bluuebunny/crossref_metadata_2025 huggingface dataset:
Effects of Endwall Suction and Blowing on Compressor Stability Enhancement
Authors: Lee N. K. W. , Greitzer E. M.
Published: June 2008

An experimental investigation was carried out to examine the effects on stall margin of flow injection into, and flow removal out of, the endwall region of an axial compressor blade row. A primary objective of the investigation was to clarify the mechanism by which casing treatment (which involves both removal and injection) suppresses stall in turbomachines. To simulate the relative motion between blade and treatment, the injection and removal took place through a slotted hub rotating beneath a cantilevered stator row. Overall performance data and detailed (time-averaged) flowfield measurements were obtained. Flow injection and removal both increased the stalling pressure rise, but neither was as effective as the wall treatment. Removal of high-blockage flow is thus not the sole reason for the observed stall margin improvement in casing or hub treatment; injection can also contribute significantly to stall suppression. The results indicate that the increase in stall pressure rise achieved with injection is linked to the streamwise momentum of the injected flow, and it is suggested that this should be the focus of further studies.
Application of Recess Vaned Casing Treatment to Axial Flow Fans
Authors: Azimian A. R. , Elder R. L. , McKenzie A. B.
Published: June 2008

The effect of applying a vaned recessed casing treatment to a single-stage axial flow fan has been investigated. The influences of the axial position of the recess relative to the rotor leading edge and other geometric modifications of the vane passage have been examined. Significant improvements in stall margin were observed without (in some builds) loss in peak efficiency. Slow and fast frequency response yawmeter probes have been used in the study to examine both the steady flow conditions and the unsteady flow caused by rotating stall.
Effect of Blade Tip Configuration on Tip Clearance Loss of a Centrifugal Impeller
Authors: Ishida M. , Ueki H. , Senoo Y.
Published: June 2008

According to the theory presented by the authors, the tip clearance loss of an un-shrouded centrifugal impeller mainly consists of two kinds of loss; one is the drag due to the leakage flow through the blade tip clearance and the other is the pressure loss to support the fluid in the thin annular clearance space between the shroud and the blade tip against the pressure gradient in the meridional plane without blades. The former is proportional to the leakage flow or the contraction coefficient of leakage flow. The authors have conducted performance tests using an impeller with 16 backward-leaning blades in three configurations of the blade tip: round edge, sharp square edge, and edge with an end-plate. The experimental tip clearance effects can be predicted by the theory assuming reasonable contraction coefficients. They are 0.91, 0.73, and 0.53 for the respective tip configurations. The impeller efficiency is improved by about 1.5 point by reducing the contraction coefficient from 0.91 to 0.53, providing that the tip clearance ratio at the exit of impeller is 0.1. More improvement is expected for an impeller with highly loaded blades where the leakage loss shares the major part of the tip clearance loss.
Blade Sweep for Low-Speed Axial Fans
Authors: Wright T. , Simmons W. E.
Published: June 2008

The available literature on aerodynamic and acoustic properties of axial fans with swept blades is presented and discussed with particular emphasis on noise mechanisms and the influence of high-intensity inlet turbulence on “excess” noise. The acoustic theory of Kerschen and Envia for swept cascades is applied to the problem of axial fan design. These results are compared to available data and a provisional model for specifying sweep angles is presented. The aerodynamic performance theory for swept-bladed rotors of Smith and Yeh is adapted for use in designing low-speed axial fans. Three prototype fans were designed using the resultant computer codes. One is a baseline fan with blade stocking lines radially oriented, and two are fans having swept blades of increasingly greater forward sweep. Aerodynamic testing shows that performance of the fans lies within a band width of about ± 2 percent of volume flow rate and pressure rise predictions in the region of design performance, effectively validating the design procedure for selection of the blading parameters. Noise testing of the fans was carried out and the results show an average noise reduction for the swept-bladed fans of about 7 dBA overall, and a reduction of pure tone noise at blade-pass frequency of about 10 dB compared to the zero-sweep baseline model, in close agreement with the theory of Kerschen and Envia.
Secondary Flow Due to the Tip Clearance at the Exit of Centrifugal Impellers
Authors: Ishida M. , Senoo Y. , Ueki H.
Published: June 2008

The velocity distribution was measured at the exit of two different types of un-shrouded centrifugal impeller under four different tip clearance conditions each; one with 20 radial blades and inducers and the other with 16 backward-leaning blades. The effect of tip clearance on input power was also measured. By increasing the tip clearance, the input power was hardly changed in the radial blade impeller and was reduced in the backward-leaning blade impeller. The velocity distribution normalized by the passage width between hub and shroud wall was hardly changed at the exit of the radial blade impeller by varying the tip clearance. On the other hand, the relative flow angle was reduced significantly and monotonously by an increase of tip clearance in the backward-leaning blade impeller. The change in input power due to the tip clearance was clearly related to the change of flow pattern at the exit of impeller due to the secondary flow. This is most likely caused by the component, normal to the blade, of the shear force to support the fluid in the clearance space against the pressure gradient in the meridional plane without blades.
Automotive Research Compressor Experience
Authors: Pampreen R. C.
Published: June 2008

