1115/96-GT-074.
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An experimental investigation was carried out on the effect of blade chordwise lean on the losses in highly loaded rectangular turbine cascades. Detail measurements include 10 traverses from the upstream to the downstream of the cascades with five-hole spherical probes. Compared with the experimental data of the conventional straight and pitchwise lean blades under the same conditions, it is shown that the effect of chordwise lean on the development of the cascade losses is similar to that of pitchwise lean. However, the chordwise lean produces smaller streamwise adverse pressure gradients near both endwalls and a smaller spanwise negative one starting from the acute angle side in the first part of the passages in chordwise lean cascade, thereby the saddle point separations and intensities of the passage vortices are weakened and the secondary vortex losses are cut down notably.

Commentary by Dr. Valentin Fuster
1996;():V001T01A031. doi:10.1115/96-GT-075.
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The benefits of wave rotor-topping in turboshaft engines, subsonic high-bypass turbofan engines, auxiliary power units, and ground power units are evaluated. The thermodynamic cycle performance is modeled using a one-dimensional steady-state code; wave rotor performance is modeled using one-dimensional design/analysis codes. Design and off-design engine performance is calculated for baseline engines and wave rotor-topped engines, where the wave rotor acts as a high pressure spool. The wave rotor-enhanced engines are shown to have benefits in specific power and specific fuel flow over the baseline engines without increasing turbine inlet temperature. The off-design steady-state behavior of a wave rotor-topped engine is shown to be similar to a conventional engine. Mission studies are performed to quantify aircraft performance benefits for various wave rotor cycle and weight parameters. Gas turbine engine cycles most likely to benefit from wave rotor-topping are identified. Issues of practical integration and the corresponding technical challenges with various engine types are discussed.

Topics: Waves , Gas turbines , Rotors
Commentary by Dr. Valentin Fuster
1996;():V001T01A032. doi:10.1115/96-GT-097.
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Design and analysis of the turbine disc cooling and secondary air systems, thermal analysis of the rotor structure and subsequent validation of the temperature levels are described for a 4.9 MW industrial gas turbine.

A general purpose air flow network solver specifically developed for GT air systems is used to analyse and optimise the disc cooling air flows and determine the interaction between disc air, rotor blade cooling air and leakage.

Analysis of the individual wheelspace cavities is undertaken accounting for windage heating, rim sealing and potential for ingestion, heat transfer levels and cooling air temperature rise. This involves both simple 1D methods based on conservation of angular momentum and CFD analysis to determine flow recirculation mechanisms in the disc wheelspace.

The thermal boundary conditions from the above are applied to a finite difference thermal analysis model of the turbine rotor to predict temperature distributions at both steady state and transient conditions.

Test validation on the engine provides air system pressures and temperatures to validate the airflow analysis and rotor temperatures from thermal paints to validate the thermal analysis.

Topics: Cooling , Turbines , Disks
Commentary by Dr. Valentin Fuster
1996;():V001T01A033. doi:10.1115/96-GT-098.
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Experimental data have shown that combustor hot streaks can lead to pressure side “hot spots” on first-stage turbine rotor blades. Although many modern turbines operate at high subsonic or transonic flow speeds, the majority of bot streak experiments and numerical simulations performed during the last decade have been for low-speed flows. The presence of shock waves in a turbine stage can significantly affect the surface temperature distributions, and a knowledge of the interaction between shock waves and combustor hot streaks may help in the turbine design process. In the present investigation, quasi-three-dimensional unsteady Navier-Stokes simulations have been performed for a high-pressure turbine operating at two vane settings. At the open-vane setting, the flow is predominantly high subsonic with no trailing-edge shock waves, and at the closed-vane setting there are trailing-edge shocks.

Topics: Transonic flow
Commentary by Dr. Valentin Fuster
1996;():V001T01A034. doi:10.1115/96-GT-099.
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An idealized 3 blade test section has been used to study tip clearance effects which occur in transonic axial turbines. At subsonic inlet conditions (Mis1 = 0.56) the flow leaves the test section supersonic (Mis2 = 1.26). The tip clearance was varied from 0 to 15% of the chord length.

Extensive laser-2-focus anemometry was used to determine the tip gap mass flow based on the velocity vectors for gaps with 6, 10 and 15% chord. At small clearances the tip gap flow is mainly influenced by the pressure drop between pressure and suction side, while for larger gaps the main flow field dominates the tip gap flow.

The variation of the blade loading with the tip clearance was measured by static pressure tappings at 50% and 90% Span. Furthermore the static pressure along the tip surface was measured for varying tip clearances.

Pitot probe traverses in the tip vortex region at different downstream positions revealed the vortex structures and vortex core evolution. For tip gaps of 3 and 6%, multiple vortices were detected which were not fully mixed downstream. The origin of these vortices moves towards the trailing edge for larger gaps.

Commentary by Dr. Valentin Fuster
1996;():V001T01A035. doi:10.1115/96-GT-100.
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The regions of laminar and turbulent flow have been investigated in a linear cascade of a high tuming HP rotor blades. Measurements of intermittency close to the blade and end wall surfaces have shown substantial areas of laminar and transitional flow. The implications for turbulence modelling are important, and Navier-Stokes computations have been performed to investigate how well transition can be modelled in such a flow. Using the intermittency data to specify transitional areas, the mixing length model of turbulence produces excellent results, although there is some sensitivity to the assumed freestream length scale. High Reynolds k-ε model results show too much turbulence and loss using the measured high inlet length scale, but the results are improved with the Kato-Launder modification. A low Reynolds number model does not seem to predict the transition effects, although more work is required with this model.

Commentary by Dr. Valentin Fuster
1996;():V001T01A036. doi:10.1115/96-GT-113.
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Three-dimensional flow field measurements are presented for a large scale transonic turbine blade cascade. Flow field total pressures and pitch and yaw flow angles were measured at an inlet Reynolds number of 1.0 × 106 and at an isentropic exit Mach number of 1.3 in a low turbulence environment. Flow field data was obtained on five pitchwise/spanwise measurement planes, two upstream and three downstream of the cascade, each covering three blade pitches. Three-hole boundary layer probes and five-hole pitch/yaw probes were used to obtain data at over 1200 locations in each of the measurement planes. Blade and endwall static pressures were also measured at an inlet Reynolds number of 0.5 × 106 and at an isentropic exit Mach number of 1.0. Tests were conducted in a linear cascade at the NASA Lewis Transonic Turbine Blade Cascade Facility. The test article was a turbine rotor with 136° of turning and an axial chord of 12.7 cm. The flow field in the cascade is highly three-dimensional as a result of thick boundary layers at the test section inlet and because of the high degree of flow turning. The large scale allowed for very detailed measurements of both flow field and surface phenomena. The intent of the work is to provide benchmark quality data for CFD code and model verification.

Commentary by Dr. Valentin Fuster
1996;():V001T01A037. doi:10.1115/96-GT-114.
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The flow through the tip clearance region of a transonic compressor rotor (NASA rotor 37) was computed and compared to aerodynamic probe and laser anemometer data. Tip clearance effects were modeled both by gridding the clearance gap and by using a simple periodicity model across the ungridded gap. The simple model was run with both the full gap height, and with half the gap height to simulate a vena-contracta effect. Comparisons between computed and measured performance maps and downstream profiles were used to validate the models and to assess the effects of gap height on the simple clearance model. Recommendations were made concerning the use of the simple clearance model. Detailed comparisons were made between the gridded clearance gap solution and the laser anemometer data near the tip at two operating points. The computed results agreed fairly well with the data but overpredicted the extent of the casing separation and underpredicted the wake decay rate. The computations were then used to describe the interaction of the tip vortex, the passage shock, and the casing boundary layer.

Commentary by Dr. Valentin Fuster
1996;():V001T01A038. doi:10.1115/96-GT-115.
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The startup process is investigated for a hypothetical four-port wave rotor, envisioned as a topping cycle for a small gas turbine engine. The investigation is conducted numerically using a multi-passage, one-dimensional CFD based wave rotor simulation in combination with lumped volume models for the combustor, exhaust valve plenum, and rotor center cavity components. The simulation is described and several startup transients are presented which illustrate potential difficulties for the specific cycle design investigated. In particular it is observed that, prior to combustor light-off, or just after, the flow through the combustor loop is reversed from the design direction. The phenomenon is demonstrated and several possible modifications techniques are discussed which avoid or overcome the problem.

Commentary by Dr. Valentin Fuster
1996;():V001T01A039. doi:10.1115/96-GT-116.
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Wave rotor cycles which utilize premixed combustion processes within the passages are examined numerically using a one-dimensional CFD-based simulation. Internal-combustion wave rotors are envisioned for use as pressure-gain combustors in gas turbine engines. The simulation methodology is described, including a presentation of the assumed governing equations for the flow and reaction in the channels, the numerical integration method used, and the modeling of external components such as recirculation ducts. A number of cycle simulations are then presented which illustrate both turbulent-deflagration and detonation modes of combustion. Estimates of performance and rotor wall temperatures for the various cycles are made, and the advantages and disadvantages of each are discussed.

Topics: Combustion , Waves , Rotors
Commentary by Dr. Valentin Fuster
1996;():V001T01A040. doi:10.1115/96-GT-117.
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Wave rotors, used in a gas turbine topping cycle, offer a potential route to higher specific power and lower specific fuel consumption. In order to exploit this potential properly, it is necessary to have some realistic means of calculating wave rotor performance, taking losses into account, so that wave rotors can be designed for good performance. This in turn requires a knowledge of the loss mechanisms. The experiment reported here was designed as a statistical experiment to identify the losses due to finite passage opening time, friction, and leakage. For simplicity, the experiment used a 3-port, flow divider, wave cycle, but the results should be applicable to other cycles. A 12” diameter rotor was used, with two different lengths, 9” and 18”, and two different passage widths, 0.25” and 0.54”, in order to vary friction and opening time. To vary leakage, moveable end-walls were provided so that the rotor to end-wall gap could be adjusted. The experiment is described, and the results are presented, together with a parametric fit to the data. The fit shows that there will be an optimum passage width for a given wave rotor, since, as the passage width increases, friction losses decrease, but opening-time losses increase, and vice-versa. Leakage losses can be made small at reasonable gap sizes.

Topics: Waves , Rotors
Commentary by Dr. Valentin Fuster
1996;():V001T01A041. doi:10.1115/96-GT-118.
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The influence of inlet hot streak temperature distortion on turbine blade heat load was explored on a transonic axial flow turbine stage test article using a three-dimensional, multi-blade row unsteady Euler code. The turbine geometry was the same as that used for a recently reported testing of hot streak influence. Emphasis was placed elucidating the physical mechanisms by which hot streaks affect turbine durability. It was found that temperature distortion significantly increases both blade surface heat load nonuniformity and total blade heat load by as much as 10–30% (mainly on the pressure surface), and that the severity of this influence is a strong function of turbine geometry and flow conditions. Three physical mechanisms were identified which drive the heat load nonuniformity — buoyancy, wake convection (the Kerrebrock-Mikolajczak effect), and rotor-stator interactions. The latter can generate significant nonuniformity of the time-averaged relative frame rotor inlet temperature distribution. Dependence of these effects on turbine design variables was investigated to shed light on the design space which minimizes the adverse effects of hot streaks.

Topics: Heat , Stress , Rotors , Turbines
Commentary by Dr. Valentin Fuster
1996;():V001T01A042. doi:10.1115/96-GT-136.
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An improved understanding of a new category of stepped labyrinth seals, which feature a new “annular groove”, was obtained. A water leakage and flow visualization test facility of very large scale (relative to a typical seal) was utilized. Flow visualization experiments using a new method and digital facilities for capturing and editing digital images from an 8 mm video were conducted. The presence of an annular groove machined into the stator land increases the leakage resistance by up to 26 percent for the cases considered here. Tracer particles show the degree of throughflow path penetration into the annular groove (i.e. serpentining), which gives the largest and the smallest leakage resistance improvement over that of the corresponding conventional stepped seal.

Commentary by Dr. Valentin Fuster
1996;():V001T01A043. doi:10.1115/96-GT-137.
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An experimental investigation was conducted to determine the geometry-leakage relationship for advanced, stepped labyrinth seals. A unique, variable-geometry water test facility was constructed and used to acquire leakage resistance measurements for two-dimensional, planar models. Flow visualization techniques were also used to assist in identifying and understanding the turbulence generating flow patterns. It was found that contoured surfaces and restrictor tooth leading-edge shapes of proper dimensions can be incorporated into the cavity geometry to reduce seal leakage. Specifically, the combination of a sloping surface and a curved surface on the rotor within the labyrinth cavity gave significant improvement.

Commentary by Dr. Valentin Fuster
1996;():V001T01A044. doi:10.1115/96-GT-139.
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Inter-turbine diffusers offer the potential advantage of reducing the flow coefficient in the following stages leading to increased efficiency. The flows associated with these ducts differ from those in simple annular diffusers both as a consequence of their high-curvature S-shaped geometry and of the presence of wakes created by the upstream turbine. Experimental data and numerical simulations clearly reveal the generation of significant secondary flows as the flow develops through the diffuser in the presence of cross-passage pressure gradients. The further influence of inlet swirl is also demonstrated. Data from experimental measurements with and without an upstream turbine are discussed and computational simulations are shown not only to give a good prediction of the flow development within the diffuser but also to demonstrate the importance of modelling the fully three-dimensional nature of the flow.

Commentary by Dr. Valentin Fuster
1996;():V001T01A045. doi:10.1115/96-GT-140.
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The main objective of this work is the analysis and the comparison between different methods utilised to solve the Moore rotating stall model. To date only simplified relations between the axial flow perturbation g and the transverse one h have been utilised and presented in literature, such as h′ = −g or the truncated Fourier series. On the contrary, in this paper the accurate relation given by the Hilbert Transform is utilised, and to improve the numerical stability of the method a new expression of the first derivative of transverse flow coefficient perturbation is proposed and utilised.

