bluuebunny/crossref_metadata_2025 huggingface dataset:
An Experimental Study of Local Wall Shear Stress, Surface Static Pressure, and Flow Visualization Upstream, Alongside, and Downstream of a Blade Endwall Corner
Authors: Abdulla A. K. , Bhargava R. K. , Raj R.
Published: June 2008

The experimental study reported in this paper was performed to acquire information on the distribution of wall shear stress and surface static pressure in a blade endwall corner. The blade endwall corner region investigated was divided into three sections: 0.4 chord length upstream of the blade leading edge, inside the endwall corner region, and one chord length downstream of the blade trailing edge. The maximum increases in the values of wall shear stress were found to exist on the endwall, in the corner region, between the blade leading edge and the location of maximum blade thickness (≈ 140 percent maximum increase, compared to its far upstream value, at x/D = 6). Surface flow visualization defined the boundaries of the vortex system and provided information on the direction and magnitude of the wall shear stress. The acquired results indicated that the observed variations of wall shear stress and surface static pressure were significantly influenced by the interaction of secondary flows with pressure gradients induced by the presence of blade curvature.
Analysis of Unsteady Compressible Viscous Layers
Authors: Power G. D. , Verdon J. M. , Kousen K. A.
Published: June 2008

The development of an analysis to predict the unsteady compressible flows in blade boundary layers and wakes is presented. The equations that govern the flows in these regions are transformed using an unsteady turbulent generalization of the Levy–Lees transformation. The transformed equations are solved using a finite difference technique in which the solution proceeds by marching in time and in the streamwise direction. Both laminar and turbulent flows are studied, the latter using algebraic turbulence and transition models. Laminar solutions for a flat plate are shown to approach classical asymptotic results for both high and low-frequency unsteady motions. Turbulent flat-plate results are in qualitative agreement with previous predictions and measurements. Finally, the numerical technique is also applied to the stator and rotor of a low-speed turbine stage to determine unsteady effects on surface heating. The results compare reasonably well with measured heat transfer data and indicate that nonlinear effects have minimal impact on the mean and unsteady components of the flow.
Secondary Flow and Loss Distribution in a Radial Compressor With Untwisted Backswept Vanes
Authors: Sipos G.
Published: June 2008

The unshrouded impeller and the vaneless diffuser of a single-stage radial compressor have been investigated at three flow rates. Three-dimensional velocities and pressures were measured at a tip speed of 84 m/s by an L2F-velocimeter, a slanted single hotwire probe, and piezoresistive pressure transducers. The measurements show that upstream of the blading the averaged meridional inlet flow angle is about 54 deg and a periodic variation of the meridional flow angle of about 25 deg occurs near the casing wall. Further, an inlet vortex in the clockwise direction appears and an initial whirl is induced. The specific work of the initial whirl corresponds to approximately 12 percent of the enthalpy losses between inlet pipe and diffuser outlet. In the beginning of the passage, the inlet vortex is suppressed and a solid body vortex in the counterclockwise direction can be observed. At the outlet, a heavy flow deceleration at the blade suction side with subsequent separation can be seen. Increasing the flow rate decreases the wake and causes a more uniform loss distribution in this area. The measured secondary vortex flow and rotary stagnation pressure gradients are compared with test results from impellers with inducer. The incidence of the investigated impeller is greater than that of the impellers with inducer, but the wake-jet outlet flows are very similar. Inlet losses could be reduced by improving incidence angles by matching the blade angles to the inlet flow angles. Smaller blade angles at the shroud would reduce or eliminate separation at the leading edge, and the resulting reduction in low-momentum fluid along the suction surface would help to avoid separation on that surface near the outlet.
Centrifugal Compressor Inlet Guide Vanes for Increased Surge Margin
Authors: Rodgers C.
Published: June 2008

This paper describes the results of compressor rig testing with a moderately high specific speed, high inducer Mack number, single-stage centrifugal compressor, with a vaned diffuser, and adjustable inlet guide vanes (IGVs). The results showed that the high-speed surge margin was considerably extended by the regulation of the IGVs, even though the vaned diffuser was apparently operating stalled. Simplified one-dimensional analysis of the impeller and diffuser performances indicated that at inducer tip Mach numbers approaching and exceeding unity, the high-speed surge line was triggered by inducer stall. Also, IGV regulation increased impeller stability. This permitted the diffuser to operate stalled, providing the net compression system stability remained on a negative slope.
A Two-Dimensional Flow Analysis Model for Designing a Nozzle-less Volute Casing for Radial Flow Gas Turbines
Authors: Owarish H. O. , Ilyas M. , Bhinder F. S.
Published: June 2008

This paper describes a two-dimensional model for flow analysis and design of a single entry nozzle-less volute casing for inward flow radial turbines. The model takes into account the cross-sectional shape of the casing by dividing it into a number of segments, which are further subdivided into control volumes. Changes in flow properties are calculated by considering the changes in momenta of fluid in the tangential and radial directions across each control volume. The model has been computerized using Fortran 77 for the IBM AT or 100 percent compatible microcomputers.
An Axial Turbobrake
Authors: Goodisman M. I. , Oldfield M. L. G. , Kingcombe R. C. , Jones T. V. , Ainsworth R. W. , Brooks A. J.
Published: June 2008

