bluuebunny/crossref_metadata_2025_split huggingface dataset:
Effects of Winglet Geometry on the Aerodynamic Performance of Tip Leakage FLow in a Turbine Cascade
Authors: Zhou Chao , Hodson Howard , Tibbott Ian , Stokes Mark 
Published: July 2013

Experimental and numerical methods were used to investigate the aerodynamic performance of a winglet tip in a linear cascade. A flat tip and a cavity tip are studied as baseline cases. The flow patterns over the three tips are studied. The flow separates over the pressure side edge. For the cavity tip and the winglet tip, vortices appear in the cavity. These vortices reduce the discharge coefficient of the tip.
The purpose of using a winglet tip is to reduce the driving pressure difference. The pressure side winglet of the winglet geometry studied in this paper has little effect in reducing the driving pressure difference. It is found that the suction side winglet reduces the driving pressure difference of the tip leakage flow near the leading edge, but increases the driving pressure difference from midchord to the trailing edge. This is also used to explain the findings and discrepancies in other studies. Compared with the flat tip, the cavity tip and the winglet tip achieve a reduction of the loss to the size of the tip gap.
The effects of the rounding of the pressure side edge of the tips were studied to simulate the effects of deterioration. As the size of the pressure side edge radius increase, the tip leakage mass flow rate and the loss increase. The improvement of the aerodynamic performance by using a winglet remains similar when comparing with a flat tip or a cavity tip with the same pressure side radius.
Analysis of Radial Migration of Hot-Streak in Swirling Flow Through HP Turbine Stage
Authors: Khanal B. , He L. , Northall J. , Adami P. 
Published: July 2013

The high pressure (HP) turbine is subject to inlet flow non-uniformities resulting from the combustor. A lean-burn combustor tends to combine temperature variations with strong swirl and, although considerable research efforts have been made to study the effects of a circumferential temperature non-uniformity (hot-streak), there is relatively little known about the interaction between the two.
This paper presents a numerical investigation of the transonic test HP stage MT1 behaviour under the combined influence of the swirl and hot-streak. The in house Rolls-Royce HYDRA numerical CFD suite is used for all the simulations of the present study. Baseline configurations with either hot-streak or swirl at the stage inlet are analyzed to assess the methodology and to identify reference performance parameters through comparisons with the experimental data. Extensive computational analyses are then carried out for the cases with hot-streak and swirl combined including both the effects of the combustor-NGV clocking and the direction of the swirl. The present results for the combined hot-streak and swirl cases reveal distinctive radial migrations of hot fluid in the NGV and rotor passages with considerable impact on the aerothermal performance. It is illustrated that the blade heat transfer characteristics and their dependence on the clocking position can be strongly affected by the swirl direction. A further computational examination is carried out on the validity of a superposition of the influences of swirl and hot-streak. It shows that the blade heat transfer in a combined swirl and hot-streak case cannot be predicted by the superposition of each in isolation.
Multi-Objective Constrained Aero-Mechanical Optimization of an Axial Compressor Transonic Blade
Authors: Astrua Pio , Piola Stefano , Silingardi Andrea , Bonzani Federico 
Published: July 2013

This paper presents a flexible and effective optimization approach to design an axial compressor transonic blade for heavy duty gas turbines. The design goals are to improve design efficiency, choke margin and off-design performance while maintaining mass flow in design point as well as structural integrity. The new blade has to provide a wide operating range and to satisfy tight geometrical constraints.
A database of aero-mechanical calculation results is obtained for three operating conditions. A number of 3D flow simulations are performed using a CFD solver with endwall boundary layer simplified model (thin layer) to reduce computational costs. The optimization process adopts a set of artificial neural networks (ANN) trained for each operating condition and a random walking search algorithm to determine the multi-objective Pareto Front. ANN enables speed up of the optimization process and allows high flexibility in choosing criteria for optimum member selection. Random walking algorithm gives a fast and effective method to predict the multi-dimensional Pareto Front.
The Efficient Computation of Transient Flow in Turbine Blade Rows Using Transformation Methods
Authors: Connell Stuart , Hutchinson Brad , Galpin Paul , Campregher Rubens , Godin Philippe 
Published: July 2013

