bluuebunny/crossref_metadata_2025 huggingface dataset:
Experimental Study of the Swirling Flow in the Volute of a Centrifugal Pump
Authors: Elholm T. , Ayder E. , Van den Braembussche R.
Published: June 2008

Detailed three-dimensional velocity distributions, corresponding to design and off-design operation, were measured in two different circumferential cross sections of a volute by means of LDV. It is shown that the swirl has a forced vortex type velocity distribution and that the location of the swirl center changes with mass flow. The throughflow velocity distribution is primarily defined by the conservation of angular momentum. A strong interaction between the throughflow and swirl velocity is observed. Flow visualization in the tongue region reveals a reversal of the velocity at the volute inlet with increasing mass flow. The pressure drop between volute outlet and inlet at low mass flow pushes extra fluid through the tongue gap and increases the mass flow in the volute. The abrupt pressure rise at high mass flow results in local return flow perturbing the flow in the outlet pipe.
The Effect of Tip Leakage Flow on Part-Load Performance of a Mixed-Flow Pump Impeller
Authors: Goto A.
Published: June 2008

The flow phenomena around the positive slope region of the head-flow characteristic were investigated experimentally on a mixed-flow pump impeller at various tip clearances for both shrouded and unshrouded cases. A positively sloped head-flow characteristic (abrupt decrease in pressure head) was caused by the onset of extensive flow separation in the impeller at the casing-suction surface corner. The corner separation in unshrouded cases appeared at a much lower flow rate than the shrouded case due to the favorable effect of the tip leakage flow, which displaced the wake region away from the corner. The interaction between the tip leakage flows and secondary flows and the formation of the wake regions in shrouded and unshrouded cases were explained based on experimental observation and computations by the Dawes three-dimensional Navier–Stokes code. In the shrouded case, the flow rate at which an abrupt decrease in pressure head appeared was lowered substantially by introducing a leakage flow through a slit made between the shroud and the blade tip. Inlet recirculation was triggered by the corner separation and developed more gradually for larger tip clearances. Both the increased loss, due to the extensive flow separation, and the decreased Euler’s head, due to the abrupt change in flow pattern caused by the inlet recirculation, were responsible for the generation of positively sloped head-flow characteristic in the unshrouded case when the tip clearance was small, while the increased loss was the primary factor in the shrouded case.
A Comparison of Two Methods for Utilizing Steam Turbine Exhaust Hood Flow Field Data
Authors: Tindell R. H. , Alston T. M.
Published: June 2008

This article describes the effects of two methods for representing the nonuniform distribution of flow properties across a steam turbine discharge annulus, on the hood loss coefficient. One method uses a mass-weighted integration of the property across the station, while the other is based on a mass-derived representative value of the property. The former has the potential for very high accuracy provided a sufficient number of points are integrated. The latter, while less accurate, is easier to apply and therefore more commonly used. The analytical modeling includes a simplistic step profile of pressure across the annulus, as well as a three-dimensional exhaust hood, flow-field simulation calculated using a Navier–Stokes code. Results show that significant errors can occur in the hood loss coefficient with the mass-derived approach. Although the analysis centers on hood loss coefficient as the performance parameter whose sensitivity is being monitored, the results highlight the pitfalls of improper application of measured data for any internal flow system.
The Simulation of Three-Dimensional Viscous Flow in Turbomachinery Geometries Using a Solution-Adaptive Unstructured Mesh Methodology
Authors: Dawes W. N.
Published: June 2008

This paper presents a numerical method for the simulation of flow in turbomachinery blade rows using a solution-adaptive mesh methodology. The fully three-dimensional, compressible, Reynolds-averaged Navier–Stokes equations with k–ε turbulence modeling (and low Reynolds number damping terms) are solved on an unstructured mesh formed from tetrahedral finite volumes. At stages in the solution, mesh refinement is carried out based on flagging cell faces with either a fractional variation of a chosen variable (like Mach number) greater than a given threshold or with a mean value of the chosen variable within a given range. Several solutions are presented, including that for the highly three-dimensional flow associated with the corner stall and secondary flow in a transonic compressor cascade, to demonstrate the potential of the new method.
A Numerical Procedure of Three-Dimensional Design Problem in Turbomachinery
Authors: Xu J. Z. , Gu C. W.
Published: June 2008

A numerical method for solving the three-dimensional aerothermodynamic design problem with some type of Mach number distributions on the blade surfaces is presented. In the usual aerothermodynamic design of a turbomachine, the three-dimensional coordinates of the blade are attained through stacking of the cascade profiles and may not ensure the desired velocity distribution. To avoid this problem, the present method will give new coordinates of the blade according to the required Mach number distribution. The method is based on the pseudostream function formulation and the treatment of the boundary conditions in the design problem is given. The numerical results show that the method is simple and useful in design.
Design Method for Subsonic and Transonic Cascade With Prescribed Mach Number Distribution
Authors: Le´onard O. , Van den Braembussche R. A.
Published: June 2008

