bluuebunny/crossref_metadata_2025 huggingface dataset:
Effect of Speed Ratio and Axial Spacing Variations on the Performance of a High Aspect Ratio, Low Speed Contra-Rotating Fan
Authors: Mistry Chetan S. , Pradeep A. M.
Published: July 2013

This paper explores the effect of speed ratio and axial spacing between high aspect ratio, low speed contra-rotating pair rotors on their aerodynamic performance. The blades were designed with a low hub-tip ratio of 0.35 and an aspect ratio of 3.0. Numerical and experimental studies are carried out on these contra-rotating rotors operating at a Reynolds number of 1.258 × 105 (based on blade chord). The first and second rotors were designed to develop a pressure rise of 1100 Pa and 900 Pa, respectively, for total mass flow rate of 6 kg/s when both operating at a design speed of 2400 rpm. The performance of the fan was evaluated based on variations of total pressure and flow angles at off-design operating conditions. The measurementsof total pressure rise, flow angles etc. are taken upstream of the first rotor and in between the two rotors and downstream of the second rotor. The performance of the contra rotating stage is mainly influenced by the axial spacing between the rotors and speed ratio of both the rotors. The study reveals that the aerodynamics of the contra-rotating stage and stall margin is significantly affected by both the speed ratio as well as the axial spacing between the rotors. It was found that with increasing the speed ratio, the strong suction generated by the second rotor, improves the stage pressure rise and stall margin. Lower axial spacing changes the flow incidence to the second rotor and thereby improves the overall performance of the stage. This however, is accompanied by an increased noise level. The performance is investigated at different speed ratios of the rotors at varying axial spacing. Detailed numerical simulations have been conducted using ANSYS CFX13© using mixing plane approach between rotors. Numerical simulations are compared with experimental results at off-design conditions. These results are validated using the experimental data. Numerical simulations are expected to provide deeper insight into the flow physics of contra-rotating rotors which may be difficult to capture experimentally.
Measurement of Aerodynamic Losses for Turbine Airfoil Cascades With Varying Pitch, Operating Under Transonic Conditions
Authors: Abraham Santosh , Panchal Kapil , Ekkad Srinath V. , Ng Wing , Lohaus Andrew S. , Malandra Anthony
Published: July 2013

The paper presents detailed experimental results of the midspan total pressure losses, secondary flow field, and static pressure measurements on two linear, high-turning turbine cascades at transonic conditions. The airfoils in the two cascades being studied are identical and their aerodynamic loading levels are varied by increasing the pitch of one cascade by 25% with respect to the other. The turbine cascades are referred to as B1-SP and B1-IP. Cascade B1-IP, with its increased pitch, has a Zweifel coefficient increased by 25%. The airfoils have a turning angle of ∼127°. Measurements are made at design and off-design conditions, at exit Mach numbers ranging from 0.71 to 0.95. The exit span of the airfoils are increased relative to the inlet span with the intention of obtaining a ratio of inlet Mach number to exit Mach number that is representative to that encountered in a real engine. This results in one end wall diverging from inlet to exit at a 13 degree angle, which simulates the required leading edge loading as seen in an engine. The objective of this study is to investigate the variation in airfoil loading distribution and the effect it has on aerodynamic performance in terms of pressure losses. Detailed loss measurements, both in the pitchwise as well as spanwise directions, are conducted at 0.1 Cax and 1.0 Cax locations downstream of the trailing edge. Results from 3D viscous numerical simulations have been used to assist the interpretation of experimental results.
Numerical Study on the Effects of Casing Treatment on Unsteadiness of Tip Leakage Flow in an Axial Compressor
Authors: Hwang Yoojun , Kang Shin-Hyoung
Published: July 2013

