bluuebunny/crossref_metadata_2025 huggingface dataset:
An Analytical and Experimental Assessment of a Diffuser Flow Phenomenon as a Precursor to Stall
Authors: Sorokes James M. , Pacheco Jorge E. , Vezier Clementine , Fakhri Syed
Published: July 2013

The paper describes an experimental and analytical study on the vaneless diffuser downstream of a high flow coefficient, high inlet relative Mach number centrifugal impeller. The diffuser flowfield exhibited a unique behavior in which the low momentum zone typically found along the shroud side of a centrifugal compressor diffuser suddenly shifted to the hub side of the diffuser just prior to the onset of diffuser rotating stall. The phenomenon was observed in the computational fluid dynamics (CFD) analyses conducted as well as in the experimental data obtained during stage validation testing. A review of the analytical and test results suggested that the phenomenon was at least partially attributable to the level of diffusion in the vaneless diffuser. Modifications made to reduce the diffusion rate were shown by CFD analysis to move the onset of the unusual shift of low momentum flow to a much lower flow rate. The modifications also increased the efficiency of the overall stage by 1.2%.
Constrained Adjoint-Based Aerodynamic Shape Optimization of a Single-Stage Transonic Compressor
Authors: Walther Benjamin , Nadarajah Siva
Published: July 2013

This paper develops a discrete adjoint formulation for the constrained aerodynamic shape optimization in a multistage turbomachinery environment. The adjoint approach for viscous, internal flow problems and the corresponding adjoint boundary conditions are discussed. To allow for a concurrent rotor/stator optimization a non-reflective adjoint mixing-plane formulation is proposed. A sequential-quadratic programming algorithm is utilized to determine an improved airfoil shape based on the objective function gradient provided by the adjoint solution. The functionality of the proposed optimization method is demonstrated by the redesign of a midspan section of a single-stage transonic compressor. The objective is to maximize the isentropic efficiency while constraining the mass flow rate and the total pressure ratio.
Turbulence Levels are High at the Combustor-Turbine Interface
Authors: Cha Chong M. , Ireland Peter T. , Denman Paul A. , Savarianandam Vivek
Published: July 2013

Turbulence measurements are made in a novel gas turbine rig facility recently used to study combustor-turbine interactions in jet engines [1]. The rig is capable of numerous area traverses surrounding engine turbine nozzle guide vanes (NGVs). The rig is unique in that complete engine hardware of the annular combustion subsystem is used to simulate the upstream flow entering the turbine. The rig operates at cold, near-atmospheric conditions. The turbulence measurements include both the turbulence intensities and lengthscales and span an area over a single combustor sector. Axial measurement planes include locations both upstream and downstream of the real engine hardware NGVs. The upstream plane corresponds to a conventional combustor-turbine interface plane. In [1], pressure, velocity, and passive scalar mixing measurements were presented along with RANS CFD predictions. Here, in addition to the newly measured turbulence quantities, large-eddy simulations (LES) are performed for the complete, coupled combustor-turbine system.

Good agreement between rig data and CFD is seen at the combustor-turbine interface, with LES yielding improved predictions over RANS. For the flow through the NGV passages, vortex visualizations of the simulated flowfields show significant differences to the classic, commonly accepted picture of Langston [2] and others [3]. The difference is attributed to the high turbulence levels created by the combustor. The impact of the limitations of the combustor-turbine rig on these findings is discussed.
Characteristics of the Tip Leakage Vortex in a Low-Speed Axial Compressor With Different Rotor Tip Gaps
Authors: Zhang Zhibo , Yu Xianjun , Liu Baojie
Published: July 2013

The detailed evolutionary processes of the tip leakage flow/vortex inside the rotor passage are still not very clear for the difficulties of investigating of them by both experimental and numerical methods. In this paper, the flow fields near the rotor tip region inside the blade passage with two tip gaps, 0.5% and 1.5% blade height respectively, were measured by using stereoscopic particle image velocimetry (SPIV) in a large-scale low speed axial compressor test facility. The measurements are conducted at four different operating conditions, including the design, middle, maximum static pressure rise and near stall conditions. In order to analyze the variations of the characteristics of the tip leakage vortex (TLV), the trajectory, concentration, size, streamwise velocity, and the blockage parameters are extracted from the ensemble-averaged results and compared at different compressor operating conditions and tip gaps. The results show that the formation of the TLV is delayed with large tip clearance, however, its trajectory moves much faster in an approximately linear way from the blade suction side to pressure side. In the tested compressor, the size of the tip gap has little effects on the scale of the TLV in the spanwise direction, on the contrary, its effects on the pitch-wise direction is very prominent. Breakdown of the TLV were both found at the near-stall condition with different tip gaps. The location of the initiation of the TLV breakdown moves downstream from the 60% chord to 70% chord as the tip gap increases. After the TLV breakdown occurs, the flow blockage near the rotor tip region increases abruptly. The peak value of the blockage effects caused by the TLV breakdown is doubled with the tip gap size increasing from 0.5% to 1.5% blade span.
Impeller-Diffuser Interaction in Centrifugal Compressors
Authors: Robinson Chris , Casey Michael , Hutchinson Brad , Steed Robin
Published: July 2013

This paper reports several CFD analyses of a centrifugal compressor stage with a vaned diffuser at high pressure ratio using different techniques to model the rotor-stator interaction. A conventional steady stage calculation with a mixing-plane type interface between the rotor and stator was used as a baseline. This simulation gave excellent agreement with the measured performance characteristics at design speed, demonstrating the ability of the particular steady simulation used to capture the essential features of the blockage interaction between the components.

