Showing papers in "Journal of Turbomachinery-transactions of The Asme in 1990"
TL;DR: In this paper, detailed measurements have been made of the transient stalling process in an axial compressor stage, where the stage is of high hub-casing ratio and stall is initiated in the rotor.
Abstract: Detailed measurements have been made of the transient stalling process in an axial compressor stage. The stage is of high hub-casing ratio and stall is initiated in the rotor. If the rotor tip clearance is small stall inception occurs at the hub, but at clearances typical for a multistage compressor the inception is at the tip. The crucial quantity in both cases is the blockage caused by the endwall boundary layer. Prior to stall, disturbances rotate around the inlet flow in sympathy with rotating variations in the endwall blockage; these can persist for some time prior to stall, rising and falling in amplitude before the final increase, which occurs as the compressor stalls. © 1990 by ASME.
278 citations
TL;DR: In this article, the effects of buoyancy and Coriolis forces on heat transfer in turbine blade internal coolant passages were investigated with a large scale, multi-pass, smooth-wall heat transfer model with both radially inward and outward flow.
Abstract: Experiments were conducted to determine the effects of buoyancy and Coriolis forces on heat transfer in turbine blade internal coolant passages. The experiments were conducted with a large scale, multi-pass, smooth-wall heat transfer model with both radially inward and outward flow. An analysis of the governing flow equations showed that four parameters influence the heat transfer in rotating passages (coolant-to-wall temperature ratio, Rossby number, Reynolds number and radius-to-passage hydraulic diameter ratio). These four parameters were varied over ranges which are typical of advanced gas turbine engine operating conditions. It was found that both Coriolis and buoyancy effects must be considered in turbine blade cooling designs and that the effect of rotation on the heat transfer coefficients was markedly different depending on the flow direction. Local heat transfer coefficients were found to decrease by as much as 60 percent and increase by 250 percent from no rotation levels. Comparisons with a pioneering stationary vertical tube buoyancy experiment showed reasonably good agreement. Correlation of the data is achieved employing dimensionless parameters derived from the governing flow equations.
234 citations
183 citations
TL;DR: In this article, a time-marching Euler calculation for 2D and quasi-3D unsteady flows in oscillating blade rows is presented, based on a finite volume scheme with cell-vertex discretization in space and 2-step Runge-Kutta integration in time.
Abstract: A time-marching Euler calculation for 2-D and quasi-3-D unsteady flows in oscillating blade rows is presented, based on a finite volume scheme with cell-vertex discretization in space and 2-step Runge-Kutta integration in time. Calculated results for unsteady flows in an oscillating flat plate cascade are in good agreement with those from two well-established linear methods, LINSUB and FINEL. The unsteady pressure distribution and aerodynamic damping calculated by the present method for a turbine blade test case (Aeroelasticity Workshop Standard Configuration No. 4 cascade) agree well with the corresponding experimental data
150 citations
TL;DR: In this paper, the rotor-stator interaction within a centrifugal pump having several vaned diffusers, under conditions of different flow coefficients and different radial gaps between the impeller blade trailing edge and the diffuser vane leading edge, was investigated.
Abstract: Steady and unsteady diffuser vane pressure measurements have been conducted with a two-dimensional test impeller, in an experimental investigation of rotor-stator interaction within a centrifugal pump having several vaned diffusers, under conditions of different flow coefficients and different radial gaps between the impeller blade trailing edge and the diffuser vane leading edge. 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. Increasing the number of diffuser vanes was found to result in a significant decrease of impeller blade pressure fluctuations.
148 citations
TL;DR: In this paper, a procedure for simulating the flow field within multistage turbomachinery, including the effects of unsteadiness, compressibility, and viscosity, is presented.
Abstract: This work outlines a procedure for simulating the flow field within multistage turbomachinery, which includes the effects of unsteadiness, compressibility, and viscosity. The associated modeling equations are the average passage equation system, which governs the time-averaged flow field within a typical passage of a blade row embedded within a multistage configuration. The results from a simulation of a low aspect ratio stage and one-half turbine will be presented and compared with experimental measurements. It will be shown that the secondary flow field generated by the rotor causes the aerodynamic performance of the downstream vane to be significantly different from that of an isolated blade row.