A complete and detailed comparison between the results of the simplified methods and the solution proposed here is presented. This comparison is extended to a wide range of geometrical and physical compressor parameters, and it allows the accuracy of simplified approaches to be tested.

Finally, a correlative approach estimating overall rotating stall effects based on the complete solution proposed here is presented. It allows rotating stall influence to be quickly and easily taken into account in several axial compressor areas (design, optimisation, active control, etc.).

Commentary by Dr. Valentin Fuster
1996;():V001T01A046. doi:10.1115/96-GT-141.
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A recently developed, time-accurate multigrid viscous solver has been extended to handle quasi-three-dimensional effects and applied to the first stage of a modern transonic compressor. Interest is focused on the inlet guide vane (IGV):rotor interaction where strong sources of unsteadiness are to be expected. Several calculations have been performed to predict the stage operating characteristics. Flow structures at various mass flow rates, from choke to near stall, are presented and discussed. Comparisons between unsteady and steady pitch-averaged results are also included in order to obtain indications about the capabilities of steady, multi-row analyses.

Commentary by Dr. Valentin Fuster
1996;():V001T01A047. doi:10.1115/96-GT-149.
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The objective of the present paper is to investigate the axial skewing effect of a turbine stator blading on its aerodynamic characteristics systematically, to find the possibility for reducing the secondary losses by using axially skewed bladings and to understand better its flow physics in turbomachinery. In the present paper a typical turbine stator blading is applied as an example to generate a set of axially skewed bladings for systematical study and illustration of their differences of aerodynamics characteristics. The paper gives the procedure for generating different forward-skewed and backward-skewed blades with the same profile sections at the same radius and the numerical method used is also described briefly. The method is based on the 3-D time-marching finite volume Navier-Stokes solution and was developed by the Institute of Engineering Thermophysics, Chinese Academy of Sciences. The turbulence model proposed by Baldwin and Lomax is used here for predicting the effective viscosity. The calculated results and their comparisons are also given in the present paper. On the basis of the analysis it is shown that the appropriate use of skewed blades gives designers another possibility to control the flow in the blade channel. By adopting forward-skewed blades to replace the straight blades can be reduced the blade loading near the leading edge and in the central part of the span. It is also found that the pressure gradient at the endwall of forward-skewed blading yields the radial force that enables to avoid the boundary layer separation from the endwall. The axial skew of blade enables to restrain the strength of the secondary flow. Therefore, the total pressure loss can be reduced. An attention should be paid: if the skewed blade for stator is chosen to be used, the radial distribution of the outlet flow angle from stator vane is required to meet the optimal incidence satisfaction to the rotor blades. Otherwise, it results in the reduction of efficiency.

Topics: Turbines , Stators
Commentary by Dr. Valentin Fuster
1996;():V001T01A048. doi:10.1115/96-GT-151.
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The three-dimensional flow in centrifugal impellers is investigated on the basis of a detailed analysis of the results of numerical simulations. In order to gain confidence in this process, an in-depth validation is performed, based on computations of Krain’s centrifugal compressor and of a radial pump impeller, both with vaneless diffusers. Detailed comparisons with available experimental data provide high confidence in the numerical tools and results. The appearance of a high loss ‘wake’ region results from the transport of boundary layer material from the blade surfaces to the shroud region and its location depends on the balance between secondary and tip leakage flows and is not necessarily connected to 3D boundary layer separation. Although the low momentum spots near the shroud can interfere with 3D separated regions, the main outcome of the present analysis is that these are two distinct phenomena. Part I of this paper focuses on the validation base of the numerical approach, based on fine mesh simulations, while Part II presents an analysis of the different contributions to the secondary flows and attempts to estimate their effect on the overall flow pattern.

Commentary by Dr. Valentin Fuster
1996;():V001T01A049. doi:10.1115/96-GT-152.
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The three-dimensional flow in centrifugal impellers is investigated on the basis of a detailed analysis of the results of numerical simulations. An in-depth validation has been performed, based on the computations of Krain’s centrifugal compressor and a radial pump impeller, both with vaneless diffusers and detailed comparisons with available experimental data, discussed in Part I, provide high confidence in the numerical tools and results. The low energy, high loss ‘wake’ region results from a balance between various contributions to the secondary flows influenced by tip leakage flows and is not necessarily connected to 3D boundary layer separation. A quantitative evaluation of the different contributions to the streamwise vorticity is performed, identifying the main features influencing their intensity. The main contributions are: the passage vortices along the end walls due to the flow turning; a passage vortex generated by the Coriolis forces proportional to the local loading and mainly active in the radial parts of the impeller; blade surface vortices due to the meridional curvature. The analysis provides an explanation for the differences in wake position under different geometries and flow conditions. A secondary flow representation is derived from the calculated 3D flow field for the two geometries validated in Part I, and the identified flow features largely confirm the theoretical analysis.

Commentary by Dr. Valentin Fuster
1996;():V001T01A050. doi:10.1115/96-GT-153.
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In order to improve the operating range of a centrifugal compressor, computer-controlled variable inlet and diffuser vanes were attached to a compressor with a pressure ratio of 2.5. Low-solidity cascade vanes capable of controlling the vane angle up to 0 degrees from the tangential direction were used for the vaned diffuser. The compressor’s overall performance was then tested using a closed-loop test stand. By automatically adjusting the diffuser vanes to the most suitable flow angle, pressure fluctuations caused by the unstable flow in the diffuser during low-flow operation of the centrifugal compressor could be suppressed, and the compressor could be operated nearly up to the shut-off flow rate without any surge. The author experimentally confirmed the critical operating range of both the impeller and diffuser at two different tip speeds and five inlet guide vane angles. Furthermore, a three-dimensional viscous flow-analysis method was applied to the impeller, and a three-dimensional momentum integral analysis method was applied to the diffuser. Then the critical operating ranges obtained in the experiments were qualitatively validated. The operating range of a centrifugal compressor under low-flow conditions, which has until now been limited because of surge, dramatically improved in this study, thereby demonstrating that it may be possible to develop a surge-free centrifugal compressor.

Commentary by Dr. Valentin Fuster
1996;():V001T01A051. doi:10.1115/96-GT-154.
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A potential flow computer model that can handle blade row interaction problems has been used to analyze the circumferential static pressure distribution at the trailing edge plane of the last rotor in an axial compressor which is produced by a downstream stator/strut system. The computer model is based on the Douglas-Neumann formulation. The code was used to design a circumferentially nonuniform stagger angle distribution for the stator that reduced the static pressure disturbance on the rotor. The predicted circumferential static pressure distribution and its resulting frequency content at the rotor trailing edge station for the baseline (uniform circumferential stagger angles) stator and for the optimized stator are compared to static pressure data and derived frequency content from engine tests of each configuration. The results show good agreement between the model predictions and the test data. The results are further confirmed by measurements of rotor strain levels with the baseline stator and with the optimized stator, which show a proportional decrease in rotor strain for the optimized stator configuration. Since incorporation of this low-cost modification, there has been no evidence of vibratory induced rotor distress, thereby improving engine reliability and maintainability and enhancing customer satisfaction.

Commentary by Dr. Valentin Fuster
1996;():V001T01A052. doi:10.1115/96-GT-155.
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Centrifugal compressors have the widest compressor application area, covering aircraft engines, small stationary gas turbines, process and refinery industries, the refrigeration industry, and turbochargers. Despite the vast literature coverage of diffuser systems for centrifugal compressors, there are not more than twenty publications in the open literature on the family of vaned diffusers known as Low Solidity Vaned Diffuser (LSVD). This is highly surprising, in light of the fact, that practically all process and refrigeration compressors manufacturers, at one time or another, have attempted to design and test LSVD. Therefore with the strong belief that any work on LSVD either theoretical or experimental will be welcomed, this paper presents the performance of two newly designed LSVD.

Comparative experimental studies on diffuser systems for centrifugal compressors, performed at the Michigan State University Turbomachinery Lab are presented. A vaneless, a conventional vaned and two low solidity vaned diffusers were tested. The results are compared for the effect of the diffuser systems on the stage performance, the maximum efficiency, and the operating range of the compressor. The effect of the vane number in low solidity vaned diffuser on the performance is also discussed.

Commentary by Dr. Valentin Fuster
1996;():V001T01A053. doi:10.1115/96-GT-156.
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A hybrid grid system that combines the Chimera overset grid scheme and an unstructured grid method is developed to study fluid flow and heat transfer problems. With the proposed method, the solid structural region, in which only heat conduction is considered, can be easily represented using an unstructured grid method. As for the fluid flow region external to the solid material, the Chimera overset grid scheme has been shown to be very flexible and efficient in resolving complex configurations. The numerical analyses require the flow field solution and material thermal response to be obtained simultaneously. A continuous transfer of temperature and heat flux is specified at the interface, which combines the solid structure and the fluid flow into an integral system. Numerical results are compared with analytical and experimental data for a flat plate and a C3X cooled turbine cascade. A simplified drum-disk system is also simulated to show the effectiveness of this hybrid grid system.

Topics: Heat transfer
Commentary by Dr. Valentin Fuster
1996;():V001T01A054. doi:10.1115/96-GT-157.
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This paper describes experimental and computational work to examine the unsteady flowfield in the vaned diffuser of a medium specific speed pump. The time periodic flowfield in the diffuser has been examined experimentally with laser optical techniques and with unsteady pressure transducers. The flow has been computed with a general purpose Navier-Stokes CFD code, whereby the unsteady effects have been simulated by a time periodic inlet profile which translates across the diffuser inlet and represents the wakes and potential interaction from the impeller. Comparisons of a steady simulation with time-average inlet conditions, an unsteady simulation with a time periodic inlet profile and the time-average of the unsteady simulation are used to validate the code developments and to examine features of the unsteady flow through the diffuser. Comparisons with experimental data identify that this simple computational model with a time periodic inlet profile is able to simulate the convection and decay of disturbances passing through the diffuser.

Commentary by Dr. Valentin Fuster
1996;():V001T01A055. doi:10.1115/96-GT-158.
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A new blade design methodology, which allows designers to react rapidly to changes in functional requirements of turbines and transition quickly from concept to design, has been discussed in this paper. This methodology reduces the design cycle time by creating a three-dimensional model of the blade, through concurrent design of multiple two-dimensional blade sections. An efficient direct design procedure has been developed by coupling direct optimization techniques with two-dimensional aerodynamic analysis codes. A method for interpreting the flowfield solution to compute the airfoil quality has been developed and is used to compute the objective function during optimization. Aerodynamic, mechanical and geometry constraints are imposed on the design to ensure that the optimized design meets feasibility requirements of all engineering disciplines. During the design optimization process, the designers’ interactions are simulated through use of rules that are based on designer heuristics. This procedure is used for the design of a high pressure, steam turbine blade; the results are discussed in this paper.

Commentary by Dr. Valentin Fuster
1996;():V001T01A056. doi:10.1115/96-GT-159.
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This study is concerned with the use of low-Reynolds-number models of turbulence transport in the computation of flows through rotating cavities. The models tested are the Launder and Sharma low-Re k-ε (L-S) and a low-Re differential second-moment closure (DSM), first used by Iacovides and Toumpanakis, both with and without the Yap correction term to the dissipation rate equation. The cases examined include rotor-stator systems without throughflow, rotor-stator systems with radial outflow, contra-rotating disc systems without throughflow and also with radial outflow, co rotating discs with radial outflow and also rotor-stator systems with radial inflow. Earlier studies have shown that, when no throughflow or when radial outflow is involved, the L-S tends to over-estimate the size of the regions over which the boundary layers remain laminar, while the zonal k-ε/l-eqn model is unable to predict partially laminarized flows. A modification to the ε equation proposed here, which in regions of low turbulence reduces the dissipation rate when the fluid is in solid body rotation, provides a simple empirical way to significantly improve the L-S predictions of partially laminarized flows through rotating cavities, to acceptable levels. The DSM model used, in some cases led to some further predictive improvements and, for rotor-stator systems without throughflow, to a significant improvement in the predicted value of the moment coefficient. The Yap length scale correction term, while in most cases it has either a beneficial or a neutral effect on the flow predictions, in cases involving radial inflow it leads to poorer predictions. Models that do not rely on wall distance thus appear more likely to have a wider range of applicability.

Commentary by Dr. Valentin Fuster
1996;():V001T01A057. doi:10.1115/96-GT-160.
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Conditionally averaged Navier-Stokes equations are used to describe transitional flow in adverse pressure gradient combined with a transport equation for the intermittency factor γ. A transport equation developped in earlier work has been modified to eliminate the use of a distance along a streamline. An extension of the correlations is proposed to determine the spot growth parameter in adverse pressure gradient. This approach is verified against flows over a flat plate with an elliptical leading edge.

Commentary by Dr. Valentin Fuster
1996;():V001T01A058. doi:10.1115/96-GT-166.
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The laminar-turbulent transition process has been documented in a concave-wall boundary layer subject to low (0.6%) free-stream turbulence intensity. Transition began at a Reynolds number, Rex (based on distance from the leading edge of the test wall), of 3.5×105 and was completed by 4.7×105. The transition was strongly influenced by the presence of stationary, streamwise, Görtler vortices. Transition under similar conditions has been documented in previous studies, but because concave-wall transition tends to be rapid, measurements within the transition zone were sparse. In this study, emphasis is on measurements within the zone of intermittent flow. Twenty-five profiles of mean streamwise velocity, fluctuating streamwise velocity, and intermittency have been acquired at five values of Rex, and five spanwise locations relative to a Görtler vortex. The mean velocity profiles acquired near the vortex downwash sites exhibit inflection points and local minima. These minima, located in the outer part of the boundary layer, provide evidence of a “tilting” of the vortices in the spanwise direction. Profiles of fluctuating velocity and intermittency exhibit peaks near the locations of the minima in the mean velocity profiles. These peaks indicate that turbulence is generated in regions of high shear, which are relatively far from the wall. The transition mechanism in this flow is different from that on flat walls, where turbulence is produced in the near-wall region. The peak intermittency values in the profiles increase with Rex, but do not follow the “universal” distribution observed in most flat-wall, transitional boundary layers. The results have applications whenever strong concave curvature may result in the formation of Görtler vortices in otherwise 2-D flows. Because these cases were run with a low value of free-stream turbulence intensity, the flow is not a replication of a gas turbine flow. However, the results do provide a base case for further work on transition on the pressure side of gas turbine airfoils, where concave curvature effects are combined with the effects of high free-stream turbulence and strong streamwise pressure gradients, for they show the effects of embedded streamwise vorticity in a flow that is free of high-turbulence effects.

