bluuebunny/crossref_metadata_2025 huggingface dataset:
Improving Aerodynamic Matching of Axial Compressor Blading Using a 3D Multistage Inverse Design Method
Authors: van Rooij M. P. C. , Dang T. Q. , Larosiliere L. M.
Published: November 2008

Current turbomachinery design systems increasingly rely on multistage CFD as a means to diagnose designs and assess performance potential. However, design weaknesses attributed to improper stage matching are addressed using often ineffective strategies involving a costly iterative loop between blading modification, revision of design intent, and further evaluation of aerodynamic performance. A scheme is proposed herein which greatly simplifies the design point blade row matching process. It is based on a three-dimensional viscous inverse method that has been extended to allow blading analysis and design in a multi-blade row environment. For computational expediency, blade row coupling is achieved through an averaging-plane approximation. The proposed method allows improvement of design point blade row matching by direct regulation of the circulation capacity of the blading within a multistage environment. During the design calculation, blade shapes are adjusted to account for inflow and outflow conditions while producing a prescribed pressure loading. Thus, it is computationally ensured that the intended pressure-loading distribution is consistent with the derived blading geometry operating in a multi-blade row environment that accounts for certain blade row interactions. The viability of the method is explored in design exercises on a 2.5-stage, highly loaded compressor.
Development of Global Optimization Method for Design of Turbine Stages
Authors: Kawagishi Hiroyuki , Kudo Kazuhiko
Published: November 2008

A new optimization method which can search for the global optimum solution and decrease the number of iterations was developed. The performance of the new method was found to be effective in finding the optimum solution for single- and multi-peaked functions for which the global optimum solution was known in advance. According to the application of the method to the optimum design of turbine stages, it was shown that the method can search the global optimum solution at approximately one seventh of the iterations of GA (Genetic Algorithm) or SA (Simulated Annealing).
Numerical Simulation of Gas Flow and Heat Transfer in Cross-Wavy Primary Surface Channel for Microtubine Recuperators
Authors: Liang H. X. , Wang Q. W. , Luo L. Q. , Feng Z. P.
Published: November 2008

Three-dimensional numerical simulation was conducted to investigate the flow field and heat transfer performance of the Cross-Wavy Primary Surface (CWPS) recuperators for microturbines. Using high-effective compact recuperators to achieve high thermal efficiency is one of the key techniques in the development of microturbine in recent years. Recuperators need to have minimum volume and weight, high reliability and durability. Most important of all, they need to have high thermal-effectiveness and low pressure-losses so that the gas turbine system can achieve high thermal performances. These requirements have attracted some research efforts in designing and implementing low-cost and compact recuperators for gas turbine engines recently. One of the promising techniques to achieve this goal is the so-called primary surface channels with small hydraulic dimensions. In this paper, we conducted a three-dimensional numerical study of flow and heat transfer for the Cross-Wavy Primary Surface (CWPS) channels with two different geometries. In the CWPS configurations the secondary flow is created by means of curved and interrupted surfaces, which may disturb the thermal boundary layers and thus improve the thermal performances of the channels. To facilitate comparison, we chose the identical hydraulic diameters for the above four CWPS channels. Since our experiments on real recuperators showed that the Reynolds number ranges from 150 to 500 under the operating conditions, we implemented all the simulations under laminar flow situations. By analyzing the correlations of Nusselt numbers and friction factors vs. Reynolds numbers of the four CWPS channels, we found that the CWPS channels have superior and comprehensive thermal performance with high compactness, i.e., high heat transfer area to volume ratio, indicating excellent commercialized application in the compact recuperators.
Experimental Study of Boundary Layer Control Through Tip Injection on Straight and Swept Compressor Blades
Authors: Roy Bhaskar , Chouhan Manish , Kaudinya Kota Venkata
Published: November 2008

Recent research on the use of three dimensional blade designs incorporating sweep and casing boundary layer control through tip injection have shown improved efficiency, stage loading and stable operating range Both methods are known to stabilize the flow at the tip trailing edge (Stall Inception Point) of the compressor blade under lower than design mass flow conditions and improve the compressor performance. This study, performed on a low speed axial flow fan, is aimed to improve the understanding and to assess the improved stall margin due to tip injection on swept and straight blades. Injection through flush mounted nozzles on the casing before the rotor blades is of practical importance because of the non-interference of the nozzles with the main flow. Benefits in terms of efficiency and operating range have already been established and swept blades were found to perform better than the baseline straight blades. The injection mechanism involves 12 symmetrically located flush mounted nozzles before the rotor blades in the casing. The number of injection nozzles used is an additional parameter of the study. Although there are manufacturing limitations in the air injection angle, two injection angles (10° and 30°) have been tried. The improvement in stall margin is more pronounced in the straight blades than in the swept blades and the 10° injection angle with 6 active injection holes gives the best performance in comparison to other combinations.
Assessment of Laminar-Turbulent Transition Models for Turbomachinery Flow Computations
Authors: Cutrone L. , De Palma P. , Pascazio G. , Napolitano M.
Published: November 2008