117 citations
TL;DR: In this article, the unsteady, thin-layer Navier-Stokes equations in two spatial dimensions are solved on a system of patched and overlaid grids for a rotor-stator configuration from an axial turbine.
Abstract: 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 a rotor-stator configuration from an axial turbine The governing equations are solved using a finite-difference, upwind algorithm that is set in an iterative, implicit framework Results in the form of pressure contours, time-averaged pressures, unsteady pressure amplitudes and phase are presented 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 The current study uses multiple rotors and stators 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
95 citations
TL;DR: In this paper, the turbulent heat transfer and friction for fully developed flow of air in a square channel in which two opposite walls are roughened with 90° full ribs, parallel and crossed full ribs with angles-of-attack (α) of 60° and 45°, 90° discrete ribs, and parallel and cross-discrete ribs with = 60°, 45° and 30°.
Abstract: Experiments have been conducted to study the turbulent heat transfer and friction for fully developed flow of air in a square channel in which two opposite walls are roughened with 90° full ribs, parallel and crossed full ribs with angles-of-attack (α) of 60° and 45°, 90° discrete ribs, and parallel and crossed discrete ribs with = 60°, 45°, and 30°. The discrete ribs are staggered in alternate rows of three and two ribs. Results are obtained for a rib height-to-channel hydraulic diameter ratio of 0.0625, a rib pitch-to-height ratio of 10, and Reynolds numbers between 10,000 and 80,000. Parallel angled discrete ribs are superior to 90° discrete ribs and parallel angled full ribs, and are recommended for internal cooling passages in gas turbine airfoils. For α = 60° and 45°, parallel discrete ribs have higher ribbed wall heat transfer, lower smooth wall heat transfer, and lower channel pressure drop than parallel full ribs. Parallel 60° discrete ribs have the highest ribbed wall heat transfer and parallel 30° discrete ribs cause the lowest pressure drop. The heat transfer and pressure drops in crossed angled full and discrete rib cases are all lower than those in the corresponding 90° and parallel angled rib cases. Crossed arrays of angled ribs have poor thermal performance and are not recommended.Copyright © 1990 by ASME
95 citations
TL;DR: In this article, the authors discuss approaches taken over many years to achieve very high loading levels in axial-flow compressors, mainly associated with aircraft turbine engines, and discuss some innovative schemes to increase diffusion limits that took place in the 1960s and 1970s.
Abstract: This paper discusses approaches taken over many years to achieve very high loading levels in axial-flow compressors. These efforts have been associated predominantly with aircraft turbine engines. The objective has been to reduce the size and weight of the powerplant, to increase its simplicity and ruggedness, and, whenever possible, to reduce cost. In the introduction, some fundamentals are reviewed that indicate that increased work per stage can only be obtained at a cost of increased Mach number, increased diffusion, or both. The earliest examples cited are some ambitious development programs of the 1950s and 1960s. Some innovative schemes to increase diffusion limits are described that took place in the 1960s and 1970s. Major advancements in dealing with higher Mach number were made in the 1980s. Finally, a few thoughts directed toward potential future developments are presented.
94 citations
TL;DR: In this article, the influence of high mainstream turbulence on leading edge film effectiveness and heat transfer coefficient was studied using a blunt body with a semicylinder leading edge with a flat afterbody.
Abstract: The influence of high mainstream turbulence on leading edge film effectiveness and heat transfer coefficient was studied. High mainstream turbulence was produced by a passive grid and a jet grid. Experiments were performed using a blunt body with a semicylinder leading edge with a flat afterbody. The mainstream Reynolds number based on leading edge diameter was about 100,000. Spanwise and streamwise distributions of film effectiveness and heat transfer coefficient in the leading edge and on the flat sidewall were obtained for three blowing ratios, through rows of holes located at ± 15 and ± 40 deg from stagnation
87 citations
TL;DR: In this article, the effect of rotation, aspect ratio, and turbulator roughness on heat transfer in rib-roughened passages was investigated in an orthogonally rotating setup to simulate the actual rotation of the cooling passages.
