bluuebunny/crossref_metadata_2025 huggingface dataset:
The Effect of Density Ratio on the Heat Transfer Coefficient From a Film-Cooled Flat Plate
Authors: Ammari H. D. , Hay N. , Lampard D.
Published: June 2008

The effect of density ratio of cooling films on the heat transfer coefficient on a flat plate is investigated using a heat-mass transfer analogy. The experimental technique employed uses a swollen polymer surface and laser holographic interferometry. A density ratio of 1.0 was achieved using air as the injectant. Density ratios of 1.38 and 1.52, representative of turbine operating conditions, were obtained by using foreign gases. The coolant fluids were injected at various blowing rates through a single normal hole or through a row of holes spaced at three-diameter intervals, and inclined at 35 or 95 deg to the mainstream direction. The experiments were conducted under isothermal conditions in a subsonic, zero mainstream pressure gradient turbulent boundary layer. The results indicated large differences in behavior between the two injection angles. For normal injection, the heat transfer coefficient at a fixed blowing parameter was insensitive to the variation of density ratio, whereas for 35 deg injection strong dependence was observed. Scaling parameters for the heat transfer data have been proposed so that use can be made of data obtained at density ratios not representative of gas turbine practice. In addition, a correlation for normal injection data has been formulated.
Local Heat Transfer in Internally Cooled Turbine Airfoil Leading Edge Regions: Part I—Impingement Cooling Without Film Coolant Extraction
Authors: Bunker R. S. , Metzger D. E.
Published: June 2008

An experimental study has been designed and performed to measure very localized internal heat transfer characteristics in large-scale models of turbine blade impingement-cooled leading edge regions. Cooling is provided by a single line of equally spaced multiple jets, aimed at the leading edge apex and exiting the leading edge region in the opposite or chordwise direction. Detailed two-dimensional local surface Nusselt number distributions have been obtained through the use of aerodynamically steady but thermally transient tests employing temperature-indicating coatings. The thin coatings are sprayed directly on the test surface and are observed during the transient with automated computer vision and data acquisition systems. A wide range of parameter combinations of interest in cooled airfoil practice are covered in the test matrix, including combinations of variations in jet Reynolds number, airfoil leading edge sharpness, jet pitch-to-diameter ratio, and jet nozzle-to-apex travel distance. Measured local Nusselt numbers at each chordwise location back from the stagnation line have been used to calculate both the spanwise average Nusselt number and spanwise Nusselt number gradient as functions of chordwise position. Results indicate general increases in heat transfer with approximately the 0.6 power of jet Reynolds number, increases in heat transfer with decreasing leading edge sharpness as well as with decreasing nozzle-to-apex distance, and increases in spanwise average heat transfer with decreasing jet pitch-to-diameter ratio. The latter increases are accompanied by increases in the spanwise gradient of the heat transfer coefficient. Comparison with available prior results of much coarser spatial resolution shows good agreement and establishes confidence in the use of the results for design purposes and as baseline results for comparison with subsequent experiments involving film cooling bleed.
Local Heat Transfer in Internally Cooled Turbine Airfoil Leading Edge Regions: Part II—Impingement Cooling With Film Coolant Extraction
Authors: Metzger D. E. , Bunker R. S.
Published: June 2008

An experimental study has been designed and performed to measure very localized internal heat transfer characteristics in large-scale models of turbine blade impingement-cooled leading edge regions that allow extraction, or bleed-off, of a portion of the internal cooling flow to provide leading edge film cooling along the blade external surface. The internal impingement air is provided by a single line of equally spaced multiple jets, aimed at the leading edge apex and generally exiting, minus the bleed-off flow, in the opposite or chordwise direction. The film coolant flow extraction takes place through two lines of holes, one each on the blade suction side and the blade pressure side, both fairly close to the airfoil leading edge. Detailed two-dimensional local surface Nusselt number distributions have been obtained through the use of aerodynamically steady but thermally transient tests employing temperature-indicating coatings. The thin coatings are sprayed directly on the test surfaces, and are observed during a test transient with automated computer vision and data acquisition systems. A wide range of parameter combinations of interest in cooled airfoil practice is covered in the test matrix, including combinations of variations in jet Reynolds number, airfoil leading edge sharpness, jet pitch-to-diameter ratio, and jet nozzle-to-apex travel distance. Measured local Nusselt numbers at each chordwise location back from the stagnation line have been used to calculate both the spanwise-average Nusselt numbers and spanwise Nusselt number gradients as functions of chordwise position. The results without film coolant extraction, presented in the Part I companion paper, are used as a basis of comparison to determine the additional effects of the film cooling bleed. Results indicate that heat transfer is primarily dependent on jet Reynolds number with smaller influences from the flow extraction rate. The results also suggest that changes in the spanwise alignment of the impingement nozzles relative to the position of the film cooling holes can cause significant variations in leading edge metal temperatures.
Impingement/Effusion Cooling: The Influence of the Number of Impingement Holes and Pressure Loss on the Heat Transfer Coefficient
Authors: Al Dabagh A. M. , Andrews G. E. , Abdul Husain R. A. A. , Husain C. I. , Nazari A. , Wu J.
Published: June 2008