Computational predictions of the transient flow in turbine blade rows are considered. Adjacent blade rows typically contain unequal numbers of blades and vanes, requiring a computation over multiple passages per row to permit application of simple periodic boundary conditions. For typical geometries, use of conventional solution methods requires computation over all or a significant portion of the wheel to ensure a time accurate solution.
The computational load is significantly reduced by methods which enable a one or two-passage solution to accurately model the full wheel (or part wheel, if applicable) solution. In this work, three methods are used: Profile Transformation, Fourier Transformation and Time Transformation.
This paper will concentrate on the evaluation of these methods on two turbine geometries. The first test case is a frozen gust analysis for a high pressure transonic turbine; the geometry includes hub and casing cavities together with a complex tip. The second test case is a low pressure turbine stage run over a range of operating points. Comparisons between the various methods and the equivalent part wheel periodic solution are made to demonstrate the accuracy and computational efficiency of the transformation methods.
PADRAM: Parametric Design and Rapid Meshing System for Complex Turbomachinery Configurations
Authors: Milli Andrea , Shahpar Shahrokh 
Published: July 2013

This paper reports the progress made in a parametric design and rapid meshing system (PADRAM) developed under two recent UK national sponsored research programs. PADRAM is designed to parametrically change the blade geometry and rapidly generate body-conformal high-quality viscous meshes. This allows speeding up the CFD loop by making the meshing process fully automatic on the basis of pre-designed templates. The geometry parameterisation is done within the mesh generator, making its integration within the optimisation loop straightforward.
The paper presents examples that demonstrate how incorporating real geometry features into PADRAM is fundamental to achieve numerical models closer to reality. This is key factor in trusting the CFD solution and making use of it to further improve current designs. It also shows that feature-based structured mesh is good for cases that need to be repeatable across sites and teams, where consistency of the mesh is crucial and quick answers required to cope with tight project deadlines.
The incorporation of additional complex geometrical features limits the applicability of the template structured approach and can be sometimes at the expense of mesh quality. In this regard, novel unstructured meshing technologies have been developed and implemented into PADRAM in order to deal with non standard and complex configurations. Two of these methods are presented in this paper (i.e., Zipper Layer and Delaunay Cavity). The paper shows how these can be used to paste together various multi-block structured meshes, hence providing the most suitable meshing topology to be used for each component separately. This makes the PADRAM code a unique meshing tool, able to combine the advantages of the template-topology approach to the flexibility of fully unstructured meshes. A number of supportive examples is reported in the paper.
Numerical and Analytical Study of the Effect of Film Cooling on HP NGV Capacity
Authors: Hambidge Chris , Povey Tom 
Published: July 2013

A numerical study has been conducted to investigate the effects of film cooling on turbine capacity in transonic HP NGVs. The study is part of a wider research program the aim of which is to quantify the sensitivity of capacity predictions to the methods of numerical prediction.
Results of two dimensional CFD are presented for the case of single and multiple film cooling rows. The capacity characteristics (with pressure ratio) for modern airfoil profiles, and the change in capacity with film cooling are investigated. The interaction between rows is shown to be weak, and possibility of using superposition techniques is discussed. The effect of turbulence model on the change in capacity with film cooling is discussed.
The CFD simulations are compared to a simple analytical two dimensional mixing model developed to aid fundamental understanding of the problem. The effect of film cooling on turbine capacity is modelled by considering a mainstream flow displaced by a partially mixed film cooling layer. The model uses fundamental cooling performance parameters which can be estimated an early stage of the engine design process, providing a means for early assessment of the correction due to film cooling.
It is argued that it may be unnecessary to simulate film cooling in capacity predictions.
Unsteady Interaction Between Annulus and Turbine Rim Seal Flows
Authors: Chilla Martin , Hodson Howard , Newman David 
Published: July 2013

In core gas turbines relatively cold air is purged through the hub gap between stator and rotor in order to seal the disc space against flow ingestion from the main annulus. Although the sealing mass flow rate is commonly very small compared to the main annulus mass flow rate, it can have significant effects on the development of the passage endwall flows and on the overall loss generation.
In this paper, the interaction between annulus and rim sealing flows is investigated using numerical simulations of a generic high-pressure turbine. At first, the numerical approach is validated by comparing the results of calculations to measurement data at the design flow conditions. Following that, results from steady and unsteady calculations are used to describe in detail the aerodynamics in overlap-type rim seals and their effects on the blade passage flow. It is found that the flow interaction at the rim seal interface is strongly influenced by the velocity deficit of the rim sealing flow relative to the annulus flow as well as by the circumferentially non-uniform pressure field imposed by the rotor blades. At typical sealing flow conditions, the flow interaction is found to be naturally unsteady, with periodical vortex shedding into the rotor passage.
Finally, the influence of the specific rim seal shape on the flow unsteadiness at the rim seal interface is investigated and the impact on turbine performance is assessed.
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.
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.
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.
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.