This paper provides a thorough comparison of different laminar-to-turbulent bypass transition models. The models are based on combinations of two transition-onset correlations and three intermittency factor models. They have been embedded in a Reynolds averaged Navier–Stokes solver employing a low-Reynolds number k–ω turbulence model. The performance of the transition models have been validated by computing three well documented incompressible flows over a flat plate, namely, test T3A, T3B, and T3C2 of ERCOFTAC SIG 10, with different free-stream conditions, the latter being characterized by non-zero pressure gradient. Finally, a more complex test case, namely the two-dimensional compressible flow through a linear turbine cascade, has been considered, for which detailed experimental data are available in the literature.
Analysis of Some Improved Blade Tip Concepts
Authors: Prakash Chander , Lee Ching-Pang , Cherry David , Wadia Aspi , Doughty Roger
Published: November 2008

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.
Studies of Wake-Induced Bypass Transition of a Flat-Plate Boundary Layer: Comparisons Between Two Transition Modes Induced by Small Sphere Wake and Thin Wire Wake
Authors: Funazaki Ken-ichi , Otsuki Takahiro
Published: November 2008

This study aims at deepening the understanding of wake-induced bypass transition process of a flat-plate boundary layer using two types of wake generating objects, which are small spheres and thin wires. Main focus is on emergence of isolated turbulent spots from the influence of the wake passage over the boundary layer. Precursors of the wake-induced turbulent spot, which have not been observed in an explicit manner in any other previous studies, are also of concern in this study. It is expected that wakes from the wires are so weak that an isolated turbulent spot may be induced by the wire wake, while the position of the spot emergence varies randomly along the wire. A multi-channel sensor with 7 hot-wire probes acquires the velocity data of the flow over the flat plate subjected to the wake passage. These velocity data reveal the spot shape and spot generation rate. In addition, the existence of Klebanoff mode in this wake-affected boundary layer is examined.
Criteria for Spike Initiated Rotating Stall
Authors: Vo Huu Duc , Tan Choon S. , Greitzer Edward M.
Published: November 2008

A computational study to define the phenomena that lead to the onset of short length-scale (spike) rotating stall disturbances has been carried out. Unsteady simulations show 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 length scale circumferential extent of spike disturbances. The scenario developed is consistent with numerical simulations as well as with 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.
Intermittency Based Unsteady Boundary Layer Transition Modeling: Implementation Into Navier-Stokes Equations
Authors: Schobeiri M. T.
Published: November 2008

This paper presents recent advances in boundary layer research that deal with an intermittency based unsteady boundary layer transition model and its implementation into the Reynolds averaged Navier-Stokes equations (RANS). RANS equations are conditioned to include the ensemble averaged unsteady intermittency function. The unsteady boundary layer transition model is based on a universal unsteady intermittency function developed earlier. It accounts for the effects of periodic unsteady wake flow on the boundary layer transition. The transition model is the result of an inductive approach analyzing the unsteady data obtained by experiments on a curved plate at zero-streamwise pressure gradient under periodic unsteady wake flow. To validate this model, systematic experimental investigations were conducted on the suction and pressure surfaces of turbine blades that were integrated into a turbine cascade test facility, which was designed for unsteady boundary layer investigations. This model is implemented into the above mentioned conditioned RANS-equations and calculation results are presented.
Aerodynamic Optimization of Industrial Fan Blade Cascades
Authors: Lotfi O. , Teixeira J. A. , Ivey P. C. , Sheard G. , Kinghorn I. R.
Published: November 2008

The paper describes the application of a Genetic Algorithm (GA) to the aerodynamic shape optimization of a low speed fan cascade. This task is accomplished by modifying the blade camber line while keeping the same blade thickness distribution and mass flow rate. A number of design examples have been studied and the behaviour of the genetic algorithm has been tested. The fitness value of the objective function is evaluated using the commercial turbomachinery CFD code CFX-TASCflow. In this work Bezier curves were used for the description of the blade camber line. Specific interfaces were developed in order to link the optimization code, the grid generator TASCgrid utilized to define the computational meshes and the commercial flow solver TASCflow employed within an automated design loop. The obtained results show that the genetic algorithm is capable of automatically improving the design of axial fan blade profiles.
Computation of Flow and Heat Transfer in Rotating Rectangular Channels (AR=4:1) With Pin-Fins by a Reynolds Stress Turbulence Model
Authors: Su Guoguang , Chen Hamn-Ching , Han Je-Chin
Published: November 2008