A iterative procedure for blade design, using a time marching procedure to solve the unsteady Euler equations in the blade-to-blade plane, is presented. A flow solver, which performs the analysis of the flow field for a given geometry, is transformed into a design method. This is done by replacing the classical slip condition (no normal velocity component) by other boundary conditions, in such a way that the required pressure or Mach number distribution may be imposed directly on the blade. The unknowns are calculated on the blade wall using the so-called compatibility relations. Since the blade shape is not compatible with the required pressure distribution, a nonzero velocity component normal to the blade wall evolves from the new flow calculation. The blade geometry is then modified by resetting the wall parallel to the new flow field, using a transpiration technique, and the procedure is repeated until the calculated pressure distribution has converged to the required one. Examples for both subsonic and transonic flows are presented and show a rapid convergence to the geometry required for the desired Mach number distribution. An important advantage of the present method is the possibility to use the same code for the design and the analysis of a blade.
Effects of Simulated Rotation on Tip Leakage in a Planar Cascade of Turbine Blades: Part I—Tip Gap Flow
Authors: Yaras M. I. , Sjolander S. A.
Published: June 2008

The paper presents further results from a continuing study on tip leakage in axial turbines. Rotation has been simulated in a linear cascade test section by using a moving-belt tip wall. Measurements were made inside the tip gap with a three-hole pressure probe for a clearance size of 3.8 percent of the blade chord. Two wall speeds are considered and the results are compared with the case of no rotation. As in other experiments, significant reduction in the gap mass flow rate is observed due to the relative motion. The detailed nature of the measurements allows the dominant physical mechanism by which wall motion affects the tip gap flow to be identified. Based on the experimental observations, an earlier model for predicting the tip gap flow field is extended to the case of relative wall motion. Part II of the paper examines the effect of the relative motion on the downstream flow field and the blade loading.
On the Propagation of Viscous Wakes and Potential Flow in Axial-Turbine Cascades
Authors: Korakianitis T.
Published: June 2008

This paper investigates the propagation of pressure disturbances due to potential-flow interaction and viscous-wake interaction from upstream blade rows in axial-turbine-blade rotor cascades. Results are obtained by modeling the effects of the upstream stator viscous wake and potential-flow fields as incoming disturbances on the downstream rotor flow field, where the computations are performed. A computer program is used to calculate the unsteady rotor flow fields. The amplitudes for the rotor inlet distortions due to the two types of interaction are based on a review of available experimental and computational data. We study the propagation of the isolated potential-flow interaction (no viscous-wake interaction), of the isolated viscous wake interaction (no potential-flow interaction), and of the combination of interactions. The discussion uses as example a lightly loaded cascade for a stator-to-rotor-pitch ratio R = 2. We examine the relative magnitudes of the unsteady forces for two different stator-exit angles. We also explain the expected differences when the stator-to-rotor pitch ratio is decreased (to R = 1) and increased (to R = 4). We offer new and previously unpublished explanations of the mechanisms of generation of unsteady forces on the rotor blades. The potential flow field of the rotor cuts into the potential flow field of the stator. After the potential-flow disturbance from the stator is cut into a rotor cascade, it propagates into the relative flow field of the rotor passage as a potential-flow disturbance superimposed on the rotor-relative flow. The potential flow field of the rotor near the leading edge and the leading edge itself cut into the wake and generate two counterrotating vortical patterns flanking the wake centerline in the passage. The vortical pattern upstream of the wake centerline generates an increase in the local pressure (and in the forces acting on the sides of the passage). The vortical pattern downstream of the wake centerline generates a decrease in the local pressure (and in the forces acting on the sides of the passage). The resulting unsteady forces on the blades are generated by the combined (additive) interaction of the two disturbances.
Effects of Stator Wakes and Spanwise Nonuniform Inlet Conditions on the Rotor Flow of an Axial Turbine Stage
Authors: Zeschky J. , Gallus H. E.
Published: June 2008

Detailed measurements have been performed in a subsonic, axial-flow turbine stage to investigate the structure of the secondary flow field and the loss generation. The data include the static pressure distribution on the rotor blade passage surfaces and radial-circumferential measurements of the rotor exit flow field using three-dimensional hot-wire and pneumatic probes. The flow field at the rotor outlet is derived from unsteady hot-wire measurements with high temporal and spatial resolution. The paper presents the formation of the tip clearance vortex and the passage vortices, which are strongly influenced by the spanwise nonuniform stator outlet flow. Taking the experimental values for the unsteady flow velocities and turbulence properties, the effect of the periodic stator wakes on the rotor flow is discussed.
Wind Tunnel Wall Effects in a Linear Oscillating Cascade
Authors: Buffum D. H. , Fleeter S.
Published: June 2008