A full annulus simulation using a transient rotor-stator interaction (TRS) method was then used at the peak efficiency point to obtain a fully unsteady reference solution, and this predicted a small increase in peak efficiency. Finally, a computationally less expensive unsteady calculation using a Time Transformation (TT) method was carried out. This gave similar results to the fully transient calculation suggesting that this is an acceptable approach to estimate unsteady blade loading from the interaction.

The impeller diffuser spacing was then reduced from 15 to 7% of the impeller tip radius using the more affordable TT approach. This identified an increase in efficiency of 1% and predicted unsteady pressure fluctuations in the impeller which were 116% higher with the closely spaced diffuser.
Interactions Between the Combustor Swirl and the High Pressure Stator of a Turbine
Authors: Giller Lucas , Schiffer Heinz-Peter
Published: July 2013

The interaction between the strongly swirling combustor outflow and the high pressure turbine nozzle guide vanes were investigated at the cascade test rig at Technische Universität Darmstadt. The test section of the rig consists of six swirl generators and five cascade vanes. The three middle vanes are equipped with film cooling holes at the leading edges. The swirler nozzles are aligned with the center of the cascade passages.

The operating settings are defined by the swirl number, the distance between the swirler nozzles and the vanes, the blowing ratio and the radial angle of the film cooling holes.

Flow field measurements using PIV downstream of the swirlers and five hole probe measurements at the inlet and outlet plane of the cascade were accomplished. Measurements using the ammonia diazo technique to determine the adiabatic film cooling effectiveness on the surface of the center cascade vane were also carried out.

It is shown that a swirling inflow leads to a strong alteration of the flow field and the losses in the passages in comparison to an axial inflow. Furthermore, the impact of the swirl on the formation of the cooling film and it’s adiabatic film cooling effectiveness is presented.
A Numerical Investigation Into the Sources of Endwall Loss in Axial Flow Turbines
Authors: Denton John , Pullan Graham
Published: July 2013

Endwall loss, often termed “secondary loss”, in axial turbines has been intensively studied for many years, despite this the physical origin of much of the loss is not really understood. This lack of understanding is a serious impediment to our ability to predict the loss and to the development of methods for reducing it. This paper aims to study the origins of the loss by interrogating the results from detailed and validated CFD calculations. The calculation method is first validated by comparing its predictions to detailed measurements in a turbine cascade. Very good agreement between the calculations and the measurements is obtained. The solution is then examined in detail to highlight the sources of entropy generation in the cascade, several different sources of loss are found to be significant.

The same blade row is then used to study the effects of the of the inlet boundary layer thickness on the loss. It is found that only the inlet boundary layer loss and the mixing loss vary greatly with inlet boundary layer thickness. Finally a complete 50% reaction stage, with identical stator and rotor blade profiles, is examined using both steady calculations, with a mixing plane model, and the time average of unsteady calculations. It is found that the endwall flow in the rotor is completely different from that in the stator. Because of this it is considered that results from endwall flow and loss measurements in cascades are of limited relevance to the endwall flow in a real turbine. The results are also used to discuss the validity of the mixing plane model.
An Explanation for Flow Features of Spike-Type Stall Inception in an Axial Compressor Rotor
Authors: Yamada Kazutoyo , Kikuta Hiroaki , Iwakiri Ken-ichiro , Furukawa Masato , Gunjishima Satoshi
Published: July 2013

The unsteady behavior and three-dimensional flow structure of spike-type stall inception in an axial compressor rotor have been investigated by experimental and numerical analyses. Previous studies have revealed that the test compressor falls into a mild stall after emergence of a spike, in which multiple stall cells, each consisting of a tornado-like vortex, are rotating. However, the flow mechanism from the spike onset to the mild stall remains unexplained. The purpose of this study is to describe the flow mechanism of a spike stall inception in a compressor.

In order to capture the transient phenomena of spike-type stall inception experimentally, an instantaneous casing pressure field measurement technique was developed, in which 30 pressure transducers measure an instantaneous casing pressure distribution inside the passage for one blade pitch at a rate of 25 samplings per blade passing period. This technique was applied to obtain the unsteady and transient pressure fields on the casing wall during the inception process of the spike stall. In addition, the details of the three-dimensional flow structure at the spike stall inception have been analyzed by a numerical approach using the detached-eddy simulation (DES).