Computations with multi-block chimera grids were performed to study the three-dimensional turbulent flow and heat transfer in a rotating rectangular channel with staggered arrays of pin-fins. The channel aspect ratio (AR) is 4:1, the pin length to diameter ratio (H/D) is 2.0, and the pin spacing to diameter ratio is 2.0 in both the stream-wise (S1/D) and span-wise (S2/D) directions. A total of six calculations have been performed with various combinations of rotation number, Reynolds number, and coolant-to-wall density ratio. The rotation number and inlet coolant-to-wall density ratio varied from 0.0 to 0.28 and from 0.122 to 0.20, respectively, while the Reynolds number varied from 10,000 to 100,000. For the rotating cases, the rectangular channel was oriented at 150 deg with respect to the plane of rotation to be consistent with the configuration of the gas turbine blade. A Reynolds-Averaged Navier-Stokes (RANS) method was employed in conjunction with a near-wall second-moment turbulence closure for detailed predictions of mean velocity, mean temperature, and heat transfer coefficient distributions.
LES of the Flow and Acoustics Generated by an Aircraft Fan Running in the Vicinity of the Ground
Authors: Moroianu Dragos¸ , Karlsson Arne , Fuchs Laszlo
Published: November 2008

An important component of the aircraft generated noise, especially ahead of it, is the fan noise created by the high velocity variation near the blades and the interaction of the rotating fan with the fluid. In order to predict the sound, the method used involves the acoustical analogy developed by Ffowcs Williams and Hawkings. Computation of the flow field is performed in the incompressible LES framework, while the noise is evaluated using a non-homogeneous wave equation. In this work the influence of the angle between the fan and ground, on the total sound spectrum, the noise generation and the noise propagation is investigated. It is found that the near field is dominated by the blade passage frequency and an upward inclination of the fan will produce a slightly different sound pressure level than a horizontal or downward inclination.
Experimental Evaluation of the Flow-Field in a State of the Art Linear Cascade With Boundary-Layer Suction
Authors: Hja¨rne Johan , Chernoray Valery , Larsson Jonas , Lo¨fdahl Lennart
Published: November 2008

This paper presents a detailed experimental evaluation of the flow field in a state of the art linear cascade with boundary-layer suction. The linear cascade was designed using a new design-process as described by Hja¨rne et al. [1]. From the measurements presented in this paper it can be concluded that all of the features of the test-rig work as expected. Hence the measurements validate the design process presented by Hja¨rne et al. in [1]. The intention with the test facility is to provide high quality benchmark cases for the flow field around aggressive designs of Low Pressure Turbine/Outlet Guide Vanes (LPT/OGV’s) to be used for validation of numerical codes. The flow quality is assessed by measuring inlet uniformity, boundary-layer heights, periodicity of the static pressure distribution around the OGV’s and uniformity of the outlet flow.
Film Cooling Effectiveness From a Single Scaled-Up Fan-Shaped Hole: A CFD Simulation of Adiabatic and Conjugate Heat Transfer Models
Authors: Silieti Mahmood , Kassab Alain J. , Divo Eduardo
Published: November 2008

This paper documents a computational investigation of the film cooling effectiveness of a 3-D gas turbine endwall with one fan-shaped cooling hole. The simulations were performed for adiabatic and conjugate heat transfer models. Turbulence closure was investigated using three different turbulence models; the realizable k-ε model, the SST k-ω model, as well as the v2–f turbulence model. Results were obtained for a blowing ratio of one, and a coolant-to-mainflow temperature ratio of 0.54. The simulations used a dense, high quality, O-type, hexahedral grid with three different schemes of meshing for the cooling hole: hexahedral-, hybrid-, and tetrahedral-topology grid. The computed flow/temperature fields are presented, in addition to local, two-dimensional distribution of film cooling effectiveness for the adiabatic and conjugate cases. Results are compared to experimental data in terms of centerline film cooling effectiveness downstream cooling-hole, the predictions with realizable k-ε turbulence model exhibited the best agreement especially in the region for (2 ≤ x/D ≤ 6). Also, the results show the effect of the conjugate heat transfer on the temperature (effectiveness) field in the film cooling hole region and, thus, the additional heating up of the cooling jet itself.
The Effect of Work Processes on the Casing Heat Transfer of a Transonic Turbine
Authors: Thorpe Steven J. , Miller Robert J. , Yoshino Shin , Ainsworth Roger W. , Harvey Neil W.
Published: November 2008