The design features and development experience of an advanced automotive gas turbine centrifugal compressor are presented. The compressor was designed with moderately low specific speed, high blade aerodynamic loading, and design point on the choke characteristic. A cascade diffuser was used; a new design approach brought about an unusual performance characteristic. The influence of impeller blade loading on performance is presented. Lessons learned are summarized.
Application of a Low-Solidity Cascade Diffuser to Transonic Centrifugal Compressor
Authors: Hayami H. , Senoo Y. , Utsunomiya K.
Published: June 2008

Low-solidity circular cascades, conformally transformed from high-stagger linear cascades of double-circular-arc vanes with solidity 0.69, were used as a part of the diffuser system of a transonic centrifugal compressor. Performance test results were compared with data of the same compressor with a vaneless diffuser. Good compressor performance and a wider flow range as well as a higher pressure ratio and a higher efficiency, superior to those with a vaneless diffuser, where the flow range was limited by choke of the impeller, were demonstrated. The test circular cascade diffusers demonstrated a good pressure recovery over a wide range of flow angles, even when the inflow Mach number to the cascade was over unity.
Experiments With a Low-Specific-Speed Partial Emission Centrifugal Compressor
Authors: Rodgers C.
Published: June 2008

This paper describes experiments that were conducted several years ago on a single-stage open-face centrifugal partial emission compressor with a nondimensional specific speed of 0.15 (19.3 dimensional). The purpose of the experiments was experimentally to evaluate the partial emission compressor as a possible candidate for a high-speed low-flow natural gas fuel pump, the alternative to which was a larger heavier reciprocating compressor. A review and re-analysis of the test data were conducted. They used updated compressor data reduction performance procedures and flow models to redefine the experimental compressor as a reference case in the low-specific-speed range of single-stage centrifugal compressors. Normalized performance characteristics are presented for the experimental compressor in terms of efficiencies, diffuser recovery, and head coefficients versus diffuser relative flow parameter, or choke margin. Peak efficiency, obtained at a De Laval number of 1.16, and nondimensional specific speed of 0.15, was 34.5 percent. Stage performance is found to be dominated by diffuser geometry and choke margin, plus windage losses.
Experimental and Theoretical Study of the Swirling Flow in Centrifugal Compressor Volutes
Authors: Van den Braembussche R. A. , Ha¨nde B. M.
Published: June 2008

Measurements of the three-dimensional flow in a simplified model of a centrifugal compressor volute at design and off-design operation are presented. A nearly constant swirl velocity is observed near the walls and a forced vortex type of flow is observed in the center. This velocity distribution is almost identical at all cross sections and all operating points. An explanation is given on how this swirl distribution results from the specific way a volute is filled with fluid. The throughflow velocity component shows a large crosswise variation. A minimum or maximum velocity is observed at the volute center depending on the operating point. A simple analytic model, based on the radial equilibrium of forces, is described. Calculations for isentropic flows reveal the relation between the swirl distribution and the large increase of throughflow velocity toward the center. This explains why volutes should be designed with negative blockage. Nonisentropic calculations, using the experimental loss distribution, correctly reproduce the measured throughflow velocity and static pressure distribution.
An Approximate Three-Dimensional Aerodynamic Design Method for Centrifugal Impeller Blades
Authors: Xiaolu Zhao , Lisen Qin
Published: June 2008

An aerodynamic design method, which is based on the Mean Stream Surface Method (MSSM), has been developed for designing centrifugal compressor impeller blades. As a component of a CAD system for centrifugal compressor, it is convenient to use the presented method for generating impeller blade geometry, taking care of manufacturing as well as aerodynamic aspects. The design procedure starts with an S2m indirect solution. Afterward from the specified S2m surface, by the use of Taylor series expansion, the blade geometry is generated by straight-line elements to meet the manufacturing requirements. Simultaneously, the fluid dynamic quantities across the blade passage can be determined directly. In terms of these results, the designer can revise the distribution of angular momentum along the shroud and hub, which are associated with blade loading, to get satisfactory velocities along the blade surfaces in order to avoid or delay flow separation.
Effects of Tip Clearance of Nozzle Vanes on Performance of Radial Turbine Rotor
Authors: Hayami H. , Senoo Y. , Hyun Y. I. , Yamaguchi M.
Published: June 2008