The instantaneous casing pressure field measurement results at the stall inception show that a low-pressure region starts traveling near the leading edge in the circumferential direction just after the spiky wave was detected in the casing wall pressure trace measured near the rotor leading edge. The DES results reveal the vortical flow structure behind the low-pressure region on the casing wall at the stall inception, showing that the low-pressure region is caused by a tornado-like separation vortex resulting from a leading-edge separation near the rotor tip. A leading-edge separation occurs near the tip at the onset of the spike stall and grows to form the tornado-like vortex connecting the blade suction surface and the casing wall. The casing-side leg of the tornado-like vortex generating the low-pressure region circumferentially moves around the leading-edge line. When the vortex grows large enough to interact with the leading edge of the next blade, the leading-edge separation begins to propagate, and then, the compressor falls into a stall with decreasing performance.
Characteristics of Tip-Leakage Flow at High Stagger-Angle Setting for Rotor Blade in an Axial Flow Fan
Authors: Nishioka Takahiro , Joko Masayoshi
Published: July 2013

Rotor-tip flow fields at high stagger-angle setting were investigated to clarify the loss generation mechanism in a high specific-speed axial-flow fan. The tip clearance flow in the cases of large and small clearances, which are 2.0% and 1.0% of the rotor tip chord length respectively, are experimentally and numerically evaluated at the maximum efficiency point and the operating limit. At the maximum efficiency point, the tip leakage vortex reached to the rotor exit in both cases of large and small tip clearances. However, the leakage vortex in the case of large tip-clearance passed closer to the pressure side of the adjacent blade than that in the case of small one. Moreover, in the case of large tip clearance, the tip leakage vortex generated the large total pressure loss in the blade passage, and the interaction between the tip leakage vortex and the wake also generated the large total pressure loss at the rotor exit. Therefore, the maximum efficiency of the rotor and the fan was lower than that in the case of small tip clearance. At the operating limit, the tip-leakage vortex extended inside the blade passage and reached to the front part of the pressure side of the next blade in the case of small tip-clearance. Moreover, the double leakage flow occurred in the case of small tip clearance. In contrast, the leakage vortex reached to the leading edge of the next blade, and the spillage of the tip leakage flow from the leading edge occurred in the case of large tip clearance. The spillage of the tip leakage flow induced the larger total pressure loss than that induced by the double leakage flow. Therefore, the pressure rise in the case of large tip clearance is lower than that in the case of small tip clearance at the operating limit. It was concluded from the experimental and numerical results at the high stagger-angle setting for rotor blade that the loss generation mechanism depended on the behavior of tip-leakage vortex and that this behavior also depended on the tip-clearance.
Aerodynamic Optimisation of High Pressure Turbines for Lean-Burn Combustion System
Authors: Shahpar Shahrokh , Caloni Stefano
Published: July 2013

Modern lean-burn combustors make use of high flow swirl to maintain flame stability. The swirling flow can persist downstream of the turbine first vane, changing the loading on the rotor, leading to a reduction in efficiency. This paper presents the results of an automatic optimisation study carried out to mitigate the effect of high swirling flow on a high pressure turbine stage. A high-fidelity CFD-based design optimisation using a multi-point approximation (response surface) method is carried-out to produce a new vane and a new rotor configuration with a significantly improved aerodynamic performance. It is demonstrated that the novel optimisation methodology can cope well with a number of near equality constraints needed for a practical design.
Linear Stability Analysis of Turbo Heat Pumps
Authors: Park Jung Chan , Kim Kil Young , Jeong Jinhee , Song Seung Jin
Published: July 2013

A typical turbo heat pump system consists of a centrifugal compressor, expansion valve, and two heat exchangers or plenums — a condenser and evaporator. Compared to a gas turbine, a turbo heat pump introduces additional complexities because it is a two-phase, closed-loop system with heat exchange using a real gas/liquid (refrigerant) as working fluid. For such systems, a new linear stability model has been developed by applying a linearized small perturbation method to the nonlinear turbo heat pump surge model. When the new linear model is applied to a compressor in an open loop system (e.g. turbocharger), results identical to those of Greitzer’s model are obtained. When applied to an operating turbo heat pump, the new model accurately predicts the surge onset point. Finally, results from a parametric study on the influence of the B parameter on surge onset are presented. Thus, the new model can be applied to predict stability of various compression systems — either open or closed-loop systems running single or two-phase working fluids with or without heat exchange.
Unsteady Rotor Hub Passage Vortex Behavior in the Presence of Purge Flow in an Axial Low Pressure Turbine
Authors: Jenny P. , Abhari R. S. , Rose M. G. , Brettschneider M. , Engel K. , Gier J.
Published: July 2013

