Showing papers in "Journal of Turbomachinery-transactions of The Asme in 1986"
TL;DR: In this article, a semi-empirical model for estimating endwall losses was developed from improved understanding of complex endwall secondary flows, acquired through flow visualization and pressure loss data for axial flow turbomachine cascades.
Abstract: The development of a semi-empirical model for estimating endwall losses is described in this paper. The model has been developed from improved understanding of complex endwall secondary flows, acquired through review of flow visualization and pressure loss data for axial flow turbomachine cascades. The flow visualization data together with detailed measurements of viscous flow development through cascades have permitted more realistic interpretation of the classical secondary flow theories for axial turbomachine cascades. The re-interpreted secondary flow theories together with integral boundary layer concepts are used to formulate a calculation procedure for predicting losses due to the endwall secondary flows. The proposed model is evaluated against data from published literature and improved agreement between the data and predictions is demonstrated.Copyright © 1986 by ASME
259 citations
TL;DR: The SIMPLE method of Patankar and Spalding and its variants such as SIMPLER, SIMPLEC and SIMPLEX are segregated methods for solving the discrete algebraic equations representing the equations of motion for an incompressible fluid flow as discussed by the authors.
Abstract: The SIMPLE method of Patankar and Spalding and its variants such as SIMPLER, SIMPLEC and SIMPLEX are segregated methods for solving the discrete algebraic equations representing the equations of motion for an incompressible fluid flow. The present paper presents the extension of these methods to the solution of compressible fluid flows within the context of a generalized segregated approach. To provide a framework for better understanding the segregated approach to solving viscous compressible fluid flows an interpretation of the role of pressure in the numerical method is presented. With this interpretation it becomes evident that the linearization of the equation for mass conservation and the approach used to solve the linearized algebraic equations representing the equations of motion are important in determining the performance of the numerical method. The relative performance of the various segregated methods are compared for several subsonic and supersonic compressible fluid flows.Copyright © 1986 by ASME
113 citations
TL;DR: A mathematical model for closing or mathematically completing the system of equations which describes the time-averaged flow field through the blade passages of multistage turbomachinery is proposed in this paper.
Abstract: A mathematical model is proposed for closing or mathematically completing the system of equations which describes the time-averaged flow field through the blade passages of multistage turbomachinery. These equations, referred to as the average-passage equation system, govern a conceptual model which has proven useful in turbomachinery aerodynamic design and analysis. The closure model is developed so as to insure a consistency between these equations and the axisymmetric through-flow equations. The closure model was incorporated into a computer code for use in simulating the flow field about a high-speed counterrotating propeller and a high-speed fan stage. Results from these simulations are presented.
105 citations
TL;DR: In this article, the off-design performance of a low-pressure turbine rotor root section in a linear cascade was investigated using pneumatic probes and surface flow visualization, and significant differences in the detail of the flow occur over the range of operating conditions investigated.
Abstract: The ability of a given blade profile to operate over a wide range of conditions is often of the utmost importance. This paper reports the off-design performance of a low-pressure turbine rotor root section in a linear cascade. Data were obtained using pneumatic probes and surface flow visualization. The effects of incidence (+9,0, -20 deg), Reynolds (1.5, 2.9, 6.0 x 10/sup 5/), pitch-chord ratio (0.46, 0.56, 0.69), and inlet boundary layer thickness (0.011, 0.022 delta*/C) are discussed. Particular attention is paid to the three dimensionality of the flow field. Significant differences in the detail of the flow occur over the range of operating conditions investigated. It is found that the production of new secondary loss is greatest at lower Reynolds numbers, positive incidence, and the higher pitch-chord ratios.
96 citations
79 citations
TL;DR: In this article, an experimental investigation was conducted to characterize a symmetrical horseshoe vortex system in front of and around a single large-diameter right cylinder centered between the sidewalls of a wind tunnel.
Abstract: An experimental investigation was conducted to characterize a symmetrical horseshoe vortex system in front of and a around a single large-diameter right cylinder centered between the sidewalls of a wind tunnel. Surface flow visualization anad surface static pressure measurements as well as extensive mean velocity and pressure measurements in and around the vortex system were acquired. The results lend new insight into the formation and development of the vortex system. Contrary to what has been assumed previously, a strong vortex was not identified in the streamwise plane of symmetry, but started a significant angular distance away from it. Rather than the multiple vortex systems reported by others, only a single primary vortex and saddle point were found. The scale of the separation process at the saddle point was much smaller than the scale of the approaching bundary layer thickness. Results of the present study not only shed light on such phenomena as the nonsymmetrical endwall flow in axial turbomachinery but can also be used as a test case for three-dimensional computational fluid mechanics computer codes.