The “Axial Turbobrake” (patent applied for) is a novel turbomachine that can be used to absorb power generated by test turbines. Unlike a compressor, there is no pressure recovery through the turbobrake. This simplifies the aerodynamic design and enables high-stage loadings to be achieved. The blades used have high-turning two-dimensional profiles. This paper describes a single-stage axial turbobrake, which is driven by the exhaust gas of the test turbine and is isolated from the turbine by a choked throat. In this configuration no fast-acting controls are necessary as the turbobrake operates automatically with the turbine flow. Tests on a 0.17 scale model show that the performance is close to that predicted by a simple two-dimensional theory, and demonstrate that the turbobrake power absorption can be controlled and hence matched to that typically produced by the first stage of a modern highly loaded transonic turbine. A full-size axial turbobrake will be used in a short-duration rotating turbine experiment in an Isentropic Light Piston Tunnel at RAE Pyestock.
Rim Sealing of Rotor–Stator Wheelspaces in the Absence of External Flow
Authors: Chew J. W. , Dadkhah S. , Turner A. B.
Published: June 2008

Sealing of the cavity formed between a rotating disk and a stator in the absence of a forced external flow is considered. In these circumstances the pumping action of the rotating disk may draw fluid into the cavity through the rim seal. Minimum cavity throughflow rates required to prevent such ingress are estimated experimentally and from a mathematical model. The results are compared with other workers’ measurements. Measurements for three different types of rim seal are reported for a range of seal clearances and for rotational Reynolds numbers up to 3 × 106. The mathematical model is found to correlate the experimental data reasonably well.
An Experimental Investigation of Nozzle-Exit Boundary Layers of Highly Heated Free Jets
Authors: Lepicovsky J.
Published: June 2008

An experimental investigation of the effects of nozzle operating conditions on the development of nozzle-exit boundary layers of highly heated air free jets is reported in this paper. The total pressure measurements in the nozzle-exit boundary layer were obtained at a range of jet Mach numbers from 0.1 to 0.97 and jet total temperatures up to 900 K. The analysis of results shows that the nozzle-exit laminar boundary-layer development depends only on the nozzle-exit Reynolds number. For the nozzle-exit turbulent boundary layer, however, it appears that the effects of the jet total temperature on the boundary-layer integral characteristics are independent from the effect of the nozzle-exit Reynolds number. This surprising finding has not yet been reported. Further, laminar boundary-layer profiles were compared with the Pohlhausen solution for a flat-wall converging channel and an acceptable agreement was found only for low Reynolds numbers. For turbulent boundary layers, the dependence of the shape factor on relative Mach numbers at a distance of one momentum thickness from the nozzle wall resembles Spence’s prediction. Finally, the calculated total pressure loss coefficient was found to depend on the nozzle-exit Reynolds number for the laminar nozzle-exit boundary layer, while for the turbulent exit boundary layer this coefficient appears to be constant.
Influence of Unsteady Effects on the Measurements in a Transonic Axial Compressor
Authors: Paulon J. , Zhang Zhifang , Jia Pingfang , Meng Jingfei
Published: June 2008

Interaction phenomena between rotor and stator are unavoidable in advanced compressors and their effects increase with the performance of the turbomachines. Until now, it was not possible to quantify the interaction effects, but with the development of three-dimensional unsteady computation codes in a complete stage, it is possible to know, in detail, the flow field through the machine and to make evident and to explain the difficulties encountered in measuring the flow parameters. A study has been conducted in this way at ONERA, on an axial transonic compressor stage. The computations have been made with a simulation of the losses; in this manner, the overall computed and measured performances of the compressor are the same. A detailed analysis of the unsteady computation results makes evident, between rotor and stator, large variations of some parameters of the flow as a function of time, but also as a function of the axial and tangential relative position of steady probes and stator blades. Unsteady measurements made on another transonic machine confirm the indications given by these computations.
The Numerical Simulation of a High-Speed Axial Flow Compressor
Authors: Mulac R. A. , Adamczyk J. J.
Published: June 2008

The advancement of high-speed axial flow multistage compressors is impeded by a lack of detailed flow field information. Recent developments in compressor flow modeling and numerical simulation have the potential to provide needed information in a timely manner. This paper, which consists of two parts, will explore this topic. The first part will address the development of a computer program to solve the viscous form of the average-passage equation system for multistage turbomachinery. Programming issues such as in-core versus out-of-core data storage and CPU utilization (parallelization, vectorization, and chaining) will be addressed. Code performance will be evaluated through the simulation of the first four stages of a five-stage, high-speed, axial flow compressor on a CRAY Y-MP8/8128 computer. The second part will address the flow physics, which can be obtained from the numerical simulation. In particular, an examination of the end-wall flow structure will be made, and its impact on blockage distribution assessed.
Turbulence Modeling for Secondary Flow Prediction in a Turbine Cascade
Authors: Cleak J. G. E. , Gregory-Smith D. G.
Published: June 2008

