Showing papers in "Journal of Turbomachinery-transactions of The Asme in 2005"

Criteria for Spike Initiated Rotating Stall

Abstract: A computational study to define the phenomena that lead to the onset of short length-scale (spike) rotating stall disturbances has been carried out. Based on unsteady simulations, we hypothesize there are two conditions necessary for the formation of spike disturbances, both of which are linked to the tip clearance flow. One is that the interface between the tip clearance and oncoming flows becomes parallel to the leading-edge plane. The second is the initiation of backflow, stemming from the fluid in adjacent passages, at the trailing-edge plane. The two criteria also imply a circumferential length scale for spike disturbances. The hypothesis and scenario developed are consistent with numerical simulations and experimental observations of axial compressor stall inception. A comparison of calculations for multiple blades with those for single passages also allows statements to be made about the utility of single passage computations as a descriptor of compressor stall.

Effect of Hole Geometry on the Thermal Performance of Fan-Shaped Film Cooling Holes

Abstract: This study evaluates the impact of typical cooling hole shape variations on the thermal performance of fan-shaped film holes. A comprehensive set of experimental test cases featuring 16 different film-cooling configurations with different hole shapes have been investigated. The shape variations investigated include hole inlet-to-outlet area ratio, hole coverage ratio, hole pitch ratio, hole length, and hole orientation (compound) angle. Flow conditions applied cover a wide range of film blowing ratios M=0.5 to 2.5 at an engine-representative density ratio DR= 1.7. An infrared thermography data acquisition system is used for highly accurate and spatially resolved surface temperature mappings. Accurate local temperature data are achieved by an in situ calibration procedure with the help of thermocouples embedded in the test plate. Detailed film-cooling effectiveness distributions and discharge coefficients are used for evaluating the thermal performance of a row of fan-shaped film holes. An extensive variation of the main geometrical parameters describing a fan-shaped film-cooling hole is clone to cover a wide range of typical film-cooling applications in current gas turbine engines. Within the range investigated, laterally averaged film-cooling effectiveness was found to show only limited sensitivity from variations of the hole geometry parameters. This offers the potential to tailor the hole geometry according to needs beyond pure cooling performance, e.g., manufacturing facilitations.

Averaging Nonuniform Flow for a Purpose

Abstract: Averaging non-uniform flow is important for the analysis of measurements in turbomachinery and gas turbines; more recently an important need for averaging arises with results of CFD. In this paper we show that there is a method for averaging which is “correct”, in the sense of preserving the essential features of the non-uniform flow, but that the type of averaging which is appropriate depends on the application considered. The crucial feature is the decision to retain or conserve those quantities which are most important in the case considered. Examples are given to demonstrate the appropriate methods to average non-uniform flows in the accounting for turbomachinery blade row performance, production of thrust in a nozzle, and mass flow capacity in a choked turbine. It is also shown that the numerical differences for different types of averaging are, in many cases, remarkably small.Copyright © 2005 by ASME

Effects of Damping and Varying Contact Area at Blade-Disk Joints in Forced Response Analysis of Bladed Disk Assemblies

Abstract: An approach is developed to analyze the multiharmonic forced response of large-scale finite element models of bladed disks taking account of the nonlinear forces acting at the contact interfaces of blade roots. Area contact interaction is modeled by area friction contact elements which allow for friction stresses under variable normal load, unilateral contacts, clearances, and interferences. Examples of application of the new approach to the analysis of root damping and forced response levels are given and numerical investigations of effects of contact conditions at root joints and excitation levels are explored for practical bladed disks.

A Review of Some Early Design Practice Using Computational Fluid Dynamics and a Current Perspective

Abstract: In the early development of gas turbines, many empirical design rules were used; for example in obtaining fluid deflection using the deviation from blading angles, in the assumption of zero radial velocities (so-called radial equilibrium) and in expressions for clearance loss (the Lakshminarayana formulas). The validity of some of these rules, and the basic fluid mechanics behind them, is examined by use of modern ideas and computational fluid dynamics (CFD) codes. A current perspective of CFD in design is given, together with a view on future developments.