Abstract: Turbine blade cooling is imperative in advanced aircraft engines. The extremely hot gases that operate within the turbine section require turbine blades to be cooled by a complex cooling circuit. This cooling arrangement increases engine efficiency and ensures blade materials a longer creep life. One principle aspect of the circuit involves serpentine internal cooling passes throughout the core of the blade. Roughening the inside surfaces of these cooling passages with turbulence promoters provides enhanced heat transfer rates from the surface. The purpose of this investigation was to study the effect of rotation, aspect ratio, and turbulator roughness on heat transfer in these rib-roughened passages. The investigation was performed in an orthogonally rotating setup to simulate the actual rotation of the cooling passages. Single-pass channels, roughened on two opposite walls, with turbulators positioned at 45 deg angle to the flow, in a criss-cross arrangement, were studied throughout this experiment. The ribs were arranged such that their pitch-to-height ratio remained at a constant value of 10. An aspect ratio of unity was investigated under three different rib blockage ratios (turbulator height/channel hydraulic diameter) of 0.1333, 0.25, and 0.3333. A channel with an aspect ratio of 2 was also investigated for a blockage ratio of 0.25. Air was flown radially outward over a Reynolds number range of 15,000 to 50,000. The rotation number was varied from 0 to 0.3. Stationary and rotating cases of identical geometries were compared. Results indicated that rotational effects are more pronounced in turbulated passages of high aspect and low blockage ratios for which a steady increase in heat transfer coefficient is observed on the trailing side as rotation number increases while the heat transfer coefficient on the leading side shows a steady decrease with rotation number. However, the all-smooth-wall classical pattern of heat transfer coefficient variation on the leading and trailing sides is not followed for smaller aspect ratios and high blockage ratios when the relative artificial roughness is high.
TL;DR: In this article, the effects of injection rate and strength of curvature on film cooling performance of gas injected through a row of holes on a convex surface was studied, and compared with film cooling of concave and flat surfaces.
Abstract: The effects of injection rate and strength of curvature on film cooling performance of gas injected through a row of holes on a convex surface is studied. Comparisons are made to film cooling of concave and flat surfaces. Three different relative strengths of curvature (ratio of radius of curvature to radius of injection hole), two density ratios (0.95 and 2.0), and a wide range of blowing rates (0.3 to 2.7) are considered. A foreign gas injection technique (mass transfer analogy) is used. The strength of curvature was controlled by varying the injection hole diameter. At low blowing rates, film cooling is more effective on the convex surface than on a flat or a concave surface. The cross stream pressure gradient present in curved flows tends to push the jet into the convex wall. As the injection rate is increased, normal and tangential jet momentum promote lift-off from the convex surface, thereby lowering performance. In contrast, previous studies show that on a concave surface, tangential jet momentum, flow instabilities, and blockage improve performance on a concave surface as blowing rate is increased.Copyright © 1990 by ASME
TL;DR: In this paper, it was shown that the base pressure and loss can be reasonably well predicted by inviscid Euler calculations, and this was later confirmed in the case of transonic turbine blades.
Abstract: 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 byut 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
TL;DR: In this paper, heat transfer and leakage loss measurements were obtained for compressible flows in typical straight-through labyrinth seals with high rotational speeds, and the experiments were an extension of earlier measurements in a stationary test facility.
Abstract: Heat transfer and leakage loss measurements were obtained for compressible flows in typical straight-through labyrinth seals with high rotational speeds. The experiments are an extension of our earlier measurements in a stationary test facility. In order to ensure direct comparisons to the original experiments, the principal dimensions of the test facility and gas dynamic parameters of the hot gas were kept similar. The new study encompasses a wide range of Taylor numbers, Reynolds numbers, and clearances between the rotating annular fins and the stationary shroud
TL;DR: In this paper, the authors 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.
Abstract: The off-design performance of axial turbines is usually predicted by calculating the incidence losses using empirical correlations. 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
TL;DR: In this paper, flat plate model tests are carried out to investigate the effect of both the boundary layer state and strailing edge geometry on the vortex shedding frequency in a mixed laminar-turbulent separation from turbine blades.
Abstract: The investigation is restricted to the subsonic domain. Flat plate model tests are carried out to investigate the effect of both the boundary layer state and strailing 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
TL;DR: In this paper, the surface roughness of turbine engine blades from F-100 and TF-39 aeroengines was measured using profilometer measurements of the boundary layer flow and heat transfer.