Predictions of secondary flow in an axial turbine cascade have been made using three different turbulence models: mixing length, a one-equation model and a k–ε mixing length hybrid model. The results are compared with results from detailed measurements, not only by looking at mean flow velocities and total pressure loss, but also by assessing how well turbulence quantities are predicted. It is found that the turbulence model can have a big influence on the mean flow results, with the mixing length model giving generally the best mean flow. None of the models give good predictions of the turbulent shear stresses in the vortex region, although the k–ε model gives quite good turbulent kinetic energy values. The one-equation model is the only one to contain a transition criterion. The importance of such a criterion is illustrated, but the present one needs development to give reliable predictions in the complex flow within a blade passage.
Unsteady Navier–Stokes Simulation of Transonic Cascade Flow Using an Unfactored Implicit Upwind Relaxation Scheme With Inner Iterations
Authors: Furukawa M. , Nakano T. , Inoue M.
Published: June 2008

An implicit upwind scheme has been developed for Navier–Stokes simulations of unsteady flows in transonic cascades. The two-dimensional, Reynolds-averaged Navier–Stokes equations are discretized in space using a cell-centered finite volume formulation and in time using the Euler implicit method. The inviscid fluxes are evaluated using a highly accurate upwind scheme based on a TVD formulation with the Roe’s approximate Riemann solver, and the viscous fluxes are determined in a central differencing manner. The algebraic turbulence model of Baldwin and Lomax is employed. To simplify grid generations, a zonal approach with a composite zonal grid system is implemented, in which periodic boundaries are treated as zonal boundaries. A new time linearization of the inviscid fluxes evaluated by Roe’s approximate Riemann solver is presented in detail. No approximate factorization is introduced, and unfactored equations are solved by a pointwise relaxation method. To obtain time-accurate solutions, 30 linear iterations are performed at each time step. Numerical examples are presented for unsteady flows in a transonic turbine cascade where periodic unsteadiness is caused by the trailing edge vortex shedding.
Three-Dimensional Navier–Stokes Computation of Turbomachinery Flows Using an Explicit Numerical Procedure and a Coupled k–ε Turbulence Model
Authors: Kunz R. F. , Lakshminarayana B.
Published: June 2008

An explicit, three-dimensional, coupled Navier–Stokes/k–ε technique has been developed and successfully applied to complex internal flow calculations. Several features of the procedure, which enable convergent and accurate calculation of high Reynolds number two-dimensional cascade flows, have been extended to three dimensions, including a low Reynolds number compressible form of the k–ε turbulence model, local time-step specification based on hyperbolic and parabolic stability requirements, and eigenvalue and local velocity scaling of artificial dissipation operators. A flux evaluation procedure, which eliminates the finite difference metric singularity at leading and trailing edges on H- and C-grids, is presented. The code is used to predict the pressure distribution, primary velocity, and secondary flows in an incompressible, turbulent curved duct flow for which CFD validation quality data are available. Also, a subsonic compressor rotor passage, for which detailed laser, rotating hot-wire, and five-hole pressure probe measurements have been made is computed. Detailed comparisons between predicted and measured core flow and near-wall velocity profiles, wake profiles, and spanwise mixing effects downstream of the rotor passage are presented for this case. It is found that the technique provides accurate and convergent engineering simulation of these complex turbulent flows.
Three-Dimensional Flow and Mixing in an Axial Flow Compressor With Different Rotor Tip Clearances
Authors: Goto A.
Published: June 2008

The effect of difference in rotor tip clearance on the mean flow fields and unsteadiness and mixing across a stator blade row were investigated using hot-wire anemometry, pressure probes, flow visualization, and the ethylene tracer-gas technique on a single-stage axial flow compressor. The structure of the three-dimensional flow fields was discussed based on results of experiments using the 12-orientation single slanted hotwire technique and spectrum analysis of velocity fluctuation. High-pass filtered measurements of turbulence were also carried out in order to confirm small-scale velocity fluctuation, which is more realistically referred to as turbulence. The span-wise distribution of ethylene gas spreading, estimated by the measured small-scale velocity fluctuation at the rotor exit, agreed quite well with that which was experimentally measured. This fact suggests the significant role of turbulence, generated within the rotor, in the mixing process across the downstream stator. The value of the maximum mixing coefficient in the tip region was found to increase linearly as the tip clearance became enlarged, starting from the value at midspan.
Inviscid-Viscous Coupled Solution for Unsteady Flows Through Vibrating Blades: Part 2—Computational Results
Authors: He L. , Denton J. D.
Published: June 2008