Detached Eddy Simulation of Film Cooling Performance on the Trailing Edge Cut-Back of Gas Turbine Airfoils

Abstract: The present study deals with the unsteady flow simulation of trailing edge film cooling on the pressure side cut-back of gas turbine airfoils. Before being ejected tangentially on the inclined cut-back surface, the coolant air passes a partly converging passage that is equipped with turbulators such as pin fins and ribs. The film mixing process on the cut-back is complicated. In the near slot region, due to the turbulators and the blunt pressure side lip, turbulence is expected to be anisotropic. Furthermore, unsteady flow phenomena like vortex shedding from the pressure side lip might influence the mixing process (i.e. the film cooling effectiveness on the cut-back surface). In the current study, three different internal cooling designs are numerically investigated starting from the steady RaNS solution, and ending with unsteady detached eddy simulations (DES). Blowing ratios M = 0.5; 0.8; 1.1 are considered. To obtain both, film cooling effectiveness as well as heat transfer coefficients on the cut-back surface, the simulations are performed using adiabatic and diabatic wall boundary conditions. The DES simulations give a detailed insight into the unsteady film mixing process on the trailing edge cut-back, which is indeed influenced by vortex shedding from the pressure side lip. Furthermore, the time averaged DES results show very good agreement with the experimental data in terms of film cooling effectiveness and heat transfer coefficients.Copyright © 2005 by ASME

The Dynamics of the Horseshoe Vortex and Associated Endwall Heat Transfer—Part I: Temporal Behavior

Abstract: Instantaneous flow topology and the associated endwall heat transfer in the leading-edge endwall region of a symmetric airfoil are presented. An experimental technique was employed that allowed the simultaneous recording of instantaneous particle image velocimetry flow field and thermochromic liquid-crystal-based endwall heat transfer data. The endwall flow is dominated by a horseshoe vortex that forms from reorganized impinging boundary layer vorticity. A corner vortex is shown to be a steady feature of the corner region, while a secondary vortex develops sporadically immediately upstream of the horseshoe vortex. The region upstream of the horseshoe vortex is characterized by a bimodal switching of the near-wall reverse flow, which results in quasi-periodic eruptions of the secondary vortex. The bimodal switching of the reverse flow in the vicinity of the secondary vortex is linked to the temporal behavior of the down-wash fluid on the leading edge of the foil. Frequency analysis of the flow field and endwall heat transfer data, taken together, indicate that the eruptive behavior associated with the horseshoe vortex occurs at a frequency that is essentially the same as the measured turbulence bursting period of the impinging turbulent endwall boundary layer.Copyright © 2005 by ASME

The Effect of Hot-Streaks on HP Vane Surface and Endwall Heat Transfer: An Experimental and Numerical Study

Abstract: Pronounced non-uniformities in combustor exit flow temperature (hot-streaks), which arise because of discrete injection of fuel and dilution air jets within the combustor and because of end-wall cooling flows, affect both component life and aerodynamics. Because it is very difficult to quantitatively predict the affects of these temperature non-uniformities on the heat transfer rates, designers are forced to budget for hot-streaks in the cooling system design process. Consequently, components are designed for higher working temperatures than the mass-mean gas temperature, and this imposes a significant overall performance penalty. An inadequate cooling budget can lead to reduced component life. An improved understanding of hot-streak migration physics, or robust correlations based on reliable experimental data, would help designers minimise the overhead on cooling flow that is currently a necessity. A number of recent research projects sponsored by a range of industrial gas turbine and aero-engine manufacturers attest to the growing interest in hot-streak physics. This paper presents measurements of surface and end-wall heat transfer rate for an HP nozzle guide vane (NGV) operating as part of a full HP turbine stage in an annular transonic rotating turbine facility. Measurements were conducted with both uniform stage inlet temperature and with two non-uniform temperature profiles. The temperature profiles were non-dimensionally similar to profiles measured in an engine. A difference of one half of an NGV pitch in the circumferential (clocking) position of the hot-streak with respect to the NGV was used to investigate the affect of clocking on the vane surface and end-wall heat transfer rate. The vane surface pressure distributions, and the results of a flow-visualisation study, which are also given, are used to aid interpretation of the results. The results are compared to two-dimensional predictions conducted using two different boundary layer methods. Experiments were conducted in the Isentropic Light Piston Facility (ILPF) at QinetiQ Farnborough, a short duration engine-size turbine facility. Mach number, Reynolds number and gas-to-wall temperature ratios were correctly modelled. It is believed that the heat transfer measurements presented in this paper are the first of their kind.© 2005 ASME