Abstract: Results are presented from profilometer measurements of the surface roughness on in-service turbine engine blades from F-100 and TF-39 aeroengines. 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
TL;DR: In this article, a simple flow and heat transfer model incorporating these features was used to estimate both tip and shroud heat transfer provided that reasonable estimates of the clearance gap size and clearance leakage flow can be made.
Abstract: Unshrouded blades of axial turbine stages move in close proximity to the stationary outer seal, or shroud, of the turbine housing. The pressure difference between the concave and convex sides of the blade drives a leakage flow through the gap between the moving blade tip and adjacent wall. This clearance leakage flow and accompanying heat-transfer are of interest because of long obvious effects on aerodynamic performance and structural durability, but understanding of its nature and influences has been elusive. Previous studies indicate that the leakage through the gap is mainly a pressure-driven flow whose magnitude is related strongly to the airfoil pressure loading distribution and only weakly, if at all, to the relative motion between blade tip and shroud. A simple flow and heat-transfer model incorporating these features can be used to estimate both tip and shroud heat transfer provided that reasonable estimates of the clearance gap size and clearance leakage flow can be made. The present work uses a numerical computation of the leakage flow to link the model to a specific turbine geometry and operating point for which a unique set of measured local tip and shroud heat fluxes are available. The resulting comparisons between the model estimates and measured heat-transfer are good. The model should thus prove useful in the understanding and interpretation of future measurements, and should additionally prove useful for providing early design estimates of the levels of tip and shroud heat transfer that need to be compensated for by active turbine cooling.Copyright © 1990 by ASME
TL;DR: In this article, a linear cascade of high-turning, low aspect ratio turbine blades has been measured in great detail at five planes within the cascade and two downstream in order to trace the generation of stagnation pressure loss in the passage.
Abstract: Flow through a linear cascade of high-turning, low aspect ratio turbine blades has been measured in great detail at five planes within the cascade and two downstream in order to trace the generation of stagnation pressure loss in the passage. Endwall shear stresses have been measured using a hot-film probe and an oil-drop viscosity balance technique
TL;DR: In this article, the spanwise averaged effectiveness and heat transfer coefficient for an inclined slot and a single row of holes in the presence of favorable, zero and adverse pressure gradients were measured.
Abstract: Film-cooling in the presence of mainstream pressure gradients typical of gas turbines has been studied experimentally on a flat plate This paper describes, measurements of the spanwise averaged effectiveness and heat transfer coefficient for an inclined slot and a single row of holes in the presence of favourable, zero and adverse pressure gradients. Acceleration parameters of K = 2.62×10−6 and - 0.22 × 10−6 were achieved at the point of injection where the freestream unit Reynolds number was held constant at Re/m = 2.7 × 107. The flow was accelerated to high Mach number and results are analysed using a superposition model of film-cooling which included the effects of viscous energy dissipation. The experimental results show the effects of pressure gradient differ between the geometries and a discussion of these results is included. The unblown turbulent boundary layer with pressure gradient were also studied. Experiments were performed using the Isentropic Light Piston Tunnel, a transient facility which enables conditions representative of those in the engine to be attained.Copyright © 1990 by ASME
TL;DR: In this paper, an experimental study of three-dimensional flow field in an annular compressor cascade with an upstream rotor has been carried out at four different incidences to the stator blade.
Abstract: An experimental study of three-dimensional flow field in an annular compressor cascade with an upstream rotor has been carried out at four different incidences to the stator blade. Blade boundary layers and the three-dimensional flow field at the exit are surveyed using a hot-wire sensor and a five-hole probe, respectively. The data on the blade boundary layer, passage flow, and separated corner flow are presented. A detailed interpretation of the effects of upstream wakes on the entire passage flow is presented and compared with the data in the absence of a rotor
TL;DR: The results of performance measurements and detailed measurements of the mean flow field at rotor inlet and rotor exit in three squirrel cage fan configurations were taken with a five-hole probe and yield total pressure, static pressure and three components of velocity.
Abstract: This paper presents the results of performance measurements and detailed measurements of the mean flow field at rotor inlet and rotor exit in three squirrel cage fan configurations. The flow-field measurements were taken with a five-hole probe and yield total pressure, static pressure and the three components of velocity. Measurements were taken for two casing throat areas and for two different rotors. For each configuration the flow field was measured for flow rates below, near and above the best-efficiency point. Flow patterns are complex and there is reverse flow through the rotor blading even at the best-efficiency operating condition. Although complex, the main features of flow behaviour can be understood. They were common to all three fan configurations.Copyright © 1989 by ASME
TL;DR: In this article, the effects of turbulent flow through rectangular straight ducts rotating in an orthogonal mode have been investigated and the degree of heat transfer augmentation on the pressure side is found to depend on the Reynolds number as well as on Rossby number.