Measurements of the overall heat transfer coefficient within an impingement/effusion cooled wall are presented. The FLUENT CFD computer code has been applied to the internal aerodynamics to demonstrate the importance of internal recirculation in the impingement gap. This generates a convective heat transfer to the impingement jet. Measurements of this heat transfer plate coefficient are presented that show it to be approximately half of the impingement/effusion heat transfer coefficient. The influence of the relative pressure loss or X/D between the impingement and effusion walls was investigated, for an effusion X/D of 4.67 and a Z of 8 mm, and shown to be only significant at high G where a reduction in h of 20 percent occurred. Increasing the number of holes N in the impingement/effusion array at a constant Z of 8 mm reduced h by 20 percent, mainly due to the higher Z/D for the smaller holes at high N. Reduced numbers of impingement holes relative to the effusion holes, in a ratio of 1 to 4, were shown to have a small influence on h with a maximum reduction in h of 20 percent at high G and a negligible effect at low G.
An Experimental Study of Turbine Vane Heat Transfer With Leading Edge and Downstream Film Cooling
Authors: Nirmalan N. V. , Hylton L. D.
Published: June 2008

This paper presents the effects of downstream film cooling, with and without leading edge showerhead film cooling, on turbine vane external heat transfer. Steady-state experimental measurements were made in a three-vane, linear, two-dimensional cascade. The principal independent parameters—Mach number, Reynolds number, turbulence, wall-to-gas temperature ratio, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio—were maintained over ranges consistent with actual engine conditions. The test matrix was structured to provide an assessment of the independent influence of parameters of interest, namely, exit Mach number, exit Reynolds number, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio. The vane external heat transfer data obtained in this program indicate that considerable cooling benefits can be achieved by utilizing downstream film cooling. The downstream film cooling process was shown to be a complex interaction of two competing mechanisms. The thermal dilution effect, associated with the injection of relatively cold fluid, results in a decrease in the heat transfer to the airfoil. Conversely, the turbulence augmentation, produced by the injection process, results in increased heat transfer to the airfoil. The data presented in this paper illustrate the interaction of these variables and should provide the airfoil designer and computational analyst with the information required to improve heat transfer design capabilities for film-cooled turbine airfoils.
Prediction of Heat Transfer Characteristics for Discrete Hole Film Cooling for Turbine Blade Applications
Authors: Tafti D. K. , Yavuzkurt S.
Published: June 2008

A two-dimensional injection model is used with a two-dimensional low Reynolds number k-ε model boundary layer code. The three-dimensional effects of the discrete hole injection process are introduced in the two-dimensional prediction scheme through an “entrainment fraction” (Υ). An established correlation between Υ and the injection parameters obtained in a previous paper is used to predict the film cooling effectiveness (η) and heat transfer coefficients for multirow injection, injection into a laminar boundary layer, and finally injection on convex curved surfaces. Predictions of η are in good agreement with experimental data for most of the cases tested. Predictions of Stanton numbers defined by St(0) and St(l) are good for low injection ratios (M) but as M increases the values are underpredicted. In spite of some shortcomings, in the authors’ opinion the present two-dimensional prediction scheme is one of the most comprehensive developed so far. It is seen that the entrainment fraction Υ is quite universal in its application to two-dimensional predictions of the discrete hole film cooling process.
Behavior of a Coolant Film With Two Rows of Holes Along the Pressure Side of a High-Pressure Nozzle Guide Vane
Authors: Arts T. , Bourguignon A. E.
Published: June 2008

The purpose of this paper is to quantify the influence on external convective heat transfer of a coolant film whose position varies along the pressure side of a high-pressure turbine nozzle guide vane. The measurements were performed in the short-duration Isentropic Light Piston Compression Tube facility of the von Karman Institute. The effects of external and internal flow are considered in terms of Mach number, Reynolds number, free-stream turbulence intensity, blowing rate, and coolant to free-stream temperature ratio. The way to evaluate these results in terms of film cooling efficiency and heat transfer coefficient is finally discussed.
Effects of Wake Passing on Stagnation Region Heat Transfer
Authors: O’Brien J. E.
Published: June 2008

An experimental study is described in which both time-averaged and time-resolved effects of wake passing were measured in a cylinder stagnation region. The experiments were carried out in an annular-flow wind tunnel, which was fitted with a spoked-wheel wake generator. The cylindrical spokes produce wakes that simulate those shed from a turbine inlet guide vane. Time-averaged heat transfer results indicate an asymmetric distribution of heat transfer coefficient about the stagnation line, with higher heat transfer coefficients on the windward side (with respect to the bar-passing direction), which corresponds to the suction side of a turbine blade. This asymmetry is also reflected in the time-resolved heat transfer results, which were obtained using a test cylinder instrumented with platinum thin-film gages. Unsteady heat flux records reveal very large positive excursions (as much as a factor of three) in instantaneous heat flux during wake passing on the windward side of the cylinder and much smaller effects on the leeward side. Hot-film records in the cylinder stagnation region were also obtained by operating the thin-film gages in the constant-temperature mode. Spectra of these hot-film records indicate that vortex shedding is a major contributor to the unsteady buffeting of the test-cylinder boundary layer at circumferential stations located at both + 60 deg and − 60 deg from the stagnation line, but makes a very small contribution on the stagnation line itself.
Phase and Time-Resolved Measurements of Unsteady Heat Transfer and Pressure in a Full-Stage Rotating Turbine
Authors: Dunn M. G.
Published: June 2008