Heat transfer and aerodynamics of turbine blade tips in a linear cascade

Abstract: Local measurements of the heat transfer coefficient and pressure coefficient were conducted on the tip and near tip region of a generic turbine blade in a five-blade linear cascade. Two tip clearance gaps were used: 1.6% and 2.8% chord. Data was obtained at a Reynolds number of 2.3 X 10 5 based on exit velocity and chord. Three different tip geometries were investigated: A flat (plain) tip, a suction-side squealer, and a cavity squealer. The experiments reveal that the flow through the plain gap is dominated by flow separation at the pressure-side edge and that the highest levels of heat transfer are located where the flow reattaches on the tip surface. High heat transfer is also measured at locations where the tip-leakage vortex has impinged onto the suction surface of the aerofoil. The experiments are supported by flow visualization computed using the CFX CFD code which has provided insight into the fluid dynamics within the gap. The suction.vide and cavity squealers are shown to reduce the heat transfer in the gap but high levels of heat transfer are associated with locations of impingement, identified using the flow visualization and aerodynamic data. Film cooling is introduced on the plain tip at locations near the pressure-side edge within the separated region and a net heat flux reduction analysis is used to quantify the performance of the successful cooling design.

On the Use of Atmospheric Boundary Conditions for Axial-Flow Compressor Stall Simulations

Abstract: This paper describes a novel way of prescribing computational fluid dynamics (CFD) boundary conditions for axial-flow compressors. The approach is based on extending the standard single passage computational domain by adding an intake upstream and a variable nozzle downstream. Such a route allows us to consider any point on a given speed characteristic by simply modifying the nozzle area, the actual boundary conditions being set to atmospheric ones in all cases. Using a fan blade, it is shown that the method not only allows going past the stall point but also captures the typical hysteresis loop behavior of compressors.

Film Cooling Effectiveness and Heat Transfer on the Trailing Edge Cut-Back of Gas Turbine Airfoils With Various Internal Cooling Designs

Abstract: The present study deals with trailing edge film cooling on the pressure side cut-back of gas turbine airfoils. Before being ejected tangentially onto the inclined cut-back surface the coolant air passes a partly converging passage that is equipped with turbulators such as pin fins and ribs. The experiments are conducted in a generic set-up and cover a broad variety of internal cooling designs. A subsonic atmospheric open-loop wind tunnel is utilized for the tests. The test conditions are characterized by a constant Reynolds number of Rehg = 250,000, a turbulence intensity of Tuhg = 7%, and a hot gas temperature of Thg = 500K. Due to the ambient temperature of the coolant, engine realistic density ratios between coolant and gas can be realized. Blowing ratios cover a range of 0.20

Computational Fluid Dynamics Study of Wake-Induced Transition on a Compressor-Like Flat Plate

Abstract: Recent experimental work has documented the importance of wake passing on the behavior of transitional boundary layers on the suction surface of axial compressor blades. This paper documents computational fluid dynamics (CFD) simulations using a commercially available general-purpose CFD solver, performed on a representative case with unsteady transitional behavior. The study implements an advanced version of a three-equation eddy-viscosity model previously developed and documented by the authors, which is capable of resolving boundary layer transition. It is applied to the test cases of steady and unsteady boundary layer transition on a two-dimensional flat plate geometry with a freestream velocity distribution representative of the suction side of a compressor airfoil. The CFD results are analyzed and compared to a similar experimental test case from the open literature. Results with the model show a dramatic improvement over more typical Reynolds-averaged Navier-Stokes (RANS)-based modeling approaches, and highlight the importance of resolving transition in both steady and unsteady compressor aerosimulations.

