bluuebunny/crossref_metadata_2025 huggingface dataset:
Aircraft Turbofan Noise
Authors: Groeneweg J. F. , Rice E. J.
Published: November 2009

Recent advances in the understanding of turbofan noise generation and suppression in aircraft engines are reviewed with particular emphasis on NASA research. The review addresses each link in the chain of physical processes which connect unsteady flow interactions with fan blades to far field noise. Mechanism identification and description, duct propagation, radiation, and acoustic suppression are discussed. Recent advances in the experimental technique of fan inflow control assure that inflight generation mechanisms are not masked by extraneous sources in static tests. Rotor blade surface pressure and wake velocity measurements aid the determination of the types and strengths of the generation mechanisms. Approaches to predicting or measuring acoustic mode content, optimizing treatment impedance to maximize attenuation, translating impedance into porous wall structure, and interpreting far field directivity patterns are illustrated by comparisons of analytical and experimental results. A persistent theme of the review is the interdependence of source and acoustic treatment design to minimize far field noise. Areas requiring further research are discussed and the relevance of aircraft turbofan results to quieting other turbomachinery installations is addressed.
A Critical Evaluation of Three Centrifugal Compressors With Pedigree Data Sets: Part 5—Studies in Component Performance
Authors: Japikse D.
Published: November 2009

A series of three centrifugal compressor data sets, each passing rigorous tests for quality, has been completed. Using this data base, single-zone and two-zone modeling is thoroughly critiqued in order to judge the appropriateness of such modeling and, hence, to set guidelines for subsequent design optimization using either single-zone or two-zone modeling.
Three-Dimensional Boundary Layer on a Compressor Rotor Blade at Peak Pressure Rise Coefficient
Authors: Lakshminarayana B. , Popovski P.
Published: November 2009

A comprehensive study of the three-dimensional turbulent boundary layer on a compressor rotor blade at peak pressure rise coefficient is reported in this paper. The measurements were carried out at various chordwise and radial locations on a compressor rotor blade using a rotating miniature V configuration hot-wire probe. The data are compared with the measurement at the design condition. Substantial changes in the blade boundary layer characteristics are observed, especially in the outer 16 percent of the blade span. The increased chordwise pressure gradient and the leakage flow at the peak pressure coefficient have a cumulative effect in increasing the boundary layer growth on the suction surface. The leakage flow has a beneficial effect on the pressure surface. The momentum and boundary layer thicknesses increase substantially from those at the design condition, especially near the outer radii of the suction surface.
Three-Dimensional Flow Field Measurements in a Radial Inflow Turbine Scroll Using LDV
Authors: Malak M. F. , Hamed A. , Tabakoff W.
Published: November 2009

The results of an experimental study of the three-dimensional flow field in a radial inflow turbine scroll are presented. A two-color LDV system was used in the measurement of three orthogonal velocity components at 758 points located throughout the scroll and the unvaned portion of the nozzle. The cold flow experimental results are presented for through-flow velocity contours and the cross velocity vectors.
Laser-Doppler Studies of the Wake-Effected Flow Field in a Turbine Cascade
Authors: Wittig S. , Dullenkopf K. , Schulz A. , Hestermann R.
Published: November 2009

The interaction between consecutive blade rows can be expected to have important effects on the heat transfer in cooled gas turbine cascades. In determining the local heat transfer under the influence of wake flow, nonintrusive optical measuring techniques were used to obtain the flow velocities and turbulence structures in the cascade inlet flow as well as along the test blade’s surface. The main purpose of the measurements is to provide accurate experimental data for the development of predictive codes. The applicability primarily of the laser-Doppler technique is discussed and problems arising from the use of laser-dual-focus anemometry are reported. In simulating the effects of wake flow, a plane airfoil was traversed in front of the cascade. Both the axial distance between the airfoil and the cascade and the position in circumferential direction were changed in discrete steps. Turbulence intensities between 1.4 and 15 percent were recorded in cold gas flow. The effects on the blade heat transfer are illustrated.
Three-Dimensional Flow in a Low-Pressure Turbine Cascade at Its Design Condition
Authors: Hodson H. P. , Dominy R. G.
Published: November 2009

This paper describes an experimental study of the three-dimensional flow within a high-speed linear cascade of low-pressure turbine blades. Data were obtained using pneumatic probes and a surface flow visualization technique. It is found that in general, the flow may be described using concepts derived from previous studies of high-pressure turbines. In detail, however, there are differences. These include the existence of a significant trailing shed vortex and the interaction of the endwall fluid with the suction surface flow. At an aspect ratio of 1.8, the primary and secondary losses are of equal magnitude.
Production and Development of Secondary Flows and Losses in Two Types of Straight Turbine Cascades: Part 1—A Stator Case
Authors: Yamamoto A.
Published: November 2009