This paper considers the effect of the rotor tip on the casing heat load of a transonic axial flow turbine. The aim of the research is to understand the dominant causes of casing heat-transfer. Experimental measurements were conducted at engine-representative Mach number, Reynolds number and stage inlet to casing wall temperature ratio. Time-resolved heat-transfer coefficient and gas recovery temperature on the casing were measured using an array of heat-transfer gauges. Time-resolved static pressure on the casing wall was measured using Kulite pressure transducers. Time-resolved numerical simulations were undertaken to aid understanding of the mechanism responsible for casing heat load. The results show that between 35% and 60% axial chord the rotor tip-leakage flow is responsible for more than 50% of casing heat transfer. The effects of both gas recovery temperature and heat transfer coefficient were investigated separately and it is shown that an increased stagnation temperature in the rotor tip gap dominates casing heat-transfer. In the tip gap the stagnation temperature is shown to rise above that found at stage inlet (combustor exit) by as much as 35% of stage total temperature drop. The rise in stagnation temperature is caused by an isentropic work input to the tip-leakage fluid by the rotor. The size of this mechanism is investigated by computationally tracking fluid path-lines through the rotor tip gap to understand the unsteady work processes that occur.
Effect of Clocking on the Heat Transfer Distribution of a Second Stator Tested in a One and a Half Stage HP Turbine
Authors: Billiard N. , Paniagua G. , De´nos R.
Published: November 2008

This paper reports detailed heat transfer measurements on the second stator of a one and a half stage HP turbine. The tests are performed under engine representative conditions. Four vane-to-vane clocking positions are investigated. The second stator inlet and outlet flowfield is monitored. The heat transfer measurements are performed with thin-film gauges around the blade profile at 15%, 50% and 85% span as well as on the hub and casing endwalls. On a time-averaged point of view, the inlet conditions of the second stator vary noticeably as a function of clocking. This variation is due to a change of relative position of the total pressure pitchwise distribution, linked to the first stator, with respect to the pitchwise variation of the static pressure, linked to the second stator. On the hub platform, the location of a zone of high heat transfer, probably due to high velocity levels, moves in conjunction with the clocking. When this zone is in the center of the passage, the time-averaged thermal load around the airfoil profile is significantly lower. On a time-resolved point of view, the rotor exit Mach number undergoes periodically short excursions in the transonic regime. The passing of these shock/potential distortion events can be clearly seen on the second stator profile. Thanks to a model of isentropic compression, it is demonstrated that the heat flux variations originate in static pressure fluctuations that are causing variations in the gas temperature. The clocking changes not only the mean levels of the heat transfer but also the intensity and the trajectory of the fluctuations.
Modeling and Evaluation of Centrifugal Compressor Performance Variations Using Probabilistic Analysis
Authors: Tecza Joseph , Menegay Peter , Koch Jay
Published: November 2008

This paper presents a methodology for analyzing the variation in compressor stage performance due to component dimensional variations for a range of flows, speeds, gas compositions and scale factors. Due to the large number of input parameters involved, a Design of Experiments (DOE) approach was used to develop key variables, and to develop response surface models of head coefficient and efficiency in terms of these key variables. These response surface-based performance models then are used for a probabilistic analysis of head coefficient and efficiency as functions of dimensional variations, for a range of compressor sizes. The variations in dimensions are expressed as probability distributions and evaluated using a Monte-Carlo integration technique. The techniques for developing the response surfaces and performing the probabilistic analysis are described, as are methods for evaluating both the effects of dimensional variation on performance and for evaluating how much dimensional variation can be tolerated before the variation exceeds established limits.
Edge Matching: Part I — Theory and Implementation
Authors: Ji L. C. , Chen J.
Published: November 2008

Theory and implementation of Edge Matching (EM) are presented in this paper. The EM concerns the spatial relation between a blade trailing edge and the leading edge of the sequent blade row. It reflects spanwise distribution of phase angles of certain performance parameter variation versus unsteady inflow. While EM theory is described first. Two methods of implementing EM, namely spanwise integration (SI) and full optimization (FO), are described. After theoretical analysis, the interaction between linear turbine cascade and upstream wake is used to illustrate EM numerically. More specific and detailed numerical investigations are presented in accompanying paper ([1]). All these suggest that EM provides a new degree of freedom (DOF) for turbomachinery design with potentials of improving aerodynamic, aeroelastic, aeroacoustic and other performances by matching phase angles of unsteady flow along the whole span.
Edge Matching: Part II — Numerical Investigations
Authors: Chen J. , Ji L. C.
Published: November 2008