Commentary by Dr. Valentin Fuster
1996;():V001T01A059. doi:10.1115/96-GT-168.
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The three-dimensional viscous flow field development in the nozzle passage of an axial flow turbine stage was measured using a “x” hot-wire probe. The measurements were carried out at one axial station on the endwall and blade surfaces and at several spanwise and pitchwise locations. Static pressure measurements and flow visualization, using a fluorescent oil technique, were also performed to obtain the location of transition and the endwall limiting streamlines. The boundary layers on the blade surface were found to be very thin and laminar, except on the suction surface downstream of 70% axial chord. Strong radial pressure gradient, especially close to the suction surface, induces strong radial flow velocities in the trailing edge regions of the blade. On the endwalls, the boundary layers were turbulent and much thicker, especially near the suction corner of the casing surface, caused by the secondary flow. The secondary flow region near the suction casing surface corner indicated the presence of the passage vortex detached from the blade surface. The boundary layer code accurately predicts the three-dimensional boundary layers on both vane surfaces in regions where the influence of secondary flow is small.

Commentary by Dr. Valentin Fuster
1996;():V001T01A060. doi:10.1115/96-GT-170.
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A mathematical model is developed for the flow field diagnosis problem in multistage axial compressors. In view of the ill-posedness of the diagnostic problem, an effective measure is adopted to transfer the diagnostic problem into a variational problem which is solved by a regularization method. Two numerical results demonstrate the rationality of the flow field diagnosis problem for the compressors running near the design point and the effectiveness of the computational method.

Commentary by Dr. Valentin Fuster
1996;():V001T01A061. doi:10.1115/96-GT-171.
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The paper presents the results of an experimental investigation on a multistage centrifugal blower, during rotating stall. The test plant allows to change the turbomachine characteristics; in this research the blower has been tested in two different configurations: two-stage and four-stage, with vaneless diffusers.

The unsteady flow field inside the blower has been measured by means of hot-wire anemometers. Three single hot-wire probes have been utilised to measure the development of the rotating stall, while a crossed hot-wire probe has been utilised to obtain the instantaneous flow field behind the impellers.

The measurements have been done at different flow rate values, including stall inception.

Commentary by Dr. Valentin Fuster
1996;():V001T01A062. doi:10.1115/96-GT-177.
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The work describes a quasi three dimensional time marching method (Euler Solver), allowing for variations in streamtube height and radius in the meridional plane, for the generation of blade shapes. The method is applicable for both compressor and turbine blades. The generation of the blade shapes is done corresponding to a prescribed pressure distribution (with or without shock waves) around the blade surfaces. The method can be used to either generate entire blade shapes or to modify regions of an existing inefficient blade shape. Viscous effects are allowed for by using an integral boundary layer growth method in the scheme. For the generation of turbine blade shapes the effects of trailing edge base pressure are also included in the method. The accuracy of the scheme has been demonstrated through various test cases. The use of the method results in vast reduction in overall design time and hence cost.

Commentary by Dr. Valentin Fuster
1996;():V001T01A063. doi:10.1115/96-GT-179.
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An aerodynamic investigation of the influence of outlet stator part (vaneless diffuser and return channel) surface roughness on aerodynamic performance of a very low flow coefficient centrifugal stage has been carried out. The stage with design inlet flow coefficient 0.007 was tested within the range of stage Mach number Mu2=0.5–1.1. Then the surface quality of outlet stator part was improved and the tests have been repeated. Aerodynamic performance and losses in both vaneless diffuser and return channel with de-swirl vanes were investigated. The values of isentropic head coefficient increased while those of loss coefficient decreased nearly in the whole range of characteristics in the stage with improved surface quality. The detailed pressure recovery in vaneless diffuser in vicinity of design point measured and calculated by the performance prediction method is compared and discussed. The nonsteady flow phenomena were also investigated. The change of dynamic stability limit by improving of surface quality was observed.

Commentary by Dr. Valentin Fuster
1996;():V001T01A064. doi:10.1115/96-GT-198.
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A blind test case for a compressor rotor (ROTOR 37) was organized by the ASME/IGTI at its 1994 meeting in order to assess the predictive capabilities of the turbomachinery CFD tools. The results from the different CFD codes showed a wide scatter which in part is due to the differences in the turbulence models that were used. In order to systematically isolate the capabilities and limitations of the turbulence models, ROTOR 37 flow is computed from the same numerical platform with three different turbulence models. These include: the Baldwin-Lomax model, the standard k-ϵ model, and an improved version of this k-ϵ model. The results from the three models are compared with the experiment. We find that with increasing model complexity the results move closer to the experiment. Several sensitivity studies are carried out to bracket the uncertainty in the computations. These include the effect of: wall boundary conditions for the turbulence models; numerical accuracy of the turbulence solver; and the effect of the inlet boundary condition for turbulence.

Commentary by Dr. Valentin Fuster
1996;():V001T01A065. doi:10.1115/96-GT-199.
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A theory is presented for calculating the fluctuations in a laminar boundary layer when the free stream is turbulent. The kinetic energy equation for these fluctuations is derived and a new mechanism is revealed for their production. A methodology is presented for solving the equation using standard boundary layer computer codes. Solutions of the equation show that the fluctuations grow at first almost linearly with distance and then more slowly as viscous dissipation becomes important. Comparisons of calculated growth rates and kinetic energy profiles with data show good agreement.

In addition, a hypothesis is advanced for the effective forcing frequency and free-stream turbulence level which produce these fluctuations. Finally, a method to calculate the onset of transition is examined and the results compared to data.

Commentary by Dr. Valentin Fuster
1996;():V001T01A066. doi:10.1115/96-GT-202.
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Theoretical analyses and experimental results are reported for two unique variable geometry techniques used with pipe diffusers to enhance off-design performance. One technique mechanically closes the diffuser throat in an unusual manner. The other allows flow recirculation to close the throat artificially while attempting to improve diffuser inlet flow characteristics.

Results clearly show that surge margin may be significantly improved by either method and that flow recirculation may offer improved efficiency.

Topics: Diffusers , Pipes , Geometry
Commentary by Dr. Valentin Fuster
1996;():V001T01A067. doi:10.1115/96-GT-203.
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Measurements of pressure distributions, profile losses and flow deviation were carried out on a planar turbine cascade in incompressible flow to assess the effects of partial roughness coverage of the blade surfaces. Spanwise oriented bands of roughness were placed at various locations on the suction and pressure surfaces of the blades. Roughness height, spacing between roughness elements, and band width were varied.

A computational method based on the inviscid/viscous interaction approach was also developed; its predictions were in good agreement with the experimental results. This indicates that good predictions can be expected for a variety of cascade and roughness configurations from any 2-dimensional analysis which couples an inviscid method with a suitable rough surface boundary-layer analysis. The work also suggests that incorporation of the rough wall skin-friction law into a 3-D Navier-Stokes code would enable good predictions of roughness effects in 3-dimensional situations.

Roughness was found to have little effect on static pressure distribution around the blades and on deviation angle, provided that it does not precipitate substantial flow separation. Roughness on the suction surface can cause large increases in profile losses; roughness height and location of the leading edge of the roughness band are particularly important. Loss increments due to pressure-surface roughness are much smaller than those due to similar roughness on the suction surface.

Commentary by Dr. Valentin Fuster
1996;():V001T01A068. doi:10.1115/96-GT-238.
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A detailed investigation of aerodynamic performance of three low-speed rear axial compressor stage bladings with the aspect ratios: 0.75, 1.0 and 1.25 was carried out. The bladings of industrial type consist of rotor, stator and outlet guide vanes. The outer and inner diameters of the stage are constant. The hub/tip ratio is 0.871 and the outer diameter is 800 mm. Stage blading is followed by an annular diffuser with outlet chamber.

The effect of blade aspect ratio on compressor stage performance was also analysed with the use of straight cascade data. This data supported the test stage experimental results. We found that the effect of aspect ratio on stage performance is not remarkable in the considered range. There are some differences at off-design conditions. The lowest value of blading efficiency was obtained in the case with the lowest aspect ratio value.

Three inlet velocity profiles were modelled with the use of lengthened inlet annulus and a screen specially designed. It was found that there is a significant effect of inlet velocity profile distortion on rear compressor stage blading performance for all aspect ratios.

Aerodynamic characteristics of compressor stage blading with annular diffuser and outlet chamber were determined. During the investigation we also removed the outlet guide vanes. Therefore the effect of swirl and inlet velocity profile could be investigated.

Commentary by Dr. Valentin Fuster
1996;():V001T01A069. doi:10.1115/96-GT-239.
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Steady state and dynamic data were acquired in a T55-L-712 compressor rig. In addition, a T55-L-712 engine was instrumented and similar data were acquired. Rig and engine stall/surge data were analyzed using modal techniques. This paper compares rig and engine preliminary results for the ground idle (approximately 60% of design speed) point. The results of these analyses indicate both rig and engine dynamic events are preceded by indications of traveling wave energy in front of the compressor face. For both rig and engine, the traveling wave energy contains broad band energy with some prominent narrow peaks and, white the events are similar in many ways, some noticeable differences exist between the results of the analyses of rig data and engine data.

Topics: Engines , Compressors
Commentary by Dr. Valentin Fuster
1996;():V001T01A070. doi:10.1115/96-GT-240.
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To protect a compressor from surge, it is necessary to accurately calculate the location of the interface between stable operation and surge. Describing the Surge Limit Interface in certain coordinate systems results in a surface which is invariant to compressor suction conditions such as temperature and molecular weight. We refer to these coordinates as invariant coordinates. We explore these invariant coordinate systems and some nearly invariant systems useful for antisurge control. Some of them are commonly used in the industry, others are quite novel. This work serves to point out the unifying basis of them all.

The applications for these methods are mainly industrial compressors. Varying molecular weight represents the main challenge since a real-time measurement for this parameter is unavailable.

We present compressor maps constructed from test data in these coordinates. The validity of this approach is well supported by these data.

Topics: Compressors
Commentary by Dr. Valentin Fuster
1996;():V001T01A071. doi:10.1115/96-GT-241.
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This paper describes an actuator placement methodology for the active control of purely one-dimensional instabilities of a seven-stage axial compressor using an air bleeding strategy. In this theoretical study, using stage-by-stage non-linear modelling based on the conservation equations of mass, momentum, and energy, a scheduling LQR (Linear Quadratic Regulator) controller is designed for several actuator locations in a compressor from the first stage to the plenum. In this controller design, the LQR weighting matrices are selected so that the associated cost function includes only air bleeding mass flow leading to the minimisation of the air bleed. The LQR cost function represents a measure of the consumption of air bleeding and can be calculated analytically using the solution of an Algebraic Riccati Equation. From analysis of the cost at different compressor stages, the location of an air bleeding actuator is selected at the stage with the minimum cost. Finally, using an ACSL simulation program, the scheduling controller has been integrated with a non-linear. stage-by-stage model and the time response of the air bleeding mass flow at different locations has been obtained to confirm the results from the analytical approach. Results are presented to show actively stabilised compressor flow beyond the surge point where the air bleed is minimised. These results also indicate the preferred location of the actuator at the compressor downstream stages for both low and high compressor speeds.

Commentary by Dr. Valentin Fuster
1996;():V001T01A072. doi:10.1115/96-GT-253.
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This paper addresses the causal link first described by Smith between the unsteady flow induced by the rotor wakes and the compressor steady-state performance. As an initial step, inviscid flow in a compressor stage is examined. First of a kind numerical simulations are carried out to show that if the rotor wakes are mixed out after (as opposed to before) the stator passage, the time-averaged overall static pressure rise is increased and the mixing loss is reduced. An analytical model is also presented and shown to agree with the numerical results; the model is then used to examine the parametric trends associated with compressor design parameters.

Commentary by Dr. Valentin Fuster
1996;():V001T01A073. doi:10.1115/96-GT-254.
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Aero compressor technology has seen a significant advancement during the last two decades. Research on Rotor-Stator stage has focused on improvements in the design as well as off-design performance. Work has also been carried out to improve the clearance losses, stage loading as well as the stability of the operation. In this connection, the work on end-bend rotor-stator and variable geometry stators is specially significant. These efforts have however, yielded marginal improvements as far as the capability of the stage to produce pressure rise and its through-flow capacity are concerned. The interest in contra-rotation has emerged with a view to achieve considerable high pressure rise per stage besides its effects on stability of the stage to rotating stall/surge suppression. Contra-rotation concept has already found its acceptability in the development of future fuel efficient gas turbine plants and aero engines.

This paper presents a review of the experimental and theoretical investigations on the aero-dynamic and aero-acoustic performance of the contra-rotating pressure stage. The areas of future work on contra-rotation are also outlined.