The present study intends to give some experimental information on secondary flows and on the associated total pressure losses occurring within turbine cascades. Part 1 of the paper describes the mechanism of production and development of the loss caused by secondary flows in a straight stator cascade with a turning angle of about 65 deg. A full representation of superimposed secondary flow vectors and loss contours is given at fourteen serial traverse planes located throughout the cascade. The presentation shows the mechanism clearly. Distributions of static pressures and of the loss on various planes close to blade surfaces and close to an endwall surface are given to show the loss accumulation process over the surfaces of the cascade passage. Variation of mass-averaged flow angle, velocity and loss through the cascade, and evolution of overall loss from upstream to downstream of the cascade are also given. Part 2 of the paper describes the mechanism in a straight rotor cascade with a turning angle of about 102 deg.
Production and Development of Secondary Flows and Losses in Two Types of Straight Turbine Cascades: Part 2—A Rotor Case
Authors: Yamamoto A.
Published: November 2009

Part 1 of this paper [1] presents the detailed mechanism of secondary flows and the associated losses occurring within a straight stator cascade with a relatively low turning angle of about 65 deg. The significant contribution of secondary flows on the loss production process was shown only near the blade suction surface downstream from the cascade throat (Z/Cax = 0.74) in which regional flows decelerated due to adverse pressure gradient. In the second part, the same experimental analysis is applied to a straight rotor cascade with a much larger turning angle of 102 deg. Flow surveys were made at 12 traverse planes located throughout the rotor cascade. The larger turning results in a similar but much stronger contribution of the secondary flows to the loss developing mechanism. Evolution of overall loss starts quite early within the cascade, and the rate of the loss growth is much larger in the rotor case than in the stator case.
Unsteady Flow Interaction Caused by Stator Secondary Vortices in a Turbine Rotor
Authors: Binder A. , Forster W. , Mach K. , Rogge H.
Published: November 2009

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.
Experimental Investigation of Boundary Layer Separation With Heated Thin-Film Sensors
Authors: Pucher P. , Go¨hl R.
Published: November 2009

The heated thin-film method was adapted to meet the requirements of investigations on boundary layer behavior in a turbine rig. Special multisensor 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.
Unducted Fan Aerodynamic Design
Authors: Smith L. H.
Published: November 2009

A method being used to design unducted fans, which are counterrotating propellers with unusually high disk loadings, is described. The quasi-three-dimensional approach employed is similar to that used for ducted fan designs; it consists of a circumferential-average flowfield analysis followed by cascade designs that include secondary flow and sweep effects. The open-tip condition enters as a significant secondary flow contribution. Blade shapes defined by this process are analyzed by a three-dimensional Euler solver and final adjustments are made as necessary. Examples of application of the method are presented to elucidate the magnitudes of the phenomena involved.
An Accurate and Efficient Euler Solver for Three-Dimensional Turbomachinery Flows
Authors: Shieh C. F. , Delaney R. A.
Published: November 2009

Accurate and efficient Euler equation numerical solution techniques are presented for analysis of three-dimensional turbomachinery flows. These techniques include an efficient explicit hopscotch numerical scheme for solution of the three-dimensional time-dependent Euler equations and an O-type body-conforming grid system. The hopscotch scheme is applied to the conservative form of the Euler equations written in general curvilinear coordinates. The grid is constructed by stacking from hub to shroud two-dimensional O-type grids on equally spaced surfaces of revolution. Numerical solution results for two turbine cascades are presented and compared with experimental data to demonstrate the accuracy of the analysis method.
Some Experiments With a Supersonic Axial Compressor Stage
Authors: Wennerstrom A. J.
Published: November 2009

Between 1970 and 1974, ten variants of a supersonic axial compressor stage were designed and tested. These included two rotor configurations, three rotor tip clearances, addition of boundary-layer control consisting of vortex generators on both the outer casing and the rotor, and the introduction of slots in the stator vanes. Design performance objectives were a stage total pressure ratio of 3.0 with an isentropic efficiency of 0.82 at a tip speed of 1600 ft/s (488 m/s). The first configuration passed only 70 percent of design flow at design speed, achieving a stage pressure ratio of 2.25 at a peak stage isentropic efficiency of 0.61. The rotor was grossly separated. The tenth variant passed 91.4 percent of design flow at design speed, producing a stage pressure ratio of 3.03 with an isentropic efficiency of 0.75. The rotor achieved a pressure ratio of 3.59 at an efficiency of 0.87 under the same conditions. Major conclusions were that design tools available today would undoubtedly permit the original goals to be met or exceeded. However, the application for such a design is currently questionable because efficiency goals considered acceptable for most current programs have risen considerably from the level considered acceptable at the inception of this effort. Splitter vanes placed in the rotor permitted very high diffusion levels to be achieved without stalling. However, viscous effects causing three-dimensional flows violating the assumption of flow confined to concentric stream tubes were so strong that a geometry optimization does not appear practical without a three-dimensional, viscous analysis. Passive boundary-layer control in the form of vortex generators and slots does appear to offer some benefit under certain circumstances.
On the Shear-Stress Integral of Turbulent Boundary Layers
Authors: Pfeil H. , Go¨ing M.
Published: November 2009