Experimental Investigation of the Aerothermal Performance of a High Blockage Rib-Roughened Cooling Channel

Abstract: The present study deals with a detailed experimental investigation of the turbulent flow inside a rib-roughened turbine blade cooling channel. The measurements are carried out in a stationary straight channel with high blockage ribs installed on one wall. The main objective is to enhance the understanding and deepen the analysis of this complex flow field with the help of highly resolved particle image velocimetry measurements. A quasi-three-dimensional view of the flow field is achieved, allowing the identification of the main time-averaged coherent structures. The combined analysis of the present aerodynamic results with available heat transfer data emphasizes the role of the mean and fluctuating flow features in the heat transfer process. In particular, the stream wise/ normal to the wall component of the Reynolds stress tensor is shown to be strictly related to the heat transfer rate on the channel surfaces. A correlation to estimate the heat transfer field from the aerodynamic data is presented for the high blockage rib roughened channel flow.

Axial Compressor Deterioration Caused by Saltwater Ingestion

Abstract: Gas turbine performance deterioration can be a major economic factor. An example is within offshore installations where a degradation of gas turbine performance can mean a reduction of oil and gas production. This paper describes the test results from a series of accelerated deterioration tests on a General Electric J85-13 jet engine. The axial compressor was deteriorated by spraying atomized droplets of saltwater into the engine intake. The paper presents the overall engine performance deterioration as well as deteriorated stage characteristics. The results of laboratory analysis of the salt deposits are presented, providing insight into the increased surface roughness and the deposit thickness and distribution. The test data show good agreement with published stage characteristics and give valuable information regarding stage-by-stage performance deterioration.

Aerodynamics of Tip Leakage Flows Near Partial Squealer Rims in an Axial Flow Turbine Stage

Abstract: This paper deals with an experimental investigation of aerodynamic characteristics of full and partial-length squealer rims in a turbine stage. Full and partial-length squealer rims are investigated separately on the pressure side and on the suction side in the Axial Flow Turbine Research Facility (AFTRF) of the Pennsylvania State University. The streamwise length of these partial squealer tips and their chordwise position are varied to find an optimal aerodynamic tip configuration. The optimal configuration in this cold turbine study is defined as the one that is minimizing the stage exit total pressure defect in the tip vortex dominated zone. A new channel arrangement diverting some of the leakage flow into the trailing edge zone is also studied. Current results indicate that the use of partial squealer rims in axial flow turbines can positively affect the local aerodynamic field by weakening the tip leakage vortex. Results also show that the suction side partial squealers are aerodynamically superior to the pressure side squealers and the channel arrangement. The suction side partial squealers are capable of reducing the stage exit total pressure defect associated with the tip leakage flow to a significant degree.

Fan-Shaped Hole Effects on the Aero-Thermal Performance of a Film-Cooled Endwall

Abstract: The present paper investigates the effects of a fan-shaped hole endwall cooling geometry on the aero-thermal performance of a nozzle vane cascade. Two endwall cooling geometries with four rows of holes were tested, for different mass flow rate ratios: the first configuration is made of cylindrical holes, whereas the second one features conical expanded exits and a reduced number of holes. The experimental analysis is mainly focused on the variations of secondary flow phenomena related to different injection rates, as they have a strong relationship with the film cooling effectiveness. Secondary flow assessment was performed through downstream 3D aerodynamic measurements, by means of a miniaturized 5-hole probe. The results show that at high injection rates, the passage vortex and the 3D effects tend to become weaker, leading to a strong reduction of the endwall cross flow and to a more uniform flow in spanwise direction. This is of course obtained at the expense of a significant increase of losses. The thermal behavior was then investigated through the analysis of adiabatic effectiveness distributions on the two endwall configurations. The wide-banded thermochromic liquid crystals (TLC) technique was used to determine the adiabatic wall temperature. Using the measured distributions of film-cooling adiabatic effectiveness, the interaction between the secondary flow vortices and the cooling jets can be followed in good detail all over the endwall surface. Fan-shaped holes have been shown to perform better than cylindrical ones: at low injection rates, the cooling performance is increased only in the front part of the vane passage. A larger improvement of cooling coverage all over the endwall is attained with a larger mass flow rate, about 1.5% of core flow, without a substantial increase of the aerodynamic losses.