Edge Matching (EM), presented in accompanying paper [1], may make it manipulable to utilize time-accurate unsteady flow simulation in daily design of multistage turbomachinery. The spatial relation between a blade trailing edge and the leading edge of the sequent blade row is taken into account in improving turbomachinery performance. This paper presents numerical investigations to further illustrate and validate the EM theory and its implementations. A single stage transonic axial compressor and a single stage steam turbine are modified with EM. Numerical results show that EM might provide an effective degree of freedom (DOF) for unsteady design of turbomachinery.
Analytical Investigation of a Low Pressure Turbine With and Without Flowpath Endwall Gaps, Seals and Clearance Features
Authors: Cherry David , Wadia Aspi , Beacock Rob , Subramanian Mani , Vitt Paul
Published: November 2008

Numerical simulations for low pressure turbine (LPT) stages of a high bypass turbofan engine are presented and discussed in this study. A smooth flowpath configuration and a flowpath configuration with endwall features consistent with the actual engine geometry were considered for the numerical analysis to demonstrate the significance of including hub and tip flowpath details for proper performance prediction and design improvement studies. Fully three-dimensional, multistage, mutiblock, viscous flow analysis methodology was applied for first three stages of a moderately loaded LPT to predict aerodynamic performance of individual components, stage and for the overall turbine. Numerical results were obtained first for the smooth endwall configuration that ignores flowpath cavities, gaps and leaks in the numerical model. Following the smooth endwall calculations, a second set of calculations was performed with hub and tip flowpath details to closely represent actual engine geometry and experimental rig hardware. The approach of using smooth endwall contours for multi stage, multi blade row computational analysis is quite common for modeling simplicity. However, as the flow features are expected to be more complex in high pressure ratio, highly loaded turbine stages of next generation aircraft engines, it is imperative that flowpath and endwall geometry details such as gaps, seals, leakage and clearance effects are included in the numerical simulation for improved component design and stage performance prediction. This study addresses this particular issue by including endwall details and quantifies performance differences between the two modeling approaches. An O-H mesh topology was utilized for the blades, wheel space cavities, labyrinth seals and clearances for better flowfield resolution and numerical accuracy. Component performance, secondary flow details of endwall cavities, seal leakage and loss features of each blade row, for individual stage and for the overall turbine stage is presented and discussed for the two sets of calculations. Computed results are compared with experimental data obtained with high speed rig testing for verification and for understanding of the flow physics.
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
Authors: Sewall Evan A. , Tafti Danesh K.
Published: November 2008

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.
An Investigation of the Effects of Film Cooling in a High-Pressure Aeroengine Turbine Stage
Authors: Chana Kam S. , Hilditch Mary A. , Anderson James
Published: November 2008

Cooling is required to enable the turbine components to survive and have acceptable life in the very high gas temperatures occurring in modern engines. The cooling air is bled from the compression system, with typically about 15% of the core flow being diverted in military engines and about 20% in civil turbofans. Cooling benefits engine specific thrust and efficiency by allowing higher cycle temperatures to be employed, but the bleed air imposes cycle penalties and also reduces the aerodynamic efficiency of the turbine blading, typically by 2–4%. Cooling research aims to develop and validate improved design methodologies that give maximum cooling effectiveness for minimum cooling flow. This paper documents external cooling research undertaken in the Isentropic Light Piston Facility at QinetiQ as part of a European collaborative programme on turbine aerodynamics and heat transfer. In Phase I, neither the ngv nor the rotor was cooled; cooling was added to the ngv only for Phase II, and to the rotor and ngv in Phase III. Coolant blowing rates and density ratios were also varied in the experiments. This paper describes the ILPF and summarises the results of this systematic programme, paying particular attention to the variation in aerofoil heat transfer with changing coolant conditions, and the effects coolant ejection has on the aerofoil’s aerodynamic performance.
Performance and Flow Characteristics of an Optimized Supercritical Compressor Stator Cascade
Authors: Song Bo , Ng Wing F.
Published: November 2008

An experimental and numerical study was performed on an optimized compressor stator cascade designed to operate efficiently at high inlet Mach numbers (M1) ranging from 0.83 to 0.93 (higher supercritical flow conditions). Linear cascade tests confirmed that low losses and high turning were achieved at normal supercritical flow conditions (0.7 < M1 < 0.8), as well as higher supercritical flow conditions (0.83 < M1 < 0.93), both at design and off-design incidences. The performance of this optimized stator cascade is better than those reported in the literature based on Double Circular Arc (DCA) and Controlled Diffusion Airfoil (CDA) blades, where losses increase rapidly for M1 > 0.83. A 2-D Navier-Stokes solver was applied to the cascade to characterize the performance and flow behavior. Good agreement was obtained between the CFD and the experiment. Experimental loss characteristics, blade surface Mach numbers, shadowgraphs, along with CFD flowfield simulations, were presented to elucidate the flow physics. It is found that low losses are due to the well-controlled boundary layer, which is attributed to an optimum flow structure associated with the blade profile. The multi-shock pattern and the advantageous pressure gradient distribution on the blade are the key reasons of keeping the boundary layer from separating, which in turn accounts for the low losses at the higher supercritical flow conditions.
A Simple Inverse Cascade Design Method
Authors: Phillipsen Bent
Published: November 2008