In radial turbines with variable nozzles, the flow downstream of the nozzles could be distorted by the leakage flow through the tip clearance of the upstream nozzle vanes. To investigate the effects of flow distortion on the performance of turbine rotors, two rotors with different numbers of blades were tested for three types of distorted velocity distribution at the rotor inlet. In the case of the 20-blade rotor with moderate blade loading, the flow distortion at the rotor inlet had a negligible effect on the rotor characteristics, and the measured data on the turbine performances agreed well with prediction. Predictions were made with a conventional one-dimensional flow model applied to the rotor flow, while a two-layer flow model was applied to the flow in the nozzle with clearance. In the case of the 10-blade rotor with heavy blade loading, however, the rotor performance was found to be sensitive to the inlet flow distortion and was considerably lower than the prediction.
Optimum Design of a Low-Pressure, Double Inflow, Radial Steam Turbine for Open-Cycle Ocean Thermal Energy Conversion
Authors: Schobeiri T.
Published: June 2008

This paper describes the thermo-fluid design concept of a low-pressure, radial inflow steam turbine for the Open-Cycle Ocean Thermal Energy Conversion (OTEC). For the optimum design consideration, the necessary theoretical tools are developed. The design concept permits the utilization of single-stage and multistage configurations for small and medium power ranges, respectively.
Simulation of Compressor Performance Deterioration Due to Erosion
Authors: Tabakoff W. , Lakshminarasimha A. N. , Pasin M.
Published: June 2008

Experimental results obtained from cascades and single-stage compressor performance tests before and after erosion were used to test a fault model to represent erosion. This model was implemented on a stage stacking program developed to demonstrate the effect of erosion in a multistage compressor. The effect of individual stage erosion on the overall compressor performance is also demonstrated.
Secondary Flows and Vortex Motion in a High-Efficiency Backswept Impeller at Design and Off-Design Conditions
Authors: Hah C. , Krain H.
Published: June 2008

The three-dimensional viscous flow field of a 4.7:1 pressure ratio backswept impeller was studied experimentally and numerically by using laser velocimetry and an advanced three-dimensional viscous code. The impeller was designed by a CAD method, and a maximum rotor efficiency of 94 percent was achieved. Both the experimental and the theoretical approach revealed comparatively smooth impeller discharge velocity profiles at all three operating conditions (design, choke, and near surge) differing widely from the well-known jet/wake-type flow pattern. The three-dimensional viscous code was used for detailed flowfield studies, i.e., secondary flows; vortex motion and tip-clearance effects were analyzed at design and off-design conditions. The comparison of experimental and numerical results indicates that the tip-clearance effect should be properly modeled to predict the impeller flow pattern properly and that optimum shape of rotor exit flow pattern can be obtained by controlling the swirling vortex motion.
Flow Field Investigation in a Low-Solidity Inducer by Laser-Doppler Velocimetry
Authors: Boccazzi A. , Perdichizzi A. , Tabacco U.
Published: June 2008

The results of an experimental investigation of the flow field within a low-solidity inducer at design and off-design flow rates are presented and discussed; particular attention is devoted to the analysis of the flow field, at the tip in front of the leading edge, for the flow rate close to the back-flow onset. The flow field was measured by means of a laser-Doppler velocimeter at four different axial positions upstream, within, and downstream of the inducer. Axial, tangential, and relative flow angle distributions, in the measuring planes, are presented for three different flow coefficients. At the lower flow rate, the plots show the presence of reverse flow in the region close to the hub downstream of the trailing edge. For the same flow rate, quite low axial velocities are detected at the tip. This is in agreement with pressure probe traverses carried out in a slightly downstream section; these measurements also show radial inward velocities of the same order of magnitude as the axial velocities. Circumferentially averaged losses were evaluated from specific work and total head rise given by pressure probes.
Experimental Investigation of Rotor-Stator Interaction in a Centrifugal Pump With Several Vaned Diffusers
Authors: Arndt N. , Acosta A. J. , Brennen C. E. , Caughey T. K.
Published: June 2008

This paper describes an experimental investigation of rotor-stator interaction in a centrifugal pump with several vaned diffusers. Steady and unsteady diffuser vane pressure measurements were made for a two-dimensional test impeller. Unsteady impeller blade pressure measurements were made for a second two-dimensional impeller with blade number and blade geometry identical to the two-dimensional impeller used for the diffuser vane pressure measurements. The experiments were conducted for different flow coefficients and different radial gaps between the impeller blade trailing edge and the diffuser vane leading edge (5 and 8 percent of the impeller discharge radius). The largest pressure fluctuations on the diffuser vanes and the impeller blades were found to be of the same order of magnitude as the total pressure rise across the pump. The largest pressure fluctuations on the diffuser vanes were observed to occur on the suction side of the vane near the vane leading edge, whereas on the impeller blades the largest fluctuations were observed to occur at the blade trailing edge. However, the dependence of the fluctuations on the flow coefficient was found to be different for the diffuser vanes and the impeller blades; on the vane suction side, the fluctuations were largest for the maximum flow coefficient and decreased with decreasing flow coefficient, whereas at the blade trailing edge, the fluctuations were smallest for the maximum flow coefficient and increased with decreasing flow coefficient. Increasing the number of the diffuser vanes resulted in a significant decrease of the impeller blade pressure fluctuations. The resulting lift on the diffuser vanes was computed from the vane pressure measurements; the magnitude of the fluctuating lift was found to be larger than the steady lift.
Design and Testing of the FT8 Gas Turbine Low-Pressure Compressor
Authors: Monhardt R. J. , Richardson J. H. , Boettcher J. M.
Published: June 2008