The paper presents an integral method to predict turbulent boundary layer behavior in two-dimensional, incompressible flow. The method is based on the momentum and moment-of-momentum integral equations and a friction law. By means of the compiled data of the 1968 Stanford Conference, the results show that the integral of the turbulent shear stress across the boundary layer, which appears in the moment-of-momentum integral equation, can be described using only two basic assumptions for all cases of flow.
Heat Transfer and Fluid Mechanics Measurements in Transitional Boundary Layers on Convex-Curved Surfaces
Authors: Wang T. , Simon T. W.
Published: November 2009

An experimental investigation of the effects of convex curvature and free-stream turbulence on momentum and heat transfer in a boundary layer undergoing natural transition was performed. The test section was designed for providing a two-dimensional boundary layer flow on a uniformly heated curved surface. The unique feature of the test wall was its flexibility to be bent to various radii of curvature R. Three cases of R = ∞, 180 cm, and 90 cm, are presented, each with two free-stream turbulence intensity levels, 0.68 and 2 percent. Mean velocity and temperature profiles and local Stanton number data were taken in a heated flow and streamwise Reynolds normal stress profiles were taken in an isothermal flow. Mild convex curvature (R = 180 cm) is shown to delay transition. Bending the test wall to a smaller radius of curvature, 90 cm, results in no significant further delay of transition. Cases with both curvature and higher free-stream disturbance effects show a pronounced dominance of the latter. Heat transfer rates and mean temperature profiles are shown to change more slowly after transition than wall shear stress values and velocity profiles. This can be observed as higher values of turbulent Prandtl number and lower values of the Reynolds-analogy factor, 2St/Cf, in the transitional and early-turbulent flows than values known to apply for mature turbulent flows. The results of this study increase the data base for development of transition prediction models which can be used in gas turbine blade design.
Mixed-Flow Compressor Stage Design and Test Results With a Pressure Ratio of 3:1
Authors: Musgrave D. S. , Plehn N. J.
Published: November 2009

This paper presents a brief history of mixed-flow compressors, possible applications, and the design and measured performance of a recently tested 3:1 pressure-ratio stage. The stage is intended to run behind a multistage axial compressor; it has an envelope radius only 9.4 percent greater than the rotor tip radius. A tandem cascade diffusing system is used to promote flow range and thus aid matching to the axial stages. Compressor maps from the rig test are presented along with additional data (from static taps and exit rakes) that characterize the behavior of various elements of the stage.
The Measurement of Boundary Layers on a Compressor Blade in Cascade: Part 1—A Unique Experimental Facility
Authors: Deutsch S. , Zierke W. C.
Published: November 2009

A unique cascade facility is described which permits the use of laser-Doppler velocimetry (LDV) to measure blade boundary layer profiles. Because of the need for a laser access window, the facility cannot reply on continuous blade pack suction to achieve two-dimensional, periodic flow. Instead, a strong suction upstream of the blade pack is used in combination with tailboards to control the flow field. The distribution of the upstream suction is controlled through a complex baffling system. A periodic, two–dimensional flow field is achieved at a chord Reynolds number of 500,000 and an incidence angle of 5 deg on a highly loaded, double circular arc, compressor blade. Inlet and outlet flow profiles, taken using five-hole probes, and the blade static-pressure distribution are used to document the flow field for use with the LDV measurements (see Parts 2 and 3). Inlet turbulence intensity is measured, using a hot wire, to be 0.18 percent. The static-pressure distribution suggests both separated flow near the trailing edge of the suction surface and an initially laminar boundary layer profile near the leading edge of the pressure surface. Probe measurements are supplemented by sublimation surface visualization studies. The sublimation studies place boundary layer transition at 64.2 ± 3.9 percent chord on the pressure surface, and indicate separation on the suction surface at 65.6 percent ± 3.5 percent chord.
Influence of the Reynolds Number on the Performance of Centrifugal Compressors
Authors: Strub R. A. , Bonciani L. , Borer C. J. , Casey M. V. , Cole S. L. , Cook B. B. , Kotzur J. , Simon H. , Strite M. A.
Published: November 2009