The design of profile sections is a key task in the axial turbomachinery design process. Both direct and inverse methods are in use for this purpose. Though it is relatively easy to fulfill mechanical and geometrical constraints with a direct method, where the complete geometry is specified, it can be difficult to obtain the desired distribution of velocity along the profile with this method. On the other hand, it can be difficult to get an acceptable geometry with an inverse method, where the velocity is prescribed and the geometry calculated. A simple inverse cascade design method, which combines the advantages of both the direct and the traditional inverse design method, is described. The theory behind this method directly shows how the prescribed loading distribution and the cascade design data, such as flow angles and blade thickness, influence the cascade geometry. Various practical aspects of the turbomachinery cascade design task and their relation to the described method are shown. The usefulness of the method is demonstrated through the recovery of the geometry of various cascades published in the literature. The method is valid for subsonic and shock free transonic turbine and compressor cascade design, but good results can also be obtained for turbine cascades with weak shocks, i.e. when shocks are not the dominating feature of the design.
Investigation of Vortex Shedding and Wake-Wake Interaction in a Transonic Turbine Stage Using Laser-Doppler-Velocimetry and Particle-Image-Velocimetry
Authors: Go¨ttlich E. , Woisetschla¨ger J. , Pieringer P. , Hampel B. , Heitmeir F.
Published: November 2008

The current paper presents a time-resolved experimental flow investigation in a highly loaded transonic gas turbine stage operating continuously under engine representative conditions. The measurement was performed with a two-component Laser-Doppler-Velocimeter (LDV) and a three-component stereoscopic Particle-Image-Velocimeter (3C-PIV). Unsteady velocity data were obtained in axis perpendicular planes (LDV) and tangential planes (3C-PIV) between stator and rotor as well as downstream of the rotor. The results of the time-resolved investigation at several radii show the vortex shedding process from the trailing edges of nozzle guide vanes and rotor blades. This vortex shedding was found to be phase locked to higher harmonics of the blade passing frequency. Pressure waves evoked by reflection of the trailing edge shocks of the vanes on the passing rotor blades interact with the boundary layers on the rear suction side of the vanes and on the rotor blade surfaces while running upstream and downstream the flow. They are responsible for this phase-locking phenomenon of the shedding vortices. At midspan, the vortices shedding from stator and rotor blades were also observed by PIV. The in-plane vorticity distribution was used to discuss the wake-wake interaction indicating that wake segments from the nozzle guide vanes were chopped by the rotor blades. These chopped segments are still visible in the distributions as a pair of counter rotating vortices. The nozzle wake segments are transported through the rotor passages by the flow, influencing the vortex street of the rotor blades as they pass by with the higher velocity of the main flow. A comparison with a numerical simulation is also given.
Influence of Blade Leading Edge Geometry on Turbine Endwall Heat(Mass) Transfer
Authors: Han S. , Goldstein R. J.
Published: November 2008

The secondary flows, including passage and other vortices in a turbine cascade cause significant aerodynamic losses and thermal gradients. Leading-edge modification of the blade has drawn considerable attention as it has been shown to reduce the secondary flows. However, the heat transfer performance of a leading-edge modified blade has not been investigated thoroughly. Since a fillet at the leading edge blade is reported to reduce the aerodynamic loss significantly, the naphthalene sublimation technique with a fillet geometry is used to study local heat (mass) transfer performance in a simulated turbine cascade. The present paper compares Sherwood number distributions on an endwall with a simple blade and a similar blade having modified leading-edge by adding a fillet. With the modified blades, a horseshoe vortex is not observed and the passage vortex is delayed or not observed for different turbulence intensities. However, near the blade trailing edge the passage vortex has gained as much strength as with the simple blade for low turbulence intensity. Near the leading edge on the pressure and the suction surface, higher mass transfer regions are observed with the fillets. Apparently the corner vortices are intensified with the leading-edge modified blade.
Cooling Performance Limits of a Turbine Blade With Potassium Evaporative Cooling
Authors: Townsend Jessica , Kerrebrock Jack , Stickler David
Published: November 2008