Modeling of Film Cooling—Part I: Experimental Study of Flow Structure

Abstract: This paper, which is Part I of a two part paper, reports on new experimental data taken in a steady flow, flat plate wind tunnel at ETH Zurich, while Part II utilizes this data for calibration and validation purpose of a new film cooling model embedded in a 3D CFD code. The facility simulates the film cooling row flow field on the pressure side of a turbine blade. Engine representative non-dimensionals are achieved, providing a faithful model at larger scale. Heating the free stream air and strongly cooling the coolant gives the required density ratio between coolant and free-stream. The three dimensional velocities are recorded using non-intrusive PIV, seeding is provided for both air streams. Two different cylindrical hole geometries are studied, with different angles. Blowing ratio is varied over a range to simulate pressure side film cooling. The three dimensional flow structures are revealed.Copyright © 2005 by ASME

PIV Maps of Tip Leakage and Secondary Flow Fields on a Low-Speed Turbine Blade Cascade With Moving End Wall

Abstract: New, detailed flow field measurements are presented for a very large low-speed cascade representative of a high-pressure turbine rotor blade with turning of 110 degrees and blade chord of 10 m Data was obtained for tip leakage and passage secondary flow at a Reynolds number of 40 × 105 , based on exit velocity and blade axial chord Tip clearance levels ranged from 0% to 168% of blade span (0% to 3% of blade chord) Particle Image Velocimetry (PIV) was used to obtain flow field maps of several planes parallel to the tip surface within the tip gap, and adjacent passage flow Vector maps were also obtained for planes normal to the tip surface in the direction of the tip leakage flow Secondary flow was measured at planes normal to the blade exit angle at locations upstream and downstream of the trailing edge The interaction between the tip leakage vortex and passage vortex is clearly defined, revealing the dominant effect of the tip leakage flow on the tip endwall secondary flow The relative motion between the casing and the blade tip was simulated using a motor-driven moving belt system A reduction in the magnitude of the under-tip flow near the endwall due to the moving wall is observed and the effect on the tip leakage vortex examinedCopyright © 2005 by ASME

An Empirical Prediction Method for Secondary Losses in Turbines: Part II — A New Secondary Loss Correlation

Abstract: A new empirical prediction method for design and off-design secondary losses in turbines has been developed. The empirical prediction method is based on a new loss breakdown scheme, and as discussed in Part I, the secondary loss definition in this new scheme differs from that in the conventional one. Therefore, a new secondary loss correlation for design and off-design incidence values has been developed. It is based on a database of linear cascade measurements from the present authors' experiments as well as cases available in the open literature. The new correlation is based on correlating parameters that are similar to those used in existing correlations. This paper also focuses on providing physical insights into the relationship between these parameters and the loss generation mechanisms in the endwall region. To demonstrate the improvements achieved with the new prediction method, the measured cascade data are compared to predictions from the most recent design and off-design secondary loss correlations (Kacker, S. C. and Okapuu, U., 1982, ASME J. Turbomach., 104, pp. 111-119, Moustapha, S. H., Kacker, S. C., and Tremblay, B., 1990, ASME J. Eng. Power, 112, pp. 267-276) using the conventional loss breakdown. The Kacker and Okapuu correlation is based on rotating-rig and engine data, and a scaling factor is needed to make their correlation predictions apply to the linear cascade environment. This suggests that there are additional and significant losses in the engine that are not present in the linear cascade environment.