A quasi-three-dimensional inviscid-viscous coupled approached has been developed for unsteady flows around oscillating blades, as described in Part 1. To validate this method, calculations for several steady and unsteady flow cases with strong inviscid-viscous interactions are performed, and the results are compared with the corresponding experiments. Calculated results for unsteady flows around a biconvex cascade and a fan tip section highlight the necessity of including viscous effects in predictions of turbomachinery blade flutter at transonic flow conditions.
A Study of Stall in a Low Hub–Tip Ratio Fan
Authors: Soundranayagam M. , Elder R. L.
Published: June 2008

An investigation has been carried out in order to define the process of rotating stall inception in a low-speed, low hub–tip ratio fan. Based on elementary cascade analysis, the fan would be expected to stall from the root. However, considerable experimental evidence indicates that tip stall is more frequently incurred. Although the analysis has been undertaken for a specific fan, it is considered to be representative of a broad range of machines. The analysis has involved two primary considerations, first the effect of streamtube contraction, which has been studied theoretically, and second, real flow effects (those not contained in the theoretical model), which have been studied experimentally. The study of streamtube contraction indicates that the root rematches to a more stable operating point, thus alleviating some of the problems in that region. The experimental investigation was undertaken on an isolated rotor, with successive build modifications to increase the likelihood of rotating stall inception at the root. It was apparent that real fluid effects tended to steepen the root pressure rise characteristic, thus enhancing the stability in that region. The performance of the fan at the tip tended to be poor, providing a pressure characteristic with a lower negative gradient than anticipated, indicating less stability than simple flow models would suggest. Hot-wire flow mapping at the rotor exit supported the overall conclusion that the rotor showed a strong reluctance to stall at the root apparently due to “centrifuging” of the boundary layer toward the tip.
Validation of a Numerical Method for Unsteady Flow Calculations
Authors: Giles M. , Haimes R.
Published: June 2008

This paper describes and validates a numerical method for the calculation of unsteady inviscid and viscous flows. A companion paper compares experimental measurements of unsteady heat transfer on a transonic rotor with the corresponding computational results. The mathematical model is the Reynolds-averaged unsteady Navier–Stokes equations for a compressible ideal gas. Quasi-three-dimensionality is included through the use of a variable streamtube thickness. The numerical algorithm is unusual in two respects: (a) For reasons of efficiency and flexibility, it uses a hybrid Navier–Stokes/Euler method, and (b) to allow for the computation of stator/rotor combinations with arbitrary pitch ratio, a novel space–time coordinate transformation is used. Several test cases are presented to validate the performance of the computer program, UNSFLO. These include: (a) unsteady, inviscid flat plate cascade flows (b) steady and unsteady, viscous flat plate cascade flows, (c) steady turbine heat transfer and loss prediction. In the first two sets of cases comparisons are made with theory, and in the third the comparison is with experimental data.
Blade Row Interaction in a Multistage Low-Pressure Turbine
Authors: Arndt N.
Published: June 2008

The objective of this work was to enhance the understanding of unsteady flow phenomena in multistage low-pressure turbines. For this purpose, hot-film probe measurements were made downstream of every rotor blade row of a five-stage low-pressure turbine. Rotor–rotor interaction and stator–rotor interaction were observed to have a profound influence on the flow through the low-pressure turbine. Interaction of rotors of different turbine stages occurred owing to the influence of the wakes shed by one rotor blade row upon the flow through the next downstream rotor blade row. This wake-induced rotor–rotor interaction resulted in strongly amplitude-modulated periodic and turbulent velocity fluctuations downstream of every rotor blade row with the exception of the most upstream one. Significantly different wake depths and turbulence levels measured downstream of every rotor blade row at different circumferential positions evidenced the effect of the circumferentially nonuniform stator exit flow upon the next downstream rotor blade row. Stator-rotor interaction also strongly influenced the overturning and the under-turning of the rotor wakes, caused by the rotor secondary flows, in the rotor endwall regions. Low rotor wake overturning and underturning, i.e., reduced rotor secondary flow influence, were observed to correlate well with low rotor wake turbulence levels.
Rotor Blade Unsteady Aerodynamic Gust Response to Inlet Guide Vane Wakes
Authors: Manwaring S. R. , Fleeter S.
Published: June 2008

A series of experiments is performed in an extensively instrumented axial flow research compressor to investigate the fundamental flow physics of wake-generated periodic rotor blade row unsteady aerodynamics at realistic values of the reduced frequency. Unique unsteady data are obtained that describe the fundamental unsteady aerodynamic gust interaction phenomena on the first-stage rotor blades of a research axial flow compressor generated by the wakes from the inlet guide vanes. In these experiments, the effects of steady blade aerodynamic loading and the aerodynamic forcing function, including both the transverse and chordwise gust components, and the amplitude of the gusts, are investigated and quantified.
Pressure Fluctuation on Casing Wall of Isolated Axial Compressor Rotors at Low Flow Rate
Authors: Inoue M. , Kuroumaru M. , Ando Y.
Published: June 2008