Commentary by Dr. Valentin Fuster
1996;():V001T01A074. doi:10.1115/96-GT-255.
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The present study addresses two aspects of the horseshoe vortex, namely its significance in the secondary flow in a turbine blade passage and the possibility of reducing its strength by an active flow mechanism, i.e the transverse injection of coolant air through a slot in a cylinder-endwall junction. The study reports on the results of two experiments in low speed wind tunnels, which employed a calibrated five-hole Pitot tube to measure the velocity vectors and the resulting secondary flowfields. The first aspect was studied in a 90° square cross section bend duct. The two horseshoe vortex legs were simulated by two half-Delta wing vortex generators. The results showed that the horseshoe vortices influence two regions of the secondary flowfield, i.e one near the passage entrance, where the pressure side leg forces a three dimensional separation of the endwall boundary layer, and the other is in the exit plane, where the coupling of the horseshoe with the passage vortex redistributes the flow with total pressure losses, without affecting the total loss, and increases the secondary kinetic energy by about 20%. For the second aspect, a rectangular bluff body, with a cylindrical leading edge, was positioned over the tunnel endwall and the transverse air injection was implemented through a thin slot, covering the 180° arc in the leading edge-endwall junction. The results showed that, for an average injection velocity equal to 35% that of the mainstream, the size and strength of the horseshoe vortex leg were reduced by nearly 60%. On the other hand, for stronger injection rates the vortex size and strength were increased.

Topics: Vortices
Commentary by Dr. Valentin Fuster
1996;():V001T01A075. doi:10.1115/96-GT-256.
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Siemens heavy duty Gas Turbines have been well known for their high power output combined with high efficiency and reliability for more than 3 decades. Offering state of the art technology at all times, the requirements concerning the cooling and sealing air system have increased with technological development over the years. In particular the increase of the turbine inlet temperature and reduced NOx requirements demand a highly efficient cooling and sealing air system. The new Vx4.3A family of Siemens gas turbines with ISO turbine inlet temperatures of 1190°C in the power range of 70 to 240 MW uses an effective film cooling technique for the turbine stages 1 and 2 to ensure the minimum cooling air requirement possible. In addition, the application of film cooling enables the cooling system to be simplified. For example, in the new gas turbine family no intercooler and no cooling air booster for the first turbine vane are needed.

This paper deals with the internal air system of Siemens gas turbines which supplies cooling and sealing air. A general overview is given and some problems and their technical solutions are discussed. Furthermore a state of the art calculation system for the prediction of the thermodynamic states of the cooling and sealing air is introduced. The calculation system is based on the flow calculation package Flowmaster (Flowmaster International Ltd.), which has been modified for the requirements of the internal air system. The comparison of computational results with measurements give a good impression of the high accuracy of the calculation method used.

Commentary by Dr. Valentin Fuster
1996;():V001T01A076. doi:10.1115/96-GT-257.
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The flow path of multistage centrifugal compressors is characterized by two 180-degree turns per stage: the inlet turning bend connecting the radial inflow return channel with the radial outflow impeller and the cross-over bend connecting the radial outflow diffuser with the return channel. Due to higher flow velocity and larger width to turning radius ratio, the turning losses are substantially higher in the inlet bend than in the return channel bend. Performance measurements were taken using different annular through-flow turning vane arrangements designed to reduce the inlet turning losses and increase the overall efficiency of the multistage centrifugal compressor. The experiments have shown consistent efficiency gains with corresponding capacity increases by adding multiple annular turning vanes in the inlet bend. The performance improvement potential of the vanes depends strongly on the positioning of these vanes in the flow passage. Based on these results, an empirical turning loss model was developed with the capability to predict the performance improvement achievable with correctly positioned single or multiple turning vanes in the impeller inlet bend area.

Commentary by Dr. Valentin Fuster
1996;():V001T01A077. doi:10.1115/96-GT-258.
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Navier-Stokes equations written in the Boussinesq’s approximation are solved numerically for simulation of the laminar and turbulent flows of the cooling air through a long gap between two co-rotating cylinders kept at different temperatures. Flow situations with Grashof numbers up to 107, are considered. Data on velocity and temperature fields as well as on local Nusselt number distributions and averaged heat transfer have been obtained for a 60° sector. It has been established that the spanwise size of buoyancy-driven vortices developing in a 60° sector channel changes significantly over the gap in case of laminar flow and remain nearly constant for turbulent flow. For the laminar regime, the cross flow intensity is approximately ten times higher than for turbulent flows at the Reynolds numbers of the order of 104. Some computations have been performed using a simplified single-cell formulation.

Commentary by Dr. Valentin Fuster
1996;():V001T01A078. doi:10.1115/96-GT-259.
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Wave rotors have been investigated over several decades in part due to their potential for increasing the maximum cycle temperature in gas turbine engines via their self-cooling mechanism. Recent activities in this field have centered on the experiments and CFD design tools developed at NASA Lewis Research Center. Because of the fundamental objectives of that program, the work to date has concentrated on wave rotors rather than wave turbines. Wave turbines differ from wave rotors in that the flow passages are curved, similar to conventional turbines, so that the unit produces net shaft power. The purpose of this paper is to present an analysis technique which is used to quantify the substantial impact which blade curvature has on the maximum gas expansion ratio, and hence on the maximum cycle temperature. Limited optimization of the overall pressure ratio allows the maximum specific power and the corresponding efficiency to be found as a function of wave turbine inlet and exit blade angles and Mach number. A potential increase in specific power of 69% and a 6.8 percentage point increase in thermal efficiency over a conventional gas-turbine engine can be achieved through the use of a wave turbine.

Topics: Engines , Waves , Design , Turbines , Blades
Commentary by Dr. Valentin Fuster
1996;():V001T01A079. doi:10.1115/96-GT-261.
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It was shown in a previous paper of the authors (1991) that jet and wake in the flow of the impeller of the centrifugal compressor are developed from the Dean’s type vortex pair formed in the curvature of the blade channel. The jet rotating against the sense of the impeller is weakened, and the wake rotating in the sense of the impeller is enhanced during travelling with the flow toward the outlet. This property is attributed to the conservation of the potential vorticity of the vortex. The experimental result obtained by Krain (1984) has confirmed this theory.

The secondary flows found by Farge and Johnson (1990) enable the determination of the vorticity of the wake at the outlet of the impeller. It amounts to 6.9 Ω and 5.8 Ω for the radial-blading and the 60°-backswept blading impeller, respectively. The intensity of the vortex jet is weakened to undetectable value for both the impellers. The patterns of these secondary flow fields are also quite different between these two kinds of impellers. Whilst that of the former is controlled by the intrinsic motion, that of the latter is governed by the relative velocity along the blades.

Furthermore, the experimental result obtained by the injection of colored dye at the impeller outlet and the measured velocity field around the impeller reveal an intense reverse flow in the radial blading impeller, travelling from the outlet toward the inlet, along the shroud. It can be shown that this reverse flow is caused by the intrinsic motion occuring in this impeller and impinging on the leading edge of the diffuser vane. As the rotating stall is introduced by the reverse flow, the low-solidity vaned diffuser, and still better the vaneless diffuser can therefore shift the stall line to a very low flow rate.

Commentary by Dr. Valentin Fuster
1996;():V001T01A080. doi:10.1115/96-GT-262.
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A broad operating range between surge and choke is so important for turbocharger compressors and many other applications that a vaneless diffuser, with its reduced efficiency, is usually adopted. With the demand for increased pressure ratio the operating range naturally reduces and techniques to extend the range are necessary. The inducer bleed slot is a technique which has been adopted in turbocharger compressors. This approach was first reported by Fisher (1988) and was described as a Map Width Enhancement slot (MWE). The flow conditions in the MWE slot and impeller inlet duct were investigated with a view to developing an improved understanding of the flow mechanisms involved as the flow rate was reduced from choice to surge. Mean temperature and pressure measurements were recorded in the MWE passage, the main inducer duct to the impeller and the inlet duct upstream of the compressor. In addition the development of flow pulsations were monitored with pressure transducers in the MWE passage, the main inducer duct and the inlet duct, together with the application of flow visualisation techniques. The transient pressure measurements showed that low frequency flow pulsations developed in the MWE passage at high flow rates. As the flow rate was reduced the low frequency pulsations disappeared and flow reversal through the MWE passage developed. It was shown that flow reversal through the MWE passage commenced at flow rates close to the peak efficiency point for the compressor.

Commentary by Dr. Valentin Fuster
1996;():V001T01A081. doi:10.1115/96-GT-263.
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This paper reports a study of the combined effects of swirl and circumferential inlet flow distortion on the flow field of an axial flow fan stage. The study involves steady state measurements of the flow field at the rotor inlet, exit and the stator exit of the single stage axial flow fan subjected to circumferential inlet flow distortion and swirl. Flow field survey was done at two flow coefficients, namely, ϕ = 0.45 and ϕ = 0.285. The flow at the inlet to the rotor was measured using a three hole pressure probe and five hole pressure probes were used at the rotor and stator exits. The study indicated that at the design flow coefficient swirl had caused deterioration of the performance in addition to that caused by distortion. In addition, the attenuation of distortion was high in the presence of swirl.

Commentary by Dr. Valentin Fuster
1996;():V001T01A082. doi:10.1115/96-GT-300.
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Knowledge of flow and heat transfer in bearing chambers is important in the design of engine oil systems. In this paper a simplified model of the oil film on the housing of a bearing chamber is presented and the results compared with other workers’ measurements. An integral approach is used and the analysis includes the effects of surface friction, heat transfer, gravity and swirl of the oil at inlet. Two-dimensionality is assumed with variations in the axial direction being neglected. The model is expected to apply at high rotational speeds where “rimming” dominates with the oil flowing around the drum in a continuous film. A similar rimming flow regime occurs in a rotating, horizontal drum partially filled with liquid and the present model is also tested against data for this problem.

Some good qualitative and quantitative agreement with measurements has been found, but significant discrepancies and uncertainties remain. Overall the results of this first attempt to analyse the bearing chamber flow are considered very encouraging.

Commentary by Dr. Valentin Fuster
1996;():V001T01A083. doi:10.1115/96-GT-305.
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A new flow measurement technique is described which allows for the non-intrusive simultaneous measurement of flow velocity, density, and viscosity. The viscosity information can be used to derive the flow field temperature. The combination of the three measured variables and the perfect-gas law then leads to an estimate of the flow field pressure. Thus, the instantaneous state of a flow field can be completely described.

Three-State anemometry (3SA), a derivative of PIV, which uses a combination of three monodisperse sizes of styrene seeding particles is proposed. A marker seeding is chosen to follow the flow as closely as possible, while intermediate and large seeding populations provide two supplementary velocity fields, which are also dependent on fluid density and viscosity. A simplified particle motion equation, for turbomachinery applications, is then solved over the whole field to provide both density and viscosity data. The three velocity fields can be separated in a number of ways. The simplest and that proposed in this paper is to dye the different populations and look through interferometric filters at the region of interest.

The two critical aspects needed to enable the implementation of such a technique are a suitable selection of the diameters of the particle populations, and the separation of the velocity fields. There has been extensive work on the seeding particle behaviour which allows an estimate of the suitable particle diameters to be made. A technique is described in this paper to allow the separation of μm range particle velocity fields through fluorescence (separation through intensity also being possible). Some preliminary results by computer simulations of a 3SA image are also presented. The particle sizes chosen were 1 μm and 5 μm tested on the near-wake flow past a cylinder to investigate viscosity only, assuming uniform flow density. The accuracy of the technique, derived from simulations of swirling flows, is estimated as 0.5% RMS for velocity, 2% RMS for the density and viscosity, and 4% RMS for the temperature estimate.

Commentary by Dr. Valentin Fuster
1996;():V001T01A084. doi:10.1115/96-GT-308.
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Heat transfer in the low pressure turbine interdisc cavity of an aero gas turbine engine with a closed rotating outer-rim and forced radially outward jetflow directed along the downstream disc-cob front face was partially investigated by experiment and theory as part of an advanced cooling design concept study. Within the interdisc cavity, several metal temperature transient and steady state measurements in the circumferential direction as well as the 2-D axisymmetric plane at rotating speeds of 1500 rpm and 7000 rpm were made. The results are based on matching the measured metal and air temperature at both speed levels as well as the transient behavior between the two speed levels to those predicted by the 2-D axisymmetric transient thermal model. A qualitative description of the 3-D nature of the flow field is given with the aid of CFD studies. The results indicate that the skewed forced jetflow produces a stronger variation to the level of heat transfer at high rotating speed. The jetflow partially penetrates outward through the cavity providing enhanced free-disc forced convection heat transfer (approximately 25%) at high rotating speed to about 70% of the downstream disc radial length only. Toward the rim subcavities, the level of heat transfer drops considerably compared to that of a free-disc and heat transfer along the hot rotating rim and colder flange surfaces are described by flat plate natural convection. The jetflow exits the cavity at the bore of the upstream disc by turning forward within the cavity, substantially reducing the level of heat transfer along the diaphragm of this disc, and providing forced convection heat transfer on the cob whose level is higher than the rest of the upstream disc represented by natural convection. At low rotating speed, a dominating mixed convection mechanism is evident throughout the interdisc cavity with significantly lesser variation in heat transfer. A critical Gr/Re2 value of 0.02 was established as the minimum where both natural and forced convection are important. The resulting behavior in local heat transfer coefficient and Nusselt number along the hot bolted rim and the discs are compared with those of previous investigators looking at heat transfer in a rotating cavity.

Commentary by Dr. Valentin Fuster
1996;():V001T01A085. doi:10.1115/96-GT-309.
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In some engines, corotating gas–turbine discs are cooled by air introduced at the periphery of the system. The air enters through holes in a stationary peripheral casing and leaves through the rim seals between the casing and the discs. This paper describes a combined computational and experimental study of such a system for a range of flowrates and for rotational Reynolds numbers of up to Reϕ = 1.5 × 106. Computations are made using an axisymmetric elliptic solver, incorporating the Launder–Sharma low–Reynolds–number k–ε turbulence model, and velocity measurements are obtained using laser–Doppler anemometry.