Experiments in a linear oscillating cascade reveal that the wind tunnel walls enclosing the airfoils have, in some cases, a detrimental effect on the oscillating cascade aerodynamics. In a subsonic flow field, biconvex airfoils are driven simultaneously in harmonic, torsion-mode oscillations for a range of interblade phase angle values. It is found that the cascade dynamic periodicity—the airfoil-to-airfoil variation in unsteady surface pressure—is good for some values of interblade phase angle but poor for others. Correlation of the unsteady pressure data with oscillating flat plate cascade predictions is generally good for conditions where the periodicity is good and poor where the periodicity is poor. Calculations based upon linearized unsteady aerodynamic theory indicate that pressure waves reflected from the wind tunnel walls are responsible for the cases where there is poor periodicity and poor correlation with the predictions.
Computational Study of Stall Flutter in Linear Cascades
Authors: Abdel-Rahim A. , Sisto F. , Thangam S.
Published: June 2008

Aeroelastic interaction in turbomachinery is of prime interest to opertors, designers, and aeroelasticans. Operation at off-design conditions may promote blade stall; eventually the stall pattern will propagate around the blade annulus. The unsteady periodic nature of propagating stall will force blade vibration and blade flutter may occur if the stall propagation frequency is entrained by the blade natural frequency. In this work a computational scheme based on the vortex method is used to simulate the flow over a linear cascade of airfoils. The viscous effect is confined to a thin layer, which determines the separation points on the airfoil surfaces. The preliminary structural model is a two-dimensional characteristic section with a single degree of freedom in either bending or torsion. A study of the relationship between the stall propagation frequency and the blade natural frequency has been conducted. The study shows that entrainment, or frequency synchronization, occurs, resulting in pure torsional flutter over a certain interval of reduced frequency. A severe blade torsional amplitude (of order 20 deg) has been computed in the entrainment region, reaching its largest value in the center of the interval. However, in practice, compressor blades will not sustain this vibration and blade failure may occur before reaching such a large amplitude. Outside the entrainment interval the stall propagation is shown to be independent of the blade natural frequency. In addition, computational results show that there is no entrainment in the pure bending mode. Rather, “de-entrainment” occurs with similar flow conditions and similar stall frequencies, resulting in blade buffeting in pure bending.
Optical Measurements of Unducted Fan Flutter
Authors: Kurkov A. P. , Mehmed O.
Published: June 2008

The paper describes a nonintrusive optical method for measuring flutter vibrations in unducted fan or propeller rotors and provides detailed spectral results for two flutter modes of a scaled unducted fan. The measurements were obtained in a high-speed wind tunnel. A single-rotor and a dual-rotor counterrotating configuration of the model were tested; however, only the forward rotor of the counterrotating configuration fluttered. Conventional strain gages were used to obtain flutter frequency; optical data provided complete phase results and an indication of the flutter mode shape through the ratio of the leading- to trailing-edge flutter amplitudes near the blade tip. In the transonic regime the flutter exhibited some features that are usually associated with nonlinear vibrations. Experimental mode shape and frequencies were compared with calculated values that included centrifugal effects.
Stall Inception in Axial Flow Compressors
Authors: Day I. J.
Published: June 2008

Studies have been conducted on two laboratory test compressors to investigate the process leading to the formation of finite amplitude rotating stall cells. The measurements were obtained from circumferential arrays of hot wires and were spatially and temporarily analyzed to show that modal perturbations are not always present prior to stall, and when present, sometimes have little direct effect on the formation of the stall cells. The measurements lead to the conclusion that the occurrence of modal perturbations, and the formation of finite amplitude stall cells, are two separate phenomena, both occurring under roughly the same conditions at the peak of the pressure rise characteristic. The measurements also underline the hitherto unsuspected importance of short length scale disturbances in the process of stall inception. Examples are given of different ways in which stall cells can develop and the conclusions are backed up with a summary of current test data from various machines around the world.
Active Control of Rotating Stall in a Low-Speed Axial Compressor
Authors: Paduano J. D. , Epstein A. H. , Valavani L. , Longley J. P. , Greitzer E. M. , Guenette G. R.
Published: June 2008

The onset of rotating stall has been delayed in a low-speed, single-stage, axial research compressor using active feedback control. Control was implemented using a circumferential array of hot wires to sense propagating waves of axial velocity upstream of the compressor. Using this information, additional circumferentially traveling waves were then generated with appropriate phase and amplitude by “wiggling” inlet guide vanes driven by individual actuators. The control scheme considered the wave pattern in terms of the individual spatial Fourier components. A simple proportional control law was implemented for each harmonic. Control of the first spatial harmonic yielded an 11 percent decrease in the stalling mass flow, while control of the first, second, and third harmonics together reduced the stalling mass flow by 23 percent. The control system was also used to measure the sine wave response of the compressor, which exhibited behavior similar to that of a second-order system.
Evaluation of Approaches to Active Compressor Surge Stabilization
Authors: Simon J. S. , Valavani L. , Epstein A. H. , Greitzer E. M.
Published: June 2008