This paper presents a discussion of the design and test phases of the Low-Pressure Compressor (LPC) used in the FT8 Gas Turbine. Design objectives, including efficiency, durability, and stall margin goals, are covered along with actual test results indicating successful accomplishment of these objectives. Unique features of the test program and an overview of the FT8 Gas Generator in comparison to the JT8D Aircraft Engine are also presented.
A Variable-Geometry Power Turbine for Marine Gas Turbines
Authors: Karstensen K. W. , Wiggins J. O.
Published: June 2008

Gas turbines have been accepted in naval surface ship applications, and considerable effort has been made to improve their fuel consumption, particularly at part-load operation. This is an important parameter for shipboard engines because both propulsion and electrical-generator engines spend most of their lives operating at off-design power. An effective way to improve part-load efficiency of recuperated gas turbines is by using a variable power turbine nozzle. This paper discusses the successful use of variable power turbine nozzles in several applications in a family of engines developed for vehicular, industrial, and marine use. These engines incorporate a variable power turbine nozzle and primary surface recuperator to yield specific fuel consumption that rivals that of medium speed diesels. The paper concentrates on the experience with the variable nozzle, tracing its derivation from an existing fixed vane nozzle and its use across a wide range of engine sizes and applications. Emphasis is placed on its potential in marine propulsion and auxiliary gas turbines.
Surface Roughness Measurements on Gas Turbine Blades
Authors: Taylor R. P.
Published: June 2008

Results are presented from profilometer measurements of the surface roughness on in-service turbine engine blades from F-100 and TF-39 aeroengines. On each blade, one roughness profile is taken in the region of the leading edge, the midchord and the trailing edge on both the pressure and suction sides for a total of six profiles. Thirty first-stage turbine blades are measured from each engine. Statistical computations are performed on these profiles and the root mean square height, skewness and kurtosis of the roughness height distribution are presented along with the correlation length of the autocorrelation function. The purpose of this work is to provide insight into the nature of surface roughness characteristics of in-service turbine blades which can be used in the development of scaled laboratory experiments of boundary layer flow and heat transfer on turbine engine blades.
The Influence of Boundary Layer State on Vortex Shedding From Flat Plates and Turbine Cascades
Authors: Sieverding C. H. , Heinemann H.
Published: June 2008

The paper aims at a better understanding of the reasons for the wide range of Strouhal numbers observed on turbine blades. The investigation is restricted to the subsonic domain. First, flat plate model tests are carried out to investigate the effect of both the boundary layer state and trailing edge geometry on the vortex shedding frequency. A particular objective of the tests is to obtain data for the very common case of a mixed laminar-turbulent separation from turbine blades. These basic tests are followed by three cascade tests with blades of very different suction side velocity distributions. Based on the experience gained from the flat plate test program, an attempt is made to interpret the Strouhal number variation with Mach number and Reynolds number, and to relate the vortex frequency change to the boundary layer state on the blade surfaces.
A Theory for Wake-Induced Transition
Authors: Mayle R. E. , Dullenkopf K.
Published: June 2008

A theory for transition from laminar to turbulent flow as the result of unsteady, periodic passing of turbulent wakes in the free stream is developed using Emmons’ transition model. Comparisons made to flat plate boundary layer measurements and airfoil heat transfer measurements confirm the theory.
Effects of Adverse Pressure Gradients on the Nature and Length of Boundary Layer Transition
Authors: Walker G. J. , Gostelow J. P.
Published: June 2008

Existing transition models are surveyed and deficiencies in previous predictions, which seriously overestimate transition length under an adverse pressure gradient, are discussed. A new model for transition in an adverse pressure gradient situation is proposed and experimental results are provided that confirm its validity. A correlation for transition length is advanced that incorporates both Reynolds number and pressure gradient effects. Under low free-stream turbulence conditions the basic mechanism of transition is laminar instability. There are, however, physical differences between zero and adverse pressure gradients. In the former case, transition occurs randomly, due to the breakdown of laminar instability waves in sets. For an adverse pressure gradient, the Tollmien–Schlichting waves appear more regularly with a well-defined spectral peak. As the adverse pressure gradient is increased from zero to the separation value the flow evolves continuously from random to periodic behavior and the dimensionless transition length progressively decreases.
Unsteady Transition in an Axial-Flow Turbine: Part 2—Cascade Measurements and Modeling
Authors: Addison J. S. , Hodson H. P.
Published: June 2008