The stationary peripheral casing creates a recirculation region: there is radial outflow in boundary layers on the discs and inflow in the core between the boundary layers. The radial extent of the recirculation region increases as the flow rate increases and as the rotational speed decreases. In the core, the radial and tangential components of velocity, Vr and Vϕ, are invariant in the axial direction, and the measured values of Vϕ conform to a Rankine–vortex flow. The agreement between the computed and measured velocities is not as good as that found for other rotating–disc systems, and deficiencies in the turbulence model are believed to be responsible.

Commentary by Dr. Valentin Fuster
1996;():V001T01A086. doi:10.1115/96-GT-352.
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Radial loads and direction of a centrifugal gas compressor containing a high specific speed mixed flow impeller and a single tongue volute, were determined both experimentally and computationally at both design and off-design conditions. The experimental methodology was developed in conjunction with a traditional ASME PTC-10 closed-loop test to determine radial load and direction. The experimental study is detailed in Part 1 of this paper (Moore and Flathers, 1996). The computational method employs a commercially available, fully-3D viscous code to analyze the impeller and the volute interaction. An uncoupled scheme was initially used where the impeller and volute were analyzed as separate models using a common vaneless diffuser geometry. The two calculations were then repeated until the boundary conditions at a chosen location in the common vaneless diffuser were nearly the same. Subsequently, a coupled scheme was used where the entire stage geometry was analyzed in one calculation, thus eliminating the need for manual iteration of the two independent calculations. In addition to radial load and direction information, this computational procedure also provided aerodynamic stage performance. The effect of impeller front face and rear face cavities was also quantified. The paper will discuss computational procedures, including grid generation and boundary conditions, as well as comparisons of the various computational schemes to experiment. The results of this study will show the limitations and benefits of Computational Fluid Dynamics (CFD) for determination of radial load, direction, and aerodynamic stage performance.

Topics: Gas compressors
Commentary by Dr. Valentin Fuster
1996;():V001T01A087. doi:10.1115/96-GT-353.
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This paper describes the development of a subscale single stage centrifugal compressor with a dimensionless specific speed (Ns) of 1.8, originally designed for full size appllcatioa as a high volume flow, low pressure ratio, gas booster compressor. The specific stage is noteworthy in that it provides a benchmark representing the performance potential of very high specific speed compressors of which limited information is found in open literature. Stage & component test performance characteristics are presented together with traverse results at the impeller exit. Traverse test results were compared with recent CFD computational predictions, for a exploratory analytical callbration of a very high specific speed impeller geometry.

The tested subscale (0.583) compressor essentially satisfied design performance expectations with an overall stage efficiency of 74% incinding, excessive exit casing losses. It was estimated that stage efficiency could be increased to 81% with exit casing losses halved.

Topics: Compressors
Commentary by Dr. Valentin Fuster
1996;():V001T01A088. doi:10.1115/96-GT-354.
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An analysis of the effectiveness of a three-dimensional shock loss model used in transonic compressor rotor design is presented. The model was used during the design of an aft-swept, transonic compressor rotor. The demonstrated performance of the swept rotor, in combination with numerical results, is used to determine the strengths and weaknesses of the model. The numerical results were obtained from a fully three-dimensional Navier-Stokes solver. The shock loss model was developed to account for the benefit gained with three-dimensional shock sweep. Comparisons with the experimental and numerical results demonstrated that shock loss reductions predicted by the model due to the swept shock induced by the swept leading edge of the rotor were exceeded. However, near the tip the loss model under-predicts the loss because the shock geometry assumed by the model remains swept in this region while the numerical results show a more normal shock orientation.

The design methods and the demonstrated performance of the swept rotor is also presented. Comparisons are made between the design intent and measured performance parameters. The aft-swept rotor was designed using an inviscid axisymmetric streamline curvature design system utilizing arbitrary airfoil blading geometry. The design goal specific flow rate was 214.7 kg/sec/m2 (43.98 lbm/sec/ft2), the design pressure ratio goal was 2.042, and the predicted design point efficiency was 94.0. The rotor tip sped was 457.2 m/sec (1500 ft/sec). The design flow rate was achieved while the pressure ratio fell short by 0.07. Efficiency was 3 points below prediction, though at a very high 91 percent. At this operating condition the stall margin was 11 percent.

Commentary by Dr. Valentin Fuster
1996;():V001T01A089. doi:10.1115/96-GT-357.
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The flow field in a preswirled cooling air supply to a turbine rotor has been investigated by means of CFD-simulations. Coefficients for system efficiency are derived. The influences of various geometrical parameters for different configurations have been correlated with the help of appropriate coefficients. For some of the most important geometrical parameters of the coverplate receiver design recommendations have been found. For the preswirl nozzles the potential of efficiency improvement by contour design is highlighted.

Commentary by Dr. Valentin Fuster
1996;():V001T01A090. doi:10.1115/96-GT-358.
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This paper describes a programme of work, largely experimental, which was undertaken with the objective of developing an improved blade profile for the low-pressure turbine in aero-engine applications.

Preliminary experiments were conducted using a novel technique. An existing cascade of datum blades was modified to enable the pressure distribution on the suction surface of one of the blades to be altered. Various means, such as shaped inserts, an adjustable flap at the trailing edge, and changing stagger were employed to change the geometry of the passage. These experiments provided boundary layer and lift data for a wide range of suction surface pressure distributions. The data was then used as a guide for the development of new blade profiles. The new blade profiles were then investigated in a low-speed cascade that included a set of moving bars upstream of the cascade of blades 10 simulate the effect of the incoming wakes from the previous blade row in a multistage turbine environment.

Results are presented for two improved profiles that are compared with a datum representative of current practice. The experimental results include loss measurements by wake traverse, surface pressure distributions, and boundary layer measurements. The cascades were operated over a Reynolds Number range from 0.7 × 105 to 4.0 × 105. The first profile is a “laminar flow” design that was intended to improve the efficiency at the same loading as the datum. The other is a more highly loaded blade profile intended to permit a reduction in blade numbers. The more highly loaded profile is the most promising candidate for inclusion in future designs. It enables blade numbers to be reduced by 20%, without incurring any efficiency penalty. The results also indicate that unsteady effects must be taken into consideration when selecting a blade profile for the low-pressure turbine.

Topics: Pressure , Turbines , Blades
Commentary by Dr. Valentin Fuster
1996;():V001T01A091. doi:10.1115/96-GT-359.
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The wakes behind turbine blade trailing edge are characterized by large scale periodic vortex patterns known as the von Karman vortex street. The failure of steady-state Navier-Stokes calculations in modeling wake flows appears to be mainly due to ignoring this type of flow instabilities. In an effort to contribute to a better understanding of the time varying wake flow characteristics behind turbine blades, VKI has performed large scale turbine cascade tests to obtain very detailed information about the steady and unsteady pressure distribution around the trailing edge of a nozzle guide vane. Tests are run at an outlet Mach number of M2,is,=0.4 and a Reynolds number of Rec = 2·106. The key to the high spatial resolution of the pressure distribution around the trailing edge is a rotatable trailing edge with an embedded miniature pressure transducer underneath the surface and a pressure slot opening of about 1.5° of the trailing edge circle. Signal processing allowed or differentiation between random and periodic pressure fluctuations. Ultra-short schlieren pictures help in understanding the physics behind the pressure distribution.

Commentary by Dr. Valentin Fuster
1996;():V001T01A092. doi:10.1115/96-GT-360.
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A nonlinear, two-dimensional, compressible dynamic model has been developed to study rotating stall/surge inception and development in high speed, multi-stage, axial flow compressors. The flow dynamics are represented by the unsteady Euler equations, solved in each interblade row gap and inlet and exit ducts as two-dimensional domains, and in each blade passage as a one-dimensional domain. The resulting equations are solved on a computational grid. The boundary conditions between domains are represented by ideal turning coupled with empirical loss and deviation correlations. Results are presented comparing model simulations to instability inception data of an eleven stage, high pressure ratio compressor operating at part-power, and the results analyzed in the context of linear modal analysis.

Commentary by Dr. Valentin Fuster
1996;():V001T01A093. doi:10.1115/96-GT-363.
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The paper describes the phenomenon of axial compressor fouling due to aerosols contained in the air. Key parameters having effect on the level of fouling are determined. A mathematical model of a progressive compressor fouling using the stage-by-stage calculation method is developed. Calculation results on the influence of fouling on the compressor performance are presented. A new index of sensitivity of axial compressors to fouling is suggested. The paper gives information about the Turbotect’s deposit cleaning method of compressor blading and the results of its application on an operating industrial gas turbine. Regular on line and off line washings of compressor flow path make it possible to maintain a high level of engine efficiency and output.

Topics: Compressors
Commentary by Dr. Valentin Fuster
1996;():V001T01A094. doi:10.1115/96-GT-364.
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An experimental and theoretical investigation has been conducted on rotordynamic forces due to non-axisymmetric turbine tip leakage effects. This paper presents an actuator disc model which describes the flow response to a finite clearance at the rotor tip. The model simplifies the flow field by assuming that the radially uniform flow splits into two streams as it goes through the rotor. The stream associated with the tip clearance, or the underturned stream, induces radially uniform unloading of the rest of the flow, called the bladed stream. Thus, a shear layer forms between the two streams. The fraction of each stream and the strength of shear layer between the two are found as functions of the turbine loading and flow parameters without resorting to empirical correlations. The results show that this model’s efficiency predictions compare favorably with the experimental data and predictions from various correlations. A companion paper builds on this analysis to yield a model of the 3-D disturbances around an offset turbine and to predict the subsequent cross forces.

Topics: Turbines , Blades , Leakage
Commentary by Dr. Valentin Fuster
1996;():V001T01A095. doi:10.1115/96-GT-365.
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This paper presents a radius scale actuator disc model which describes the flow response to a whirling/spinning rotor in an unshrouded turbine. At each azimuth, the upstream-downstream flow variables are matched by the results from a steady blade scale analysis presented in a companion paper, with allowance for mass storage in the stator-rotor Pregion. The new model can accurately predict the magnitude of both direct and cross excitation forces as well as their breakdown into work extraction and pressure effects. The trends versus the mean flow coefficient and interblade distance are predicted. While underpredicted, a trend versus mean rotor tip clearance height is also indicated. Thus, the new model captures the dominant physical effects caused by a whirling/spinning rotor in an unshrouded turbine.

Topics: Turbines , Leakage
Commentary by Dr. Valentin Fuster
1996;():V001T01A096. doi:10.1115/96-GT-370.
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This paper presents a new technique for precursor identification in high speed compressors. The technique is a pseudo-correlation integral method referred to as the correlation method. To provide a basis for comparison, the traveling wave energy technique, which has been used extensively to study pre-stall data, is also briefly presented and applied. The correlation method has a potential advantage over the traveling wave energy method because it uses a single sensor for detection. It also requires no predisposition about the expected behavior of the data to detect “changes” in the behavior of the compressor. Both methods are used in this study to identify stall procursive events in the pressure fluctuations measured from circumferential pressure transducers located at the front face of the compressor rig. The correlation method successfully identified stall formation or changes in the compressor dynamics from data captured from four different configurations of a NASA Lewis single stage high speed compressor while it was transitioned from stable operation into stall. This paper includes an exposition on the use of nonlinear methods to identify stall precursors, a description of the methodologies used for the study, information on the NASA high speed compressor rig and experimental data acquisition, and results from the four compressor configurations. The experimental results indicate that the correlation method provides ample warning of the onset of rotating stall at high speed, in some tests on the order of 2000 rotor revolutions. Complementary features of the correlation method and the traveling wave energy method are discussed, and suggestions for future developments are made.

Commentary by Dr. Valentin Fuster
1996;():V001T01A097. doi:10.1115/96-GT-372.
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CFD codes capable of utilizing multi-block grids provide capability to analyze the complete geometry of centrifugal compressors including, among others, multiple splitter rows, tip clearance, blunt trailing edges, fillets, and slots between moving and stationary surfaces. Attendant with this increased capability is potentially increased grid setup time and more computational overhead — CPU time and memory requirements — with the resultant increase in “wall clock” time to obtain a solution. If the increase in “difficulty” of obtaining a solution significantly improves the solution from that obtained by modeling the features of the tip clearance flow or the typical bluntness of a centrifugal compressor’s trailing edge, then the additional burden is worthwhile. However, if the additional information obtained is of marginal use then modeling of certain features of the geometry may provide reasonable solutions for designers to make comparative choices when pursuing a new design. In this spirit a sequence of grids were generated to study the relative importance of modeling versus detailed gridding of the tip gap and blunt trailing edge regions of the NASA large low speed centrifugal compressor for which there is considerable detailed internal laser anemometry data available for comparison.

The results indicate: 1) There is no significant difference in predicted tip clearance mass flow rate whether the tip gap is gridded or modeled. 2) Gridding rather than modeling the trailing edge results in better predictions of some flow details downstream of the impeller, but otherwise appears to offer no great benefits. 3) The pitchwise variation of absolute flow angle decreases rapidly up to 8% impeller radius ratio and much more slowly thereafter. Although some improvements in prediction of flow field details are realized as a result of analyzing the actual geometry there is no clear consensus that any of the grids investigated produced superior results in every case when compared to the measurements. However, if a multi-block code is available it should be used as it has the propensity for enabling better predictions than a single block code which requires modeling of certain geometry features. If a single block code must be used some guidance is offered for modeling those geometry features which can’t be directly gridded.

Commentary by Dr. Valentin Fuster
1996;():V001T01A098. doi:10.1115/96-GT-389.
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The flow in vaneless diffusers with large width-to-radius ratios is analyzed by using three-dimensional boundary-layer theory. The variations of the wall shear angle in the layer and the separation radius of the turbulent boundary layer versus various parameters are calculated and compared with experimental data. The effect of the separation point on the performance of vaneless diffusers and the mechanism of rotating stall are discussed. It is concluded that when the flow rate becomes very low, the reverse flow zone on the diffuser walls extends toward the entry region of diffusers. When the rotating jet-wake flow with varying total pressure passes through the reverse flow region near the impeller outlet, rotating stall is generated. The influences of the radius ratio on the reverse flow occurrence as well as on the overall performance are also discussed.