This paper presents detailed phase-resolved heat-flux data obtained on rotor blades and a comparison of simultaneously obtained time-resolved heat-flux and static pressure data obtained on the stationary shroud of a Garrett TFE 731-2 HP full-stage rotating turbine. A shock tube is used to generate a short-duration source of heated and pressurized air and platinum thin-film gages are used to obtain heat-flux measurements. Blade results are presented at several selected blade locations. Shroud surface pressure and heat-flux time histories are presented for comparable locations relative to the blade position. For these measurements, the turbine was operating at the design flow function, the design stage pressure ratio, and at 100 percent corrected speed.
Aerodynamics of Cooling Jets Introduced in the Secondary Flow of a Low-Speed Turbine Cascade
Authors: Bario F. , Leboeuf F. , Onvani A. , Seddini A.
Published: June 2008

The aerodynamic behavior of cold discrete jets in a cold secondary flow is investigated. Configurations including single jets and rows of jets are studied. These jets are introduced through the side wall of a low-speed nozzle turbine cascade. The experimental setup and the jet behavior are fully described. The effects of location with respect to the blades, mass flow ratio, yaw, and incidence angles on the aerodynamics of single jets are investigated. The influence of neighboring jets is detailed in the case of multiple jet configurations. The interaction with the secondary flow is presented. The local pressure and velocity fields, trajectories, and visualizations are discussed. The measuring apparatus includes a five-hole probe and a hot wire for intermittency measurements.
Experimental Determination of Stator Endwall Heat Transfer
Authors: Boyle R. J. , Russell L. M.
Published: June 2008

Local Stanton numbers were experimentally determined for the endwall surface of a turbine vane passage. A six-vane linear cascade having vanes with an axial chord of 13.81 cm was used. Results were obtained for Reynolds numbers based on inlet velocity and axial chord between 73,000 and 495,000. The test section was connected to a low-pressure exhaust system. Ambient air was drawn into the test section; inlet velocity was controlled up to a maximum of 59.4 m/s. The effect of the inlet boundary layer thickness on the endwall heat transfer was determined for a range of test section flow rates. The liquid crystal measurement technique was used to measure heat transfer. Endwall heat transfer was determined by applying electrical power to a foil heater attached to the cascade endwall. The temperature at which the liquid crystal exhibited a specific color was known from a calibration test. Lines showing this specific color were isotherms, and because of uniform heat generation they were also lines of nearly constant heat transfer. Endwall static pressures were measured, along with surveys of total pressure and flow angles at the inlet and exit of the cascade.
Characteristics of Partial Length Circular Pin Fins As Heat Transfer Augmentors for Airfoil Internal Cooling Passages
Authors: Arora S. C. , Abdel-Messeh W.
Published: June 2008

Pin fins are commonly used as heat transfer augmentors for internal cooling of turbine airfoils. These pins may extend from one wall to the other or may be segmented to meet specific requirements of removing the airfoil’s varying heat load. Three configurations of the partial pins were tested in a channel with aspect ratio 25:1 and the results are compared with those for the full pins. The array average heat transfer rate decreases linearly with increasing gap and is bounded by the value for full pins at one end and that for the smooth channel at the other. However, the local distribution of the Nusselt number and the average for each of the two walls depends on the configuration of the partial pins. The friction factor was lower for partial pins than for the full pins and also decreased with increasing gap. For the configuration with all partial pins on one wall, the friction factor was found to be the lowest, with no change in the corresponding heat transfer rate from a wall with pins.
Highly Loaded Axial Flow Compressors: History and Current Developments
Authors: Wennerstrom A. J.
Published: June 2008

This paper discusses approaches taken over many years to achieve very high loading levels in axial-flow compressors. These efforts have been associated predominantly with aircraft turbine engines. The objective has been to reduce the size and weight of the powerplant, to increase its simplicity and ruggedness, and, whenever possible, to reduce cost. In the introduction, some fundamentals are reviewed that indicate that increased work per stage can only be obtained at a cost of increased Mach number, increased diffusion, or both. The earliest examples cited are some ambitious development programs of the 1950s and 1960s. Some innovative schemes to increase diffusion limits are described that took place in the 1960s and 1970s. Major advancements in dealing with higher Mach number were made in the 1980s. Finally, a few thoughts directed toward potential future developments are presented.
Aerodynamic Performance of a Scale-Model, Counterrotating Unducted Fan
Authors: Sullivan T. J.
Published: June 2008

The aerodynamic performance of a scale-model, counterrotating unducted fan has been determined and the results are discussed herein. Experimental investigations were conducted using the scale-model propulsor simulator and uniquely shaped fan blades designed by the General Electric Company under contract to NASA. The blades, designed for a high disk loading at Mach 0.72, 35,000 ft altitude max climb condition, are aft-mounted on the simulator in a “pusher” configuration. Data are compared with analytical predictions at the design point and show good agreement.
Three-Dimensional Relief in Turbomachinery Blading
Authors: Wadia A. R. , Beacher B. F.
Published: June 2008