The paper presents an experimental and computational study of the unsteady behavior of the rotor hub passage vortex in an axial low-pressure turbine. Different flow structures are identified as having an effect on the size, strength, shape, position and the unsteady behavior of the rotor hub passage vortex. The aim of the presented study is to analyze and quantify the sensitivities of the different flow structures and to investigate their combined effects on the rotor hub passage vortex. Particular attention is paid to the effect of the rim seal purge flow and of the unsteady blade row interaction. The rotor under investigation has non-axisymmetric end walls on both hub and shroud and is tested at three different rim seal purge flow injection rates. The rotor has separated pressure sides at the operating point under investigation. The non-dimensional parameters of the tested turbine match real engine conditions. The 2-sensor Fast Response Aerodynamic Probe (FRAP) technique and the Fast Response Entropy Probe (FENT) systems developed by ETH Zurich are used in this experimental campaign. Time-resolved measurements of the unsteady pressure, temperature and entropy fields between the rotor and stator blade rows are taken and analyzed. Furthermore, the results of URANS simulations are compared to the measurements and the computations are also used to detail the flow field. The experimental results show a 30% increase of the maximum unsteadiness and a 4% increase of the loss in the hub passage vortex per percent of injected rim seal cooling flow. Compared to a free stream particle, the rim seal purge flow was found to do 60% less work on the rotor.
Investigation of 3D Internal Flow Using New Flux-Reconstruction High Order Method
Authors: Lu Yi , Yuan Xin , Dawes W. N.
Published: July 2013

The new godunov-type high order method via flux-reconstruction, are capable of unifiying several popular methods including the discontinuous Galerkin, staggered-grid multi-domain method, or the spectral difference/spectral volume methods into a single family. Compared to discontinuous galerkin method, this new method retain the arbitrary order and compact properties but written in differential form and without mass matrix. The flux-reconstrcution method using general unstructured 2D/3D mesh including triangles, quadrilaterals, tetrahedrons, pyramids, prisms and hexahedrons are implemented. Compared to discontinuous galerkin method, FR is differential form without any quandature, so it runs faster and is easier to implement on GPU to achieve higher speed up ratio. The present paper investigates three different solving methods for this new method in order to cut down the hugh computational cost and memory requirement for fully implicit methods. Explicit multi-stage Runge-Kutta, nonlinear Lower-Upper Symmetry Gauss-Seidel (LU-SGS) and generalized minimal residual (GMRES) with matrix-free preconditioning methods are implemented and all these three methods use p-multigrid to smooth low-frequency errors on lower order. The results of 2D steady-state external/internal viscous flows are presented and the convergence properties for different solving method are compared. Efficiency and robustness are improved through symmetric Gauss-Seidel (SGS) iterations as preconditioning, a remarkable feature of the present GMRES+SGS method is that the storage of the full Jacobian matrix can be eliminated and only its diagonal stored. A laplacian artificial viscosity is tested to capture shock in one element and performs well for this new high order method, an L2 projection method is used to improve the perfomance of aliasing error in conservative form. The one equation Spalart-Allmaras turbulence model is unsed to solve the Raynolds average navier-stokes equations. MPI is used for parallel simulations for ralatively complex 3D internal flow when using high order method. Compared to other high order methods, results in this paper show that Flux-Reconstruction methods with efficient solving method performs well for 2D/3D general unstructured mesh in wide range of reynorlds number, and is also potential for “real geometry” simulations.
Effects of Non-Uniform Combustor Exit Flow on Turbine Aerodynamics
Authors: Pyliouras Stavros , Schiffer Heinz-Peter , Janke Erik , Willer Lars
Published: July 2013

Very-low NOx combustion concepts require a high swirl number of the flow in the combustion chamber to allow for lean burn combustion. This article deals with the influence of the resulting combustor exit swirl on the turbine aerodynamics of the first stage. This investigation is based on numerical simulations. According to the literature research additional insight into combustor-turbine interaction is achieved by taking into account a fully two dimensional inlet boundary condition. Up to now published results on combustor-turbine interaction were mostly restricted to the inhomogeneous temperature distribution at the turbine inlet. The investigations are carried out on a real engine geometry — the E3E Core 3/2 — a research project of Rolls-Royce Deutschland on lean combustion. Calculations are conducted by means of the Rolls-Royce plc code Hydra. The swirled inlet boundary condition is further scaled to test rig conditions to check for the transferability between the test rig and the real engine geometry. The results show a significant impact of the inhomogeneous turbine inflow on the stage efficiency and the thermal load. The optimization potential due to the clocking position of the combustor swirl is analyzed. The impact on the secondary flow mechanisms is analyzed with a novel visualization technique. A frequency spectrum analysis is carried out to investigate the effects of the 2D inlet boundary condition on the rotor row.
Numerical Research on the Flow Field and Performance of a Ram-Rotor and a Scrampressor
Authors: Yang Ling , Zhong Jingjun , Han Ji-ang
Published: July 2013