Commentary by Dr. Valentin Fuster
1996;():V001T01A099. doi:10.1115/96-GT-397.
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Probe blockage effects are presented for transonic flow through a calibration wind–tunnel as well as through a guide vane row in a three–stage model turbine.

Accurate experimental data from measurements in a transonic turbine are needed for the verification of CFD results. The accuracy of etatic pressure measurements in transonic turbine stages is severely affected by the pressure probe stem disturbing the surrounding flow–field. These disturbance effects are present during calibration procedures in wind–tunnels, as well as during measurements in–between turbomachinery blade rows. Therefore, the phenomenon associated with this blockage effect must be investigated for both procedures.

The influence of the blockage ratios on the static pressure readings of the four–hole wedge probe during the calibration procedure is investigated for two different wind–tunnels. The aim is to measure the blockage effects on the blade passage flow which are produced by a pneumatic pressure probe immersed in the flow between two adjacent blade rows. In order to measure these effects, two stator blades are instrumented with static pressure taps along the blade chord, as well as along the blade span. During the investigations, the radial and circumferential positions of the probes relative to the blade channel are varied. Pressure probe readings of two four–hole wedge probes with different stem diameters are compared as well as correlated to the static pressure readings of the stator blade pressure taps. The apparent deviations of the different readings are discussed.

Commentary by Dr. Valentin Fuster
1996;():V001T01A100. doi:10.1115/96-GT-400.
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A simple, robust numerical algorithm to localize moving boundary points and to interpolate unsteady solution variables across 2-D, arbitrarily overset computational grids is presented. Overset grids are allowed to move in time relative to each other. The intergrid boundary points are localized in terms of three grid points on the donor grid by a directional search algorithm. The parameters of the search algorithm give the interpolation weights at the localized boundary point. The method is independent of numerical solution algorithms and may easily be implemented on any 2-D, single block flow solver to make it a multi-block, zonal solver with arbitrarily overset computational grids. Computational results and comparisons with single grid solutions are presented for flows through a compressor cascade and over an airfoil undergoing a flapping motion. Excellent agreement is obtained against the single grid solutions.

Commentary by Dr. Valentin Fuster
1996;():V001T01A101. doi:10.1115/96-GT-404.
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It is difficult to measure flow patterns within rotating elements of a torque converter due to the complicated construction. Therefore, the numerical calculation is considered to be an effective tool to know the internal flow. Three-dimensional incompressible turbulent flow within a pump impeller of an automotive torque converter was analyzed numerically at three different speed ratios, 0.02, 0.4 and 0.8 under the same inlet boundary condition. The speed ratio was defined as the ratio of rotating speed of the turbine impeller to that of the pump. The governing equations using the k-ε model in the physical component tensor form were solved with a boundary-fitted coordinate system fixed on a rotating impeller. The solution algorithm was the SIMPLE method applied to the curvilinear coordinate system. The computed results were compared with those obtained experimentally by an oil film flow visualization technique for the pressure, suction, core and shell surfaces. Moreover, the results at three different speed ratios were examined in detail in order to clarify the behavior of secondary flow patterns. The computed results showed good agreement with the experimental results and clarified the behavior of the complicated flow patterns. The secondary flow patterns were strongly influenced by the correlation between the intensities of the Corinlis force (COF) and the centrifugal force due to the passage curvature in the meridional plane (CMF).

Commentary by Dr. Valentin Fuster
1996;():V001T01A102. doi:10.1115/96-GT-405.
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The goal of the paper is to describe wake parameters of wakes from turbine cascades in compressible flow especially in planes where the leading edge of the following blade row would be located.

The content of this paper is twofold: Data from experiments described in this paper and published data by different authors is used. Based on this data a theoretical approach is derived which describes the wake growth and the recovery of the velocity deficit. The theory is based on similarity assumptions. The derived equations depend on simple and readily available parameters such as overall losses, exit angle and Mach number. The paper follows the assumption that wakes can be described by similarity solutions. Conservation laws in an integral formulation are used to derive equations that describe the wake behavior downstream of the trailing edge of turbomachinery blades in compressible flows. The behavior of wakes will be described in terms of boundary layer dimensions (displacement and momentum thickness). In compressible turbine flows, the influence of the inviscid flow field is of great importance. In this paper an approach to take this influence into account when determining the behavior of the wake is presented.

Commentary by Dr. Valentin Fuster
1996;():V001T01A103. doi:10.1115/96-GT-409.
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Detailed measurements, with a two-component laser-Doppler velocimeter and a thermal anemometer were made near the suction surface leading edge of Controlled-Diffusion airfoils in cascade. The Reynolds number was near 700,000, Mach number equal to 0.25, and freestream turbulence was at 1.5% ahead of the cascade.

It was found that there was a localized region of high turbulence near the suction surface leading edge at high incidence. This turbulence amplification is thought to be due to the interaction of the free-shear layer with the freestream inlet turbulence. The presence of the local high turbulence affects the development of the short laminar separation bubble that forms very near the suction side leading edge of these blades. Calculations indicate that the local high levels of turbulence can cause rapid transition in the laminar bubble allowing it to reattach as a short “non-burst” type.

The high turbulence, which can reach point values greater than 25% at high incidence, is the reason that leading edge laminar separation bubbles can reattach in the high pressure gradient regions near the leading edge. Two variations for inlet turbulence intensity were measured for this cascade. The first is the variation of maximum inlet turbulence with respect to inlet-flow angle; and the second is the variation of leading edge turbulence with respect to upstream distance from the leading edge of the blades.

Commentary by Dr. Valentin Fuster
1996;():V001T01A104. doi:10.1115/96-GT-410.
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An experimental investigation of a separation bubble on a C4 leading edge plate at an incidence in a low turbulence free stream at six Reynolds numbers, is reported. The long separation bubble, formed at the leading edge, has a short laminar and transitional zone followed by a long turbulent zone. The increase in Reynolds number reduced the laminar and transitional part significantly, but its effect on the length of the separation bubble is marginal till the transition starts at the separation point. The peak intermittency factor, which occurs at the centre of the shear layer, follows the universal intermittency distribution curve. The spot production rate for the separated flows are several orders of magnitude higher than that for the attached boundary layers. The transition process is initiated by the amplification of the instability waves in the shear layer similar to the natural mode of transition. At high Reynolds numbers, the onset of transition is likely to take place at the separation point. At lower chord Reynolds numbers, the separation to onset Reynolds number and the spot production rate parameter are functions of the separation momentum thickness Reynolds number. The free stream turbulence intensity has a strong influence on the spot production rate. New correlations for transition in the leading edge separation bubbles are proposed based on all the available intermittency measurements in the leading edge separation bubbles.

Commentary by Dr. Valentin Fuster
1996;():V001T01A105. doi:10.1115/96-GT-411.
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This paper presents the computation of the flow around a controlled diffusion compressor cascade. Features associated with by-pass transition close to the leading edge — including laminar leading-edge separation — contribute significantly to the evolution of the boundary layer on the blade surface. Previous studies have demonstrated that conventional k-ε models, based on linear or non-linear Boussinesq stress-strain relations, are able to capture by-pass transition in simple shear, but are unable to resolve transitional features in complex strain, like the leading-edge separation bubble, which is of considerable influence to the suction-side flow at high inlet angle. Here, the k-ω turbulence model has been implemented in a nonstaggered, finite-volume based segregated Reynolds-Averaged Navier-Stokes solver. We demonstrate that this model, if properly sensitized to the generation of turbulence by irrotational strains, is capable of capturing the laminar leading-edge separation bubble. The real flow around the leading edge is laminar and the transition is only provoked on the reattachment region. Additional investigation of transition in a flat-plate boundary layer development has also produced reasonably promising results.

Commentary by Dr. Valentin Fuster
1996;():V001T01A106. doi:10.1115/96-GT-412.
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A new, specially-developed high-frequency-response pressure probe was used to measure the unsteady flow in the interaction region between the pump and the turbine in a hydrodynamic torque converter. In order to reduce the probe diameter, a single-hole, single-sensor cylindrical probe (⌀=1.33mm) was developed, to replace the standard multi-hole probe. The smaller the probe the higher the accuracy in unsteady flow. Therefore this is an improvement over three-hole probe. Three-hole probe measurements were simulated by recording data in three different angular positions. The time variable velocity vectors were determined using the probe’s calibration coefficients and the knowledge of the rotor positions (measured by angle-encoders) for every measurement value. During the data processing, a double ensemble averaging was carried out, taking into account the positions of the pump and the turbine.

Commentary by Dr. Valentin Fuster
1996;():V001T01A107. doi:10.1115/96-GT-415.
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Detailed measurements have been performed in a low pressure axial flow compressor stage to investigate the structure of the secondary flow field and the three-dimensional wake decay at different axial locations before and behind the rotor. The three dimensional flow field upstream and downstream of the rotor and on the centerline of the stator blade passage have been sampled periodically using a straight and a 90 degree triple-split fiber probe. Radial measurements at 39 radial stations were carried out at chosen axial positions in order to get the span-wise characteristics of the unsteady flow. Taking the experimental values of the unsteady flow velocities and turbulence properties, the effects of the rotor blade wake decay and secondary flow on the blade row spacing and stator passage flow at different operating conditions are discussed. For the normal operating point, the component of radial turbulent intensities in the leakage-flow mixing region is found to be much higher than the corresponding axial and tangential components. But for a higher value of the flow coefficient the relations are different.The results of the experiments show that triple-split fiber probes, straight and 90 degree measurements, combined with the ensemble average technique are a very useful method for the analysis of rotor flow in turbomachinery. Tip clearance vortex, secondary flow near the hub and radial flow in the wake, turbulent intensity and Reynolds stresses and also the decay of the rotor wakes can be obtained by this method.

Commentary by Dr. Valentin Fuster
1996;():V001T01A108. doi:10.1115/96-GT-418.
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This paper presents the application of a three-dimensional Navier-Stokes finite element code (NS3D) in the context of turbomachinery rotor-stator multistage interaction. A mixing-plane approach is used, in which boundary conditions at a common interface plane between adjacent blade rows are iteratively adjusted to yield a flow satisfying the continuity, momentum and energy conservation equations, in an average sense.

To further improve the solutions, a mesh adaptation technique then redistributes the mesh points of the structured grid within each component, according to an a posteriori error estimate based on the Hessian of the local flow solution. This matrix of second derivatives controls both the magnitude and direction of the required mesh movement at each node, which is then implemented using a spring analogy.

The methodology is demonstrated for the second stage compressor of the UTRC large-scale rotating rig at an rpm of 650 and a flow coefficient of 0.51, and for a two-stage compressor of a turboprop engine running at 45,000 rpm. The results compare well to the experimental data and illustrate the potential of the approach.

Commentary by Dr. Valentin Fuster
1996;():V001T01A109. doi:10.1115/96-GT-419.
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This paper summarizes the research on transonic turbine vane wake flows carried out in a Transonic Planar Cascade at the National Research Council of Canada between 1987 and 1995. The cascade used in the research is a large scale, continuously operating, inflow facility which was developed to study both flow phenomena and aerodynamics of turbine vanes. Research was initially directed at investigating the apparent redistribution of total temperature (energy) from the centre to the edge of the vane wake. This redistribution was found to be a real physical phenomenon that correlated extremely well with wake total pressure distributions indicating that the mechanism which redistributes the energy also has a direct effect on the losses associated with the vane wake. Both phenomena were found to be a function of Mach number in that the losses and the time-averaged total temperature difference between the centre and edge of the wake increased to a maximum as the Mach number approached 0.95 and then decreased by half as the Mach number was raised to 1.3.

Following Kurosaka et al (1987) conclusion that the redistribution of energy behind a circular cylinder was caused by the vortices which were shed from the trailing edge an additional experimental program was initiated to study the details of the unsteady vane wakes by the use of high speed schlieren photographs and high frequency response transducers to measure unsteady static and total pressures. This research has confirmed that over the range of Mach numbers, from low to high subsonic, a von Karman vortex street is shed continuously from the vanes. What this research has revealed is that as the transonic regime is traversed, the von Karman vortex street still occurs but only as one of a number of different and highly transient vortex shedding patterns. This breakdown of the stable von Karman vortex street is associated with the migration of the origin of the vortices from the trailing edge of the vane to the nodal point formed by the trailing edge shock waves and the confluence of the two trailing edge shear layers.

The cause of the redistribution of energy within the wake and of the high wake losses is the shedding of a continuous von Karman vortex street from the vane. In the subsonic flow regime the presence of a stable von Karman vortex street leads to high wake losses due to the depression of the base pressure while the redistribution of total temperature (energy) is caused by the combined pumping action of the vortices. The strength of the vortices and the above phenomena increase with increasing Mach number. As the transonic flow regime is encountered and the supersonic flow regime entered the coherent structures in the wake become unstable, less von Karman-like, and occur less frequently and the origin of the vortices migrates downstream. This leads to a significant elevation in base pressure and redistribution of energy, thereby implying that significant gains in engine stage efficiency can be realized by the destabilization or elimination of the von Karman vortex street from the vane wake.