A low speed axial compressor with casing treatment of axial slots was numerically investigated. Time-accurate numerical calculations were performed to simulate unsteady flow in the rotor tip region and the effects of casing treatment on the flow. Since the compressor rotor had a large tip clearance, it was found that the tip leakage flow had an inherent unsteady feature that was not associated with rotor rotation. The unsteadiness of the tip leakage flow was induced by changes in the blade loading due to the pressure distribution formed by the tip leakage flow. This characteristic is called rotating instability or self-induced unsteadiness. The frequency of the flow oscillation was found to decrease as the flow rate was reduced. On the other hand, as expected, the operating range was improved by casing treatment, as shown by calculations in good agreement with the experimentally measured data. The unsteadiness of the tip leakage flow was alleviated by the casing treatment. The interaction between the flow in the tip region and the re-circulated flow through the axial slots was observed in detail. The removal and injection of flow through the axial slots were responsible not only for the extension of the operating range but also for the alleviation of the unsteadiness. Analyses of instantaneous flow fields explained the mechanism of the interaction between the casing treatment and the unsteady oscillation of the tip leakage flow. Furthermore, the effects of changes in the amount of re-circulation and the location of the removal and injection flow on the unsteadiness of the tip leakage flow were examined.
Endwall Profiling With Tip Clearance Flows
Authors: Hilfer Michael , Ingram Grant , Hogg Simon
Published: July 2013

This paper describes the design of a non-axisymmetric profiled endwall with tip clearance flow. Previous designs have assumed that the interaction between the tip clearance flow and the endwall has been negligible and this paper allows this assumption to be tested.

A profiled endwall was designed, built and tested in a low speed linear cascade. The design was conducted using the Durham endwall design system and this paper provides details of the first experimental validation of this system. Overall the design system was found to perform successfully.

Many other researchers have conducted endwall design, test and build exercises and the key advance here was to examine the effect of profiled endwalls on tip clearance flow. To this end a tip clearance was added to both the CFD and the cascade during the design process. The results show that even with an aspect ratio of two the endwall produces noticeable changes on the tip clearance flow. This does not however necessarily prevent a successful profiled endwall from being implemented in the cascade.
Experimental Studies of Leading Edge Contouring Influence on Secondary Losses in Transonic Turbines
Authors: Saha Ranjan , Fridh Jens , Fransson Torsten , Mamaev Boris I. , Annerfeldt Mats
Published: July 2013

An experimental study of the hub leading edge contouring using fillets is performed in an annular sector cascade to observe the influence of secondary flows and aerodynamic losses. The investigated vane is a three dimensional gas turbine guide vane (geometrically similar) with a mid-span aspect ratio of 0.46. The measurements are carried out on the leading edge fillet and baseline cases using pneumatic probes. Significant precautions have been taken to increase the accuracy of the measurements. The investigations are performed for a wide range of operating exit Mach numbers from 0.5 to 0.9 at a design inlet flow angle of 90°. Data presented include the loading, fields of total pressures, exit flow angles, radial flow angles, as well as profile and secondary losses. The vane has a small profile loss of approximately 2.5% and secondary loss of about 1.1%. Contour plots of vorticity distributions and velocity vectors indicate there is a small influence of the vortex-structure in endwall regions when the leading edge fillet is used. Compared to the baseline case the loss for the filleted case is lower up to 13% of span and higher from 13% to 20% of the span for a reference condition with Mach no. of 0.9. For the filleted case, there is a small increase of turning up to 15% of the span and then a small decrease up to 35% of the span. Hence, there are no significant influences on the losses and turning for the filleted case. Results lead to the conclusion that one cannot expect a noticeable effect of leading edge contouring on the aerodynamic efficiency for the investigated 1st stage vane of a modern gas turbine.
Reynolds Number Effects in a Low Pressure Turbine
Authors: Hura Harjit S. , Joseph John , Halstead Dave E.
Published: July 2013

This paper reports the results of an experimental and analytical study of Reynolds number (Re) effect on Low Pressure Turbine (LPT) performance. LPTs suffer both a loss in efficiency and flow capacity at altitude due to thicker boundary layers and higher viscous losses. Boundary layer separation can occur in highly loaded and/or high lift designs. The magnitude of the effect is stronger for smaller engines being designed for regional jets which may have cruising altitudes above 50K feet. There is a general lack of knowledge about performance degradation in commercial LPTs under these conditions.

A test program was undertaken in a low pressure 3 stage axial turbine to quantify the effect of low Re on efficiency and flow lapse rate. Rig inlet pressures were varied from 0.27E+5 N/m2 (4 psia) to 4.40E+5 N/m2 (65 psia) to achieve over a 15 fold variation in Re. The chord based average Re varied from 30000 to 500000. Efficiency and flow function lapse of over 5% was measured. The fall off was non-linear with a rapid loss occurring at Re below 100000.