The design methods of typical supersonic aircraft intakes and shock wave compression technology have been applied to ram-rotor, a new attractive compression system. A ram-rotor is a typical structure including the compression ramp, the throat and the subsonic diffuser; a scrampressor is similar to ram-rotor, the only different is that scrampressor has no subsonic diffuser. Base on the preparatory work, it has been found that these two structures have different advantages respectively. So, in this paper, the three dimensional Reynolds-averaged Navier-Stokes equations and the Spalart-Allmaras turbulent model are used to simulate numerically the flow field of the ram-rotor and the scrampressor at the design and at the off-design conditions. The back pressure and rotational speed are mainly considered which may affect the flow field and the total performance. It has been found that back pressure can not have influence on the flow field before the throat outlet obviously. With increasing of the back pressure, the position of the flow separation zone and shock train move forwards to the inlet. The rotational speed changes the shock wave structure of the ram-rotor and scrampressor evidently. With the rotational speed increasing, each shock wave moves to the outlet and the shock wave number decreases. The ram-rotor and scrampressor structure is similar, except the ram-rotor flow structure has a large flow separation zone after the throat outlet. The compression capability of the ram-rotor is higher than that of the scrampressor. The total performance of the scrampressor is better than the ram-rotor.
Advanced Numerical Methods for the Prediction of Tonal Noise in Turbomachinery: Part II—Time-Linearized Methods
Authors: Frey Christian , Ashcroft Graham , Kersken Hans-Peter , Weckmüller Christian
Published: July 2013

This is the second part of a series of two papers on unsteady CFD methods for the numerical simulation of aerodynamic noise generation and propagation. It focuses on the application of linearized RANS methods to turbomachinery noise problems. The convective and viscous fluxes of an existing URANS solver are linearized and the resulting unsteady linear equations are transfered into the frequency domain, thereby simplifying the solution problem from unsteady time-integration to a complex linear system. The linear system is solved using a parallel, preconditioned GMRES method with restarts. In order to prescribe disturbances due to rotor stator interaction a so-called gust boundary condition is implemented. Using this inhomogeneous boundary condition one can compute the generation of the acoustic modes and their near field progagation. The application of the time-linearized methods to a modern high-bypass ratio fan is investigated. The tonal fan noise predicted by the time-linearized solver is compared to numerical results presented in the first part and to measurements.
Influence of Working Fluid Composition on the Performance of Turbomachinery in Semi-Closed Gas Turbine Cycles
Authors: Canepa Roberto , Satta Antonio
Published: July 2013

This paper focuses on the relationship between the working fluid composition and the turbomachinery performance, notably on the case of flue gas recirculation. This aspect has been investigated by means of a semi-empirical radial equilibrium code for axial flow turbomachinery analysis. The mathematical model of the code, already described in previous works by the authors, is based on the assumption of steady, adiabatic and axisymmetric flow and, in this paper, has been extended to take into account the gas composition. The performance of an 1.5 stages axial-flow compressor and of a 4 stages turbine has been investigated under various working fluid compositions as expected from cycles with flue gas recirculation and the obtained results, which have been found to be consistent with those of a commercially available CFD code, have been compared with those which could be obtained with a traditional cycle working fluid composition. This analysis has allowed to prove how this modification of the working fluid composition through the turbomachinery leads to very small performance alterations.
Instability Characteristic of a Single-Stage Centrifugal Compressor Exposed to Dry and Wet Gas
Authors: Grüner Trond G. , Bakken Lars E.
Published: July 2013

An experimental investigation was conducted to determine the instability and performance characteristics of a single-stage centrifugal compressor exposed to wet gas. The compressor was tested at different rotational speeds and low gas mass fractions (GMFs) in an open-loop test facility with ambient air and water. The stage consisted of a shrouded impeller with a vaneless diffuser surrounded by a symmetrical circular volute with increasing cross-sectional area. Liquid was uniformly injected into the impeller eye through multiple nozzles mounted in the inlet pipe.

High-response dynamic pressure sensors flush-mounted in the diffuser were used to identify instability inception and evolution. Changes in the instability and pressure ratio characteristics at different GMFs and rotational speeds were revealed. Analysis with reference to dry gas was performed. Visual observation of the wet gas surge process at the inlet is described.

Results and analysis obtained from the experimental investigation on wet gas instability are presented. The investigation contributed to knowledge concerning the impact of wet gas on the instability and performance characteristics.
URANS Analysis of Wake-Induced Effects in High-Lift, Low Reynolds Number Cascade Flows
Authors: Pacciani Roberto , Marconcini Michele , Arnone Andrea , Bertini Francesco
Published: July 2013