Commentary by Dr. Valentin Fuster
1996;():V001T01A110. doi:10.1115/96-GT-420.
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The complex three-dimensional flow field in an axial-flow impeller incorporating high-Reynolds-number k-ε turbulence model is studied in this paper. The fully three-dimensional Reynolds averaged Navier-Stokes equations are solved. A computational procedure has been developed for predicting three-dimensional incompressible separated turbulent flows in the impeller. The SIMPLE-like algorithm is used. Convective terms are approximated with higher-order upstream-weighted approximations and a TVD-type MUSCL scheme. Physical covariant velocity components are selected as dependent solving variables. The non-orthogonal boundary-fitted coordinate system and collocated grid arrangement are also employed. Rhie and Chow’s momentum interpolation method is adopted to eliminate the non-physical pressure and velocity oscillations. Periodic boundary condition and moving wall boundary condition are considered to simulate truthfully the turbulent flow field in impeller.

Two types of axial-flow impellers are computed. The first one is designed by ordinary method and the other is a improved design that has been considered with eliminating flow separation and viscous vortex in the first design. The computed results show that the fully tree-dimensional turbulent flows computation can efficiently predict three-dimensional separated flows and viscous vortex in axial-flow impeller and vaneless clearance. Using the program, a designer can improve passage geometric design to enhance the performance of the fan.

Commentary by Dr. Valentin Fuster
1996;():V001T01A111. doi:10.1115/96-GT-421.
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The time-averaged flow in the wake of a model of a turbine blade was surveyed using a three-hole pressure probe; boundary layer traverses were also carried out using a flattened pilot probe. Loss coefficients were derived from the mass-weighted deficit in stagnation pressure. Results showing the progression of loss with streamwise distance along the surface and the wake of the model are presented. It was found that the loss generated in the wake comprised one-third of the profile loss when a well-developed vortex street was present in the wake. This proportion was reduced by increasing the thickness of the suction surface boundary layer, and by simulating the deviation that occurs in a real turbine blade. In both cases the strength of the vortex street was also shown to have been reduced.

Topics: Turbine blades , Wakes
Commentary by Dr. Valentin Fuster
1996;():V001T01A112. doi:10.1115/96-GT-422.
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A simple model was developed to simulate axial flow compressor performance deterioration due to blade erosion. The simulation at both design and off-design conditions is based on a mean line row by row model, which incorporates the effects of blade roughness and tip clearance. The results indicate that the increased roughness reduces the pressure ratio as well as the adiabatic efficiency of the compressor at all speeds with the largest influence at 100% speed. Increased tip clearance has a more pronounced effect on the compressor adiabatic efficiency and a lesser effect on the pressure ratio. According to the obtained results the loss in compressor performance due to erosion increases with increased blade loading.

Commentary by Dr. Valentin Fuster
1996;():V001T01A113. doi:10.1115/96-GT-440.
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A framework for analyzing the nonlinear dynamic behavior of flexibly-bladed turbomachines is presented. The analytical description is based on matching a two dimensional, incompressible flow field across a semi-actuator disk representation of a flexible rotor and a rigid stator. The aerodynamic loading on the rotor is derived using control volume formulations applied to discrete blade passages allowing consideration of finite interblade phase angles. Depending on operating parameters, the model exhibits behaviors classified as surge, rotating stall, and stall flutter which are qualitatively consistent with experimentally observed results.

The formulation provides a tractable, nonlinear, state-space description of the dynamics responsible for surge, rotating stall, flutter, and their interaction. An analysis is performed for system parameters representative of a fan in a modern high-bypass ratio aeroengine. The behavior of the operability limiting instability modes is examined using time simulations, eigenvalues plots and two-parameter stability diagrams.

Commentary by Dr. Valentin Fuster
1996;():V001T01A114. doi:10.1115/96-GT-441.
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This work deals with a series of experiments on the influence of the blade pitch on the rotating stall phenomenon in an industrial variable pitch, low-speed axial flow fan with low hub-to-tip ratio.

Two simple hot wires were used to detect the rotating stall. One in the absolute frame and the other in the relative frame rotating with the rotor. The rotating stall features were determined, ranging from the non-existence in the whole flow range with the lowest pitch tested to one and two flow cells with the greatest pitch.

Then, a triple hot wire, calibrated by a direct method, was used to measure the absolute flow field upstream and downstream from the rotor, before and during rotating stall for five distinct blade pitches. These measurements allow us to characterize different rotating stall structures.

To understand the phenomena better, some tests were carried out in the relative frame, with the probe rotating with the rotor. An intermediate blade pitch with a single rotating cell was selected and measurements were taken at three radial positions. Velocity maps for all these measurements are presented.

Topics: Wire , Axial flow
Commentary by Dr. Valentin Fuster
1996;():V001T01A115. doi:10.1115/96-GT-444.
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A new boundary layer transition model is presented which relates the velocity fluctuations near the wall to the formation of turbulent spots. A relationship for the near wall velocity frequency spectra is also established, which indicates an increasing bias towards low frequencies as the skin friction coefficient for the boundary layer decreases. This result suggests that the dependence of transition on the turbulent length scale is greatest at low freestream turbulence levels. This transition model is incorporated in a conventional boundary layer integral technique and is used to predict eight of the ERCOFTAC test cases. Three of these test cases are for nominally zero pressure gradient and the remaining five are for a pressure distribution typical of an aft loaded turbine blade. The model is demonstrated to predict the development of the boundary layer through transition reasonably accurately for all the test cases. The sensitivity of start of transition to the turbulent length scale at low freestream turbulence levels is also demonstrated.

Topics: Boundary layers
Commentary by Dr. Valentin Fuster
1996;():V001T01A116. doi:10.1115/96-GT-450.
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A new concept for annular diffuser design permits a layout of axial multi–diffusers with the potential to achieve the same pressure rise as single–annular diffusers, but over half the length. This is possible by using the full range of geometrical variability to gain a symmetrical loading for all flow channels. The optimal geometry with respect to the radial splitter position and the diffuser struts indicates strong dependence on the outflow of the leading turbine row, at the same time influencing this outflow by means of sharp wall angles. Moreover, an advantage can be derived from the additive pressure rise of the equalization process behind the diffuser. Therefore, it is no longer appropriate to treat the diffuser as an isolated component in the design process, but also to take into account the whole diffusion system including the turbine, diffuser and ducts. CFD turned out to be an adequate tool for this complex task.

Topics: Diffusers , Design , Turbines
Commentary by Dr. Valentin Fuster
1996;():V001T01A117. doi:10.1115/96-GT-466.
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The paper deals with results of a model aerodynamic research of flow in regions of a steam turbine stage with relatively long blades. The flow at design conditions attains transonic velocities and is investigated experimentally in transonic wind tunnel and numerically by means of 3D Euler equations for nonstationary flow. Profile cascades of five different sections of rotor blading were tested and detail flow patterns have been obtained. The experimental data are compared with results of calculations. The study of the operational flow conditions is performed. The results prove that both the tip and the root of the rotor blading are extremely loaded and they are the decisive elements for efficient operation. The middle sections have special profile forms and their aerodynamic design and test data are presented.

Commentary by Dr. Valentin Fuster
1996;():V001T01A118. doi:10.1115/96-GT-475.
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Results from scale-model experiments and industrial gas turbine tests show that strut vortex shedding in an annular exhaust diffuser can effectively be modified by adding tapered chord to the struts. The struts are bluff bodies at full-speed, no-load conditions, when inlet swirl is close to 60°. Data from wind tunnel tests show that wake Strouhal number is 0.47, larger than that expected for an isolated cylinder wake. This value of Strouhal number agrees with those measured in full-scale exhaust diffusers. Wind tunnel tests showed that a strut with tapered chord most effectively reduced wake amplitudes and shifted shedding frequency. The tapered strut was also effective in reducing shedding amplitude in a scale-model diffuser. Finally, gas turbine tests employing a tapered strut showed significant reductions in unsteady pressure and noise. A major benefit of strut taper is a reduction of noise by uncoupling of vortex shedding from acoustic resonant response.

Commentary by Dr. Valentin Fuster
1996;():V001T01A119. doi:10.1115/96-GT-483.
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Mid-span losses in the NRC transonic turbine cascade peak at an exit Mach number (M2) of ∼1.0 and then decrease by ∼40% as M2 is increased to the design value of 1.16. Since recent experimental results suggest that the decrease may be related to a reduction in the intensity of trailing edge vortex shedding, both steady and unsteady quasi-3D Navier-Stokes simulations have been performed with a highly refined (unstructured) grid to determine the role of shedding. Predicted shedding frequencies are in good agreement with experiment, indicating the blade boundary layers and trailing edge separated free shear layers have been modelled satisfactorily, but the agreement for base pressures is relatively poor, probably due largely to false entropy created downstream of the trailing edge by numerical dissipation. The results emphasize the importance of accounting for the effect of vortex shedding on base pressure and loss.

Commentary by Dr. Valentin Fuster
1996;():V001T01A120. doi:10.1115/96-GT-484.
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Compressor stall was simulated in the Low Speed Cascade Wind Tunnel at the Turbopropulsion Laboratory. The test blades were of controlled-diffusion design with a solidity of 1.67, and stalling occurred at 10 degrees of incidence above the design inlet air angle. All measurements were taken at a flow Reynolds number, based on chord length, of 700 000.

Laser-sheet flow visualization techniques showed that the stalling process was unsteady and occurred over the whole cascade. Detailed laser-Doppler-velocimetry measurements over the suction side of the blades showed regions of continuous and intermittent reverse flow. The measurements of the continuous reverse flow region at the leading edge were the first data of their kind in the leading edge separation bubble. The regions of intermittent reverse flow, measured with laser Doppler velocimeter, corresponded to the flow visualization studies. Blade surface pressure measurements showed a decrease in normal force on the blade as would be expected at stall. Data is presented in a form which characterizes the unsteady positive and negative velocities about their mean, for both the continuous reverse flow regions and the intermittent reverse flow regions.

Commentary by Dr. Valentin Fuster
1996;():V001T01A121. doi:10.1115/96-GT-486.
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The development of the unsteady suction side boundary layer of a highly loaded LP turbine blade has been investigated in a rectilinear cascade experiment. Upstream rotor wakes were simulated with a moving-bar wake generator. A variety of cases with different wake-passing frequencies, different wake strength and different Reynolds-numbers were tested. Boundary layer surveys have been obtained with a single hot-wire probe. Wall shear stress has been investigated with surface-mounted hot-film gauges. Losses have been measured.

The suction surface boundary layer development of a modern highly loaded LP turbine blade is shown to be dominated by effects associated with unsteady wake-passing. Whereas without wakes the boundary layer features a large separation bubble at a typical cruise Reynolds-number, the bubble was largely suppressed if subjected to unsteady wake-passing at a typical frequency and wake strength. Transitional patches and becalmed regions, induced by the wake, dominated the boundary layer development. The becalmed regions inhibited transition and separation and are shown to reduce the loss of the wake-affected boundary layer.

An optimum wake-passing frequency exists at cruise Reynolds-numbers. For a selected wake-passing frequency and wake-strength, the profile loss is almost independent of Reynolds-number. This demonstrates a potential to design highly loaded LP turbine profiles without suffering large losses at low Reynolds-numbers.

Commentary by Dr. Valentin Fuster
1996;():V001T01A122. doi:10.1115/96-GT-487.
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Laminar separation bubbles are commonly observed on turbomachinery blades and therefore require effective methods for their prediction. Therefore, a newly developed transition model by Gostelow et al. (1995) is incorporated into an upwind-biased Navier-Stokes code to simulate laminar-turbulent transition in the boundary layer. A study of the influence of the two adjustable parameters of the model, the transition onset location and the spot generation rate, is conducted and it is found that it can predict laminar separation bubbles, measured on a NACA 0012 airfoil. Additional results are presented for separation bubbles in an annular compressor cascade.

Commentary by Dr. Valentin Fuster
1996;():V001T01A123. doi:10.1115/96-GT-489.
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Measurements are presented of the calmed region behind triggered wave packets and turbulent spots under a controlled diffusion adverse pressure gradient in a wind tunnel. Similar measurements are also presented from the stator blades of an axial flow compressor, where turbulent spots are induced by the passing of rotor wakes. The purpose is to gain an appreciation of turbulent spot behavior under a strong adverse pressure gradient as a foundation for the more accurate modeling of spots and their environment in predictions of transitional boundary layer flows. Under an adverse pressure gradient the calmed region behind the spot is extensive; its interaction with the surrounding turbulent layer is complex and is dependent on whether the surrounding natural boundary layer is laminar or turbulent. Some insights are gleaned concerning the behavior of the calmed region which will subsequently be used in attempts to model the calmed region. Although these fundamental investigations of the calmed region have been extensive much remains to be understood.

Topics: Turbulence
Commentary by Dr. Valentin Fuster
1996;():V001T01A124. doi:10.1115/96-GT-494.
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The interaction of wakes shed by a moving bladerow with a downstream bladerow causes unsteady flow. The meaning of the freestream stagnation pressure and stagnation enthalpy in these circumstances has been examined using simple analyses, measurements and CFD. The unsteady flow in question arises from the behaviour of the wakes as so-called negative-jets. The interactions of the negative-jets with the downstream blades lead to fluctuations in static pressure which in turn generate fluctuations in the stagnation pressure and stagnation enthalpy. It is shown that the fluctuations of the stagnation quantities created by unsteady effects within the bladerow are far greater than those within the incoming wake. The time-mean exit profiles of the stagnation pressure and stagnation enthalpy are affected by these large fluctuations. This phenomenon of energy separation is much more significant than the distortion of the time-mean exit profiles that is caused directly by the cross-passage transport associated with the negative-jet, as described by Kerrebrock and Mikolajczak. Finally, it is shown that if only time-averaged values of loss are required across a bladerow, it is nevertheless sufficient to determine the time-mean exit stagnation pressure.

Topics: Wakes , Turbines , Blades
Commentary by Dr. Valentin Fuster
1996;():V001T01A125. doi:10.1115/96-GT-505.
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Within the scope of a compressor research program on endwall flows a highly loaded linear compressor cascade of NACA 65 profiles was investigated.

