Integrated Outlet Guide Vane Design for an Aggressive S-Shaped Compressor Transition Duct

2012 ◽  
Vol 135 (1) ◽  
Author(s):  
A. D. Walker ◽  
A. G. Barker ◽  
J. F. Carrotte ◽  
J. J. Bolger ◽  
M. J. Green
Keyword(s):  
Gas Turbines ◽  
Static Pressure ◽  
Pressure Field ◽  
Guide Vane ◽  
Integrated Design ◽  
Axial Compressor ◽  
Test Facility ◽  
Speed Test ◽  
Duct Geometry ◽  
System Length

Within gas turbines the ability to design shorter aggressive S-shaped ducts is advantageous from a performance and weight saving perspective. However, current design philosophies tend to treat the S-shaped duct as an isolated component, neglecting the potential advantages of integrating the design with the upstream or downstream components. In this paper, such a design concept is numerically developed in which the upstream compressor outlet guide vanes are incorporated into the first bend of the S-shaped duct. Positioning the vane row within the first bend imparts a strong radial gradient to the pressure field within the vane passage. Tangential lean and axial sweep are employed such that the vane geometry is modified to exactly match the resulting inclined static pressure field. The integrated design is experimentally assessed and compared to a conventional nonintegrated design on a fully annular low speed test facility incorporating a single stage axial compressor. Several traverse planes are used to gather five-hole probe data which allow the flow structure to be examined through the rotor, outlet guide vane and within the transition ducts. The two designs employ almost identical duct geometry, but integration of the vane row reduces the system length by 21%. Due to successful matching of the static pressure field, the upstream influence of the integrated vane row is minimal and the rotor performance is unchanged. Similarly, the flow development within both S-shaped ducts is similar such that the circumferentially averaged profiles at duct exit are almost identical, and the operation of a downstream component would be unaffected. Overall system loss remains nominally unchanged despite the inclusion of lean and sweep and a reduction in system length. Finally, the numerical design predictions show good agreement with the experimental data thereby successfully validating the design process.

Integrated OGV Design for an Aggressive S-Shaped Compressor Transition Duct

2011 ◽  
Author(s):  
A. D. Walker ◽  
A. G. Barker ◽  
J. F. Carrotte ◽  
J. J. Bolger ◽  
M. J. Green
Keyword(s):  
Gas Turbines ◽  
Static Pressure ◽  
Pressure Field ◽  
Guide Vane ◽  
Integrated Design ◽  
Axial Compressor ◽  
Test Facility ◽  
Speed Test ◽  
Duct Geometry ◽  
System Length

Within gas turbines the ability to design shorter aggressive S-shaped ducts is advantageous from a performance and weight saving perspective. However, current design philosophies tend to treat the S-shaped duct as an isolated component, neglecting the potential advantages of integrating the design with the upstream or downstream components. In this paper such a design concept is numerically developed in which the upstream compressor outlet guide vanes are incorporated into the first bend of the S-shaped duct. Positioning the vane row within the first bend imparts a strong radial gradient to the pressure field within the vane passage. Tangential lean and axial sweep are employed such that the vane geometry is modified to exactly match the resulting inclined static pressure field. The integrated design is experimentally assessed and compared to a conventional non-integrated design on a fully annular low speed test facility incorporating a single stage axial compressor. Several traverse planes are used to gather five-hole probe data which allow the flow structure to be examined through the rotor, outlet guide vane and within the transition ducts. The two designs employ almost identical duct geometry, but integration of the vane row reduces the system length by 21%. Due to successful matching of the static pressure field, the upstream influence of the integrated vane row is minimal and the rotor performance is unchanged. Similarly the flow development within both S-shaped ducts is similar such that the circumferentially averaged profiles at duct exit are almost identical, and the operation of a downstream component would be unaffected. Overall system loss remains nominally unchanged despite the inclusion of lean and sweep and a reduction in system length. Finally, the numerical design predictions show good agreement with the experimental data thereby successfully validating the design process.

An Advanced Single-Stage Turbine Facility for Investigating Non-Axisymmetric Contoured Endwalls in the Presence of Purge Flow

2019 ◽  
Author(s):  
Robin R. Jones ◽  
Oliver J. Pountney ◽  
Bjorn L. Cleton ◽  
Liam E. Wood ◽  
B. Deneys J. Schreiner ◽  
...  
Keyword(s):  
Gas Turbines ◽  
Static Pressure ◽  
Optical Measurement ◽  
Guide Vane ◽  
Single Stage ◽  
Test Facility ◽  
Pressure Distributions ◽  
Mainstream Flow ◽  
Purge Flow ◽  
Nozzle Guide

Abstract In modern gas turbines, endwall contouring (EWC) is employed to modify the static pressure field downstream of the vanes and minimise the growth of secondary flow structures developed in the blade passage. Purge flow (or egress) from the upstream rim-seal interferes with the mainstream flow, adding to the loss generated in the rotor. Despite this, EWC is typically designed without consideration of mainstream-egress interactions. The performance gains offered by EWC can be reduced, or in the limit eliminated, when purge air is considered. In addition, EWC can result in a reduction in sealing effectiveness across the rim seal. Consequently, industry is pursuing a combined design approach that encompasses the rim-seal, seal-clearance profile and EWC on the rotor endwall. This paper presents the design of, and preliminary results from a new single-stage axial turbine facility developed to investigate the fundamental fluid dynamics of egress-mainstream flow interactions. To the authors’ knowledge this is the only test facility in the world capable of investigating the interaction effects between cavity flows, rim seals and EWC. The design of optical measurement capabilities for future studies, employing volumetric velocimetry and planar laser induced fluorescence are also presented. The fluid-dynamically scaled rig operates at benign pressures and temperatures suited to these techniques and is modular. The facility enables expedient interchange of EWC (integrated into the rotor bling), blade-fillet and rim-seals geometries. The measurements presented in this paper include: gas concentration effectiveness and swirl measurements on the stator wall and in the wheel-space core; pressure distributions around the nozzle guide vanes at three different spanwise locations; pitchwise static pressure distributions downstream of the nozzle guide vane at four axial locations on the stator platform.