This work is the result of an investigation based on numerous test data supplied by major compressor manufacturers in USA and in Europe. The main objective of the work is to propose improved formulae for the correction of the efficiency, the head, and the flow as influenced by the Reynolds number variation between workshop tests and specified conditions, carried out with the same machine. Tests on hand have shown that a sufficiently good correlation between measured and predicted values can be obtained with the proposed formulae. In addition a proposal is made for the allowable range, taking into account the inherent limitations for accurate testing at low Reynolds numbers. As a conclusion to this study it is recommended that the proposed formulae and allowable range be reviewed by the ASME, the ISO, or any other appropriate associations for adoption in revised test codes for centrifugal compressors.
Prediction of Film Cooling by a Row of Holes With a Two-Dimensional Boundary-Layer Procedure
Authors: Scho¨nung B. , Rodi W.
Published: November 2009

The present paper describes predictions of film cooling by a row of holes. The calculations have been performed by a two-dimensional boundary-layer code with special modifications that account for the basically three-dimensional, elliptic nature of the flow after injection. The elliptic reverse-flow region near the injection is leapt over and new boundary-layer profiles are set up after the blowing region. They take into account the oncoming boundary layer as well as the characteristics of the injected jets. The three dimensionality of the flow, which is very strong near the injection and decreases further downstream, is modeled by so-called dispersion terms, which are added to the two-dimensional boundary-layer equations. These terms describe additional mixing by the laterally nonuniform flow. Information on the modeling of the profiles after injection and of the dispersion terms has been extracted from three-dimensional fully elliptic calculations for specific flow configurations. The modified two-dimensional boundary-layer equations are solved by a forward-marching finite-volume method. A coordinate system is used that stretches with the growth of the boundary layer. The turbulent stresses and heat fluxes are obtained from the k-ε turbulence model. Results are given for flows over flat plates as well as for flows over gas turbine blades for different injection angles, relative spacings, blowing rates, and injection temperatures. The predicted cooling effectiveness and heat transfer coefficients are compared with experimental data and show generally fairly good agreement.
A Trace Gas Technique to Study Mixing in a Turbine Stage
Authors: Joslyn H. D. , Dring R. P.
Published: November 2009

An experimental technique to study mixing in a turbine stage is demonstrated. An axisymmetric, radial temperature profile at the inlet to the first stator of a large-scale, low-speed, single-stage, axial flow turbine model is simulated with a radial trace gas concentration distribution. Mixing or redistribution of the inlet profile by three-dimensional aerodynamic mechanisms (other than temperature-driven mechanisms) is determined from trace gas concentration measurements made in both the stationary and rotating frames of reference at various locations through the turbine. The trace gas concentration contours generated are consistent with flow pitch angle measurements made downstream of the first stator and with surface flow visualization on the rotor airfoil and the hub endwall. It is demonstrated that this trace gas technique is well suited to quantify many aspects of the redistribution and diffusion of an inlet temperature profile as it is convected through a turbine stage.
Time-Averaged Heat-Flux Distributions and Comparison With Prediction for the Teledyne 702 HP Turbine Stage
Authors: Dunn M. G. , Chupp R. E.
Published: November 2009

Time-averaged heat-flux distributions are reported for the vane and blade of the Teledyne CAE 702 HP full-stage rotating turbine. 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. The thin-film gages were concentrated on the midspan region from the leading edge to near the trailing edge. The blade contained two contoured inserts wtih gages spaced very close together so that the leading edge distribution could be resolved. The NGV and blade results are compared with predictions obtained using a flat-plate technique, an eddy-diffusing model (STAN 5), and a k–ε model. The results of the comparison between data and prediction suggest that: (a) first, the vane data are bounded by the turbulent flat plate and the fully turbulent STAN 5 prediction. For the vane, the k–ε prediction is in relatively good agreement with the STAN 5 prediction and (b) secondly, the blade data are acceptably predicted by the k–ε prediction on both the pressure and the suction surfaces. The STAN 5 fully turbulent calculation for the blade falls above the data (essentially in agreement with the turbulent flat-plate calculation) and the STAN 5 fully laminar falls substantially below the data. With the exception of the pressure loadings and the geometry, the code inputs used for these predictions were identical to those previously used to predict the Garrett TFE 731-2 HP turbine and the Garrett LART HP turbine.
A Method for Aerodynamic Design of Blades in Quasi-Three-Dimensional Calculation of Turbomachines
Authors: Wang Zhengming
Published: November 2009

A special inverse problem is formulated in which the shape of the mean streamline and the circumferential thickness distribution of the profile are given. On the basis of the series expansion method on a selected streamline, in quasi-three-dimensional aerodynamic design, the blade profile thickness is automatically fulfilled by computer. Six radial sections of a turbine blade are designed by this method.
Prediction of Turbulent Source Flow Between Corotating Disks With an Anisotropic Two-Equation Turbulence Model
Authors: Shirazi S. A. , Truman C. R.
Published: November 2009