Recent work has shown that compression systems can be actively stabilized against the instability known as surge, thereby realizing a significant gain in system mass flow range. Ideally, this surge stabilization requires only a single sensor and a single actuator connected by a suitable control law. Almost all research to date has been aimed at proof of concept studies of this technique, using various actuators and sensor combinations. In contrast, the work reported here can be regarded as a step toward developing active control into a practical technique. In this context, the paper presents the first systematic definition of the influence of sensor and actuator selection on increasing the range of stabilized compressor performance. The results show that proper choice of sensor as well as actuator crucially affects the ability to stabilize these systems, and that, overall, those actuators most closely coupled to the compressor (as opposed to the plenum or throttle) appear most effective. In addition, the source of the disturbances driving the system (for example, unsteady compressor pressure rise or unsteady combustor heat release) has a strong influence on control effectiveness, as would be expected for a controls problem of this type. This paper both delineates general methodologies for the evaluation of active compressor stabilization strategies and quantifies the performance of several approaches that might be implemented in gas turbine engines.
Inviscid-Viscous Coupled Solution for Unsteady Flows Through Vibrating Blades: Part 1—Description of the Method
Authors: He L. , Denton J. D.
Published: June 2008

An efficient coupled approach between inviscid Euler and integral boundary layer solutions has been developed for quasi-3-D unsteady flows induced by vibrating blades. For unsteady laminar and turbulent boundary layers, steady correlations are adopted in a quasi-steady way to close the integral boundary layer model. This quasi-steady adoption of the correlations is assessed by numerical test results using a direct solution of the unsteady momentum integral equation. To conduct the coupling between the inviscid and viscous solutions for strongly interactive flows, the unsteady Euler and integral boundary layer equations are simultaneously time-marched using a multistep Runge–Kutta scheme, and the boundary layer displacement effect is accounted for by a first order transpiration model. This time-resolved coupling method converges at conditions with considerable boundary layer separation.
Low Aspect Ratio Transonic Rotors: Part 1—Baseline Design and Performance
Authors: Law C. H. , Wadia A. R.
Published: June 2008

The analytical design and experimental test of a single-stage transonic axial-flow compressor are described. This design is the baseline of a compressor design study in which several blade design parameters have been systematically varied to determine their independent effects on compressor performance. The baseline design consisted of ruggedizing an existing compressor design that demonstrated outstanding aerodynamic performance, to correct some undesirable aeromechanical characteristics. The design study was performed by varying only one design parameter at a time, keeping the other design variables as close as possible to the baseline design. Specific design parameters of interest were those for which very few data were available to determine their sensitivity on compressor performance. This paper describes the baseline compressor design and its experimental performance. A detailed definition and flow analysis of the baseline design test point (used as the basis for all subsequent design variations) are provided.
Low Aspect Ratio Transonic Rotors: Part 2—Influence of Location of Maximum Thickness on Transonic Compressor Performance
Authors: Wadia A. R. , Law C. H.
Published: June 2008

Transonic compressor rotor performance is sensitive to variations in several known design parameters. One such parameter is the chordwise location of maximum thickness. This article reports on the design and experimental evaluation of two versions of a low aspect ratio transonic rotor that had the location of the tip blade section maximum thickness moved forward in two increments from the nominal 70 percent to 55 and 40 percent chord length, respectively. The original hub characteristics were preserved and the maximum thickness location was adjusted proportionately along the span. Although designed to satisfy identical design speed requirements, the experimental results reveal significant variation in the performance of the rotors. At design speed, the rotor with its maximum thickness located at 55 percent chord length attains the highest peak efficiency among the three rotors but has lowest flow rollback relative to the other two versions. To focus on current ruggedization issues for transonic blading (e.g., bird and ice ingestion), detailed comparison of test data and analysis to characterize the aerodynamic flow details responsible for the measured performance differences were confined to the two rotors with the most forward location of maximum thickness. A three-dimensional viscous flow analysis was used to identify the performance-enhancing features of the higher efficiency rotor and to provide guidance in the interpretation of the experimental measurements. The computational results of the viscous analysis show that the difference in performance between the two rotors can be attributed to the higher shock losses that result from the increased leading edge “wedge angle” as the maximum thickness is moved closer to the leading edge. The test data and the three-dimensional viscous analysis also reveal that the higher efficiency rotor achieves the same static pressure rise potential and loading at a higher flow level than its less efficient counterpart and this is responsible for its resulting lower flow rollback and apparent loss in stall margin. Comparison of the peak efficiencies attained by the two rotors described in this article with the baseline ruggedized rotor performance presented in part 1 of this paper suggests the existence of an optimum maximum thickness location at 55 to 60 percent chord length for such low aspect ratio transonic rotors.
A Statistical Approach to the Experimental Evaluation of Transonic Turbine Airfoils in a Linear Cascade
Authors: Shelton M. L. , Gregory B. A. , Doughty R. L. , Kiss T. , Moses H. L.
Published: June 2008