Detailed study of the three-dimensional flow field was carried out for three different cases of tip-clearance including zero clearance with a variation of blade loading, blade height, and inlet boundary layer thickness. Using a small five hole probe measurements have been performed inside and downstream of the blade passage. Additional information about the formation and development of the passage and tip-clearance vortices is obtained from static pressure tappings at midspan and on the endwall, and a surface flow visualization technique.

The cascade performance is presented in terms of turning angle and loss coefficient. It is found that in addition to the frequently investigated effects of tip-clearance and blade loading, the displacement thickness of the inlet boundary layer has a significant influence on the radial distribution of the losses and the outlet angle. However, the overall loss behavior remains almost unaffected by the inlet boundary layer. Only for low values of the aspect ratio around one, zero tip-clearance, and high values of incidence the influence of the aspect ratio becomes important. In this case the endwall flow regions from both blade ends are linked.

The experimental results provide an extended basis for the improvement of the known correlations on aerodynamic losses and flow angle deviation as well as for the validation of 3D-Navier-Stokes calculations. An improved approach for loss correlations is presented in this paper.

Commentary by Dr. Valentin Fuster
1996;():V001T01A126. doi:10.1115/96-GT-506.
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This paper describes the structure of the tip clearance flow in a low speed isolated compressor rotor. Pneumatic cobra probes are radially traversed upstream and downstream of the blade row and the time averaged total pressure losses across the blade row calculated. The increase in pressure losses due to the tip clearance flow is clearly seen.

The nature of the tip losses is investigated further using a unique 3D laser transit anemometer to measure velocities and turbulence levels. A 3D representation of the resulting flow field is then constructed using the experimentally measured velocity vectors. With the aid of ‘stream particles’ released into this flow field a vortex structure is then visualised. A section through the path of this vortex assists in showing its development through the blade row.

Due to the co-location of this vortex and the total pressure losses in the passage, it is this vortex which is believed to be responsible for the excess total pressure losses in the tip region.

Commentary by Dr. Valentin Fuster
1996;():V001T01A127. doi:10.1115/96-GT-507.
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Results of numerical simulations conducted for a high pressure compressor rotor with two different levels of tip clearance are presented. A three-dimensional, steady, Reynolds-Averaged Navier-Stokes code was utilized to perform the computations. The simulations were executed over a range of flow coefficients by specifying different axisymmetric radial profiles in static pressure downstream of the rotor. In this manner, the effect of the downstream stator row was approximated using a simple, circumferentially averaged, radial pressure profile as the boundary condition behind the rotor. The back pressure profiles utilized were those deduced from inviscid flow computations for two different stator designs: (1) a conventional radial stator, and (2) a three-dimensional “bowed” stator. Results of the rotor simulations with nominal tip clearance show that the boundary condition induced by the bowed stator causes a 2% decrease in rotor pressure rise capability, and a 9% increase in rotor loss as compared with the conventional stator. In addition, as the tip clearance is increased to twice the nominal value, the rotor loss grows at a rate 25% higher for the rotor subjected to the bowed stator pressure profile. Accompanying this is a dramatic reduction in rotor speedline slope and pressure rise capability. Analysis of the simulations shows these effects to be linked to the response of the rotor tip clearance vortex to the exit pressure profile set by the downstream stator. These results indicate the need to accurately model the effects of the radial variation in static pressure imposed by the downstream airfoil rows.

Topics: Pressure , Rotors , Stators
Commentary by Dr. Valentin Fuster
1996;():V001T01A128. doi:10.1115/96-GT-508.
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The three-dimensional viscous flow characteristics and the complex vortex system downstream of the rotor of an industrial exial fan have been determined by an experimental investigation using hot-wire anemometer. Single-wire slanted and straight type probes have been rotated about the probe axis using a computer controlled stepper motor. Measurements have been taken at four planes behind the blade trailing edge. The results show the characteristics of the relative flow as velocity components, secondary flow and kinetic energy defect. Turbulence intensity and Reynolds stress components in the leakage vortex area are also presented. The evolution of the leakage vortex flow during the decay process has also been evaluated in terms of dimension, position and intensity.

Commentary by Dr. Valentin Fuster
1996;():V001T01A129. doi:10.1115/96-GT-509.
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The problem of determining the optimal configuration of a cooled gas-turbine blade is approached by an entropy minimization technique proposed in previous works by the same authors. The present paper describes the application of the same line of thought to a more complex (and realistic) pseudo-optimization procedure, in which the objective function is again the global entropy generation rate, but two integral constraints are added to the original formulation: the maximum blade temperature (weak constraint) and the overall enthalpy drop of the working fluid in the blade passage (strong constraint). The discontinuous optimization procedure is presented here in an application which resembles a trial-and-error technique, but can be rigorously and formally described and implemented [12].

As a “zero configuration”, a realistic 2-D geometry is considered, and the thermo-fluiddynamic field around it is computed via a standard finite-element code. Then, the entropy generation rates in the blade/fluid system are calculated, and the value of the overall enthalpy drop of the gas as well as the value and location of the maximum blade temperature are recorded. Keeping all other parameters fixed (in particular, maintaining the same cooling air flowrate), the geometry of the blade is slightly “perturbed”, by introducing arbitrary modifications in the blade profile, the number and location of cooling holes, etc. Again, the velocity and temperature fields are computed, and inlet conditions are tuned so that the overall enthalpy drop remains approximately constant and the blade maximum temperature does not exceed a certain assigned value.

An “optimal” configuration is found, which is affected by the minimal entropy generation rate, while abiding to the imposed constraints.

The procedure is demonstrated on a realistic blade profile, and is shown to produce a better performing cascade, at least in this 2-D simulation. The extension to 3-D problems is — in principle — straightforward (but see Section 3 for further comments).

Commentary by Dr. Valentin Fuster
1996;():V001T01A130. doi:10.1115/96-GT-512.
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This paper first proposes a new control strategy, stabilization on the sense of practical stability, to actively stabilize the axial compression systems regardless their instability modes. Then, the theory of practical stability is applied to analyze the practical stability of the pure surge system which is a compression system without any nonaxisymmetric disturbances that grow into rotating stall. This analysis reveals that the upper limit of the amplitude in which the system surges is determined by the nonlinearity of the compressor characteristic. Thus, a nonlinear controller is designed to control the trajectory of the pure surge system by modifying the nonlinear terms in the expression for the compressor characteristic. Numerical simulation shows that the nonlinear controller can effectively shrink the size of the trajectory of the pure surge system, and therefore stabilize the system in the sense of practical stability.

Topics: Stability , Surges
Commentary by Dr. Valentin Fuster
1996;():V001T01A131. doi:10.1115/96-GT-514.
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The design and testing of multi-channel data transmission systems for a transient turbine test rotor are presented. Two multi-channel systems were required for the conditioning and transmission of the electrical signals from thin film gauges and un-packaged embedded pressure transducers which are to be used for the measurement of rotor blade surface heat transfer and pressure respectively. All measurements will be taken from the same rotor at design Mach and Reyoolds numbers. Signal conditioning and amplification will be performed in the rotating frame to solve the problems associated with accurate sigoal transmission. Also, the use of a multi-chanoel array of light emitting diodes for data transmission is described which allows accurate signal reconstruction. Both systems have been initially stress screened and are shown to have an adequate bandwidth for resolving wake blade interactions.

Commentary by Dr. Valentin Fuster
1996;():V001T01A132. doi:10.1115/96-GT-543.
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The results from the area traverse measurements of the unsteady total temperature using a high response aspirating probe downstream of the second stator of a three stage axial flow compressor are presented. The measurements were conducted at the peak efficiency operating point. The unsteady total temperature data is resolved into deterministic and unresolved components. Hub and casing regions have high levels of unsteadiness and consequently high levels of mixing. These regions have significant levels of shaft resolved and unresolved unsteadiness. Comparisons are made between the total temperature and the total pressure data to examine the rotor 2 wake characteristics and the temporal variation of the stator exit flow. Isentropic efficiency calculations at the midpitch location show that there is about a 4% change in the algebraically averaged efficiency across the blades of the second rotor and if all the rotor 2 blades were behaving as a “best” blade, the improvement in efficiency would be about 1.3%. An attempt is made to create a composite flow field picture by correlating the unsteady velocity data with temperature and pressure data.

Commentary by Dr. Valentin Fuster
1996;():V001T01A133. doi:10.1115/96-GT-544.
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An existing three dimensional Navier-Stokes flow solver with an explicit Runge-Kutta algorithm and a low Reynolds number k-ε turbulence model has been modified in order to simulate turbomachinery flows in a more efficient manner. The solver has been made to converge more rapidly through use of the mutligrid technique. Stability problems associated with use of multigrid in conjunction with two equation turbulence models are addressed and techniques to alleviate instability are investigated. Validation for the new code was performed with a transonic turbine cascade tested by Perdichizzi. In the fully three dimensional turbulent cascade, real convergence (i.e. CPU time) was improved nearly two times the original code. Robustness was enhanced with the full multigrid initialization procedure. The same test case was then used to perform a series of simulations that investigated the effect of different exit Mach numbers on secondary flow features. This permitted an in depth study into the mechanisms of secondary flow formation and secondary losses at high Mach numbers. In this cascade, it was found that secondary losses and secondary flow deviation, which are fairly constant in incompressible flows with similar geometries, underwent a large reduction in the compressible flow range. The structure of the trailing edge shock system and the reduced endwall boundary layer at supersonic exit conditions were shown to be very significant in reducing the amount of secondary flow and losses.

Commentary by Dr. Valentin Fuster
1996;():V001T01A134. doi:10.1115/96-GT-546.
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A 3D viscous solver has been used to model the flow in the stator of a highly loaded single-stage transonic fan. The fan has a very high level of aerodynamic loading at the hub, which results in a severe hub endwall stall. Prediction of the flow at the 100% speed, peak efficiency condition has been carried out and comparisons are made with experiment, including stator exit traverses and fixed blade surface pressure tappings and flow visualisation. Comparisons are also made with an analysis of the rotor and stator rows using the DRA S1-S2 method.

The 3D predictions show good qualitative agreement with measurements in all regions of the flow field. Quantitatively the flow away from the hub region agreed the best. The general trends of the severe hub endwall stail were predicted, although the shape and size did not match experiment exactly. The S1-S2 system was unable to predict the hub endwall stall, since it arises from fully 3D flow effects.

Commentary by Dr. Valentin Fuster
1996;():V001T01A135. doi:10.1115/96-GT-547.
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The effects of circumferential distortions in inlet total pressure on the flow field in a low-aspect-ratio, high-speed, high-pressure-ratio, transonic compressor rotor are investigated in this paper. The flow field was studied experimentally and numerically with and without inlet total pressure distortion. Total pressure distortion was created by screens mounted upstream from the rotor inlet. Circumferential distortions of 8 periods per revolution were investigated at two different rotor speeds. The unsteady blade surface pressures were measured with miniature pressure transducers mounted in the blade. The flow fields with and without inlet total pressure distortion were analyzed numerically by solving steady and unsteady forms of the Reynolds-averaged Navier-Stokes equations. Steady three-dimensional viscous flow calculations were performed for the flow without inlet distortion while unsteady three-dimensional viscous flow calculations were used for the flow with inlet distortion. For the time-accurate calculation, circumferential and radial variations of the inlet total pressure were used as a time-dependent inflow boundary condition. A second-order implicit scheme was used for the time integration. The experimental measurements and the numerical analysis are highly complementary for this study because of the extreme complexity of the flow field. The current investigation shows that inlet flow distortions travel through the rotor blade passage and are convected into the following stator. At a high rotor speed where the flow is transonic, the passage shock was found to oscillate by as much as 20% of the blade chord, and very strong interactions between the unsteady passage shock and the blade boundary layer were observed. This interaction increases the effective blockage of the passage, resulting in an increased aerodynamic loss and a reduced stall margin. The strong interaction between the passage shock and the blade boundary layer increases the peak aerodynamic loss by about one percent.

Commentary by Dr. Valentin Fuster
1996;():V001T01A136. doi:10.1115/96-GT-549.
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Early in 1995, Westinghouse decided to seize an opportunity to speed up the implementation of a design upgrade for the 501FA Row 2 vane segment. The planned upgrade of this vane segment had been in the conceptual design stage for a few months. Normally, this type of modification would be planned and implemented over many months. In order to achieve the desired performance goals and meet the customer’s delivery requirements this project would have to move very quickly. The project time frame required that from the conceptual drawing to start of production be no more than eight weeks.

To achieve the accelerated schedule, Westinghouse decided to team with the part machining suppliers to develop a game plan that allowed for true concurrent engineering. The resulting plan was to complete the part design at the same time that all the machining tooling and manufacturing process development was being completed. As a result of combining the resources of the turbine manufacturer with those of the machining vendors, it was possible to implement a complicated part upgrade in a matter of weeks not months.

By utilizing 3-D computer models to define the part configuration as well as to build the machining tools and develop the machining process, the team was able to meet the challenge. The final design was optimized for performance as well as ease of manufacturing. This paper describes the triumphs as well as some of the problems that the team encountered along the way to delivering the final engine hardware.

Topics: Design
Commentary by Dr. Valentin Fuster
1996;():V001T01A137. doi:10.1115/96-GT-550.
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A two-dimensional (θ,z) Navier-Stokes solver for multi-port wave rotor flow simulation is described. The finite-volume form of the unsteady thin-layer Navier-Stokes equations arc integrated in time on multi-block grids that represent the stationary inlet and outlet ports and the moving rotor passages of the wave rotor. Computed results are compared with three-port wave rotor experimental data. The model is applied to predict the performance of a planned four-port wave rotor experiment. Two-dimensional flow features that reduce machine performance and influence rotor blade and duct wall thermal loads are identified. The performance impact of rounding the inlet port wall, to inhibit separation during passage gradual opening, is assessed.

Commentary by Dr. Valentin Fuster

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