An anisotropic form of a low-Reynolds-number two-equation turbulence model has been implemented in a numerical solution for incompressible turbulent flow between corotating parallel disks. Transport equations for turbulent kinetic energy and dissipation rate were solved simultaneously with the governing equations for the mean-flow variables. Comparisons with earlier mixing-length predictions and with measurements are presented. Good agreement between the present predictions and the measurements of velocity components and turbulent kinetic energy was obtained. The low-Reynolds-number two-equation model was found to model adequately the near-wall region as well as the effects of rotation and streamline divergence, which required ad hoc assumptions in the mixing-length model.
Numerical Prediction of Turbulent Flow in Rotating Cavities
Authors: Morse A. P.
Published: November 2009

Predictions of the isothermal, incompressible flow in the cavity formed between two corotating plane disks and a peripheral shroud have been obtained using an elliptic calculation procedure and a low turbulence Reynolds number k–ε model for the estimation of turbulent transport. Both radial inflow and outflow are investigated for a wide range of flow conditions involving rotational Reynolds numbers up to ∼106. Although predictive accuracy is generally good, the computed flow in the Ekman layers for radial outflow often displays a retarded spreading rate and a tendency to laminarize under conditions that are known from experiment to produce turbulent flow.
Flow Characteristics of Long Orifices With Rotation and Corner Radiusing
Authors: McGreehan W. F. , Schotsch M. J.
Published: November 2009

The discharge coefficient of an orifice is a function of both geometric and flow effects, such as inlet corner radius, orifice length, inlet velocity orientation, and Reynolds number. Loss characteristics are available for each of these effects; however, a need exists for a reasonable method of predicting the discharge coefficient with an arbitrary combination of the above effects. Presented is a technique for calculating the discharge coefficient for arbitrary geometry with results compared to test data.
Vibration Amplitudes of Mistuned Blades
Authors: Afolabi D.
Published: November 2009

Due to the mistuning effect, the nominally identical blades on a rotor are forced to vibrate with greatly unequal amplitudes under certain circumstances. It is, therefore, desirable to have the capability of predicting the highest responding blades so that such blades may be instrumented during engine tests. However, the predictions of various investigators in this regard are apparently inconsistent. Usually, the inconsistency is attributed to differences in mathematical models. By using modal analysis, it is shown that the various results are really not in conflict, but merely reflect the local and contrasting features in the global characteristics of a typical bladed disk. Good agreement is obtained when the results of this study, which is based on a model with structural coupling, are compared with those of other investigators utilizing models with aerodynamic and structural coupling. It is concluded that if the primary resonance being excited is not in a “crossover” zone, the highest responding blades are most likely to be those with extreme mistuning.
Verification of Compressor Data Accuracy by Uncertainty Analysis and Testing Methods
Authors: Poti N. D. , Rabe D. C.
Published: November 2009

A transonic compressor designed and instrumented by the General Electric Company was recently evaluated during several test programs in the Compressor Research Facility at Wright-Patterson Air Force Base. An analytical uncertainty analysis was performed for the individual measurements, as well as the calculated performance parameters. During the tests, the experimental values of precision uncertainty were compared to the analytical predictions. The spatial variation of the individual discharge measurements was evaluated for its effect on the uncertainty of the measured parameters and calculated performance. Through the comparison of the analytical and measured uncertainty and the evaluation of spatial variation, a final uncertainty of the measured performance is presented, and the overall quality of the test results is assessed. Through these data reviews, a more accurate understanding of the performance of the test compressor is obtained.
Application of Viscous Flow Computations for the Aerodynamic Performance of a Backswept Impeller at Various Operating Conditions
Authors: Hah C. , Bryans A. C. , Moussa Z. , Tomsho M. E.
Published: November 2009

Three-dimensional flowfields in a centrifugal impeller with backswept discharge at various operating points have been numerically investigated with a three-dimensional viscous flow code. Numerical results and experimental data were compared for the detailed flowfields and overall performance of the impeller at three operating conditions (optimum efficiency, choke, and near-surge conditions). The comparisons indicate that for engineering applications the numerical solution accurately predicts various complex real flow phenomena. The overall aerodynamic performance of the impeller is also well predicted at design and off-design conditions.
Experimental Investigations of Flows Through a Plane Cascade at Large Angles of Attack With Separations
Authors: Gu Chuan-Gang , Luo Lai-Qin , Miao Yong-Miao
Published: November 2009

Measurements of various parameters in the flows through a cascade at different angles of attack have been performed. The parameters, such as Reynolds stresses (uu/U2, vv/U2, and uv/U2), pressure distribution on the blade surface, velocity distribution in the blade passage, position of the separation point, and so on, are measured at a large angle of attack with separations. In addition, the development of wake is also investigated. A new formula with second-order accuracy has been developed to analyze hot-wire signals in flows with high turbulence intensity. The hot-wire data are compared with those by conventional measurement techniques and by flow visualizations. The results are satisfactory.
Interaction Mechanisms Between Tip Leakage Flow and the Passage Vortex in a Linear Turbine Rotor Cascade
Authors: Yamamoto A.
Published: November 2009