The leading edge region of turbomachinery blading in the vicinity of the endwalls is typically characterized by an abrupt increase in the inlet flow angle and a reduction in total pressure associated with endwall boundary layer flow. Conventional two-dimensional cascade analysis of the airfoil sections at the endwalls indicates large leading edge loadings, which are apparently detrimental to the performance. However, experimental data exist that suggest that cascade leading edge loading conditions are not nearly as severe as those indicated by a two-dimensional cascade analysis. This discrepancy between two-dimensional cascade analyses and experimental measurements has generally been attributed to inviscid three-dimensional effects. This article reports on two and three-dimensional calculations of the flow within two axial-flow compressor stators operating near their design points. The computational results of the three-dimensional analysis reveal a significant three-dimensional relief near the casing endwall that is absent in the two-dimensional calculations. The calculated inviscid three-dimensional relief at the endwall is substantiated by airfoil surface static pressure measurements on low-speed research compressor blading designed to model the flow in the high-speed compressor. A strong spanwise flow toward the endwall along the leading edge on the suction surface of the airfoil is responsible for the relief in the leading edge loading at the endwall. This radial migration of flow results in a more uniform spanwise loading compared to that predicted by two-dimensional calculations.
Tip Leakage Losses in a Linear Turbine Cascade
Authors: Dishart P. T. , Moore J.
Published: June 2008

An investigation of tip leakage flow and its effects on loss production was performed on a large-scale linear turbine cascade having a tip gap measuring 2.1 percent of the blade height. The flow exiting the tip gap was measured to determine the losses incurred within the tip gap and the secondary kinetic energy due to tip leakage. Additional measurements, 40 percent of an axial chord downstream of the blade trailing edges, showed the development of the leakage flow and the overall cascade losses. At the downstream location, the additional loss due to tip leakage was found to be the sum of the measured loss at the tip gap exit plane and the amount of tip gap secondary kinetic energy that had been dissipated by that downstream location.
Development of the Tip-Leakage Flow Downstream of a Planar Cascade of Turbine Blades: Vorticity Field
Authors: Yaras M. , Sjolander S. A.
Published: June 2008

The paper presents detailed measurements of the tip-leakage flow emerging from a planar cascade of turbine blades. Four clearances of from 1.5 to 5.5 percent of the blade chord are considered. Measurements were made at the trailing edge plane, and at two main planes 1.0 and 1.56 axial chord lengths downstream of the cascade. The results give insight into several aspects of the leakage flow, including the size and strength of the leakage vortex in relation to the size of the tip gap and the bound circulation of the blade, and the evolution of the components of vorticity as the vortex diffuses laterally downstream of the blade row. The vortex was found to have largely completed its roll-up into a nearly axisymmetric structure even at the trailing edge of the cascade. As a result, it was found that the vortex could be modeled surprisingly well with a simple model based on the diffusion of a line vortex.
Secondary Loss Generation in a Linear Cascade of High-Turning Turbine Blades
Authors: Harrison S.
Published: June 2008

Flow through a linear cascade of high-turning, low aspect ratio turbine blades has been measured in great detail at five planes within the cascade and two downstream in order to trace the generation of stagnation pressure loss in the passage. Five-hole probes were used in the main flow, but as it is important to resolve the boundary layers accurately, three-hole and single flattened probes were used near the endwall and blade surfaces, respectively. Endwall shear stresses have been measured using a hot-film probe and an oil-drop viscosity balance technique. Numerical predictions and simple aerodynamic models are used in conjunction with the experimental data to estimate the relative importance of different loss mechanisms, including boundary layer shear stresses and mixing processes.
Comparison of Steady and Unsteady Secondary Flows in a Turbine Stator Cascade
Authors: Hebert G. J. , Tiederman W. G.
Published: June 2008

The effect of periodic rotor wakes on the secondary flow structure in a turbine stator cascade was investigated. A mechanism simulated the wakes shed from rotor blades by passing cylindrical rods across the inlet to a linear cascade installed in a recirculating water flow loop. Velocity measurements showed a passage vortex, similar to that seen in steady flow, during the time associated with undisturbed fluid. However, as the rotor wake passed through the blade row, a large crossflow toward the suction surface was observed in the midspan region. This caused the development of two large areas of circulation between the midspan and endwall regions, significantly distorting and weakening the passage vortices.
Inlet Skew and the Growth of Secondary Losses and Vorticity in a Turbine Cascade
Authors: Walsh J. A. , Gregory-Smith D. G.
Published: June 2008

This paper presents results of an experimental investigation into the effects of inlet skew on the flow field of a large-scale axial flow turbine cascade. The results are presented in terms of the development of the streamwise vorticity since, in classical terms, the streamwise vorticity generates the transverse velocity components that cause the generation of the secondary losses. Inlet skew is shown to have a profound effect on the distribution and magnitude of the generated losses. A number of correlations for the secondary losses are compared with the measured values and it is shown that the correlations are not adequate for accurate loss prediction purposes.
Mach Number Effects on Secondary Flow Development Downstream of a Turbine Cascade
Authors: Perdichizzi A.
Published: June 2008