A new method of turbine blade cooling, the Return Flow Cascade, has been developed in which vaporization of a liquid metal such as potassium is used to maintain the blade surface at a nearly uniform temperature. Turbine blades cooled using this technology have lower blade temperature levels compared to that available with conventional air cooling, potentially resulting in higher firing temperatures or a choice of a wider range of materials for the hot gas path. The detailed operation of the Return Flow Cascade is described including fluid mechanics and heat transfer phenomena that occur at high heat flux and radial acceleration levels characteristic of modem gas turbine engines. The performance limits of the Return Flow Cascade are identified by the development of a theoretical model that estimates the performance of the system for a range of operating conditions found in the experimental test rig and in an actual gas turbine engine. These limits include vapor choking in the internal blade passages, pool boiling limits in the blade, and surface tension restriction of liquid flow. Cascade initiation limits predicted by the internal vapor choking model are in good agreement with experimental results from testing performed at the Massachusetts Institute of Technology.
Effect of Tip and Pressure Side Coolant Injection on Heat Transfer Distributions for a Plane and Recessed Tip
Authors: Nasir Hasan , Ekkad Srinath V. , Bunker Ronald S.
Published: November 2008

The present study investigates the effects of coolant injection on adiabatic film effectiveness and heat transfer coefficients from a plane and recessed tip of a HPT first stage rotor blade. Three cases where coolant is injected from (a) five orthogonal holes located along the camber line, (b) seven angled holes located near the blade tip along the pressure side and (c) combination cases when coolant is injected from both tip and pressure side holes were studied. The pressure ratio (inlet total pressure to exit static pressure for the cascade) across the blade row was 1.2, and the experiments were run in a blow-down test rig with a four-blade linear cascade. The Reynolds number based on cascade exit velocity and axial chord length was 8.61×105 and the inlet and exit Mach number were 0.16 and 0.55, respectively. A transient infrared (IR) technique was used to measure adiabatic film effectiveness and heat transfer coefficient simultaneously for three blowing ratios of 1.0, 2.0, and 3.0. For all the cases, gap-to-blade span ratio of 1% was used. The depth-to-blade span ratio of 0.0416 was used for the recessed tip. Pressure measurements on the shroud were also taken to characterize the leakage flow and understand the heat transfer distributions. For tip injection, when blowing ratio increases from 1.0 to 2.0, film effectiveness increases for both plane and recessed tip. At blowing ratio 3.0, lift off is observed for both cases. In case of pressure side coolant injection and for plane tip, lift off is observed at blowing ratio 2.0 and reattachments of jets are observed at blowing ratio 3.0. But, almost no effectiveness is observed for squealer tip at all blowing ratios with pressure side injection. For combination case, very high effectiveness is observed at blowing ratio 3.0 for both plane and recessed blade tip. It appears that for this high blowing ratio, coolant jets from the tip hit the shroud first and then reattach back on to the blade tip. For tip injection, as blowing ratio increases heat transfer coefficient decreases for both plane and recessed tip. In case of pressure side coolant injection and for plane tip, film injection reduced heat transfer coefficient along the pressure side. Minimal effect is observed for recessed tip at all blowing ratios. For combination case, very high heat transfer coefficient is observed at blowing ratio 3.0 for both plane and recessed blade tip. It appears that for this high blowing ratio, coolant jets from the tip hit the shroud first and then reattach back on to the blade tip.
Intermittent Behavior of the Separated Boundary Layer Along the Suction Surface of a Low Pressure Turbine Blade Under Periodic Unsteady Flow Conditions
Authors: O¨ztu¨rk B. , Schobeiri M. T. , Ashpis David E.
Published: November 2008

The paper experimentally and theoretically studies the effects of periodic unsteady wake flow and aerodynamic characteristics on boundary layer development, separation and re-attachment along the suction surface of a low pressure turbine blade. The experiments were carried out at Reynolds number of 110,000 (based on suction surface length and exit velocity). For one steady and two different unsteady inlet flow conditions with the corresponding passing frequencies, intermittency behavior were experimentally and theoretically investigated. The current investigation attempts to extend the intermittency unsteady boundary layer transition model developed in previously to the LPT cases, where separation occurs on the suction surface at a low Reynolds number. The results of the unsteady boundary layer measurements and the intermittency analysis were presented in the ensemble-averaged, and contour plot forms. The analysis of the boundary layer experimental data with the flow separation, confirms the universal character of the relative intermittency function which is described by a Gausssian function.
Numerical Study of Embedded Supersonic Compressor Stages
Authors: Kempf Severin G. , Guillot Stephen , Ng Wing Fai , Wellborn Steven R. , Chriss Randall M.
Published: November 2008