In order to study the loss generation mechanisms due to the tip-leakage flow in turbine rotor passages, extensive traverse measurements were made of the three-dimensional flows in a low-speed linear cascade for various tip-clearance sizes and for various cascade inlet flow angles (or incidences). Effects of the leakage flow on the cascade downstream flow fields and interactions between the leakage flow and the passage vortices are discussed in detail based on the traverse measurements and flow-visualization tests in terms of secondary flows and the associated losses. Other traverses were also performed of the tip-casing endwall flows both inside and outside the tip-clearance gap using a micro five-hole pitot tube to reveal the axial development of the interaction throughout the cascade passage. Overall loss characteristics of the present high-turning cascade with blunt leading and trailing edges are obtained and compared with those predicted by the Ainley–Mathieson method.
Navier–Stokes Solution for Steady Two-Dimensional Transonic Cascade Flows
Authors: Kwon O. K.
Published: November 2009

A robust, time-marching Navier–Stokes solution procedure based on the explicit hopscotch method is presented for solution of steady, two-dimensional, transonic turbine cascade flows. The method is applied to the strong conservation form of the unsteady Navier–Stokes equations written in arbitrary curvilinear coordinates. Cascade flow solutions are obtained on an orthogonal, body-conforming “O” grid with the standard k–ε turbulence model. Computed results are presented and compared with experimental data.
Shock Embedding Discontinuous Solution of Elliptic Equation for Inverse Problem of Transonic S2 Flow
Authors: Chen Hongji , Wu Chung-Hua
Published: November 2009

In the case of the inverse problem of S2 surface flow in a transonic turbomachine, if the meridional component of the relative velocity is subsonic, the equation governing the flow is elliptic. Through the use of a proper conservative form of the stream-function principal equation and embedding the shock relations into the principal equation, the transonic flow over the whole S2 surface containing the discontinuity of the passage shock can be calculated. A computer code employing this method has been programmed. The algorithm, which is only a little different from the subsonic code, is simple, accurate, and reliable. This program is particularly useful in the solution of the three-dimensional transonic flow in a fan or compressor through iterative computation of transonic S1 and S2 flows, when the former is carried out by the recently developed method of “separate-region calculation with shock fitting.” The new computer code is used to calculate the S2 flow in a typical transonic rotor and the difference between the result obtained by this method and that by the commonly used one, in which the shock discontinuity is not taken into account, can be clearly seen.
The Design of Shock-Free Compressor Cascades Including Viscous Boundary-Layer Effects
Authors: Hassan A. A.
Published: November 2009

A numerical procedure has been developed for the design of shock-free transonic compressor cascades with an allowance for viscous effects, providing that the boundary layer is fully attached over the blade. The method described combines, in an iterative process, a modified inviscid hodograph-based inverse-design algorithm (CIDA), originally developed by the author for the design of shock-free airfoils, and the inverse boundary-layer algorithm (LTBLCEQL) of Miner et al. [22]. In the numerical procedure, the inviscid subsonic and supersonic regions of the flow are decoupled allowing the solution of either an elliptic or hyperbolic-type partial differential equation for the full stream function. For the subcritical portion of the flow, the inviscid calculation is performed in a computational plane which is obtained through a sequence of conformal and numerical transformations of the two-sheeted hodograph plane. For the supercritical portion, a characteristic calculation is carried out in the hodograph plane. The results are then mapped back to the physical plane to determine the inviscid blade configuration. Viscous effects are then incorporated via the boundary-layer displacement surface concept. The boundary-layer algorithm incorporates a two-layer eddy viscosity turbulence model and allows for gradual, rather than instantaneous, transition to turbulence. Two examples of shock-free compressor blades are given to demonstrate the capabilities of the numerical coupling procedure.
A Full Quasi-Three-Dimensional Calculation of Flow in Turbomachines
Authors: Ribaut M.
Published: November 2009

The problem of calculating a three-dimensional flow with three families of two-dimensional solutions is considered. From a throughflow solution and several blade-to-blade solutions the boundary values of a set of transverse solutions are obtained. The resulting overdetermined boundary condition problem is solved by means of a compatibility function defining the divergence of the transverse velocity field. The influence of the latter on the blade-to-blade solution is formulated in two different ways and compared with the experiment. The new method should considerably improve the prediction of three-dimensional cascade flow presenting a large amount of streamwise vorticity in the meridional direction.
Experimental Investigation of Secondary Flow and Mixing Downstream of Straight Turbine Cascades
Authors: Mobarak A. , Khalafallah M. G. , Osman A. M. , Heikal H. A.
Published: November 2009