In aircraft engine design (and in other applications), small improvements in turbine efficiency may be significant. Since analytical tools for predicting transonic turbine losses are still being developed, experimental efforts are required to evaluate various designs, calibrate design methods, and validate CFD analysis tools. However, these experimental efforts must be very accurate to measure the performance differences to the levels required by the highly competitive aircraft engine market. Due to the sensitivity of transonic and supersonic flow fields, it is often difficult to obtain the desired level of accuracy. In this paper, a statistical approach is applied to the experimental evaluation of transonic turbine airfoils in the VPI & SU transonic cascade facility in order to quantify the differences between three different transonic turbine airfoils. This study determines whether the measured performance differences between the three different airfoils are statistically significant. This study also assesses the degree of confidence in the transonic cascade testing process at VPI & SU.
Aerodynamic Performance of a Transonic Low Aspect Ratio Turbine Nozzle
Authors: Moustapha S. H. , Carscallen W. E. , McGeachy J. D.
Published: June 2008

This paper presents detailed information on the three-dimensional flow field in a realistic turbine nozzle with an aspect ratio of 0.65 and a turning angle of 76 deg. The nozzle has been tested in a large-scale planar cascade over a range of exit Mach numbers from 0.3 to 1.3. The experimental results are presented in the form of nozzle passage Mach number distributions and spanwise distribution of losses and exit flow angle. Details of the flow field inside the nozzle passage are examined by means of surface flow visualization and Schlieren pictures. The performance of the nozzle is compared to the data obtained for the same nozzle tested in an annular cascade and a stage environment. Excellent agreement is found between the measured pressure distribution and the prediction of a three-dimensional Euler flow solver.
Separated Flow in a Low-Speed Two-Dimensional Cascade: Part II—Cascade Performance
Authors: Yocum A. M. , O’Brien W. F.
Published: June 2008

This study was conducted for the purpose of providing a more fundamental understanding of separated flow in cascades and to provide performance data for fully stalled blade rows. Cascades of a single blade geometry and a solidity of unity were studied for three stagger angles and the full range of angle of attack extending well into the stalled flow regime. The Reynolds number was also varied for a limited number of cases. Results from velocity and pressure measurements made in the cascade and the overall cascade performance evaluated from these measurements are presented. In addition, results from a numerical simulation of the flow through a cascade of flat plate airfoils are used to illustrate further the effects of blade stagger and to define the correct limits for the cascade performance. The results indicate that the slope of the total pressure loss versus angle of attack curve for the flow immediately downstream of the cascade is steeper for cascades with greater stagger. The normal force coefficient was found to increase to a peak value near the angle of attack where full leading edge stall first occurs. A further increase in angle of attack results in a decline in the normal force coefficient. The peak value of the normal force coefficient is greater and occurs at a higher angle of attack for the cascades with smaller stagger.
Investigation of Tip Clearance Phenomena in an Axial Compressor Cascade Using Euler and Navier–Stokes Procedures
Authors: Kunz R. F. , Lakshminarayana B. , Basson A. H.
Published: June 2008

Three-dimensional Euler and full Navier–Stokes computational procedures have been utilized to simulate the flow field in an axial compressor cascade with tip clearance. An embedded H-grid topology was utilized to resolve the flow physics in the tip gap region. The numerical procedure employed is a finite difference Runge-Kutta scheme. Available measurements of blade static pressure distributions along the blade span, dynamic pressure and flow angle in the cascade outlet region, and spanwise distributions of blade normal force coefficient and circumferentially averaged flow angle are used for comparison. Several parameters that were varied in the experimental investigations were also varied in the computational studies. Specifically, measurements were taken and computations were performed on the configuration with and without: tip clearance, the presence of an endwall, inlet endwall total pressure profiles and simulated relative casing rotation. Additionally, both Euler and Navier–Stokes computations were performed to investigate the relative performance of these approaches in reconciling the physical phenomena considered. Results indicate that the Navier–Stokes procedure, which utilizes a low Reynolds number k–ε model, captures a variety of important physical phenomena associated with tip clearance flows with good accuracy. These include tip vortex strength and trajectory, blade loading near the tip, the interaction of the tip clearance flow with passage secondary flow, and the effects of relative endwall motion. The Euler computation provides good but somewhat diminished accuracy in resolution of some of these clearance phenomena. It is concluded that the level of modeling embodied in the present approach is sufficient to extract much of the tip region flow field information useful to designers of turbomachinery.
Numerical Simulation of Compressor Endwall and Casing Treatment Flow Phenomena
Authors: Crook A. J. , Greitzer E. M. , Tan C. S. , Adamczyk J. J.
Published: June 2008