Part 1 of this paper reanalyzed previously published measurements from the rotor of a low-speed, single-stage, axial-flow turbine, which highlighted the unsteady nature of the suction surface transition process. Part 2 investigates the significance of the wake jet and the unsteady frequency parameter. Supporting experiments carried out in a linear cascade with varying inlet turbulence are described, together with a simple unsteady transition model explaining the features seen in the turbine.
Compressor Blade Boundary Layers: Part 2—Measurements With Incident Wakes
Authors: Dong Y. , Cumpsty N. A.
Published: June 2008

This paper follows directly from Part 1 by the same authors and describes measurements of the boundary layer on a supercritical-type compressor blade with wakes from a simulated moving upstream blade row convected through the passage. (The blades and the test facilities together with the background are described in Part 1). The results obtained with the wakes are compared to those with none for both low and high levels of inlet turbulence. The transition process and boundary layer development are very different in each case, though the overall momentum thickness at the trailing edge is fairly similar. None of the models for transition is satisfactory when this is initiated by moving wakes.
The Measurement of Boundary Layers on a Compressor Blade in Cascade: Part 4—Flow Fields for Incidence Angles of −1.5 and −8.5 Degrees
Authors: Zierke W. C. , Deutsch S.
Published: June 2008

Measurements, made with laser Doppler velocimetry, about a double-circular-arc compressor blade in cascade are presented for −1.5 and −8.5 deg incidence angles and a chord Reynolds number near 500,000. Comparisons between the results of the current study and those of our earlier work at a 5.0 deg incidence are made. It is found that in spite of the relative sophistication of the measurement techniques, transition on the pressure surface at the −1.5 deg incidence is dominated by a separation “bubble” too small to be detected by the laser Doppler velocimeter. The development of the boundary layers at −1.5 and 5.0 deg is found to be similar. In contrast to the flow at these two incidence angles, the leading edge separation bubble is on the pressure surface for the −8.5 deg incidence. Here, all of the measured boundary layers on the pressure surface are turbulent—but extremely thin—while on the suction surface, a laminar separation/turbulent reattachment bubble lies between roughly 35 percent and 60 percent chord. This bubble is quite thin, and some problems in interpreting the backflow data are discussed.
Viscous Flow in a Controlled Diffusion Compressor Cascade With Increasing Incidence
Authors: Elazar Y. , Shreeve R. P.
Published: June 2008

A detailed two-component LDV mapping of the flow through a controlled diffusion compressor cascade at low Mach number ( ~ 0.25) and Reynolds number of about 7 × 105, at three inlet air angles from design to near stall, is reported. It was found that the suction-side boundary layer reattached turbulent after a laminar separation bubble, and remained attached to the trailing edge even at the highest incidence, at which losses were 3 to 4 times the minimum value for the geometry. Boundary layer thickness increased to fill 20 percent of the blade passage at the highest incidence. Results for pressure-side boundary layer and near-wake also are summarized. Information sufficient to allow preliminary assessment of viscous codes is tabulated.
An Improved Incidence Losses Prediction Method for Turbine Airfoils
Authors: Moustapha S. H. , Kacker S. C. , Tremblay B.
Published: June 2008

The off-design performance of axial turbines is usually predicted by calculating the incidence losses using empirical correlations. Periodic review and improvement to these prediction methods, to reflect recent turbine designs and test results, are essential for the accurate assessment of losses in turbine airfoils. The purpose of the present work is to evaluate existing turbine incidence loss correlations, and present an improved prediction method for profile and secondary losses at off-design conditions which correlates better with the available experimental results. The incidence losses are shown to be a function of leading edge diameter, pitch, aspect ratio and channel convergence.
The Trailing Edge Loss of Transonic Turbine Blades
Authors: Denton J. D. , Xu L.
Published: June 2008

Trailing edge loss is one of the main sources of loss for transonic turbine blades, contributing typically 1/3 of their total loss. Transonic trailing edge flow is extremely complex, the basic flow pattern is understood but methods of predicting the loss are currently based on empirical correlations for the base pressure. These correlations are of limited accuracy. Recent findings that the base pressure and loss can be reasonably well predicted by inviscid Euler calculations are justified and explained in this paper. For unstaggered choked blading, it is shown that there is a unique relationship between the back pressure and the base pressure and any calculation that conserves mass, energy and momentum should predict this relationship and the associated loss exactly. For realistic staggered blading, which operates choked but with subsonic axial velocity, there is also a unique relationship between the back pressure and the base pressure (and hence loss) but the relationship cannot be quantified without knowing a further relationship between the base pressure and the average suction surface pressure downstream of the throat. Any calculation that conserves mass, energy and momentum and also predicts this average suction surface pressure correctly will again predict the base pressure and loss. Two-dimensional Euler solutions do not predict the suction surface pressure exactly because of shock smearing but nevertheless seem to give reasonably accurate results. The effects of boundary layer thickness and trailing edge coolant ejection are considered briefly. Coolant ejection acts to reduce the mainstream loss. It is shown that suction surface curvature downstream of the throat may be highly beneficial in reducing the loss of blades with thick trailing edges operating at high subsonic or low supersonic outlet Mach numbers.
An Experimental Assessment of the Influence of Downstream Conditions on the Performance of a Transonic Turbine Nozzle of High Turning
Authors: Schaub U. W. , Williamson R. G. , Huot J.-P. , Moustapha S. H.
Published: June 2008