The pressure fluctuations on the casing wall of two axial flow compressor rotors with various tip clearances have been analyzed by the use of two kinds of correlation functions. The behavior of the pressure fluctuation varies depending on tip clearance and blade solidity. In the case of small tip clearance, the nature of disturbances becomes random as the flow rate is reduced to a stall condition. For moderate tip clearance, coherent-structured disturbances appear intermittently at low flow rate. They appear more frequently as the solidity is increased and the flow rate becomes lower. For large tip clearance, the coherent structured disturbances exist even at considerably higher flow rates. Corresponding to these features, there are peculiar patterns in the correlation designated as “phase-locked correlation functions.”
Active Suppression of Rotating Stall and Surge in Axial Compressors
Authors: Day I. J.
Published: June 2008

This paper reports on an experimental program in which active control was successfully applied to both rotating stall and surge in a multistage compressor. Two distinctly different methods were used to delay the onset of rotating stall in a four-stage compressor using fast-acting air injection valves. The amount of air injected was small compared to the machine mass flow, the maximum being less than 1.0 percent. In some compressor configurations modal perturbations were observed prior to stall. By using the air injection valves to damp out these perturbations, an improvement of about 4.0 percent in stall margin was achieved. The second method of stall suppression was to remove emerging stall cells by injecting air in their immediate vicinity. Doing this repeatedly delayed the onset of stall, giving a stall margin improvement of about 6.0 percent. Further studies were conducted using a large plenum downstream of the compressor to induce the system to surge rather than stall. The resulting surge cycles were all found to be initiated by rotating stall and therefore the stall suppression systems mentioned above could also be used to suppress surge. In addition, it was possible to arrest the cyclical pulsing of a compressor already in surge.
An Investigation of Turbulence Modeling in Transonic Fans Including a Novel Implementation of an Implicit k–ε Turbulence Model
Authors: Turner M. G. , Jennions I. K.
Published: June 2008

An explicit Navier–Stokes solver has been written with the option of using one of two types of turbulence model. One is the Baldwin–Lomax algebraic model and the other is an implicit k–ε model which has been coupled with the explicit Navier–Stokes solver in a novel way. This type of coupling, which uses two different solution methods, is unique and combines the overall robustness of the implicit k–ε solver with the simplicity of the explicit solver. The resulting code has been applied to the solution of the flow in a transonic fan rotor, which has been experimentally investigated by Wennerstrom. Five separate solutions, each identical except for the turbulence modeling details, have been obtained and compared with the experimental results. The five different turbulence models run were: the standard Baldwin–Lomax model both with and without wall functions, the Baldwin–Lomax model with modified constants and wall functions, a standard k–ε model, and an extended k–ε model, which accounts for multiple time scales by adding an extra term to the dissipation equation. In general, as the model includes more of the physics, the computed shock position becomes closer to the experimental results.
Three-Dimensional Navier–Stokes Computations of Transonic Fan Flow Using an Explicit Flow Solver and an Implicit κ–ε Solver
Authors: Jennions I. K. , Turner M. G.
Published: June 2008

Computational fluid dynamics (CFD) has become a powerful ally of the experimental test facility in revealing the flow physics of some highly complex flows. For certain classes of flow, CFD has reached maturity and is therefore being increasingly used in industry by designers. This paper is intended to show current transonic prediction capability at GE Aircraft Engines in terms of a recently developed three-dimensional Navier–Stokes code. The flow simulations addressed are concerned with transonic fan design and illustrate those issues that are important to designers such as tip leakage flow, shock boundary layer interaction, boundary layer growth, and account of internal solid bodies such as part-span shrouds and engine splitters. In this respect, three successively more complex Navier–Stokes simulations representative of modern fans—NASA Rotor 67, GE/Wennerstrom Rotor 4, and the GE/NASA E3 fans—are considered in this paper.
A Comparison of the Measured and Predicted Flow Field in a Modern Fan-Bypass Configuration
Authors: Goyal R. K. , Dawes W. N.
Published: June 2008

A three-dimensional viscous Navier–Stokes flow solver was used to predict core and bypass rotor performance and radial flow characteristics of a 4.6:1 bypass ratio, single-stage fan. The three-dimensional flow solver can handle several blade rows simultaneously and has the capability to include a downstream splitter. Results of the analysis are compared with experimental data obtained during rig testing of a modern high bypass single-stage turbofan in which rotor performance for both bypass and core streams was measured.
Hierarchical Development of Three Direct-Design Methods for Two-Dimensional Axial-Turbomachinery Cascades
Authors: Korakianitis T.
Published: June 2008