A URANS analysis of unsteady effects induced by incoming wakes in high-lift, low-Reynolds-number cascade flows has been carried out using a novel, transition-sensitive, turbulence model. It is based on the coupling of two additional transport equations, one for the so-called laminar kinetic energy (LKE) and one for a turbulence indicator function, with a low Reynolds number formulation of the Wilcox k-ω model. Two high-lift cascades (T106C and T2), recently tested at the von Kármán Institute in the framework of the two European research projects UTAT (Unsteady Transition in Axial Turbomachines) and TATMo (Turbulence and Transition Modelling for Special Turbomachinery Applications), were considered for the present study. The analyzed Reynolds number values span the whole range typically encountered in aero-engines low-pressure turbines operations, and both steady and periodically unsteady inflow conditions were considered. A detailed comparison between measurements and computations, in terms of blade surface isentropic Mach number distributions and cascade lapse rates will be presented and discussed. Results with the proposed model show its ability to predict the major effects of passing wakes on the boundary layer development and loss characteristics of high-lift cascades operating in LP-turbine conditions.
Active Flow Control Techniques on a Stator Compressor Cascade: A Comparison Between Synthetic Jet and Plasma Actuators
Authors: De Giorgi Maria Grazia , Traficante Stefania , De Luca Carla , Bello Daniela , Ficarella Antonio
Published: July 2013

In this work a CFD analysis is applied to study the suppression of the boundary layer separation into a highly-loaded subsonic compressor stator cascade, by different active flow control techniques. Active flow control techniques have the potential to delay separation and to increase the pressure ratio. In particular three different techniques have been applied: the actuation by steady jet, by zero net mass flux Synthetic Jet (SJA) and by plasma actuator.

Several works have investigated the use of synthetic jet and plasma actuators on the airfoil, but only few studies have compared the effect of these devices.

Concerning the synthetic jet actuator, a suction/blowing type boundary condition is used, imposing a prescribed sinusoidal velocity depending on velocity amplitude, jet frequency and jet angle of ejection with respect to the wall.

Concerning the plasma actuation, the effect is modeled into numerical flow solvers by adding the paraelectric force that represents the plasma force into the momentum equation. The plasma, generated by Dielectric Barrier Discharge, acts as a momentum source to the boundary layer allowing it to remain attached throughout a larger portion of the airfoil. The time-averaged body force component, acting on the fluid, depends on the frequency and on the applied voltage, the charge density, the electrical field and the dimensional properties of the actuator, like width of the electrodes and gap between the electrodes. Using this numerical model, the effect of plasma actuators to suppress the flow separation over the blade has been investigated, increasing the turbo-machinery performance too.

Finally, the comparison between the different actuation devices shows that, reducing the secondary flow structures, each actuation technique beneficially affects the performance of the stator compressor cascade, even if in the steady jet the costs are relevant.
Experimental and Numerical Investigation of Tip Leakage Flow in a 1 1/2 Stage Turbine Rig Comparing Flat and Cavity-Squealer Tip Geometries
Authors: Kegalj Martin , Schmid Gregor , Wartzek Fabian , Schiffer Heinz-Peter , Willer Lars
Published: July 2013

Tip leakage loss is one of the main loss mechanisms in shroudless high-pressure turbines. Efficiency reducing phenomena are the reduction of mass turned by the blade, and the mixing of the tip leakage and passage flow. Applying a cavity squealer tip is a promising technique to reduce the tip leakage loss. The study investigates the influence of a cavity squealer on the flow field of a 3D high-pressure turbine.

The 1 1/2 stage turbine test rig is equipped with steady and unsteady measurement techniques, like five-hole probe, 2-component hot wire, oil & dye visualization, unsteady wall pressure probes and particle-image-velocimetry measurements. To improve the flow phenomena interpretation, unsteady numerical simulations were carried out using a commercial solver.
Experimental and Analytical Surge Cycle Analysis of a High Pressure Aero Engine Compressor
Authors: Waniczek Phillip , Schoenenborn Harald , Jeschke Peter
Published: July 2013

The unsteady flow field during surge of the front rotor of an eight-stage axial aero engine compressor has been investigated experimentally and analytically. For that purpose, two newly designed multi-sensor probes are installed up- and downstream of the first rotor. Surge experiments are conducted at four different speed lines (75–93% speed) covering a wide range of the compressor map and measurements have been taken at two different channel heights (50% and 70% span). The results show that the flow field varies extremely during surge up- and downstream of the rotor. In contrast to the flow at the rotor leading edge, which is nearly independent of the rotor speed, the flow at the rotor trailing edge is highly dependent of the rotor speed. Therefore, the performance of the rotor during surge is dependent on the reverse through-flow of the stators. At low speeds the flow passes the stators without any changes in the flow direction. If speed is increased the reverse flow is guided more and more by the stators. These different flow conditions have a direct impact on the process of energy conversion of the rotor during the surge event. The incoming reverse flow at the rotor trailing edge impinges on the blade from the suction surface side at lower speeds and turns to the pressure surface side when speed is increased. Hence, the deviation and specific work grow.