The present work was conducted to extend information derived in a previous study where it was found that the performance of a transonic turbine nozzle of high turning angle was significantly affected by the design and operation of the downstream rotor and the detailed characteristics of the nozzle flow field. Measurements obtained in an operating stage environment were compared with information from simplified test arrangements. In particular, it was possible to make direct comparisons of static pressure distributions at the vane root as affected by downstream conditions. Detailed exit flow surveys indicated areas susceptible to the influence of rotor presence. Probe interaction effects also were assessed in a special series of tests conducted over a representative range of Mach numbers. Finally, experiments were performed with a perforated plate simulating rotor presence, in an attempt to determine the limitations of such simplified test arrangements in the generation of pertinent data. The results shed further light on a complex subject, and have been aimed at practical validation of test arrangements and techniques.
The Use of Circumferentially Varying Stagger Guide Vanes in an Axial Flow Pump or Compressor
Authors: Horlock J. H.
Published: June 2008

An actuator disk analysis is given of the flow through a guide vane and rotor combination. It is shown that changes in total pressure across the rotor are in general related to circumferential variations in guide vane outlet angle. In particular known variations in inlet total pressure may be eliminated by suitable circumferential changes in guide vane stagger.
Inlet Distortion Generated Periodic Aerodynamic Rotor Response
Authors: Manwaring S. R. , Fleeter S.
Published: June 2008

Fundamental inlet distortion-generated rotor blade row unsteady aerodynamics, including the effects of both the detailed aerodynamic forcing function for the first time and steady loading are experimentally investigated in an extensively instrumented axial-flow research compressor. A two-per-rev forcing function with three gust amplitude ratios is generated. On the rotor blade pressure surface, the unsteady pressure nondimensionalization compresses the magnitude data with mean flow incidence angle. This is not the case on the higher camber suction surface. These pressure surface unsteady data are primarily affected by the steady loading level, whereas the suction surface unsteady data are a function of the steady loading level and distribution as well as the gust amplitude ratio. In addition, a design inlet distortion blade surface unsteady pressure correlation is considered.
Visualization Studies in Rotating Disk Cavity Flows
Authors: Nordquist J. , Abrahamson S. , Wechkin J. , Eaton J.
Published: June 2008

An experimental study was performed in a simplified turbine disk cavity consisting of a single disk rotating near a stationary flat plate, bounded radially by an axial seal fixed to the plate. The disk Reynolds number was 6.2 × 105. Cooling flow was supplied axially at a dimensionless radius of 0.33. Flow visualization showed boundary layers on both the rotating and stationary disks with the core between the layers rotating in solid body motion. At intermediate cooling flow rates, large vortical structures aligned with the disk spin axis and spanning the core were observed.
Numerical Simulation of the Two-Dimensional Viscous Compressible Flow in Blade Cascades Using a Solution-Adaptive Unstructured Mesh
Authors: Side´n G. L. D. , Dawes W. N. , Albra˚ten P. J.
Published: June 2008

An explicit finite element procedure has been coupled with an automatic generation procedure for mesh-adaptive steady-state simulations of two-dimensional viscous compressible flows in cascades. Turbulence is modeled by a two-layer algebraic eddy viscosity model. Results show good behavior in comparison with measurements and results of a conventional H-mesh viscous flow solver. Computed loss approaches measured loss as the mesh is refined. Currently, the unstructured solver suffers in efficiency terms because the automatic mesh generator tends to produce inefficient equilateral triangles in the regions of shock waves and boundary layers where stretched elements would be more appropriate. This means that, at least for the Navier–Stokes equations, the unstructured approach is not yet competitive with conventional structured techniques. Nevertheless, this will change once the key advantages of geometric flexibility and user-independent solutions force rapid development.
Throughflow Theory for Nonaxisymmetric Turbomachinery Flow: Part I—Formulation
Authors: Dring R. P. , Oates G. C.
Published: June 2008

Throughflow theory has been limited in its applicability and in its accuracy by the fact that it has not historically been cast in a form that can account for the non-axisymmetries that naturally arise in turbomachinery flow due to the presence of finite numbers of rotor and stator airfoils. The attempt to circumvent this limitation by the introduction of an aerodynamic blockage factor has been demonstrated in earlier work to produce fundamental inconsistencies in the calculation, which lead to significant errors in the regions of the flow where the nonaxisymmetries are severe. The formulation in Part I of the present work is a derivation of a system of through-flow equations for nonaxisymmetric flow. A benchmark database is used in Part II to provide input to the calculation and to help identify the dominant terms. It is demonstrated that the dominant effect of nonaxisymmetry is contained in two terms that relate the total pressure of the averaged flow to the mass-averaged total pressure. It also is demonstrated that the present formulation produces a result that is more accurate than that obtained with the historical blockage-based formulation.
A Parametric Study of Radial Turbomachinery Blade Design in Three-Dimensional Subsonic Flow
Authors: Ghaly W. S.
Published: June 2008