The direct and inverse blade-design iterations for the selection of isolated airfoils and gas turbine blade cascades are enormously reduced if the initial blade shape has performance characteristics near the desirable ones. This paper presents the hierarchical development of three direct blade-design methods of increasing utility for generating two-dimensional blade shapes. The methods can be used to generate inputs to the direct- or inverse-blade-design sequences for subsonic or supersonic airfoils for compressors and turbines, or isolated airfoils. The examples included for illustration are typical modern turbine cascades, and they have been designed by the direct method exclusively. The first method specifies the airfoil shapes with analytical polynomials. It shows that continuous curvature and continuous slope of curvature are necessary conditions to minimize the possibility of flow separation, and to lead to improved blade designs. The second method specifies the airfoil shapes with parametric fourth-order polynomials, which result in continuous-slope-of-curvature airfoils, with smooth Mach number and pressure distributions. This method is time consuming. The third method specifies the airfoil shapes by using a mixture of analytical polynomials and mapping the airfoil surfaces from a desirable curvature distribution. The third method provides blade surfaces with desirable performance in very few direct-design iterations. In all methods the geometry near the leading edge is specified by a thickness distribution added to a construction line, which eliminates the leading edge overspeed and laminar-separation regions. The blade-design methods presented in this paper can be used to improve the aerodynamic and heat transfer performance of turbomachinery cascades, and they can result in high-performance airfoils in very few iterations.
Calculation of Three-Dimensional Boundary Layers on Rotor Blades Using Integral Methods
Authors: Karimipanah M. T. , Olsson E.
Published: June 2008

The important effects of rotation and compressibility on rotor blade boundary layers are theoretically investigated. The calculations are based on the momentum integral method and results from calculations of a transonic compressor rotor are presented. Influence of rotation is shown by comparing the incompressible rotating flow with the stationary one. Influence of compressibility is shown by comparing the compressible rotating flow with the incompressible rotating one. Two computer codes for three-dimensional laminar and turbulent boundary layers, originally developed by SSPA Maritime Consulting AB, have been further developed by introducing rotation and compressibility terms into the boundary layer equations. The effect of rotation and compressibility on the transition have been studied. The Coriolis and centrifugal forces that contribute to the development of the boundary layers and influence its behavior generate crosswise flow inside the blade boundary layers, the magnitude of which depends upon the angular velocity of the rotor and the rotor geometry. The calculations show the influence of rotation and compressibility on the boundary layer parameters. Momentum thickness and shape factor increase with increasing rotation and decrease when compressible flow is taken into account. For skin friction such effects have inverse influences. The different boundary layer parameters behave similarly on the suction and pressure sides with the exception of the crossflow angle, the crosswise momentum thickness, and the skin friction factor. The codes use a nearly orthogonal streamline coordinate system, which is fixed to the blade surface and rotates with the blade.
Measurement and Prediction of Tip Clearance Flow in Linear Turbine Cascades
Authors: Heyes F. J. G. , Hodson H. P.
Published: June 2008

This paper describes a simple two-dimensional model for the calculation of the leakage flow over the blade tips of axial turbines. The results obtained from calculations are compared with data obtained from experimental studies of two linear turbine cascades. One of these cascades has been investigated by the authors and previously unpublished experimental data are provided for comparison with the model. In each of the test cases examined, excellent agreement is obtained between the experimental and predicted data. Although ignored in the past, the importance of pressure gradients along the blade chord is highlighted as a major factor influencing the tip leakage flow.
Experimental Study on the Three-Dimensional Flow Within a Compressor Cascade With Tip Clearance: Part I—Velocity and Pressure Fields
Authors: Kang S. , Hirsch C.
Published: June 2008

Experimental results from a study of the three-dimensional flow in a linear compressor cascade with stationary endwall at design conditions are presented for tip clearance levels of 1.0, 2.0, and 3.3 percent of chord, compared with the no-clearance case. In addition to five-hole probe measurements, extensive surface flow visualizations are conducted. It is observed that for the smaller clearance cases a weak horseshoe vortex forms in the front of the blade leading edge. At all the tip gap cases, a multiple tip vortex structure with three discrete vortices around the midchord is found. The tip leakage vortex core is well defined after the midchord but does not cover a significant area in traverse planes. The presence of the tip leakage vortex results in the passage vortex moving close to the endwall and the suction side.
Measurement of the Three-Dimensional Tip Region Flow Field in an Axial Compressor
Authors: Stauter R. C.
Published: June 2008

A two-color, five-beam LDV system has been configured to make simultaneous three-component velocity measurements of the flow field in a two-stage axial compressor model. The system has been used to make time-resolved measurements both between compressor blade rows and within the rotating blade passages in an axial compressor. The data show the nature and behavior of the complex, three-dimensional flow phenomena present in the tip region of a compressor as they convect downstream. In particular, the nature of the tip leakage vortex is apparent, being manifested by high blockage as well as the expected vortical motion. The data indicate that the radial flows associated with the tip leakage vortex begin to decrease while within the rotor passage, and that they temporarily increase aft of the passage.
An Investigation of Factors Influencing the Calibration of Five-Hole Probes for Three-Dimensional Flow Measurements
Authors: Dominy R. G. , Hodson H. P.
Published: June 2008