3D CFD analysis was conducted in parallel to predict overall performance but also understand loss details within the blade rows. Measured inlet profiles of total pressure, temperature and air angle, and exit static pressure were used as boundary conditions. Leakages and purge flows were modeled as source terms. A turbulence transition model and wall integration grids was used. The CFD results corroborate the test findings on the overall efficiency and flow capacity lapse rate. Analysis of blade row performance at high and low Re shows a sharp increase in profile loss at low Re.
Effects of Boundary Layer Ingestion on the Aero-Acoustics of Transonic Fan Rotors
Authors: Defoe J. J. , Spakovszky Z. S.
Published: July 2013

The use of boundary-layer-ingesting, embedded propulsion systems can result in inlet flow distortions where the interaction of the boundary layer vorticity and the inlet lip causes horseshoe vortex formation and the ingestion of streamwise vortices into the inlet. A previously-developed body-force-based fan modeling approach was used to assess the change in fan rotor shock noise generation and propagation in a boundary-layer-ingesting, serpentine inlet. This approach is employed here in a parametric study to assess the effects of inlet geometry parameters (offset-to-diameter ratio and downstream-to-upstream area ratio) on flow distortion and rotor shock noise.

Mechanisms related to the vortical inlet structures were found to govern changes in the rotor shock noise generation and propagation. The vortex whose circulation is in the opposite direction to the fan rotation (counter-swirling vortex) increases incidence angles on the fan blades near the tip, enhancing noise generation. The vortex with circulation in the direction of fan rotation (co-swirling vortex) creates a region of subsonic relative flow near the blade tip radius which decreases the sound power propagated to the far-field.

The parametric study revealed that the overall sound power level at the fan leading edge is set by the ingested streamwise circulation, and that for inlet designs in which the streamwise vortices are displaced away from the duct wall, the sound power at the upstream inlet plane increased by as much as 9 dB. By comparing the far-field noise results obtained to those for a conventional inlet, it is deduced that the changes in rotor shock noise are predominantly due to the ingestion of streamwise vorticity.
The Effect of Airfoil Thickness on the Efficiency of LP Turbines
Authors: Torre Diego , Vázquez Raúl , Armanañzas Leyre , Partida Fernando , García-Valdecasas Guillermo
Published: July 2013

The effect of airfoil thickness on the efficiency of Low Pressure (LP) Turbines has been investigated experimentally in a multistage turbine high-speed rig. The rig consists of three stages of a state of the art LP turbine. The stages are characterized by a very high hade angle, reverse cut-off design, very high lift and very high aspect ratio airfoils.

Two different sets of stators have been designed and tested. The first set of stators is made of airfoils with a thickness to chord ratio around 10% along the span with exception of a small areas close to the endwalls. In those areas, the thickness has been increased above the previous value to reduce the secondary flows. These types of airfoils have been referred in the literature as “spoon” airfoils.

The second set of stators has been designed to have the same spanwise distribution of pressure coefficient (Cp) on the suction surface than the first set. However, the thickness to chord ratio was increased along the span up to values around 20% to rise the velocity of the flow and to remove any separation bubble on the pressure side. The resulting shape of the profiles is representative of “hollow” airfoils.

The velocity triangles, chord distribution, leading and trailing edge locations and flowpath have been maintained between both sets. They have been tested with the same blades and at the same operating conditions with the intention of determining the impact of the profile thickness on the overall efficiency. The turbine characteristics: sensitivity to speed, specific work, Reynolds number and purge flows have been obtained for both sets.

The comparison of the results suggests that the efficiency of both types of airfoils exhibit the same behaviour, no significant differences in the results can be distinguished.
Experimental Investigation of the Unsteady Flow Field Downstream of a Counter-Rotating Two-Spool Turbine Rig
Authors: Lengani Davide , Santner Cornelia , Spataro Rosario , Paradiso Berardo , Göttlich Emil
Published: July 2013

The paper presents an experimental investigation of the unsteady flow field in the two-spool counter-rotating transonic turbine at Graz University of Technology. The test setup consists of a high pressure (HP) stage, a diffusing mid turbine frame with turning struts (TMTF) and a shrouded low pressure (LP) rotor. The two rotors are mounted on mechanically independent shafts in order to provide engine-representative operating conditions. The rig was designed in cooperation with MTU Aero Engines and Volvo Aero within the EU project DREAM (ValiDation of Radical Engine Architecture SysteMs).