The purpose of this paper is to investigate the flow field downstream of turbine cascades of low aspect ratio, often used in vehicles and small turbomachines. Experimental investigation was carried out to study the flow downstream of three sets of turbine cascades having the same blade turning angle of about 83 deg but different profiles. The total energy losses were measured at several planes downstream of the cascade of blades in order to determine the changes in gross secondary flow loss coefficient and the growth of the mixing loss with distance downstream. Influence of inlet boundary layer thickness, aspect ratio, and exit Mach number on the nature of the flow at the exit plane of the cascade and total energy loss were studied. The tests were performed with four values of aspect ratio: 1.16, 0.8, 0.5, and 0.25. Some new correlations were deduced that predict energy loss coefficients as a function of distance downstream, aspect ratio, and exit Mach number as well as the upstream boundary layer thickness. The test results compare well with other published correlations.
Influence of Deposit on the Flow in a Turbine Cascade
Authors: Bo¨lcs A. , Sari O.
Published: November 2009

An experimental study on a gas turbine cascade operating under transonic flow conditions is presented. The flow is compared for airfoil shapes corresponding to the design geometry and the geometry taken from a rotor blade, in an industrial gas turbine burning heavy oil, after a few thousand hours of operation. Steady-state data have been obtained in a linear cascade over a range of isentropic exit Mach numbers from 0.6 to 1.6. The flow field was determined by static pressure measurements on the side walls up- and downstream of the cascade, on one side wall in the blade passage, and on the blade surface at midspan. Furthermore, the flow was visualized by the methods of Schlieren and laser holography. The results show that the choked flow conditions are reached at different steady-state isentropic outlet Mach numbers for the two blade shapes. The deposit, typical for a gas turbine, does not however significantly modify the boundary layer separation point. The flow visualization indicates that the shock wave fluctuations have not been significantly influenced by the important roughness and thickness of the deposit. The experimental results on the two cascades are also compared with two-dimensional time-marching calculations after Denton. In the subsonic regime, good agreement was found for the “clean” blade. For the profile with deposit, the flow cannot be correctly predicted by the time-marching calculation, even in subsonic flow condition. The sonic line calculated by the numerical model under transonic outlet conditions (0.9 < M2S < 1.20) does not agree with the laser holography measurements for either of the two cascades.
Numerical Analysis of Airfoil and Cascade Flows by the Viscous/Inviscid Interactive Technique
Authors: Hwang C. J. , Jiang F. L. , Hsieh J. M. , Chang S. B.
Published: November 2009

A viscous-inviscid interaction calculation is performed to study the steady, two-dimensional, incompressible/subsonic compressible, attached and separated flows for isolated airfoils and airfoil cascades. A full-potential code was coupled with a laminar/transition/turbulent finite difference code using the semi-inverse method. For the potential flow, the finite element method is employed and the circulation is considered as an unknown parameter. In order to handle the problem efficiently, an automatic grid generation technique is necessary. For the incompressible flow, the solution can be achieved without iteration. However, for the compressible flow, a “pseudotime integral” is used to find a steady-state solution. To understand the viscous effect, the boundary layer equations are solved by the implicit, finite difference method. For the turbulent flow, the algebraic eddy-viscosity formulation of Cebeci and Smith is used. The location of transition from laminar to turbulent flow is predicted or specified by empirical data correlations. The transitional region is taken into account by an empirical intermittency factor. With regard to separated flows, the FLARE approximation and inverse method are introduced. In order to evaluate the present solution procedure, the numerical results are compared to the theoretical and experimental data given in other papers and reports.
A New Variational Finite Element Computation for the Aerodynamic Inverse Problem in Turbines With Long Blades
Authors: Qin Ren , Yi Fengming , Wang Hongguang
Published: November 2009

In this paper a new finite element computation for the aerodynamic inverse problem in turbines with long blades is proposed. The influences of blade thickness, blade force, and other factors on aerodynamic parameters may be taken into account. The computational results are compared with those of the finite difference method in noncurvilinear coordinates. They indicate that this method has an appropriate accuracy, good suitability to the complex boundaries, and good convergence. This new finite element computation is a good method for aerodynamic design in turbines with long blades. By use of this method the influence of the various design factors (curved and leaned blades, blade thickness, etc.) on the distributions of the aerodynamic parameters can be studied.
Effects of Embedded Vortices on Film-Cooled Turbulent Boundary Layers
Authors: Ligrani P. M. , Ortiz A. , Joseph S. L. , Evans D. L.
Published: November 2009