The effects of Reynolds number, Mach number, and turbulence on the calibrations of commonly used types of five-hole probe are discussed. The majority of the probes were calibrated at the exit from a transonic nozzle over a range of Reynolds numbers (7 × 103 < Re < 80 × 103 based on probe tip diameter) at subsonic and transonic Mach numbers. Additional information relating to the flow structure were obtained from a large-scale, low-speed wind tunnel. The results confirmed the existence of two distinct Reynolds number effects. Flow separation around the probe head affects the calibrations at relatively low Reynolds numbers while changes in the detailed structure of the flow around the sensing holes affects the calibrations even when the probe is nulled. Compressibility is shown to have little influence upon the general behavior of these probes in terms of Reynolds number sensitivity but turbulence can affect the reliability of probe calibrations at typical test Reynolds numbers.
On the Conservation of Rothalpy in Turbomachines
Authors: Lyman F. A.
Published: June 2008

The conditions under which rothalpy is conserved are investigated by means of the energy and moment-of-momentum equations for unsteady flow of a viscous, compressible fluid. Differential and integral equations are given for the total enthalpy and rothalpy in both stationary and rotating coordinates. From the equations in rotating coordinates it is shown that rothalpy may change due to: (1) pressure fluctuations caused by flow unsteadiness in the rotating frame; (2) angular acceleration of the rotor; (3) work done by viscous stresses on the relative flow in the rotating frame; (4) work done by body forces on the relative flow; (5) changes in entropy due to viscous dissipation and heat transfer. Conclusions of this investigation are compared with those of previous authors, some of whom have stated that rothalpy is conserved even in viscous flows. A modified Euler’s turbomachine equation, which includes viscous effects, is derived and errors in previous derivations noted.
Investigations on a Radial Compressor Tandem-Rotor Stage With Adjustable Geometry
Authors: Josuhn-Kadner B. , Hoffmann B.
Published: June 2008

A radial compressor stage has been investigated mainly experimentally for aerodynamic stage optimization. The rotor (πt = 3.9) consists of a profiled axial inducer and a conventionally designed radial impeller. Inducer and impeller can be locked at different circumferential positions relative to each other, thus forming a tandem wheel with adjustable geometry. Conventional and Laser-2-Focus system measurements for the tandem rotor and the stage were performed at different operating points to study the influence of the circumferential clearance geometry between inducer and impeller with respect to compressor characteristics and performance. Furthermore, three-dimensional Navier–Stokes calculations are being developed at design point condition to analyze the flow field. A small influence of the inducer adjustment on the rotor characteristics is observed. The maximum rotor efficiency of 93.5 percent varies in a range of less than 1 percent depending on the different inducer positions.
Design and Rotor Performance of a 5:1 Mixed-Flow Supersonic Compressor
Authors: Mo¨nig R. , Elmendorf W. , Gallus H. E.
Published: June 2008

In consideration of further jet-engine developments required by applications for supersonic travel aircraft, airbreathing propulsion of space vehicles, or only the improvement of conventional high-performance turbo-engines, highly loaded supersonic compressors seem to meet the future demands. Particularly mixed-flow compressor stages with moderate supersonic rotor and stator inlet flow reveal the potential of high pressure rise and mass flow as well as favorable performance characteristics and efficiency. The first part of this paper presents analytical considerations for mixed-flow supersonic compressors with strong shock waves. This theoretical approach proves to be essential besides established design tools in order to ensure safe rotor and stage operation in accordance with the design objectives. In this context, the conditions for shock wave stabilization within a diagonal rotor passage are discussed in detail for design and off-design rotational speeds. The main part of this paper, however, presents the results and flow analysis obtained by extensive experimental investigations of the designed mixed-flow compressor rotor. The investigations were restricted to operation without stator in order to strictly separate rotor performance from rotor-stator interactions. The results reveal the design goals to be met in general. Mass flow, total pressure rise, and efficiency in particular show a good agreement with the design properties for near-surge operation at design and off-design conditions.
Investigation of Rotor Blade Roughness Effects on Turbine Performance
Authors: Boynton J. L. , Tabibzadeh R. , Hudson S. T.
Published: June 2008

The cold air test program was completed on the SSME (Space Shuttle Main Engine) HPFTP (High-Pressure Fuel Turbopump) turbine with production nozzle vane rings and polished coated rotor blades with a smooth surface finish of 30 μin. (0.76 μm) rms (root mean square). The smooth blades were polished by an abrasive flow machining process. The test results were compared with the air test results from production rough-coated rotor blades with a surface finish of up to 400 μin. (10.16 μm) rms. Turbine efficiency was higher for the smooth blades over the entire range tested. Efficiency increased 2.1 percentage points at the SSME 104 percent RPL (Rated Power Level) conditions. This efficiency improvement could reduce the SSME HPFTP turbine inlet temperature by 57 R (32 K), increasing turbine durability. The turbine flow parameter increased and the midspan outlet swirl angle became more axial with the smooth rotor blades.
Unsteady Boundary-Layer Transition in Flow Periodically Disturbed by Wakes
Authors: Orth U.
Published: June 2008