In addition to the surge experiments simulations of the surge events are conducted with a 1D code called SYSQ3D. The simulations and experiments match well and underline the capability of the new multi-sensor probes to accurately measure the flow patterns during surge.
Nozzle Guide Vane Cooling System Optimization
Authors: Girardeau Julian , Pardo Frederic , Pailhès Jérôme , Nadeau Jean-Pierre
Published: July 2013

The authors would like to address improvements on cooling system optimization within a turboshaft Nozzle Guide Vane (NGV). Designing high performance cooling systems able to preserve the life duration of the NGV can lead to significant aerodynamic losses. Theses losses jeopardize the performance of the whole engine. In the same time, a low efficiency cooling system may affect engine Maintenance Repair and Overhaul (MRO) costs as component life decreases. To help turbine designers, the authors studied a vane and searched for an optimal cooling design by means of an evolutionary algorithm. The associated objective function is based on satisfaction indexes, using Harrington’s desirability curves and Antonsson’s aggregation functions. Evaluation and optimization methods will be presented as well as optimized designs.
Calibration of Gas Turbine Blade Temperature Predictions Using Surrogate Models
Authors: McFarland John M. , Musgrove Grant O. , Chang Sung Yong , Ransom David L.
Published: July 2013

Actual gas turbine performance and component life at specific engine installations is highly dependent on the actual operating conditions, since not all engines are operated in the same manner. Due to the variability in turbine operation, it may be prudent to evaluate the operation of hot section components for turbine inlet conditions that are specific to a single installation. However, determining the actual turbine inlet conditions can be a difficult and expensive process that is usually only done on test bed gas turbines. This paper presents a method to determine turbine inlet conditions using a model calibration approach. Two-stage CFD and thermal analyses are developed to predict blade temperature. By varying the model inputs, computational predictions of blade temperature are calibrated to blade interdiffusion zone thickness measurements. In order to speed up calculations, surrogate models are used in place of the full-scale analysis codes during the calibration analysis. The result of the study is a prediction of the turbine inlet profile necessary to obtain the best agreement between predicted and measured blade temperatures.
Gas Turbine Blade Stress and Temperature Sensitivity to Turbine Inlet Profile and Cooling Flow
Authors: Kim Beom Soo , Kim Bum Shin , Choi Woosung , Musgrove Grant O. , McFarland John , Fierro Frank , Ransom David L.
Published: July 2013

Actual operating conditions in the hot section of a gas turbine vary from the design condition due to factors such as geographic location, component wear, and fuel composition. Turbine design practices typically use a conservative approach that requires checking the sensitivity of operating parameters such as turbine inlet profiles, cooling flows, and heat transfer correlations on component temperatures and stresses. In most cases, a sensitivity check is limited to analyzing the bounds of a range of values for only a few input parameters, whereby the inputs that produce the most conservative results are carried through the remainder of the analysis. For flow path components, however, multiple inputs must be evaluated over a range of values due to the interaction of the hot gas flow field and internal cooling systems. The study presented in this paper uses a probabilistic approach to develop surrogate models to evaluate the sensitivity of a set of operating parameters on the predicted blade temperatures and stresses. Commercially available software is utilized to predict blade temperatures and stresses for the first two stages of an industrial gas turbine. The operating parameters define the blade cooling flow and the shape and values of the turbine inlet profiles of total temperature and total pressure. The results of the study show the spatially resolved sensitivity of the operating parameters on blade temperature and stress distributions.
The Effect of Acoustic Excitation on Boundary Layer Separation of a Highly Loaded LPT Blade
Authors: Bernardini Chiara , Benton Stuart , Bons Jeffrey P.
Published: July 2013

An experimental investigation of the effect of acoustic excitation on the boundary layer development of a highly loaded low-pressure turbine blade at low-Reynolds number is investigated. The aim of this work is to study the effect of excitation at select frequencies on separation which could give indications about active flow control exploitation. The front-loaded L2F blade is tested in a low-speed linear cascade. The uncontrolled flow presents a separation bubble on the suction surface at Reynolds numbers below 40,000. For these conditions, the instability of the shear layer is documented using hot-wire anemometry. A loudspeaker upstream of the cascade is directed towards the passage inlet section. A parametric study on the effect of amplitude and frequency is carried out. The effect of the excitation frequency is observed to delay separation for a range of frequencies. However, the control authority of sound is found to be most effective at the fundamental frequency of the shear layer. The amplitude of perturbation is significant in the outcome of control until a threshold value is reached. PIV measurements allow a deeper understanding of the mechanisms leading to the reduction of separation. Data has been acquired with a low inlet turbulence level (<1%) in order to provide a cleaner environment which magnifies the effects of the excitation frequency, and with an increased turbulence intensity level of 3% which is representative of more typical engine values. Integrated wake loss values are also presented to evaluate the effect on blade performance.
Multidisciplinary Multipoint Optimization of a Transonic Turbocharger Compressor
Authors: Van den Braembussche R. A. , Alsalihi Z. , Verstraete T. , Matsuo A. , Ibaraki S. , Sugimoto K. , Tomita I.
Published: July 2013

A transonic centrifugal compressor for turbocharger applications has been redesigned by means of a multidisciplinary multipoint optimization system composed of: a 3D Navier-Stokes solver, a Finite Element stress Analyzer, a Genetic Algorithm and an Artificial Neural Network. The latter makes use of a database, containing the geometry and corresponding performance of previously analyzed impellers and allows a considerable reduction in computational effort. The performance of every new geometry is verified by a 3D Navier-Stokes solver. A Finite Element Analysis verifies the mechanical integrity of the impeller.