The results of an investigation of the three-dimensional flow downstream of a transonic turbine cascade are presented. The investigation was carried out for a wide range of Mach numbers, extending from M2is = 0.2 up to 1.55. Measurements were made in five planes at different axial locations downstream of the trailing edge (covering more than one chord length), by using a miniaturized five-hole probe especially designed for transonic flows. The results are presented in terms of local loss coefficient, vorticity, and secondary velocity plots; these plots give a detailed picture of the secondary flow development downstream of the cascade and show how flow compressibility influences the vortex configuration. As Mach number increases, the passage vortex is found to migrate toward the endwall and secondary flow effects are more confined in the endwall region. The pitchwise mass averaged loss and flow angle distributions along the blade height appear to be affected by the expansion ratio; at high Mach number both underturning and overturning angles are found to be smaller than in low velocity flows. Overall losses, vorticity, and secondary kinetic energy versus Mach number are also presented and discussed.
Three-Dimensional Separated Flow Field in the Endwall Region of an Annular Compressor Cascade in the Presence of Rotor-Stator Interaction: Part 1—Quasi-Steady Flow Field and Comparison With Steady-State Data
Authors: Schulz H. D. , Gallus H. E. , Lakshminarayana B.
Published: June 2008

An experimental study of three-dimensional flow field in an annular compressor cascade with an upstream rotor has been carried out at four different incidences to the stator blade. Blade boundary layers and the three-dimensional flow field at the exit are surveyed using a hot-wire sensor and a five-hole probe, respectively. The data on the blade boundary layer, passage flow, and separated corner flow are presented. The upstream rotor wake has a major influence on the transition, laminar separation bubble, extent of wall/corner flow separation, aerodynamic losses, secondary flow, and three-dimensional flow inside the passage. A detailed interpretation of the effects of upstream wakes on the entire passage flow is presented and compared with the data in the absence of a rotor.
Three-Dimensional Separated Flow Field in the Endwall Region of an Annular Compressor Cascade in the Presence of Rotor-Stator Interaction: Part 2—Unsteady Flow and Pressure Field
Authors: Schulz H. D. , Gallus H. E. , Lakshminarayana B.
Published: June 2008

An experimental study of the unsteady three-dimensional flow and pressure field in an annular compressor cascade with an upstream rotor has been carried out at several incidences to the stator blade. The distributions of the unsteady pressures at the blade surfaces are measured using fast response Kulite sensors. The unsteady blade boundary layers and the passage flow are measured with a hot-wire sensor. A detailed interpretation of the magnitude of unsteady pressures, phase angle differences, unsteady blade boundary layers, and wake transport through the stator passage is presented and analyzed in the paper. The unsteady pressures are found to be dominant near the blade leading edge. Substantially higher pressure fluctuations occur in this region as well as on the edge of the corner flow separation region.
Modeling Unsteady Transition and Its Effects on Profile Loss
Authors: Hodson H. P.
Published: June 2008

This paper considers the effects of wake interactions on the transition processes of turbomachine blade boundary layers. A simple model of unsteady transition is proposed, which is then used to identify a relationship between a new reduced frequency parameter and the profile loss of a blade row that is subjected to unsteady inflow. The value of this new parameter is also used to identify the nature of the boundary layer development on the blade surfaces. The influence of other parameters on the transition process is also discussed. The model is then extended to deal with the more general case. The validity of the models is demonstrated by a comparison with a correlation of the effects of wake-generated unsteadiness on profile loss that was originally proposed by Speidel. The effects of unsteady inflow on four idealized turbine blades are considered.
On the Mechanism of Dangerous Blade Vibration Due to Blade Flow Interactions on Centrifugal Compressors
Authors: Haupt U. , Jin D. F. , Rautenberg M.
Published: June 2008

Severe blade flow interactions at part-load operation conditions were investigated on a centrifugal compressor with a vaned diffuser leading to material stresses beyond the allowable values. By means of a number of measurement and analysis techniques it could be found, that a stationary periodic pressure field is produced on the circumference by the vibrating blade itself, which is induced at resonance conditions by the peripheral pressure nonuniformity due to the outlet tube. This peripheral pressure field of an integer wave number intensifies the blade resonance excitation from downstream leading to an additivity effect between wave amplitude and blade displacement. The significant role in this mechanism plays the reverse flow near the corner shroud/suction side in the impeller, occurring at part-load operation, which is controlled by the interaction of the tip angle of the vibrating blade and the flow angle at that location. It could be demonstrated that this dangerous blade vibration, in addition, is the source of a shift of the surge line toward higher mass flow, reducing the compressor operating range considerably in this operating zone.
Self-Excited Oscillation of Transonic Flow Around an Airfoil in Two-Dimensional Channels
Authors: Yamamoto K. , Tanida Y.
Published: June 2008

A self-excited oscillation of transonic flow in a simplified cascade model was investigated experimentally, theoretically and numerically. The measurements of the shock wave and wake motions, and the unsteady static pressure field predict a closed-loop mechanism, in which the pressure disturbance that is generated by the oscillation of boundary layer separation propagates upstream in the main flow and forces the shock wave to oscillate, and then the shock oscillation disturbs the boundary layer separation again. A one-dimensional analysis confirms that the self-excited oscillation occurs in the proposed mechanism. Finally, a numerical simulation of the Navier–Stokes equations reveals the unsteady flow structure of the reversed flow region around the trailing edge, which induces the large flow separation to bring about the antiphase oscillation.
Unsteady Aerodynamic Damping Measurement of Annular Turbine Cascade With High Deflection in Transonic Flow
Authors: Kobayashi H.
Published: June 2008