Heat Transfer and Film-Cooling Measurements on a Stator Vane With Fan-Shaped Cooling Holes

Abstract: In a typical gas turbine engine, the gas exiting the combustor is significantly hotter than the melting temperature of the turbine components. The highest temperatures in an engine are typically seen by the turbine inlet guide vanes. One method used to cool the inlet guide vanes is film cooling, which involves bleeding comparatively low-temperature, high-pressure air from the compressor and injecting it through an array of discrete holes on the vane surface. To predict the vane surface temperatures in the engine, it is necessary to measure the heat transfer coefficient and adiabatic film-cooling effectiveness on the vane surface. This study presents heat transfer coefficients and adiabatic effectiveness levels measured in a scaled-up, two-passage cascade with a contoured endwall. Heat transfer measurements indicated that the behavior of the boundary layer transition along the suction side of the vane showed sensitivity to the location of film-cooling injection, which was simulated through the use of a trip wire placed on the vane surface. Single-row adiabatic effectiveness measurements without any upstream blowing showed jet lift-off was prevalent along the suction side of the airfoil. Single-row adiabatic effectiveness measurements on the pressure side, also without upstream showerhead blowing, indicated jet lifted-off and then reattached to the surface in the concave region of the vane. In the presence of upstream showerhead blowing, the jet lift-off for the first pressure side row was reduced, increasing adiabatic effectiveness levels.

The Dynamics of the Horseshoe Vortex and Associated Endwall Heat Transfer—Part II: Time-Mean Results

Abstract: Time-mean endwall heat transfer and flow-field data in the endwall region are presented for a turbulent juncture flow formed with a symmetric bluff body. The experimental technique employed allowed the simultaneous recording of instantaneous particle image velocimetry flow field data, and thermochromic liquid-crystal-based endwall heat transfer data. The time-mean flow field on the symmetry plane is characterized by the presence of primary (horseshoe), secondary, tertiary, and corner vortices. On the symmetry plane the time-mean horseshoe vortex displays a bimodal vorticity distribution and a stable-focus streamline topology indicative of vortex stretching. Off the symmetry plane, the horseshoe vortex grows in scale, and ultimately experiences a bursting, or breakdown, upon experiencing an adverse pressure gradient. The time-mean endwall heat transfer is dominated by two bands of high heat transfer, which circumscribe the leading edge of the bluff body. The band of highest heat transfer occurs in the corner region of the juncture, reflecting a 350% increase over the impinging turbulent boundary layer. A secondary high heat-transfer band develops upstream of the primary band, reflecting a 250% heat transfer increase, and is characterized by high levels of fluctuating heat load. The mean upstream position of the horseshoe vortex is coincident with a region of relatively low heat transfer that separates the two bands of high heat transfer.

Turbulent Transport in Pin Fin Arrays: Experimental Data and Predictions

Abstract: The objective of this research has been to experimentally investigate the fluid dynamics of pin fin arrays in order to clarify the physics of heat transfer enhancement and uncover problems in conventional turbulence models. The fluid dynamics of a staggered pin fin array have been studied using hot wire anemometry with both single and x-wire probes at array Reynolds numbers of 3000; 10,000; and 30,000. Velocity distributions off the endwall and pin surface have been acquired and analyzed to investigate turbulent transport in pin fin arrays. Well resolved 3-D calculations have been performed using a commercial code with conventional two-equation turbulence models. Predictive comparisons have been made with fluid dynamic data. In early rows where turbulence is low, the strength of shedding increases dramatically with increasing in Reynolds numbers. The laminar velocity profiles off the surface of pins show evidence of unsteady separation in early rows. In row three and beyond laminar boundary layers off pins are quite similar. Velocity profiles off endwalls are strongly affected by the proximity of pins and turbulent transport. At the low Reynolds numbers, the turbulent transport and acceleration keep boundary layers thin. Endwall boundary layers at higher Reynolds numbers exhibit very high levels of skin friction enhancement. Well resolved 3-D steady calculations were made with several two-equation turbulence models and compared with experimental fluid mechanic and heat transfer data. The quality of the predictive comparison was substantially affected by the turbulence model and near wall methodology.Copyright © 2005 by ASME

Effect of midpassage gap, endwall misalignment, and roughness on endwall film-cooling