A numerical study is presented of the flow in the endwall region of a compressor blade row, in conditions of operation with both smooth and grooved endwalls. The computations are first compared to velocity field measurements in a cantilevered stator/rotating hub configuration to confirm that the salient features are captured. Computations are then interrogated to examine the tip leakage flow structure since this is a dominant feature of the endwall region. In particular, the high blockage that can exist near the endwalls at the rear of a compressor blade passage appears to be directly linked to low total pressure fluid associated with the leakage flow. The fluid dynamic action of the grooved endwall, representative of the casing treatments that have been most successful in suppressing stall, is then simulated computationally and two principal effects are identified. One is suction of the low total pressure, high blockage fluid at the rear of the passage. The second is energizing of the tip leakage flow, most notably in the core of the leakage vortex, thereby suppressing the blockage at its source.
Blade Loading and Shock Wave in a Transonic Circular Cascade Diffuser
Authors: Hayami H. , Sawae M. , Nakamura T. , Kawaguchi N.
Published: June 2008

A low-solidity circular cascade, conformally transformed from a high-stagger linear cascade of double-circular-arc vanes with solidity 0.69, was tested as a part of diffuser systems of a transonic centrifugal compressor and the static pressures were measured around a vane of the cascade and on the side wall between cascade vanes in detail. The blade loading of cascade vane was discussed by integrating the pressure distribution around the vane. The experimental data for lift-coefficient of vane were almost on a single straight line with positive gradient against angle-of-attack over a wide range of inflow Mach number and inflow angle. The maximum lift coefficient reached about 1.5 and the vane worked well to the surge condition of the compressor. The structure of shock wave was also discussed by drawing a contour map of the flow Mach number between cascade vanes. The normal shock wave was observed on the suction surface of vane and it moved upstream along the suction surface with the decrease in inflow angle. The vane did not fall in stall even though the Mach number upstream of the shock wave was over 1.4.
Aerodynamic Design of Turbomachinery Blading in Three-Dimensional Flow: An Application to Radial Inflow Turbines
Authors: Yang Y. L. , Tan C. S. , Hawthorne W. R.
Published: June 2008

A computational method based on a theory for turbomachinery blading design in three-dimensional inviscid flow is applied to a parametric design study of a radial inflow turbine wheel. As the method requires the specification of swirl distribution, a technique for its smooth generation within the blade region is proposed. Excellent agreements have been obtained between the computed results from this design method and those from direct Euler computations, demonstrating the correspondence and consistency between the two. The computed results indicate the sensitivity of the pressure distribution to a lean in the stacking axis and a minor alteration in the hub/shroud profiles. Analysis based on a Navier–Stokes solver shows no breakdown of flow within the designed blade passage and agreement with that from a design calculation; thus the flow in the designed turbine rotor closely approximates that of an inviscid one. These calculations illustrate the use of a design method coupled to an analysis tool for establishing guidelines and criteria for designing turbomachinery blading.
Hot Streaks and Phantom Cooling in a Turbine Rotor Passage: Part 2—Combined Effects and Analytical Modeling
Authors: Roback R. J. , Dring R. P.
Published: June 2008

This paper presents experimental documentation and analytical correlations demonstrating the effects of hot streak accumulation and phantom cooling on turbine rotor airfoil surface temperature. In particular, results are shown that quantify the impact of (1) a nonuniform temperature profile at the entrance of a turbine due to combustor-generated hot and cold streaks, and (2) cooling air discharged from the trailing edge of the upstream stator. In Part 1 of this paper, experimental results were shown for a range of controlling variables to identify where streak accumulation and phantom cooling were most likely to be strongest. These variables include streak-to-free-stream density ratio, streak injection location, and coolant-to-free-stream density and velocity ratios. In Part 2 of this paper, experimental results will be shown for the combined effects of hot streak and stator coolant on the adiabatic recovery temperature on the rotor. An analytical model is also developed to correlate the experimental results documented in Parts 1 and 2 of the paper.
An Analysis System for Blade Forced Response
Authors: Chiang Hsiao-Wei D. , Kielb R. E.
Published: June 2008