Unsteady aerodynamic forces acting on oscillating blades of a transonic annular turbine cascade were investigated in both aerodynamic stable and unstable domains, using a Freon gas annular cascade test facility. In the facility, whole blades composing the cascade were oscillated in the torsional mode by a high-speed mechanical drive system. In the experiment, the reduced frequency K was changed from 0.056 to 0.915 with a range of outlet Mach number M2 from 0.68 to 1.39, and at a constant interblade phase angle. Unsteady aerodynamic moments obtained by two measuring methods agreed well. Through the moment data the phenomenon of unstalled transonic cascade flutter was clarified as well as the significance of K and M2 for the flutter. The variation of flutter occurrence with outlet flow velocity in the experiments showed a very good agreement with theoretical analysis.
Optical Measurement of Unducted Fan Blade Deflections
Authors: Kurkov A. P.
Published: June 2008

A nonintrusive optical method for measuring unducted fan (or propeller) blade deflections is described and evaluated. The measurement does not depend on blade surface reflectivity. Deflection of a point at the leading edge and a point at the trailing edge in a plane nearly perpendicular to the pitch axis is obtained with a single light beam generated by a low-power, helium-neon laser. Quantitative analyses are performed from taped signals on a digital computer. Averaging techniques are employed to reduce random errors. Measured static deflections from a series of high-speed wind tunnel tests of a counterrotating unducted fan model are compared with available predicted deflections, which also are used to evaluate systematic errors.
The Aerodynamics of an Oscillating Cascade in a Compressible Flow Field
Authors: Buffum D. H. , Fleeter S.
Published: June 2008

Fundamental experiments are performed in the NASA Lewis Research Center Transonic Oscillating Cascade Facility to investigate and quantify the aerodynamics of a cascade of bioconvex airfoils executing torsion mode oscillations at realistic reduced frequency values. Both steady and unsteady airfoil surface pressures are measured at two inlet Mach numbers, 0.65 and 0.80. and two incidence angles, 0 and 7 deg, with the harmonic torsional airfoil cascade oscillations at realistic high reduced frequency and unsteady data obtained at several interbladephase angle values. The time-variant pressures are analyzed by means of discrete Fourier transform techniques, with these unique data compared with predictions from a linearized unsteady cascade model. The experimental results indicate that the interblade phase angle has a major effect on the chordwise distributions of the airfoil surface unsteady pressure, with the effects of reduced frequency, incidence angle, and Mach number somewhat less significant.
Annular Cascade Study of Low Back-Pressure Supersonic Fan Blade Flutter
Authors: Kobayashi H.
Published: June 2008

Low back-pressure supersonic fan blade flutter in the torsional mode was examined using a controlled-oscillating annular cascade test facility. Precise data of unsteady aerodynamic forces generated by shock wave movement, due to blade oscillation, and the previously measured data of chordwise distributions of unsteady aerodynamic forces acting on an oscillating blade, were joined and, then, the nature of cascade flutter was evaluated. These unsteady aerodynamic forces were measured by direct and indirect pressure measuring methods. Our experiments covered a range of reduced frequencies based on a semichord from 0.0375 to 0.547, six interblade phase angles, and inlet flow velocities from subsonic to supersonic flow. The occurrence of unstalled cascade flutter in relation to reduced frequency, interblade phase angle, and inlet flow velocity was clarified, including the role of unsteady aerodynamic blade surface forces on flutter. Reduced frequency of the flutter boundary increased greatly when the blade suction surface flow became transonic flow. Interblade phase angles that caused flutter were in the range from 40 to 160 deg for flow fields ranging from high subsonic to supersonic. Shock wave movement due to blade oscillation generated markedly large unsteady aerodynamic forces which stimulated blade oscillation.
Blade Vibration With Nonlinear Tip Constraint: Model Development
Authors: Wagner L. F. , Griffin J. H.
Published: June 2008

Turbine blades having integrally machined tip shrouds, with associated gaps between adjacent shrouds, often exhibit unusual vibratory responses with significant differences between the amplitudes and frequencies of individual blades on the same stage. These differences result from unavoidable variations in the shroud gaps causing, for large enough excitation, nonlinear constraint at the blade tips which varies from blade to blade. This study shows that the blade stresses cannot be adequately represented by the type of single-degree-of-freedom models that are often used for dynamic impact studies, but require the participation of higher frequency beam-type modes. The extension of the resulting beam model to multi-degree-of-freedom systems will allow the study of the “gap mistuning” phenomenon for practical bladed disks.
Effect of Weak Swirling Flow on Film Cooling Performance
Authors: Gau C. , Hwang W. B.
Published: June 2008

Experiments have been performed in a large circular pipe to study and obtain the film cooling effectivenesses with the presence of weak swirling flow in the mainstream. The swirling flow is generated by a flat vane swirler situated upstream. Cooling film is injected from an annular slot formed by the pipe wall and the circular cover plate. The radial temperature distribution measurements at several axial locations were used to infer the rate of mixing of film jet with swirling flow. The swirl number, which increases with turbulence intensity and swirl velocity in the mainstream, can significantly increase the mixing rate of film jet with swirl flow and decrease the film cooling effectiveness. During the course of the experiments, the blowing ratio ranged from 0.5 to 1.75 and the swirl number ranged from 0 to 0.6. Correlation equations for the film cooling effectiveness, which account for the effect of swirling flow, are obtained.
Predictions of Aerodynamic Performance of Wells Turbines From Aerofoil Data
Authors: Raghunathan S. , Setoguchi T. , Kaneko K.
Published: June 2008