Heat transfer effects of longitudinal vortices embedded within film-cooled turbulent boundary layers on a flat plate were examined for free-stream velocities of 10 m/s and 15 m/s. A single row of film-cooling holes was employed with blowing ratios ranging from 0.47 to 0.98. Moderate-strength vortices were used with circulating-to-free stream velocity ratios of −0.95 to −1.10 cm. Spatially resolved heat transfer measurements from a constant heat flux surface show that film coolant is greatly disturbed and that local Stanton numbers are altered significantly by embedded longitudinal vortices. Near the downwash side of the vortex, heat transfer is augmented, vortex effects dominate flow behavior, and the protection from film cooling is minimized. Near the upwash side of the vortex, coolant is pushed to the side of the vortex, locally increasing the protection provided by film cooling. In addition, local heat transfer distributions change significantly as the spanwise location of the vortex is changed relative to film-cooling hole locations.
Cold Flow Turbine Rig Tests of the Original and Redesigned Compressor Turbines of an Industrial Gas Turbine Engine
Authors: Diakunchak I. S.
Published: November 2009

This paper describes the results of cold flow turbine rig tests carried out on the original and redesigned compressor turbines of an industrial gas turbine engine. Some details of the aerodynamic design of the latest variant, a brief description of the advanced technology design methods used in this design, and a description of the test facility are included. Bulk stage performance and detail rotor exit radial-circumferential traverse results are presented. These test results demonstrate that the design point stage efficiency of the redesigned compressor turbine is about six percentage points higher than that of the original design.
Evaluation of Turbine Field Acceptance Tests: A Bayesian Approach
Authors: Bury K. V. , Spurr K. J. W.
Published: November 2009

A Bayesian error analysis is developed for treating systematic and random errors of field test data from acceptance tests of large hydraulic turbine installations. A practical procedure is suggested for allocating test points among power levels for a specified statistical precision of efficiency determination.
The Measurement and Formation of Tip Clearance Loss
Authors: Bindon J. P.
Published: November 2009

The detailed development of tip clearance loss from the leading to trailing edge of a linear turbine cascade was measured and the contributions made by mixing, internal gap shear flow, and endwall/ secondary flow were identified, separated, and quantified for the first time. Only 13 percent of the overall loss arises from endwall/secondary flow and of the remaining 87 percent, 48 percent is due to mixing and 39 percent is due to internal gap shear. All loss formation appears to be dominated by phenomena connected with the gap separation bubble. Flow established within the bubble by the pressure gradient separates as the gradient disappears and most of the internal loss is created by the entrainment of this separated fluid. When this high-loss leakage wake enters the mainstream, it separates due to the suction corner pressure gradient to create virtually all the measured mixing loss. It is suggested that the control of tip clearance loss by discharge coefficient reduction actually introduces loss. Performance improvements may result from streamlined tip geometries that optimize the tradeoff between entropy production and flow deflection.
Measurement and Modeling of the Gas Turbine Blade Transition Process as Disturbed by Wakes
Authors: LaGraff J. E. , Ashworth D. A. , Schultz D. L.
Published: November 2009

Heat transfer measurements have been made on a transonic turbine blade undergoing natural transition and with a simulation of the effect of NGV wake interactions. The use of wide bandwidth heat transfer instrumentation permits the tracking of individual unsteady events that were identified as being due to either the impinging wakes or to the turbulent spots occurring within the transition process. Trajectories of these events as seen by the blade surface instrumentation have been measured. Numerical models have been developed for the effects of both types of turbulent activity. The convection of the wake through the passage is predicted, allowing for estimations of the expected times for which the boundary layer is disturbed by the wake fluid. The new model for the random generation and subsequent growth and convection of the turbulent spots produces a time-resolved prediction of the intermittent heat transfer signals by use of a time-marching procedure. By superimposing the two numerical models it is possible to simulate the measured instantaneous heat transfer characteristics.
Boundary-Layer Flows in Rotating Cavities
Authors: Ong C. L. , Owen J. M.
Published: November 2009

A rotating cylindrical cavity with a radial outflow of fluid provides a simple model of the flow between two corotating air-cooled gas-turbine disks. The flow structure comprises a source region near the axis of rotation, boundary layers on each disk, a sink layer on the peripheral shroud, and an interior core of rotating inviscid fluid between the boundary layers. In the source region, the boundary layers entrain fluid; outside this region, nonentraining Ekman-type layers are formed on the disks. In this paper, the differential boundary-layer equations are solved to predict the velocity distribution inside the entraining and nonentraining boundary layers and in the inviscid core. The equations are discretized using the Keller-box scheme, and the Cebeci–Smith eddy-viscosity model is used for the turbulent-flow case. Special problems associated with reverse flow in the nonentraining Ekman-type layers are successfully overcome. Solutions are obtained, for both laminar and turbulent flow, for the “linear equations” (where nonlinear inertial terms are neglected) and for the full nonlinear equations. These solutions are compared with earlier LDA measurements of the radial and tangential components of velocity made inside a rotating cavity with a radial outflow of air. Good agreement between the computations and the experimental data is achieved for a wide range of flow rates and rotational speeds.