The geometrical description of the impeller has been extended to better adapt the inducer part of the impeller to transonic flows. The splitters are no longer copies of the full blades but specially designed for minimum losses and equal mass flow on both sides. The blade thickness and number of blades are unchanged because defined by robustness and inertia considerations.

The operating range is guaranteed by a two-step optimization procedure. The first one provides information allowing a modification of the inlet section to guarantee the required choking mass flow and a more accurate prediction of the boundary conditions for the Navier-Stokes analysis of the modified impeller. The second one predicts the performance curve of the new geometry for which the choking mass flow is known.

It is shown how these extensions of the optimization method have led to a considerable improvement of the efficiency and corresponding pressure ratio, while respecting the surge and choking limits without increase of the stress level.
Gradient Enhanced Surrogate Models Based on Adjoint CFD Methods for the Design of a Counter Rotating Turbofan
Authors: Backhaus Jan , Aulich Marcel , Frey Christian , Lengyel Timea , Voß Christian
Published: July 2013

This paper studies the use of adjoint CFD solvers in combination with surrogate modelling in order to reduce the computational cost of the optimization of complex 3D turbomachinery components. The method is applied to a previously optimized counter rotating turbofan, with a shape parameterized by 104 CAD parameters.

Through random changes on the reference design, a small number of design variations are created to serve as training samples for the surrogate models. A steady RANS solver and its discrete adjoint are then used to calculate objective function values and their corresponding sensitivities. Kriging and neural networks are used to build surrogate models from the training data. To study the impact of the additional information provided by the adjoint solver, each model is trained with and without the sensitivity information. The accuracy of the different surrogate model predictions is assessed by comparison against CFD calculations.

The results show a considerable improvement of the fitness function approximation when the sensitivity information is taken into account. Through a gradient based optimization on one of the surrogate models, a design with higher isentropic efficiency at the aerodynamic design point is created. This application demonstrates that the improved surrogate models can be used for design and optimization.
Hub Injection Flow Control in a Turbine Exhaust Diffuser
Authors: Volkmer Silke , Kuschel Benjamin , Hirschmann Andreas , Schatz Markus , Casey Michael , Montgomery Matthew
Published: July 2013

Downstream of a modern combined cycle gas turbine there is generally an abrupt end of the turbine inner hub wall leading to a backward facing step with separation and flow recirculation on the hub axis. This can lead to a strong hub wake with high losses in the center of the diffuser. The interaction between the end of the hub and the downstream axial diffuser can have a significant influence on diffuser flow field and hence on overall gas turbine performance. To minimize the hub wake the flow can be guided to the axis by reducing the diameter of the hub. But this can still lead to separation on the extended hub and to a strong hub wake. This paper describes experimental and numerical investigations of the use of flow injection to control the flow separation on the hub and to guide the main flow towards the axis of the diffuser to reduce the hub wake.

A new test rig has been built at the Institute of Thermal Turbomachinery and Machinery Laboratory of the University of Stuttgart, Germany (ITSM). This rig features an active flow control device based on hub flow injection with a Coanda-effect to influence the hub wake flow. In addition to experimental studies, steady CFD (computational fluid dynamics) simulations have been performed for this configuration and validated using the experimental data.

Test results and flow field computations confirm that the flow control device tested can have a beneficial influence on diffuser performance at specific conditions. If it is the casing flow which separates then hub injection has little effect. If the casing flow is attached, and a hub separation limits the performance, then hub jet flow injection can be highly beneficial. The results also indicate that the mass flow injected in the hub jet control device has to be carefully balanced.
Sound Generation and Propagation in a Centrifugal Pump With the Finite Volume EIF-Approach
Authors: Michels Thilo , Markiewicz Marian , von Estorff Otto
Published: July 2013

The development of new methods in the field of numerical aeroacoustics is one of the current research interests. For this kind of approaches, highly efficient methods are necessary. Aerodynamic noise prediction codes in use today are typically based on Lighthill analogies. In this paper an alternative approach is developed and implemented in a three dimensional CFD finite volume code. The method is based on the Expansion about Incompressible Flow (EIF) technique, which was proposed first by Hardin and Pope in 1994. Based on the solution of the incompressible flow, acoustic radiation is obtained in a compressible, inviscid fluid. The advantage of this technique, as compared to unsteady RANS simulations, is that small acoustic perturbations can easily be separated from the fluctuations of the pressure field of the bulk flow, which are some orders of magnitude greater. In comparison to the acoustic analogy approach, the method extends the region of applications towards the moderate Mach numbers (up to Ma = 0.6). Moreover, the flow and acoustic effects are resolved on different length scales. This makes the computations more efficient. The EIF method accounts both for the sound radiation and scattering. In the presented paper a further development towards three dimensional simulations and numerical implementation for centrifugal pumps is presented.