Data from Wells turbine tests are correlated with two-dimensional aerofoil data. The correlations show that the performance of the Wells turbine can be predicted from wind-tunnel data of aerofoils.
Investigation on the Power Generation and Electrolysis Behavior of Ni-YSZ∕YSZ∕LSM Cell in Reformate Fuel
Authors: Fan Chuan-gang , Iida Tatsuya , Murakami Kota , Matsui Toshiaki , Kikuchi Ryuji , Eguchi Koichi
Published: June 2008

The bigger drop of open circuit voltage (OCV) for a degraded or damaged cell unit will make it consume the electricity generated by other cells of the solid oxide fuel cell stack to electrolyze the exhaust gas. In this article, the degradation behaviors of Ni-YSZ∕YSZ (0.5mmthick)∕La0.6Sr0.4MnO3−δ cell sample caused by exhaust of hydrocarbon fuel were studied by solid oxide electrolysis cell method. I‐V curve and impedance spectra measurements were carried out to investigate the polarization feature of the electrodes during power generation and electrolysis at 1000°C with gas mixtures of H2–CO2, CO–CO2, and H2–H2O at the Ni-YSZ working electrode (WE) as well as oxygen at La0.6Sr0.4MnO3−δ counterelectrode. In case of H2–CO2 couple, It was found that the increase of CO2 only depressed the OCV due to the increase of oxygen partial pressure (PO2) at the WE side, and the characteristics of the cell were more influenced by the H2O produced during power generation. In the case of CO–CO2, more CO produced during electrolysis would lead to the carbon deposition, which can result in the increase of polarization resistance by the reduction of the active site (triple phase boundary). In the case of H2–H2O couple, the polarization contribution was different for both power generation and the electrolysis. The former likely contributed from concentration polarization since its nature of producing H2O, and the latter mainly contributed to the active site polarization as compared with that from concentration polarization due to its nature of consuming H2O.
Reactivity of LaNi0.6Fe0.4O3 With Samaria Doped Ceria
Authors: Arai Hajime , Chiba Reiichi , Komatsu Takeshi , Orui Himeko , Sugita Satoshi , Tabata Yoshitaka , Nozawa Kazuhiko , Watanabe Kimitaka , Arakawa Masayasu , Sato Kazunori
Published: June 2008

La Ni 0.6 Fe 0.4 O 3 (LNF) is one of the promising cathodes for solid oxide fuel cells, but reacts with a zirconia-based electrolyte. To prevent this undesirable reaction, a ceria phase has been introduced in between the LNF cathode and electrolyte. On the other hand, the ceria phase itself could react with lanthanum-based perovskite oxides. We examined the reactivity of LNF and Ce0.8Sm0.2O2−δ (samaria doped ceria (SDC)) in this study. The mixtures of LNF and SDC were sintered at temperatures between 1123K and 1623K and the resultants were analyzed by X-ray diffraction together with the Rietveld analysis. We also measured the activity of electrochemical cells with a LNF-SDC composite layer in between the LNF cathode and zirconia-based electrolyte. The lattice parameters of each phase are clarified and a possible reaction scheme is proposed. The cell activity was high, but was influenced by the sintering temperature of the composite. Both chemical stability and physical property of the cathode can affect the cell activity.
Development of Microtubular SOFCs
Authors: Suzuki Toshio , Funahashi Yoshihiro , Yamaguchi Toshiaki , Fujishiro Yoshinobu , Awano Masanobu
Published: June 2008

Microtubular solid oxide fuel cells (SOFCs) have successfully demonstrated their advantages over conventional (planar) SOFCs, such as high thermal stability during rapid heat cycling and large electrode area per volume, which enables one to realize SOFC systems applicable to portable devices and auxiliary power units for automobile. In this study, the fabrication method of the microtubular SOFCs was examined. Shrinkage behavior and the microstructure of electrolyte/anode as a function of sintering temperature were shown and correlated with densification of the electrolyte during cosintering process.
Ni–Fe Alloy-Supported Intermediate Temperature SOFCs Using LaGaO3 Electrolyte Film for Quick Startup
Authors: Ishihara Tatsumi , Yan Jingwang , Enoki Makiko , Okada Sachio , Matsumoto Hiroshige
Published: June 2008

Intermediate temperature solid oxide fuel cells (SOFCs), which are highly tolerant against a thermal cycle, are studied by using the Ni–Fe porous alloy substrate prepared by an in situ reduction. It was found that Ni–Fe alloy exhibits high activity against anodic reaction and suitable compatibility with LaGaO3 electrolyte. The electrolyte film of La0.9Sr0.1Ga0.8Mg0.2O3 (LSGM) and Sm0.2Ce0.8O2 (SDC) bilayer with 5μm thickness was successfully prepared on the dense NiO–Fe2O3 composite anode. After a in situ reduction, the dense plate of NiO–Fe2O3 was changed to the porous Ni–Fe alloy substrate; however, the LSGM film can keep the dense state. The prepared Ni–Fe alloy that supported LSGM cell demonstrated the maximum power densities of 0.9W∕cm2 and 0.4W∕cm2 at 873K and 573K. After heating up to 873K within 540s, there is no crack formed on the film and almost the theoretical open circuit voltage was exhibited. In addition, the maximum power density of 400mW∕cm2 was achieved at 773K. After the thermal cycling, the decrease in the maximum power density was not large, and this suggests that the film is still gas tight and highly tolerant against the thermal cycle. Quick start characteristics of the metal support SOFC could expand the SOFC application to the electric source of a mobile-field-like automobile.
La
0.6