A two-sensor fast response aerodynamic pressure probe (2S-FRAP) has been employed to provide time-resolved aerodynamic area traverses downstream of the LP turbine. Such measurement allows estimating the total deterministic unsteadiness transported through the duct. In particular, it has been possible to isolate the structures induced by each rotor by means of the encoders mounted on the two shafts. A global ensemble averaged field, which takes into account the rotor-rotor interactions, is also provided. The time resolved distributions of the flow quantities are then discussed in details. The perturbations due to the HP rotor in terms of velocity and flow angle are negligible in this downstream plane. Indeed, the largest fluctuations of velocity are due to the TMTF-LP rotor interaction, they occur in the wake and secondary flows of the TMTF. Large fluctuations of static and total pressure are instead due to both rotors to the same extent.
The Sensitivity of 2D Compressor Incidence Range to In-Service Geometric Variation
Authors: Goodhand Martin N. , Miller Robert J. , Lung Hang W.
Published: July 2013

The loss mechanisms which control 2D incidence range are discussed with an emphasis on determining which real in-service geometric variations will have the largest impact. For the majority of engine compressor blades (Minlet>0.55) both the negative and positive incidence limits are controlled by supersonic patches. It is shown that these patches are highly sensitive to the geometric variations close to, and around the leading edge. The variations used in this study were measured from newly manufactured as well as ex-service blades. Over most the high pressure compressor considered, it was shown that manufacture variations dominated.

The first part of the paper shows that, despite large geometric variations (∼10% of leading edge thickness), the incidence range responded in a linear way. The result of this is that the geometric variations have little effect on the mean incidence range of a row of blades. In the second part of the paper a region of the design space is identified where non-linear behavior can result in a 10% reduction in positive incidence range. The mechanism for this is reported and design guidelines for its avoidance offered. In the final part of the paper, the linear behavior at negative incidence and the transonic nature of the flow is exploited to design a robust asymmetric leading edge with a 5% increase in incidence range.
Investigation of the Laminar Separation-Induced Transition With the γ-Re~θt Transition Model on Low-Pressure Turbine Rotor Blades at Steady Conditions
Authors: Babajee Jayson , Arts Tony
Published: July 2013

The present study focuses on the assessment and the validation of the Langtry and Menter γ-Re~θt correlation-based transition model, recently implemented in the RANS code elsA from ONERA. The test cases are two Low-Pressure Turbine (LPT) rotor blades, with the same compressible Zweifel loading coefficient but different load distributions. The corresponding experimental results are provided by the von Karman Institute. The outlet isentropic Reynolds number, based on blade chord and outlet isentropic velocity, ranges from 60000 to 250000 in order to investigate the complex separation-induced transition phenomenon occurring at low Reynolds number cruise condition. The turbulence intensity is the natural freestream turbulence of the facility (0.9%). The numerical test campaign gives good predictions, particularly when accurately tailoring the trailing edge and wake regions of the mesh. The numerical mass-averaged kinetic losses are in good agreement with the experimental ones. Moreover, the flow topology parameters (the transition onset, the transition end and the separation) show good agreement in comparison to the experimental results and correlations from the open literature. One is even able to detect separation-induced transition with reattachment before the trailing edge at the lowest Reynolds number. However, the authors stress the turbulence Reynolds number effect on the prediction of separation-induced transition for strong diffusion LPT blades and particularly when the flow is subjected to a long bubble or even an open separation.
Origins and Structure of Spike-Type Rotating Stall
Authors: Pullan G. , Young A. M. , Day I. J. , Greitzer E. M. , Spakovszky Z. S.
Published: July 2013

In this paper we describe the structures that produce a spike-type route to rotating stall and explain the physical mechanism for their formation. The descriptions and explanations are based on numerical simulations, complemented and corroborated by experiments. It is found that spikes are caused by a loss of pressure rise capability in the rotor tip region, due to flow separation resulting from high incidence. The separation gives rise to shedding of vorticity from the leading edge and the consequent formation of vortices that span between the suction surface and the casing. As seen in the rotor frame of reference, near the casing the vortex convects toward the pressure surface of the adjacent blade. The approach of the vortex to the adjacent blade triggers a separation on that blade so the structure propagates. The above sequence of events constitutes a spike. The simulations show shed vortices over a range of tip clearances including zero. The implication is that they are not part of the tip clearance vortex, in accord with recent experimental findings. Evidence is presented for the existence of these vortex structures immediately prior to spike-type stall and, more strongly, for their causal connection with spike-type stall inception. Data from several compressors are shown to reproduce the pressure and velocity signature of the spike-type stall inception seen in the simulations.
Boundary Conditions for Vortex CFD Simulations
Authors: McLelland Grant , di Cugno Domenico , MacManus David G. , Pachidis Vassilios
Published: July 2013