A frequent cause of turbomachinery blade failure is excessive resonant response. The most common excitation source is the nonuniform flow field generated by inlet distortion, wakes and/or pressure disturbances from adjacent blade rows. The standard method for dealing with this problem is to avoid resonant conditions using a Campbell diagram. Unfortunately, it is impossible to avoid all resonant conditions. Therefore, judgments based on past experience are used to determine the acceptability of the blade design. A new analysis system has been developed to predict blade forced response. The system provides a design tool, over and above the standard Campbell diagram approach, for predicting potential forced response problems. The incoming excitation sources are modeled using a semi-empirical rotor wake/vortex model for wake excitation, measured data for inlet distortion, and a quasi-three-dimensional Euler code for pressure disturbances. Using these aerodynamic stimuli, and the blade’s natural frequencies and mode shapes from a finite element model, the unsteady aerodynamic modal forces and the aerodynamic damping are calculated. A modal response solution is then performed. This system has been applied to current engine designs. A recent investigation involved fan blade response due to inlet distortion. An aero mechanical test had been run with two different distortion screens. The resulting distortion entering the fan was measured. With this as input data, the predicted response agreed almost exactly with the measured response. In another application, the response of the LPT blades of a counterrotating supersonic turbine was determined. In this case the blades were excited by both a wake and a shock wave. The shock response was predicted to be three times larger than that of the wake. Thus, the system identified a new forcing function mechanism for supersonic turbines. This paper provides a basic description of the system, which includes: (1) models for the wake excitation, inlet distortion, and pressure disturbance; (2) a kernel function solution technique for unsteady aerodynamics; and (3) a modal aeroelastic solution using strip theory. Also, results of the two applications are presented.
The Combined Closed Form-Perturbation Approach to the Analysis of Mistuned Bladed Disks
Authors: Mignolet M. P. , Lin C.-C.
Published: June 2008

A two-step method is presented for the determination of reliable approximations of the probability density function of the forced response of a randomly mistuned bladed disk. Under the assumption of linearity, an integral representation of the probability density function of the blade amplitude is first derived. Then, deterministic perturbation techniques are employed to produce simple approximations of this function. The adequacy of the method is demonstrated by comparing several approximate solutions with simulation results.
Coupled Three-Dimensional Aeroelastic Stability Analysis of Bladed Disks
Authors: Gerolymos G. A.
Published: June 2008

In the present work an algorithm for the coupled aeromechanical computation of three-dimensional compressor cascades vibrating in a traveling-wave mode is presented and applied to the determination of aeroelastic stability of a transonic fan rotor. The initial vibratory modes are computed using a finite-element structural analysis code. The unsteady flow field response to blade vibration is estimated by numerical integration of the three-dimensional unsteady Euler equations. Coupling relations are formulated in the frequency domain, using a mode-modification technique, based on modal projection. The vibratory mode is updated at the end of the aerodynamic simulation of each period, and the updated mode is used for the simulation of the next period. A number of results illustrate the method’s potential.
Bifurcation Analysis of Surge and Rotating Stall in Axial Flow Compressors
Authors: Abed E. H. , Houpt P. K. , Hosny W. M.
Published: June 2008

The surge and rotating stall post-instability behaviors of axial flow compressors are investigated from a bifurcation-theoretic perspective, using a model and system data presented by Greitzer (1976a). For this model, a sequence of local and global bifurcations of the nonliner system dynamics is uncovered. This includes a global bifurcation of a pair of large-amplitude periodic solutions. Resulting from this bifurcation are a stable oscillation (“surge”) and an unstable oscillation (“anti-surge”). The latter oscillation is found to have a deciding significance regarding the particular post-instability behavior experienced by the compressor. These results are used to reconstruct Greitzer’s (1976b) findings regarding the manner in which post-instability behavior depends on system parameters. Although the model does not directly reflect nonaxisymmetric dynamics, use of a steady-state compressor characteristic approximating the measured characteristic of Greitzer (1976a) is found to result in conclusions that compare well with observation. Thus, the paper gives a convenient and simple explanation of the boundary between surge and rotating stall behaviors, without the use of more intricate models and analyses including nonaxisymmetric flow dynamics.
Predictions of Three-Dimensional Steady and Unsteady Inviscid Transonic Stator/Rotor Interaction With Inlet Radial Temperature Nonuniformity
Authors: Saxer A. P. , Giles M. B.
Published: June 2008

Numerical predictions of three-dimensional inviscid, transonic steady and periodic unsteady flow within an axial turbine stage are analyzed in this paper. As a first case, the unsteady effects of the stator trailing edge shock wave impinging on the downstream rotor are presented. Local static pressure fluctuations up to 60 percent of the inlet stagnation pressure are observed on the rotor suction side. The second case is an analysis of the rotor-relative radial secondary flow produced by a spanwise parabolic nonuniform temperature profile at the stator inlet. The generation of local hot spots is observed on both sides of the rotor blade behind the passing shock waves. The magnitude of the unsteady stagnation temperature fluctuations is larger than the time-averaged rotor inlet disturbance. In both cases, steady, unsteady, and time-averaged solutions are presented and compared. From these studies, it is concluded that the steady-state solution in static pressure matches well with the time-averaged periodic unsteady flow field. However, for the stagnation temperature distribution only the trend of the time-averaged solution is modeled in the steady-state solution.
Detailed Measurements of Local Heat Transfer Coefficient and Adiabatic Wall Temperature Beneath an Array of Impinging Jets
Authors: Van Treuren K. W. , Wang Z. , Ireland P. T. , Jones T. V.
Published: June 2008