74 citations
TL;DR: In this paper, the authors demonstrate that it is the boundary layer transition behavior that determines the losses on turbine cascades and that an enormous effort in measuring technique is required in order to define the location of transition from cascade experiments very accurately.
Abstract: The optimization of the blade surface velocity distribution is promising for a reduction of turbine cascade losses. Theoretical and experimental investigations on three turbine cascades with the same blade loading show the important influence of the blade pressure gradient and the free-stream turbulence on the loss behavior. The results presented demonstrate that it is the boundary layer transition behavior that determines the losses on turbine cascades. An enormous effort in measuring technique is required in order to define the location of transition from cascade experiments very accurately.
68 citations
TL;DR: In this paper, it has been shown that a small clearance between the stator blade and the stationary hub increases the efficiency of axial flow compressor stator compared to the case with no clearance.
Abstract: Measurements have been performed in a low speed high reaction single stage axial compressor. Data obtained within and downstream of the rotor, when correlated with the results of other investigations, provide a link between the existence of suction surface-hub corner separations, their associated loss mechanisms and blade loading. Within the stator, it has been shown that introducing a small clearance between the stator blade and the stationary hub increases the efficiency of the stator compared to the case with no clearance. Oil flow visualisation indicated that the leakage reduced the extensive suction surface-hub corner separation that would otherwise exist. A tracer gas experiment showed that the large radial shifts of the surface streamlines indicated by the oil flow technique were only present close to the blade. The investigation demonstrates the possible advantages of including hub clearance in axial flow compressor stator blade rows.Copyright © 1986 by ASME
67 citations
TL;DR: In this paper, the authors describe time-average heat-flux distributions obtained for the blade of a Garrett TFE 731-2 hp full-stage rotating turbine with and without injection.
Abstract: This paper describes time-average heat-flux distributions obtained for the blade of a Garrett TFE 731-2 hp full stage rotating turbine. Blade measurements were obtained both with and without injection. The injected gas was supplied from a separate reservoir and was directed into the turbine gas path via nozzle guide vane (NGV) pressure surface slots located at approximately 63 percent of the wetted distance. Blade heat-flux measurements were performed fro two different injection gas temperatures, T/sub c/T/sub o/ = 0.53 and T/sub c/T/sub o/ - 0.82. A shock tube is used as a short-duration source of heated air to which the turbine is subjected and thin-film gages are used to obtain the heat-flux measurements. Results are presented along the blade in the flow direction at 10, 50, and 90 percent span for both the pressure and suction surfaces. A sufficient number of measurements were obtained to also present spanwise distributions. At approximately the 50 percent span location, two contoured inserts containing closely spaced gages were installed in the blade so that the leading-edge region distribution could be resolved in detail. The blade results are compared with predictions obtained using a flat-plate technique and with predictions obtained using a version of STANmore » 5.« less
67 citations
TL;DR: In this paper, the results of a numerical simulation of the time-averaged inviscid flow field through the blade rows of a multiblade row turboprop configuration are presented.
Abstract: The results of a numerical simulation of the time-averaged inviscid flow field through the blade rows of a multiblade row turboprop configuration are presented. The governing equations are outlined along with a discussion of the solution procedure and coding strategy. Numerical results obtained from a simulation of the flow field through a modern high-speed turboprop will be shown.
66 citations
TL;DR: In this paper, an analytical technique for obtaining the time-resolved heat flux of a turbine blade is applied to the case of a TFE 731-2 hp full-stage rotating turbine.
Abstract: An analytical technique for obtaining the time-resolved heat flux of a turbine blade is applied to the case of a TFE 731-2 hp full-stage rotating turbine. In order to obtain the heat flux values from the thin film gage temperature histories, a finite difference procedure is used to solve the heat equation with variable thermal properties. After setting out the data acquisition and analysis procedures, their application is illustrated for three midspan locations on the blade and operation at the design flow function. Results demonstrate that the magnitude of the heat flux fluctuation due to vane-balde interaction is large by comparison to the time-averaged heat flux at all investigated locations; FFT of a portion of the heat flux record illustrates that the dominant frequencies occur at the wake-cutting frequency and its harmonics.
TL;DR: In this paper, the influence du jeu a l'extremite des aubes sur leur charge dans le cas d'une grille d'aubes lineaires de turbine.
Abstract: Influence du jeu a l'extremite des aubes sur leur charge dans le cas d'une grille d'aubes lineaires de turbine
TL;DR: The mid-span section of a previously reported controlled-diffusion compressor stator has been experimentally evaluated in cascade as mentioned in this paper, where the blade chord length is 12.7 cm, aspect ratio is 2.0, and solidity is 1.67.