The aim of this research is to identify a method of prescribing boundary conditions for CFD studies of vortex ingestion. The work illustrates the use of a validated vortex model which has been shown to give good agreement with experimental data. The model is used to generate the vortex velocity and pressure distributions. Prior knowledge of the vortex strength, core radius, Vatistas model constant, and maximum axial velocity perturbation is required. The model has been generalized, which allows any vortex location on the inlet boundary to be specified. Two examples are given. The first case is for a convecting wing-tip vortex. The second is for a ground vortex prescribed at the inlet of the compressor system. The ground vortex model is based on experimental data as well as previous validated CFD studies. A preliminary assessment of the effect of ground vortex ingestion on the performance of NASA Rotor 67 is carried out using the boundary conditions proposed. Furthermore, a sensitivity analysis on the boundary condition components was completed. The use of such boundary conditions can reduce the resource requirements and enables the independent specification of a vortical flow distortion for controlled parametric investigations of intake or compressor sensitivity studies.
The Practical Design of Turbocharger Turbines Using a Throughflow-Based Optimisation Process
Authors: Cox Graham D.
Published: July 2013

The modern trends in automotive turbocharger applications are towards the boosting of smaller internal combustion engines and more advanced systems including two-stage, turbo-compounding and hybrid electric-motor assist. Off-the-shelf turbochargers will become a smaller share of the market and the choice of major parameters for the compressor and turbine, e.g. speed and diameter, will fall outside of the manufacturer’s knowledge base. The selection of the compressor and turbine may even be independent. The only certainty is that the turbomachinery will have to be small, cheap and efficient.

To provide some guidance to the turbine designer, this paper presents the results of a study in which practical designs have been generated to cover the range of conceivable parameters, presented in non-dimensional terms to provide general applicability. All the designs are generated using a throughflow-based optimisation system in which the candidate geometries are assessed against mechanical as well as aerodynamic and inertia targets.

Analysis of the results gives clues to the form of the basic empiricism that would be of use in the preliminary design of automotive turbocharger turbines.
Design Procedure for a Turgo Type Turbine Using a Three-Dimensional Potential Flow
Authors: Clarembaux Correa Jorge Luis , De Andrade Jesús , Noguera Ricardo , Croquer Sergio , Jeanty Freddy , Asuaje Miguel
Published: July 2013

A useful methodology in the design of a Turgo Type Turbine (TTT) has been accomplished through the study of a particular three dimensional potential flow, known as Rankine Ovoids. The obtained streamlines solution for this flow was modified implementing several algorithms in order to select a suitable flow profile that could be adapted as a prediction of the flow passing through the buckets of a Turgo runner. Afterwards, the selected profile was incorporated with other geometric parameters, which were based on the hydrodynamic and geometric conditions presented in a TTT, in the design methodology proposed for this investigation. In addition, the equations to calculate power and efficiency of this kind of turbines are included. The global efficiency in the 3DT methodology was reported to be 80.8% for the designed TTT, which it is an expected value for this kind of turbines.
CREATE: Advanced CFD for HPC Performance Improvement
Authors: Pesteil A. , Cellier D. , Domercq O. , Perrot V. , Boniface J. C.
Published: July 2013

The present paper provides an overview of technological evolutions aimed at improving the aerodynamic performance of Snecma’s High Pressure Compressors. Several concepts were investigated under its CREATE compressor research program, involving an extensive simulation effort. An overview of the computational approaches involved in the evaluation and selection of innovative and most promising concepts will be given in the paper. The main topics dealt with are:

1) Aeromechanical optimization of airfoils and flow path:

In recent years, great efforts have been made to improve the aerodynamic design of airfoils. Among them, optimization methods have been progressively implemented in the design process with an increased complexity logic. The latest methods used at Snecma involve multi-objective, multi-parameters aeromechanical optimization including mean camber line, stacking axis, flow path contouring and more. This work is illustrated by two practical examples.

2) Vortex generators:

In order to control the flow, vortex generators can be forecasted as a promising step forward. The goal is to create exogenous vorticity that will counter-balance the endogenous vorticity. Thus they appear as a tool to reduce losses and improve stability in highly loaded turbomachinery devices such as modern high pressure compressors. This section of the paper will give an overview of the dedicated numerical simulations completed.