A numerical case study of a multistage, highly-loaded, relative supersonic compressor is presented. The purpose of the investigation was to highlight the changing shock structure while throttling the compressor and to give insight into possible compressor instabilities. The computational fluid dynamic (CFD) study was conducted with the NASA code ADPAC, utilizing the mixing-plane assumption for the boundary condition between adjacent, relatively-rotating blade rows. A steady, five-blade-row, numerical simulation using the Baldwin-Lomax turbulence model was performed, creating several constant speed lines. The results show that the shock structure in the downstream rotor isolates the upstream rotor from the exit conditions until the shock detaches from the leading edge. The shock structure in the upstream rotor then moves, changing the conditions for the downstream rotor. This continues as the compressor is throttled until the shock in the upstream rotor detaches from the leading edge. CFD indicates that this event causes a rapid drop in the mass flow rate, creating a mismatch between stage-one and stage-two that results in compressor instability.
An Empirical Prediction Method for Secondary Losses in Turbines: Part I — A New Loss Breakdown Scheme and Penetration Depth Correlation
Authors: Benner M. W. , Sjolander S. A. , Moustapha S. H.
Published: November 2008

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.
Characteristics of Pressure, Flow, and Heat Transfer in Rotating Internal Cooling Passages With Attached-Detached Ribs
Authors: Liou Tong-Miin , Hwang Yu-Sian , Chen Meng-Yu
Published: November 2008

Measurements of flow field, heat transfer coefficient, and wall static pressure distribution were made in a rotating two-pass square duct with detachment of 90° ribs from the first-pass leading wall and second-pass trailing wall as well as attachment of 90° ribs onto the other two opposite walls. Laser-Doppler velocimetry was used to measure the local flow velocity. The ribs were square in cross-section and their detached-distance/height ratio was 0.38. The rib-height/duct-height ratio and the pitch/rib-height ratio were 0.136 and 10, respectively. Duct Reynolds number was fixed at 1×104 and rotating number ranged from 0 to 0.2. Results are documented in terms of the main flow development, cross-stream secondary flow structure, the distributions of the pressure coefficient, the variation of friction factor with Ro, and passage averaged Nusselt number ratios under a constant flow rate and a constant pumping power condition. For CFD reference, the periodic fully developed rotating flow condition is attained after the 6th rib pair in the first pass. In addition, the relationships between the regional averaged Nusselt number, transverse and convective mean velocity component, and turbulent kinetic energy are addressed. Using these relationships the general superiority of heat transfer enhancement of the attached-detached 90° ribs arrangement over the attached-attached one and detached-detached one can be reasonably illustrated. Simple expressions are obtained to correlate the friction factor with Ro, which are lacking in the published literature. The respective contributions of the ribs and passage rotation on the passage friction loss are identified.
Separation-Bubble-Transition Measurements on a Low-Re Airfoil Using Particle Image Velocimetry
Authors: McAuliffe B. R. , Yaras M. I.
Published: November 2008

This paper presents experimental results on separation-bubble transition at low Reynolds number and low freestream turbulence, measured on an airfoil using particle image velocimetry (PIV). The two-dimensional PIV measurements have been performed over the suction surface of a low-Reynolds-number airfoil in a water tow-tank facility. Reynolds numbers, based on airfoil chord length and towing speed, of 40,000 and 65,000 have been examined at various angles of incidence, providing a range of streamwise pressure distributions and transitional separation-bubble geometries. The types of bubbles observed range from a short and thick bubble with separation near the leading edge of the airfoil, to a long and thin bubble with separation far downstream of the suction peak. The PIV measurements facilitate visualization of the vortex dynamics associated with separation-bubble transition. The growth of instability waves within the separated shear layer and eventual breakdown into turbulence is documented through the instantaneous vector fields. For all cases examined, large-scale vortex shedding and multiple reverse-flow zones are observed in the reattachment region. A technique for estimating the location of transition onset based on statistical turbulence quantities is presented, and comparisons are made to existing transition models.
Numerical Validation of Characteristic and Linearized Unsteady Boundary Conditions for Non-Integer Blade Ratios in a Non-Linear Navier Stokes Solver
Authors: Mokulys Thomas , Dewhurst Stephen C. , Abhari Reza S.
Published: November 2008

Unsteady single passage methods for turbomachinery applications are becoming a more and more mature method in predicting unsteady blade-to-blade aerodynamics in a more efficient and accurate way compared to multipassage computations with scaled airfoil geometries. The outlined work presents results of a fourier-based method for phase-lagged boundary conditions using a usual conservative approach at the interfaceplane and compares it to a new characteristic treatment as well as a linearized single row approach. The new method provides results of similar accuracy with improved stability of the algorithm and faster convergence times.