Abstract: The mid-span section of a previously reported controlled-diffusion compressor stator has been experimentally evaluated in cascade. Measurements are taken over a range of incidence angles for blade chord Reynolds numbers from 470,000 to 690,000. Blade chord length is 12.7 cm, aspect ratio is 2.0, and solidity is 1.67. Measurements include conventional cascade performance parameters as well as blade surface pressures. Computations are made for the inviscid flow field, surface boundary layers, and loss for several of the blade inlet angle conditions, are compared against corresponding data.
TL;DR: In this article, a method is shown which, with the help of time-distance diagrams, made it possible to attribute the turbulence increase to the breakdown of the secondary vortices.
Abstract: Nonintrusive measurements near and within the rotor of a cold-air turbine showed a sudden increase of turbulence energy when the wake portion of the incoming fluid entered the rotor. It has been suggested that this was due to the cutting of the passage vortices and trailing-edge shed vortices which emerge from the stator row. Since these secondary vortices are located very close to the stator wakes, it was very difficult to distinguish between the effects of shed vortex and passage vortex cutting on turbulence intensification. In the present paper, a method is shown which, with the help of time–distance diagrams, made it possible to attribute the turbulence increase to the breakdown of the secondary vortices. Further, the time–distance diagrams made it possible to locate the origin of turbulence production and follow the spreading of the highly turbulent flow regions through the rotor channel.
TL;DR: Extension des elements d'aubage de type CDA (Controlled Diffusion Airfoil) a la region d'ecoulement de la paroi d'extremite afin d'obtenir une conception de Diffusion Controlee de la seconde generation optimisee Augementation de 1,5% du rendement and de 30% de la possibilite de fonctionnement sans pompage par rapport a la premiere generation as mentioned in this paper.
Abstract: Extension des elements d'aubage de type CDA (Controlled Diffusion Airfoil) a la region d'ecoulement de la paroi d'extremite afin d'obtenir une conception de Diffusion Controlee de la seconde generation optimisee Augementation de 1,5% du rendement et de 30% de la possibilite de fonctionnement sans pompage par rapport a la premiere generation
TL;DR: In this paper, a technique which enables measurements of the surface heat transfer rate to be made using thin-film gages deposited on a vitreous enamel-coated metal model is described.
Abstract: The paper describes a technique which enables measurements of the surface heat transfer rate to be made using thin-film gages deposited on a vitreous enamel-coated metal model. It is intended that this will have particular application in rotating turbine test rigs, where it offers considerable advantages over present techniques. These include ease of manufacture, instrumentation, durability, and lack of interference with the basic flow. The procedures for gage calibration and measurement processing are outlined, and the results of wind tunnel tests which confirm that the method is both practical and accurate are described.
TL;DR: In this paper, a two-dimensional blunt body with a circular leading edge is described and measurements of free-stream and boundary-layer velocity distributions are presented and indicate that a small separation "bubble" existed where the leading edge joined the body.
Abstract: Experiments for flow about a two-dimensional blunt body with a circular leading edge are described. Measurements of the free-stream and boundary-layer velocity distributions are presented and indicate that a small separation “bubble” existed where the leading edge joined the body. In particular, it was found that the laminar leading edge boundary layer separated and reattached shortly downstream as a turbulent boundary layer with a low-momentum thickness Reynolds number. Heat transfer measurements around the body are also presented and show almost an order of magnitude increase across the bubble. Downstream of the bubble, however, the heat transfer could be correlated by a slightly modified turbulent flat plate equation using the separation point as the virtual origin of the heated turbulent boundary layer.
TL;DR: In this paper, a systematic study of the influence of different parameters on the cooling effectiveness has been carried out and 27 test cases have been calculated, varying the injection angle (α = 10° / 45° / 90°), the relative spacing (s/D = 1.5/3/5) and the blowing rate (M = 0.5 /1/2) for the same main-stream conditions.
Abstract: The present paper describes three-dimensional calculations of film cooling by injection from a single row of holes. A systematic study of the influence of different parameters on the cooling effectiveness has been carried out. 27 test cases have been calculated, varying the injection angle (α = 10° / 45° / 90°), the relative spacing (s/D = 1.5/3/5) and the blowing rate (M = 0.5/1/2) for the same main-stream conditions. The governing 3D equations are solved by a finite volume method. The turbulent stresses and heat fluxes are obtained from a k-e model modified to account for non-isotropic eddy viscosities and diffusivities. Examples of predicted velocity and temperature distributions are presented and compared with available experimental data. For all the test cases, the laterally averaged cooling effectiveness is given. On the whole, the agreement with experiments is fairly good, even though there are discrepancies about details in some of the cases. The influence of the individual parameters on the film cooling effectiveness is predicted correctly in all cases. This influence is discussed in some detail and the parameter combination with the best overall cooling performance is identified.Copyright © 1985 by ASME
TL;DR: In this paper, the authors propose experimentale de l'utilisation de la courbure du bord d'attaque du stator for ameliorer les performances d'un compresseur a ecoulement axial.