Abstract: This work is concerned with fully-developed constant-density turbulent flow through rectangular straight ducts rotating in an orthogonal mode. Ducts of both square and 2:1 aspect ratio cross-sections have been examined. For the square duct, predictions have been performed for Reynolds numbers of 33,500 and 97,000 and for the 2:1 aspect ratio duct the computations were carried out for a Reynolds number of 33,500. Values of the inverse Rossby number (Ro = ΩD/Wb) ranged from 0.005 to 0.2. Except in the immediate vicinity of the wall, the standard high-Reynolds-number version of the k-e model is used to account for the effects of turbulence. Across the near-wall sublayer the damping of turbulence is modelled through a low-Reynolds-number one-equation model.Low rotational speeds cause the formation of a pair of symmetric streamwise vortices. At higher rotational speeds, flow instabilities on the pressure side lead to transition to a more complex four-vortex structure. The transition point depends on both the cross-sectional geometry and the flow Reynolds number. Moreover, over a range of Rossby number, either two- or four-vortex solutions are possible depending upon initial conditions.The rotation leads to significant differences between the values of friction factor and Nusselt number on the suction and pressure surfaces of the duct. The degree of heat transfer augmentation on the pressure side is found to depend on the Reynolds number as well as on Rossby number. In contrast, heat-transfer attenuation on the suction side is only Rossby-number dependent.Copyright © 1990 by ASME
TL;DR: In this article, the results of an investigation of the three-dimensional flow downstream of a transonic turbine cascade are presented for a wide range of Mach numbers, extending from M 2is = 0.2 up to 1.55.
Abstract: The results of an investigation of the three-dimensional flow downstream of a transonic turbine cascade are presented. The investigation was carried out for a wide range of Mach numbers, extending from M 2is =0.2 up to 1.55. Measurements were made in five planes at different axial locations downstream of the trailling edge by using a miniaturized five-hole probe especially designed for transonic flows
TL;DR: In this paper, the aerodynamic and acoustic properties of axial fans with swept blades are discussed with particular emphasis on noise mechanisms and the influence of high-intensity inlet turbulence on "excess" noise.
Abstract: 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.
TL;DR: In this article, the 3D viscous flowfield of a 4.7:1 pressure ratio backswept impeller was studied experimentally and numerically by using laser velocimetry and an advanced 3-dimensional viscous code.
Abstract: The 3-D viscous flowfield of a 4.7:1 pressure ratio backswept impeller was studied experimentally and numerically by using laser velocimetry and an advanced 3-D viscous code. The impeller was designed by a CAD method, and a maximum rotor efficiency of 94% 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 3-D 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.Copyright © 1989 by ASME
TL;DR: In this paper, a non-dimensional design of a radial inflow turbine rotor is presented for any specified power ratio, with the objective of minimising the inlet and discharge Mach numbers so that the passage losses are minimised.
Abstract: A procedure is described which develops the non-dimensional design of a radial inflow turbine rotor. The design is developed, for any specified non-dimensional power ratio, with the objective of minimising the inlet and discharge Mach numbers so that the passage losses are minimised. Initially state of the art efficiencies are assumed but are later modified through the specification of empirical losses. The resultant non-dimensional design can be transformed to absolute dimensions through the specification of the inlet stagnation conditions and the mass flow rate of the working fluid.Copyright © 1989 by ASME
TL;DR: In this article, 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.
Abstract: 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 clarification of 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, as injection can also contribute significantly to stall suppression. The results also indicate that the increase in stall pressure rise 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.Copyright © 1989 by ASME
TL;DR: In this paper, the effect of the individual stage erosion on the overall compressor performance was also demonstrated, and a fault model was implemented on a stage stacking program developed to demonstrate the effects of erosion in a multistage compressor.
Abstract: Experimental results obtained from cascades and one 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 the individual stage erosion on the overall compressor performance is also demonstrated.Copyright © 1989 by ASME