3) Clocking:

Numerous studies are related to the benefit drawn from turbine clocking on turbomachinery performance. However, fewer examples of successful compressor clocking exist. The recent capability of computational fluid dynamics tools to reduce the computational effort necessary to investigate such an issue (by the use of harmonic balance methodology) gives the opportunity for a renewed evaluation.

4) Optimization of shroud leakage flow with main flow:

A strong interaction exists with the secondary flows originated in the inner flow path cavities. Coupled main flow path and cavities aerodynamic simulations were conducted to improve the relevance of the computations and understand the mechanisms involved.

5) Tandem bladings and splitters in axial rotors:

The last aspect of the study was focused on dual blading concepts. After a brief review of the literature, some simulations were carried out to explore the relevance of such concepts from the viewpoint of modern high pressure compressors performance improvement.
Steady State and Transient CFD Studies on Aerodynamic Performance Validation of a High Pressure Turbine
Authors: Murari Sridhar , Sunnam Sathish , Liu Jong S.
Published: July 2013

With the advent of fast computers and availability of less costly memory resources, computational fluid dynamics (CFD) has emerged as a powerful tool for the design and analysis of flow and heat transfer of high pressure turbine stages. CFD gives an insight in to flow patterns that are difficult, expensive or impossible to study using experimental techniques. However, the application of CFD depends on its accuracy and reliability. This requires the CFD code to be validated with laboratory measurements to ensure its predictive capacity.

In the continual effort to improve analysis and design techniques, Honeywell has been investigating in the use of CFD to predict the aerodynamic performance of a high pressure turbine. Reynolds Averaged Navier Stokes (RANS), unsteady models like detached eddy simulation (DES), large eddy simulation (LES), and Scale Adaptive Simulation (SAS) are used to predict the aerodynamic performance of a high pressure turbine. Mixing plane approach is used to address the flow data transport across the stationary interface in RANS simulation. The film holes on blade surface and end walls for all the analysis are modeled by using actual film hole modeling technique.

The validation is accomplished with the test results of a high pressure turbine, Energy Efficient Engine (E3). The aerodynamic performance data at design point, typical off-design points are taken as test cases for the validation study. One dimensional performance parameters such as corrected mass flow rate, total pressure ratio, cycle efficiency, and two dimensional spanwise distributions of total pressure, total temperature and flow angle that are obtained from CFD results are compared with test data. Streamlines and flow field results at different measurement planes are presented to understand the aerodynamic behavior.
An Experimental Study of Loss Sources in a Fan Operating With Continuous Inlet Stagnation Pressure Distortion
Authors: Gunn E. J. , Tooze S. E. , Hall C. A. , Colin Y.
Published: July 2013

The viability of Boundary Layer Ingesting (BLI) engines for future aircraft propulsion is dependent on the ability to design robust, efficient engine fan systems for operation with continuously distorted inlet flow. A key step in this process is to develop an understanding of the specific mechanisms by which an inlet distortion affects the performance of a fan stage. In this paper, detailed full-annulus experimental measurements of the flow field within a low-speed fan stage operating with a continuous 60-degree inlet stagnation pressure distortion are presented. These results are used to describe the three-dimensional fluid mechanics governing the interaction between the fan and the distortion and to make a quantitative assessment of the impact on loss generation within the fan. A 5.3 percentage point reduction in stage total-to-total efficiency is observed as a result of the inlet distortion. The reduction in performance is shown to be dominated by increased loss generation in the rotor due to off-design incidence values at its leading edge, an effect which occurs throughout the annulus despite the localised nature of the inlet distortion. Increased loss generation in the stator row is also observed due to flow separations that are shown to be caused by whirl angle distortion at rotor exit. By addressing these losses, it should be possible to achieve improved efficiency in BLI fan systems.
Meanline Modeling of Ported Shroud Turbocharger Compressor
Authors: Qiu Xuwen , Krivitzky Eric M. , Bollweg Peter
Published: July 2013

The requirements for higher fuel economy and better diesel and gasoline engines demand a wider range in turbocharger compressor operation. Ported shroud compressor housing is one of the most commonly used techniques for compressor map width enhancement. Although the general mechanism of such a flow feature is well understood, there are no readily available design tools to guide the engineers at the preliminary design stage. Designers have had to rely on three-dimensional (3D) CFD tools to sort out many design variables, but these tools can be prohibitively expensive.