A transient method of measuring the local heat transfer under an array of impinging jets has been developed. The use of a temperature-sensitive coating consisting of three encapsulated thermochromic liquid crystal materials has allowed the calculation of both the local adiabatic wall temperature and the local heat transfer coefficient over the complete surface of the target plate. The influence of the temperature of the plate through which the impingement gas flows on the target plate heat transfer has been quantified. Results are presented for a single in-line array configuration over a range of jet Reynolds numbers.
Shroud Heat Transfer Measurements From a Rotating Cavity With an Axial Throughflow of Air
Authors: Long C. A. , Tucker P. G.
Published: June 2008

The paper discusses measurements of heat transfer obtained from the inside surface of the peripheral shroud. The experiments were carried out on a rotating cavity, comprising two 0.985-m-dia disks, separated by an axial gap of 0.065 m and bounded at the circumference by a carbon fiber shroud. Tests were conducted with a heated shroud and either unheated or heated disks. When heated, the disks had the same temperature level and surface temperature distribution. Two different temperature distributions were tested; the surface temperature either increased, or decreased with radius. The effects of disk, shroud, and air temperature levels were also studied. Tests were carried out for the range of axial throughflow rates and speeds: 0.0025 ≤ m ≤ 0.2 kg/s and 12.5 ≤Ω≤ 125 rad/s, respectively. Measurements were also made of the temperature of the air inside the cavity. The shroud Nusselt numbers are found to depend on a Grashof number, which is defined using the centripetal acceleration. Providing the correct reference temperature is used, the measured Nusselt numbers also show similarity to those predicted by an established correlation for a horizontal plate in air. The heat transfer from the shroud is only weakly affected by the disk surface temperature distribution and temperature level. The heat transfer from the shroud appears to be affected by the Rossby number. A significant enhancement to the rotationally induced free convection occurs in the regions 2≤Ro≤4 and Ro≥20. The first of these corresponds to a region where vortex breakdown has been observed. In the second region, the Rossby number may be sufficiently large for the central throughflow to affect the shroud heat transfer directly. Heating the shroud does not appear to affect the heat transfer from the disks significantly.
M. J. Hartmann Memorial Session Paper: NASA/GE Fan and Compressor Research Accomplishments
Authors: Smith L. H.
Published: June 2008

Fan and compressor research projects carried out at GE Aircraft Engines under NASA sponsorship are described in this paper. Four 1400-fps-tip-speed rotors designed with different airfoil shapes were found to have comparable stall lines but different efficiency trends. A stator placed behind one of these affected its performance somewhat. Adjustments of variable camber inlet guide vanes placed ahead of a 1500fps stage were found to affect its pumping capability without much affecting its stall line. For the Quiet Engine Program (QEP), two 1160-fps fans and one 1550-fps fan were tested. Development of the high-speed fan revealed the effects on performance of airfoil shape and part-span shroud blockage. The 950-fps variable-pitch fan for the Quiet Clean Short-haul Experimental Engine (QCSEE) demonstrated reverse thrust capabilities and a novel method of avoiding large core inlet pressure losses during reverse thrust operation. The 1350-fps Energy Efficient Engine (E3) fan demonstrated excellent performance with a novel quarter-stage arrangement that eliminated the need for interspool bleed while giving good dirt removal potential. The E3 compressor program employed Low Speed Research Compressor tests to identify the most appropriate blading type. High-speed rig tests and engine tests were then used to develop this 23:1-spool-pressure-ratio compressor. Research on casing boundary layer control through bleeding and blowing led to the discovery that irregular casing geometries usually give stall line enhancements even without auxiliary air circuits. Some of the resulting casing treatment research is reported herein. Instances in which NASA-sponsored research has affected GE Aircraft Engine products are pointed out.
Flow Physics Leading to System Instability in a Centrifugal Pump
Authors: Wo A. M. , Bons J. P.
Published: June 2008

The off-design performance in a centrifugal pump is investigated experimentally. The objective is to identify flow features that lead to the onset of surge as the fundamental pumping system instability. Results show that there are primarily two reasons for the onset of surge as the flow is reduced in the pump studied: (a) adverse flow in the tongue region, and (b) destabilizing effect of the pipe diffuser. The former is due to premature diffusion of the flow entering the tongue region, which is manifested by increased flow recirculation through the tongue/impeller gap and flow separation on the volute outer side-wall opposite the tongue. These effects in the tongue region flow coupled with the destabilizing behavior of the pipe diffuser lead to the eventual unstable pump operation.