Boundary layers on turbomachinery blades develop in a flow that is periodically disturbed by the wakes of upstream blade cascades. These wakes have a significant effect upon laminar-turbulent boundary-layer transition. In order to study these effects, detailed velocity measurements using hot-wire probes were performed within the boundary-layer of a plate in flow periodically disturbed by wakes produced by bars moving transversely to the flow. The measurements were evaluated using the ensemble-averaging technique. The results show how the wake disturbance enters the boundary-layer and leads to a turbulent patch, which grows and is carried downstream. In favorable pressure gradients, transition due to wake turbulence occurred much earlier than predicted by linear stability theory. Between two wakes, laminar becalmed regions were observed far beyond the point at which the undisturbed boundary-layer was already turbulent.
Calculation of Three-Dimensional Unsteady Flows in Turbomachinery Using the Linearized Harmonic Euler Equations
Authors: Hall K. C. , Lorence C. B.
Published: June 2008

An efficient three-dimensional Euler analysis of unsteady flows in turbomachinery is presented. The unsteady flow is modeled as the sum of a steady or mean flow field plus a harmonically varying small perturbation flow. The linearized Euler equations, which describe the small perturbation unsteady flow, are found to be linear, variable coefficient differential equations whose coefficients depend on the mean flow. A pseudo-time time-marching finite-volume Lax-Wendroff scheme is used to discretize and solve the linearized equations for the unknown perturbation flow quantities. Local time stepping and multiple-grid acceleration techniques are used to speed convergence. For unsteady flow problems involving blade motion, a harmonically deforming computational grid, which conforms to the motion of the vibrating blades, is used to eliminate large error-producing extrapolation terms that would otherwise appear in the airfoil surface boundary conditions and in the evaluation of the unsteady surface pressure. Results are presented for both linear and annular cascade geometries, and for the latter, both rotating and nonrotating blade rows.
Numerical Solutions for Unsteady Subsonic Vortical Flows Around Loaded Cascades
Authors: Fang J. , Atassi H. M.
Published: June 2008

A frequency domain linearized unsteady aerodynamic analysis is presented for three-dimensional unsteady vortical flows around a cascade of loaded airfoils. The analysis fully accounts for the distortion of the impinging vortical disturbances by the mean flow. The entire unsteady flow field is calculated in response to upstream three-dimensional harmonic disturbances. Numerical results are presented for two standard cascade configurations representing turbine and compressor bladings for a reduced frequency range from 0.1 to 5. Results show that the upstream gust conditions and blade sweep strongly affect the unsteady blade response.
Discrete-Jet Film Cooling: A Comparison of Computational Results With Experiments
Authors: Leylek J. H. , Zerkle R. D.
Published: June 2008

Large-scale computational analyses have been conducted and results compared with experiments to understand coolant jet and crossflow interaction in discrete-jet film cooling. Detailed three-dimensional elliptic Navier–Stokes solutions, with high-order turbuence modeling, are presented for film cooling using a new model enabling simultaneous solution of fully coupled flow in plenum, film-hole, and cross-stream regions. Computations are carried out for the following range of film cooling parameters typically found in gas turbine airfoil applications: single row of jets with a film-hole length-to-diameter ratio of 1.75 and 3.5; blowing ratio from 0.5 up to 2; coolant-to-crossflow density ratio of 2; streamwise injection angle of 35 deg; and pitch-to-diameter ratio of 3. Comparison of computational solutions with experimental data give good agreement. Moreover, the current results complement experiments and support previous interpretations of measured data and flow visualization. The results also explain important aspects of film cooling, such as the development of complex flow within the film-hole in addition to the well-known counterrotating vortex structure in the cross-stream.
Effects of Free-Stream Turbulence and Adverse Pressure Gradients on Boundary Layer Transition
Authors: Gostelow J. P. , Blunden A. R. , Walker G. J.
Published: June 2008

Boundary layer measurements are presented through transition for six different free-stream turbulence levels and a complete range of adverse pressure gradients for attached laminar flow. Measured intermittency distributions provide an excellent similarity basis for characterizing the transition process under all conditions tested when the Narasimha procedure for determining transition inception is used. This inception location procedure brings consistency to the data. Velocity profiles and integral parameters are influenced by turbulence level and pressure gradient and do not provide a consistent basis. Under strong adverse pressure gradients transition occurs rapidly and the velocity profile has not fully responded before the completion of transition. The starting turbulent layer does not attain an equilibrium velocity profile. A change in pressure gradient from zero to even a modest adverse level is accompanied by a severe reduction in transition length. Under diffusing conditions the physics of the transition process changes and the spot formation rate increases rapidly; instead of the “breakdown in sets” regime experienced in the absence of a pressure gradient, transition under strong adverse pressure gradients is more related to the amplification and subsequent instability of the Tollmien-Schlichting waves. Measurements reveal an exponential decrease in transition length with increasing adverse pressure gradient; a less severe exponential decrease is experienced with increasing turbulence level. Correlations of transition length are provided that facilitate its prediction in the form of suitable length parameters including spot formation rate.