Sr
0.4

Co
0.2

Fe
0.8

O
3

–
Ag

Composite Cathode for Intermediate-Temperature Solid Oxide Fuel Cells
Authors: Sakitou Y. , Hirano A. , Imanishi N. , Takeda Y. , Liu Y. , Mori M.
Published: June 2008

The composite electrode of La0.6Sr0.4Co0.2Fe0.8O3 (LSCF) and Ag was examined as an air electrode for intermediate-temperature solid oxide fuel cells. Two types of LSCF/Ag composite electrodes were prepared. One was made by firing the printed sheet of LSCF-GDC (10mol%Gd2O3 doped CeO2)-Ag powder mixture on the GDC/8mol%Y2O3 doped zirconia (YSZ) electrolyte, and the other was prepared by firing the LSCF electrode on GDC/YSZ infiltrating a AgNO3 solution in the pores of the electrode. In both cases, the cathode polarization overpotentials were reduced by the addition of Ag into LSFC. The power density of anode support cells with the LSCF electrode was increased from 0.28W∕cm2to0.34W∕cm2 at 600°C by infiltrating a AgNO3 solution into the electrode.
Development of Evaluation Technologies for Microtubular SOFCs Under Pressurized Conditions
Authors: Hashimoto S. , Nishino H. , Liu Y. , Asano K. , Mori M. , Funahashi Y. , Fujishiro Y.
Published: June 2008

We developed evaluation technologies for microtubular solid oxide fuel cells under pressurized conditions. The pressurized cell evaluation system for the single cell was produced. The chamber temperature of the evaluation system can be controlled up to 750°C, and the maximum chamber pressure is 0.8MPa. It was possible to manually control the pressure difference between air and fuel gas within ±3kPa during the pressure increase. The hard sealing technique was introduced for the evaluation under pressurized conditions. Using two different types of commercial inorganic ceramic adhesives, the gas leakage was controlled at approximately 2%. Differential pressure control between fuel and air is effective for the stable open circuit voltage and power generation. The power generation under pressurized conditions was successful at 650°C, and the pressurized effect was clearly confirmed.
Development and Characterization of Cathode-Supported SOFCs by Single-Step Cofiring Fabrication for Intermediate Temperature Operation
Authors: Liu Yu , Hashimoto Shin-ichi , Takei Katsuhito , Mori Masashi , Funahashi Yoshihiro
Published: June 2008

In this study, a single-step cofiring through the extrusion molding and wet-ceramic coating technique was developed to fabricate a cathode-supported microtubular cell. The cell is consisting of a Ce0.9Gd0.1O1.95 (GDC) electrolyte with a NiO–GDC anode on a porous La0.6Sr0.4Co0.2Fe0.8O3−δ∕GDC tube (460μm wall thickness and 2.26mm diameter). Densification of the ceria membrane (thickness <20μm) was successful by cosintering the laminated thin electrolyte and the anode with the cathode at 1200°C. Compared with that fabricated by the conventional two-step cofiring process, the cell showed an improved performance due to the increased anode sintering temperature, which leads to an improved anode∕electrolyte interfacial property. The cell having 2cm tube length fed with humidified 30vol%H2–Ar (3% H2O) produced the maximum power densities of 0.09Wcm−2, 0.08Wcm−2, and 0.05Wcm−2, at 600°C, 550°C, and 500°C, respectively.
The Development of Thermally Stable Sealing Glass in the BaO–B2O3–SiO2 System for Planar SOFC Applications
Authors: Peng Lian , Zhu Qingshan
Published: June 2008

The purpose of the present paper is to study the influence of glass composition on the thermal stability in the SiO2–B2O3–BaO system, and three glasses were consequently investigated. Although Glass A has a coefficient of thermal expansion (CTE) that shows the best match with those of anode and electrolyte materials of solid oxide fuel cells (SOFCs), the thermal stability of Glass A is quite poor, where after being heat treated at 800°C for only 8h, the CTE of the glass increased more than 24%. The change of the CTE value was mainly attributed to the fast crystallization that formed high CTE value phases such as BaB2O4 and Ba2Si3O8. In order to improve the thermal stability, BaO in Glass A was replaced by B2O3 (Glass B) and SiO2 (Glass C). It was found that the decrease in the BaO content improved the thermal stability of the resultant glasses. Glass B showed less than 8% change of the CTE during annealing time at 800°C, while Glass C exhibited superior long-term thermal stability, where the change of the CTE was within the equipment detection limit after being heat treated for 300h at 800°C. The good thermal stability of Glass C was believed to be due to the formation of a more compact glass network after the substitution as compared with that of Glass A. The good thermal stability makes Glass C attractive to be used as the sealing material for SOFC applications.