Abstract: To maintain acceptable turbine airfoil temperatures, film cooling is typically used whereby coolant, extracted from the compressor, is injected through component surfaces. In manufacturing a turbine, the first stage vanes are cast in either single airfoils or double airfoils. As the engine is assembled, these singlets or doublets are placed in a turbine disk in which there are inherent gaps between the airfoils. The turbine is designed to allow outflow of high-pressure coolant rather than hot gas ingestion. Moreover, it is quite possible that the singlets or doublets become misaligned during engine operation. It has also become of interest to the turbine community as to the effect of corrosion and deposition of particles on component heat transfer. This study uses a large-scale turbine vane in which the following two effects are investigated: the effect of a midpassage gap on endwall film cooling and the effect of roughness on endwall film cooling. The results indicate that the midpassage gap was found to have a significant effect on the coolant exiting from the combustor-turbine interface slot. When the gap is misaligned, the results indicate a severe reduction in the film-cooling effectiveness in the case where the pressure side endwall is below the endwall associated with the suction side of the adjacent vane.

Analysis of Some Improved Blade Tip Concepts

Abstract: Over tip leakage in high-pressure turbines contributes to aerodynamic losses and migration of hot gasses towards the tip resulting in increased thermal distress. Consequently, turbine designers continue to search for improved blade tip concepts that offer the promise of reducing tip leakage. The present paper deals with the CFD analyses of some such tip configurations. The geometries, patented by GE, are variants of a conventional squealer tip and include (i) a pressure side tip shelf with vertical squealer tip wall, and (ii) a pressure side tip shelf with an inclined squealer tip wall. It is found that the inclined shelf results in separation of flow leaking over the tip resulting in reduced leakage and improved efficiency. The inclined shelf also shows a reduced efficiency derivative with clearance.Copyright © 2005 by ASME

Separation and Transition Control on an Aft-Loaded Ultra-High-Lift LP Turbine Blade at Low Reynolds Numbers: Low-Speed Investigation

Abstract: An experimental study was conducted to improve the performance of an aft-loaded ultra-high-lift low-pressure turbine blade known as U2 at low Reynolds numbers. This was achieved by manipulation of the laminar-turbulent transition process on the suction surface. The U2 profile was designed to meet the targets of reduced cost, weight and fuel burn of aircraft engines. The studies were conducted on both low-speed and high-speed experimental facilities under the unsteady flow conditions with upstream passing wakes. The current paper presents the low-speed investigation results. On the smooth suction surface, the incoming wakes are not strong enough to suppress the separation bubble due to the strong adverse pressure gradient on the suction surface and the low wake passing frequency, which allows the separation between the wakes more time to re-establish. Therefore, the profile losses of this ultra-high-lift blade are not as low as conventional or high-lift blades at low Reynolds numbers even in unsteady flows. Two different types of passive separation control devices, i.e. surface trips and air jets, were investigated to further improve the blade performance. The measurement results show that the profile losses can be further reduced to the levels similar to those of the high-lift and conventional blades due to the aft-loaded nature of this ultra-high-lift blade. Detailed surveys of the blade surface boundary layer developments showed that the loss reduction was due to the suppression of the separation underneath the wakes, the effect of the strengthened calmed region and the smaller separation bubble between wakes.Copyright © 2005 by ASME

Experimental and Numerical Study of Impingement on an Airfoil Leading Edge With and Without Showerhead and Gill Film Holes

Abstract: This experimental investigation deals with impingement on the leading-edge of an airfoil with and without showerhead film holes and its effects on heat transfer coefficients on the airfoil nose area as well as the pressure and suction side areas. A comparison between the experimental and numerical results are also made. The tests were run for a range of flow conditions pertinent to common practice and at an elevated range of jet Reynolds numbers (8000–48000). The major conclusions of this study were: a) the presence of showerhead film holes along the leading edge enhances the internal impingement heat transfer coefficients significantly, and b) while the numerical predictions of impingement heat transfer coefficients for the no-showerhead case were in good agreement with the measured values, the case with showerhead flow was underpredicted by as much as 30% indicating a need for a more elaborate turbulence modeling.Copyright © 2005 by ASME

An Empirical Prediction Method for Secondary Losses in Turbines—Part I: A New Loss Breakdown Scheme and Penetration Depth Correlation