An aerodynamic design method is described and used to implement a parametric study of radial turbomachinery blade design in three-dimensional subsonic flow. Given the impeller hub and shroud, the number of blades and their stacking position, the design method gives the detailed blade shape, flow, and pressure fields that would produce a prescribed tangentially averaged swirl schedule. The results from that study show that decreasing the number of blades increases the blade wrap, and that the blade loading is strongly affected by the rate of change of mean swirl along the mean streamlines. The results also show that the blade shape and the pressure field are rather sensitive to the prescribed mean swirl schedule, which suggests that, by carefully tailoring the swirl schedule, one might be able to control the blade shape and the pressure field and hence secondary flow.
A Three-Dimensional Inverse Method for Turbomachinery: Part I—Theory
Authors: Borges J. E.
Published: June 2008

There are surprisingly few inverse methods described in the literature that are truly three dimensional. Here, one such method is presented. This technique uses as input a prescribed distribution of the mean swirl, i.e., radius times mean tangential velocity, given throughout the meridional section of the machine. In the present implementation the flow is considered inviscid and incompressible and is assumed irrotational at the inlet to the blade row. In order to evaluate the velocity field inside the turbomachine, the blades (supposed infinitely thin) are replaced by sheets of vorticity, whose strength is related to the specified mean swirl. Some advice on the choice of a suitable mean swirl distribution is given. In order to assess the usefulness of the present procedure, it was decided to apply it to the design of an impeller for a low-speed radial-inflow turbine. The results of the tests are described in the second part of this paper.
A Three-Dimensional Inverse Method for Turbomachinery: Part II—Experimental Verification
Authors: Borges J. E.
Published: June 2008

The performance of an impeller of a low-speed radial-inflow turbine, designed using a three-dimensional inverse technique, was evaluated experimentally. This performance was compared with that achieved by a rotor typical of the present technology. Besides measuring overall quantities, in special efficiency, some traverses of flow velocity were carried out. The results of the tests showed that the new design had a peak total-to-static efficiency 1.4 points better than the conventional build. The traverses indicated that the level of swirl at exhaust of the new impeller was only half as big as that for the conventional rotor, in spite of the fact that both impellers were designed to have zero swirl at outlet. It is also shown that the rotor loss for the new impeller is considerably lower than for the conventional wheel. This research points to the desirability of using a three-dimensional inverse method for the design of turbomachines with significant three-dimensional flows.
Multi-Airfoil Navier–Stokes Simulations of Turbine Rotor–Stator Interaction
Authors: Rai M. M. , Madavan N. K.
Published: June 2008

An accurate numerical analysis of the flows associated with rotor–stator configurations in turbomachinery can be extremely helpful in optimizing the performance of turbomachinery. In this study the unsteady, thin-layer, Navier–Stokes equations in two spatial dimensions are solved on a system of patched and overlaid grids for an axial-turbine rotor–stator configuration. The governing equations are solved using a finite-difference, upwind algorithm that is set in an iterative, implicit framework. Results are presented in the form of pressure contours, time-averaged pressures, unsteady pressures, amplitudes, and phase. The numerical results are compared with experimental data and the agreement is found to be good. The results are also compared with those of an earlier study, which used only one rotor and one stator airfoil. The current study uses multiple rotor and stator airfoils and a pitch ratio that is much closer to the experimental ratio. Consequently, the results of this study are found to be closer to the experimental data.
The Extension and Application of Three-Dimensional Time-Marching Analyses to Incompressible Turbomachinery Flows
Authors: Walker P. J. , Dawes W. N.
Published: June 2008

Conventional time-marching flow solvers perform poorly when integrating compressible flow equations at low Mach number levels. This is shown to be due to unfavorable interaction between long-wavelength errors and the inflow and outflow boundaries. Chorin’s method of artificial compressibility is adopted to extend the range of Denton’s inviscid flow solver and Dawes’ three-dimensional Navier–Stokes solver to zero Mach number flows. The paper makes a new contribution by showing how to choose the artificial acoustic speed systematically to optimize convergence rate with regard to the error wave–boundary interactions. Applications to a turbine rotor and generic water pump geometry are presented.
Three-Dimensional Solutions for Inviscid Incompressible Flow in Turbomachines
Authors: Abdallah S. , Smith C. F.
Published: June 2008