Abstract: Etude experimentale de l'utilisation de la courbure du bord d'attaque du stator pour ameliorer les performances d'un compresseur a ecoulement axial. Examen de l'aerodynamique de trois conditions de moyeu de stator (diametre interieur)
TL;DR: In this paper, a mathematical model was developed to predict the enhanced coupled bending-torsion unstalled supersonic flutter stability due to alternate circumferential spacing aerodynamic detuning of a turbomachine rotor.
Abstract: A mathematical model developed to predict the enhanced coupled bending-torsion unstalled supersonic flutter stability due to alternate circumferential spacing aerodynamic detuning of a turbomachine rotor. The translational and torsional unsteady aerodynamic coefficients are developed in terms of influence coefficients, with the coupled bending-torsion stability analysis developed by considering the coupled equations of this aerodynamic detuning on coupled bending-torsion unstalled supersonic flutter as well as the verification of the modeling are then demonstrated by considering an unstable 12 bladed rotor, with Verdon's uniformly spaced Cascade B flow geometry as a baseline. However, with the elastic axis and center of gravity at 60 percent of the chord, this type of aerodynamic detuning has a minimal effect on stability. For both uniform and nonuniform circumferentially space rotors, a single degree of freedom torsion mode analysis was shown to be appropriate for values of the bending-torsion natural frequency ratio lower than 0.6 and higher 1.2. When the elastic axis and center of gravity are not coincident, the effect of detuning on cascade stability was found to be very sensitive to the location of the center of gravity with respect to the elastic axis. In addition, it was determined that when the center of gravity was forward of an elastic axis located at midchord, a single degree of freedom torsion model did not accurately predict cascade stability.
TL;DR: In this paper, the authors adapted the heated thin-film method to meet the requirements of investigations on boundary layer behavior in a turbine rig and developed special multi-sensor probes of vaporized nickel on a polyimide foil.
Abstract: The heated thin-film method was adapted to meet the requirements of investigations on boundary layer behaviour in a turbine rig. Special multi-sensor probes of vaporized nickel on a polyimide foil were developed and applied to the vanes.Basic experiments with an airfoil in a free stream were carried out and a reliable interpretation of the thin-film results was found by comparison with pressure distribution, flow visualization and laser measurements. It can be shown that this measuring device is a suitable method for the investigation of separation bubbles and boundary layer transition.Copyright © 1986 by ASME
TL;DR: Trous de petit diametre pour le refroidissement par film as discussed by the authors, transfert de chaleur par convection sur la paroi, et.
Abstract: Trous de petit diametre pour le refroidissement par film. Transfert de chaleur par convection sur la paroi
TL;DR: In this article, particle trajectories are calculated by direct integration of the particle equations of motion through a compressor impeller with two different size splitters, and particle collisions with the blade, hub and casing surfaces are determined.
Abstract: Gas turbine engines operating in areas where the atmosphere is polluted by small solid particles are examples of machines operating under particulate two phase flow conditions. The presence of solid particles in the working media leads to erosion and loss of engine power. In the present paper, particle trajectories are calculated by direct integration of the particle equations of motion through a compressor impeller with two different size splitters. Particle collisions with the blade, hub and casing surfaces are determined. Erosion caused by the particle collisions with the compressor surfaces are calculated.Copyright © 1986 by ASME
TL;DR: Experimental results from cascades and turbine rigs are presented in this article, where the boundary layer development in the cascade, on the guide vanes and on rotor blades with typically skewed boundary layers are compared to that in the turbine.
Abstract: Cascade testing tries to simulate the actual flow conditions encountered in a turbine. However, it is neither possible to reproduce the free stream turbulence structure of the turbomachinery, nor the periodic wake effects of upstream blade rows. The usual understanding is that the latter in particular results in a significantly different behaviour of the boundary layer in the engine.Experimental results from cascades and turbine rigs are presented. Grid generated free stream turbulence structure is compared to that in the turbine. Measurements of the profile pressure distribution, flush mounted hot films and flow visualization were used for the interpretation of the test results. Some observations of the boundary layer development in the cascade, on the guide vanes and on rotor blades with typically skewed boundary layers are shown indicating essentially similar behaviour in all cases.Copyright © 1986 by ASME