Abstract: Despite its wide use in meanline analyses, the conventional loss breakdown scheme is based on a number of assumptions that are known to be physically unsatisfactory. One of these assumptions states that the loss generated in the airfoil surface boundary layers is uniform across the span. The loss results at high positive incidence presented in a previous paper (Benner et al. [1]) indicate that this assumption causes the conventional scheme to produce erroneous values of the secondary loss component. A new empirical prediction method for secondary losses in turbines has been developed, and it is based on a new loss breakdown scheme. In the first part of this two-part paper, the new loss breakdown scheme is presented. Using data from the current authors’ off-design cascade loss measurements (Benner et al. [1]), it is shown that the secondary losses obtained with the new scheme produce a trend with incidence that is physically more reasonable. Unlike the conventional loss breakdown scheme, the new scheme requires a correlation for the spanwise penetration depth of the passage vortex separation line at the trailing edge. One such correlation exists (Sharma and Butler [2]); however, it was based on a small database. An improved correlation for penetration distance has been developed from a considerably larger database, and it is detailed in this paper.Copyright © 2005 by ASME

Large Eddy Simulation of Flow and Heat Transfer in the Developing Flow Region of a Rotating Gas Turbine Blade Internal Cooling Duct With Coriolis and Buoyancy Forces

Abstract: The problem of accurately predicting the flow and heat transfer in the ribbed internal cooling duct of a rotating gas turbine blade is addressed with the use of large eddy simulations (LES). Four calculations of the developing flow region of a rotating duct with ribs on opposite walls are used to study changes in the buoyancy parameter at a constant rotation rate. The Reynolds number is 20,000, the rotation number is 0.3, and the buoyancy parameter is varied between 0.00, 0.25, 0.45, and 0.65. Previous experimental studies have noted that leading wall heat transfer augmentation decreases as the buoyancy parameter increases with low buoyancy, but heat transfer then increases with high buoyancy. However, no consistent physical explanation has been given in the literature. The LES results from this study show that the initial decrease in augmentation with buoyancy is a result of larger separated regions at the leading wall. However, as the separated region spans the full pitch between ribs with an increase in buoyancy parameter, it leads to increased turbulence and increased entrainment of mainstream fluid, which is redirected toward the leading wall by the presence of a rib. The impinging mainstream fluid results in heat transfer augmentation in the region immediately upstream of a rib. The results obtained from this study are in very good agreement with previous experimental results.

Effects of Surface-Roughness Geometry on Separation-Bubble Transition

Abstract: This paper presents measurements of separation-bubble transition over a range of surfaces with randomly distributed roughness elements. The tested roughness patterns represent the typical range of roughness conditions encountered on in-service turbine blades. Through these measurements, the effects of size and spacing of the roughness elements, and the tendency of the roughness pattern toward protrusions or depressions (skewness), on the inception location and rate of transition are evaluated. Increased roughness height, increased spacing of the roughness elements, and a tendency of the roughness pattern toward depressions (negative skewness) are observed to promote earlier transition inception. The observed effects of roughness spacing and skewness are found to be small in comparison to that of the roughness height. Variation in the dominant mode of instability in the separated shear layer is achieved through adjustment of the streamwise pressure distribution. The results provide examples for the extent of interaction between viscous and inviscid stability mechanisms.Copyright © 2005 by ASME

Experimental Investigation of Local Heat Transfer Distribution on Smooth and Roughened Surfaces Under an Array of Angled Impinging Jets

Abstract: Measurements of the local heat transfer distribution on smooth and roughened surfaces under an array of angled impinging jets are presented. The test rig is designed to simulate impingement with crossflow in one direction. Jet angle is varied between 30, 60, and 90 deg as measured from the target surface, which is either smooth or randomly roughened. Liquid crystal video thermography is used to capture surface temperature data at five different jet Reynolds numbers ranging between 15,000 and 35,000. The effect of jet angle, Reynolds number, gap, and surface roughness on heat transfer and pressure loss is determined along with the various interactions among these parameters.