A primitive variable formulation is used for the solution of the incompressible Euler equation. In particular, the pressure Poisson equation approach using a nonstaggered grid is considered. In this approach, the velocity field is calculated from the unsteady momentum equation by marching in time. The continuity equation is replaced by a Poisson-type equation for the pressure with Neumann boundary conditions. A consistent finite-difference method, which insures the satisfaction of a compatibility condition necessary for convergence, is used in the solution of the pressure equation on a nonstaggered grid. Numerical solutions of the momentum equations are obtained using the second-order upwind differencing scheme, while the pressure Poisson equation is solved using the line successive overrelaxation method. Three turbo-machinery rotors are tested to validate the numerical procedure. The three rotor blades have been designed to have similar loading distributions but different amounts of dihedral. Numerical solutions are obtained and compared with experimental data in terms of the velocity components and exit swirl angles. The computed results are in good agreement with the experimental data.
Axial Flow Compressor Design Optimization: Part I—Pitchline Analysis and Multivariable Objective Function Influence
Authors: Massardo A. , Satta A.
Published: June 2008

The design of an axial flow compressor stage has been formulated as a nonlinear mathematical programming problem with the objective of minimizing the aerodynamic losses and the weight of the stage, while maximizing the compressor stall margin. Aerodynamic as well as mechanical constraints are considered in the problem formulation. A method of evaluating the objective function and constraints of the problem with a pitchline analysis is presented. The optimization problem is solved by using the penalty function method in which the Davidon-Fletcher-Powell variable metric minimization technique is employed. Designs involving the optimization of efficiency, weight of the stage, and stall margin are presented and the results discussed with particular reference to a multivariable objective function.
Axial Flow Compressor Design Optimization: Part II—Throughflow Analysis
Authors: Massardo A. , Satta A. , Marini M.
Published: June 2008

A new technique is presented for the design optimization of an axial-flow compressor stage. The procedure allows for optimization of the complete radial distribution of the geometry, since the variables chosen to represent the three-dimensional geometry of the stage are coefficients of suitable polynomials. Evaluation of the objective function is obtained with a throughflow calculation, which has acceptable speed and stability qualities. Some examples are given of the possibility to use the procedure both for redesign and, together with what was presented in Part I, for the complete design of axial-flow compressor stages.
Unsteady Lifting Surface Theory for a Rotating Cascade of Swept Blades
Authors: Kodama H. , Namba M.
Published: June 2008

A lifting surface theory is developed to predict the unsteady three-dimensional aerodynamic characteristics for a rotating subsonic annular cascade of swept blades. A discrete element method is used to solve the integral equation for the unsteady blade loading. Numerical examples are presented to demonstrate effects of the sweep on the blade flutter and on the acoustic field generated by interaction of rotating blades with a convected sinusoidal gust. It is found that increasing the sweep results in decrease of the aerodynamic work on vibrating blades and also remarkable reduction of the modal acoustic power of lower radial orders for both forward and backward sweeps.
Surface Injection Effect on Mass Transfer From a Cylinder in Crossflow: A Simulation of Film Cooling in the Leading Edge Region of a Turbine Blade
Authors: Karni J. , Goldstein R. J.
Published: June 2008

A naphthalene sublimation technique is used to study the effect of surface injection on the mass (heat) transfer from a circular cylinder in crossflow. Using a heat/mass transfer analogy the results can be used to predict film cooling effects in the leading edge region of a turbine blade. Air injection through one row of circular holes is employed in the stagnation region of the cylinder. Streamwise and spanwise injection inclinations are studied separately, and the effects of blowing rate and injection location relative to the cylinder front stagnation line are investigated. Streamwise injection produces significant mass transfer increases downstream of the injection holes, but a relatively small increase is observed between holes, normal to the injection direction. The mass transfer distribution, measured with spanwise injection through holes located near the cylinder front stagnation line, is extremely sensitive to small changes in the injection hole location relative to stagnation. When the centers of the spanwise injection holes are located 5 deg or more from the stagnation line, the holes lie entirely on one side of the stagnation line and the injection affects the mass transfer only on that side of the cylinder, approaching the pattern observed with streamwise injection.
Effects of an Embedded Vortex on Injectant From a Single Film-Cooling Hole in a Turbulent Boundary Layer
Authors: Ligrani P. M. , Williams W.
Published: June 2008

Effects of embedded longitudinal vortices on heat transfer in turbulent boundary layers with injection from a single film-cooling hole are described. These results were obtained at a free-stream velocity of 10 m/s, with a film-cooling hole inclined 30 deg to the horizontal and a blowing ratio of about 0.50. The ratio of vortex core diameter to injection hole diameter was 1.58, and the ratio of circulation to injection velocity time hole diameter was about 3.16. Coolant distributions and spatially resolved heat transfer measurements indicate that injection hole centerlines must be at least 2.9–3.4 vortex core diameters away from the vortex center in the lateral direction to avoid significant alterations to wall heat transfer and distributions of film coolant. Under these circumstances, protection from film cooling is evident at least up to 55 hole diameters downstream of injection. When the injection hole is closer to the vortex center, secondary flows convect most injectant into the vortex upwash and thermal protection from film cooling is destroyed for streamwise locations from the injection hole greater than 17.5 hole diameters.