This paper explains how to develop a ported shroud compressor model on top of a commercial meanline compressor design program. The model considers some basic parameters, such as bleed location and geometry, which drive the recirculation or bypass flow through the bleed channel. The effects of the secondary flow on the compressor performance, such as pressure rise, efficiency, and stall and choke margins are also analyzed. The model prediction is validated with CFD simulation and test data.
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.
Novel Performance Prediction of a Transonic 4.5 Stage Compressor
Authors: Schmitz Andreas , Aulich Marcel , Schönweitz Dirk , Nicke Eberhard
Published: July 2013

Computing capacities have grown exponentially in recent years and 3D-Navier-Stokes methods were developed widely. However it is still not feasible to design a multi-stage compressor directly in three dimensions. Instead, compressor design starts with 1D-design. In accordance with this approach, basic parameters such as the number of stages and stage pressure ratios are determined. In the following 2D-design, the geometry of the flow channel and the main parameters of the blade geometries can be determined. Afterwards in the 3D-design, unsteady and 3D-flow-effects are considered and the design optimized accordingly. Therefore, it is virtually impossible to correct conceptual faults in the 3D-design phase. Thus a robust and reliable 2D-Throughflow-solver including a performance prediction for modern airfoil geometries is necessary. So far there is no efficient methodology known which predicts the performance for all kinds of airfoil geometries, as it would be necessary in a 2D-Throughflow optimization process. In [1, 2] a novel methodology was presented, which is able to predict the performance for a large number of airfoil geometries accurately. This method is based on a large airfoil database which is used to train a surrogate model for airfoil performance prediction. The scope of this work is to validate and to document the progress of this new approach. In Schmitz et al. [1] it was validated on rotor 1 of the 4.5 stage transonic test compressor DLR-RIG250 of the Institute of Propulsion Technology. In this work all 4.5 stages were calculated at different speedlines and different vane positions. The results of the S2-solver are compared to experimental data and 3D-CFD calculations, obtained using the DLR in-house solver TRACE.
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.
Investigation of Streamwise and Transverse Instabilities on Swept Cylinders and Implications for Turbine Blading
Authors: Gostelow J. P. , Rona A. , De Saint Jean M. , Garrett S. J. , McMullan W. A.
Published: July 2013

The starting point for this investigation was the observation of robust streamwise streaks in flow visualization on the suction surfaces of blades in a turbine cascade at subsonic and transonic speeds. The spanwise wavelength of an array of streamwise vortices had been predicted by Kestin and Wood and is here confirmed experimentally. Observations of streaks on unswept turbine blades and on circular cylinders confirmed these earlier predictions, providing a firm basis for referencing the new measurements of vortical behavior. The observations made it clear that the boundary layers are highly three dimensional. In this paper observations of streamwise and transverse instabilities on swept circular cylinders, over a range of inclinations, are presented. The circular cylinder is a canonical case and observations relate the streamwise vorticity of the unswept case to the more aggressive crossflow instability, at high sweep angles, studied by Poll. Introducing sweep brings consideration of a wide range of instabilities. Prominent is crossflow instability resulting from the inflectional behavior of a three-dimensional boundary layer.
Numerical Assessment of Unsteady Flow Effects on a Nozzled Turbocharger Turbine
Authors: Padzillah M. H. , Rajoo S. , Martinez-Botas R. F.
Published: July 2013

In order to extract maximum amount of energy possible from the automotive reciprocating engine exhaust gas, the turbocharger usually installed closely downstream the exhaust valve thus exposing it to highly pulsating flow conditions. This condition induces highly complex flow field within the turbocharger stage and significantly impact its performance characteristics which is not fully understood. The main objective of this paper is to provide understanding of unsteady flow feature using a Computational Fluid Dynamics (CFD) approach validated with experimental data. Despite focusing on unsteady feature of the flow, this research also emphasizes the importance of accurately modelled geometry in the early section of the paper. A steady state validation against experimental data is performed prior to unsteady calculations. The effect of different phase shifting methods is described and the relationship of instantaneous efficiency with incidence angle is established. In the final section of this paper, the turbocharger stage is sectioned where its instantaneous performance is evaluated individually in each section. The unsteady simulation is performed at fixed 30 000 RPM with 20Hz pulsing flow.
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.