An experimental, Aerodynamic Evaluation of Design Choices for a Low-Pressure Compressor Transition Duct

2021 ◽  
pp. 1-42
Author(s):  
A Duncan Walker ◽  
Ian Mariah ◽  
Chris Hall
Keyword(s):  
Guide Vane ◽  
Integrated Design ◽  
Axial Compressor ◽  
Low Pressure ◽  
Flow Coefficient ◽  
Low Flow ◽  
Test Facility ◽  
Minimum Length ◽  
Complex Flow ◽  
Fully Integrated

Abstract The S-shaped duct which transfers flow from the low-pressure fan to the engine core in large civil turbofans presents a challenging problem. Aerodynamically it has a spatially and temporarily varying inlet flow combined with a complex flow field which develops under the combined influence of pressure gradients and streamline curvature. It must also accommodate the transfer of structural loads and services across the main gas path. This necessitates the use of structural vanes which can compromise aerodynamics, introduce unwanted component interactions, and erode performance. This must all be achieved with minimum length/weight and without flow separation. This paper presents a comprehensive aerodynamic evaluation of three design options for a transition duct containing (i) a long-chord, structural compressor outlet guide vane, (ii) a aerodynamically optimal but non-structural outlet guide vane in conjunction with a small number of load bearing struts and (iii) a fully integrated outlet guide vane and strut design. Evaluation was performed using a low-speed test facility incorporating a 1½ stage axial compressor and engine representative transition duct. Measured data suggest that all the options were viable. However, the aerodynamic vane and discrete struts produced the lowest system loss with the other two options being comparable. The performance of the structural vane was sensitive to off-design conditions producing a notable increase in loss at a low flow coefficient. The optimized aerodynamic vanes were much less sensitive to off-design conditions whilst the fully integrated design showed only very small changes in loss.

The Influence of Fan Root Flow on the Aerodynamics of a Low-Pressure Compressor Transition Duct

2019 ◽  
Author(s):  
A. D. Walker ◽  
I. Mariah ◽  
D. Tsakmakidou ◽  
H. Vadhvana ◽  
C. Hall
Keyword(s):  
Gas Turbines ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Pressure Profile ◽  
Low Pressure ◽  
Secondary Flows ◽  
Test Facility ◽  
Inlet Flow ◽  
Flat Profile ◽  
Bypass Ratio

Abstract To reduce fuel-burn and CO2 emissions from aero gas turbines there is a drive towards very-high bypass ratio and smaller ultra-high-pressure ratio core engine technologies. However, this makes the design of the ducts connecting various compressor spools more challenging as the higher required radius change increases their aerodynamic loading. This is exacerbated for the duct which feeds the engine core as it must accept the relatively low-quality flow produced by the fan root. This is characterised by a hub-low pressure profile and large secondary flow structures which will inevitably increase loss and the likelihood of flow separation. Additionally, the desire for shorter, lighter nacelles means that the engine intake may be unable provide a uniform inlet flow to the fan when the aircraft is at an angle of attack or subject to cross winds. Any inlet distortion this generates can also further degrade the quality of the flow entering the core of the engine. This paper uses a combination of experiments and CFD to examine the effects of the inlet flow on the aerodynamics of an engine section splitter and transition duct designed to feed the low-pressure spool of a high bypass ratio turbofan. A fully annular test facility incorporating a 1½ stage axial compressor was used to compare the system performance of a rotor that produced a nominally flat profile with one that had a notably hub deficient flow. A RANS CFD model, employing a mixing plane between the rotor and Engine Section Stator (ESS) and a Reynolds Stress turbulence model, was then validated and used to further investigate the effects of increased inlet boundary layer thickness and bulk swirl distortion at rotor inlet. Overall, changes to the inlet condition were seen to have a surprisingly small effect on the flow at duct exit — i.e. the flow presented to the downstream compressor. Changes to the inlet did, however, generate increased secondary flows and degrade the performance of the ESS. This resulted in notably increased total pressure loss; in excess of 12% for the hub-low inlet and in excess of 30% at high inlet swirl where the flow in the ESS separated. However, the increased ESS wake structures, and the enhanced mixing, delayed separation in the duct suggesting that, overall the design was reasonably robust, albeit with a significant penalty in system loss.

A New Test Facility to Investigate Film Cooling on a Non-Axisymmetric Contoured Turbine Endwall: Part I — Introduction and Aerodynamic Measurements

2015 ◽  
Author(s):  
Franz Puetz ◽  
Johannes Kneer ◽  
Achmed Schulz ◽  
Hans-Joerg Bauer
Keyword(s):  
Heat Transfer ◽  
Flow Field ◽  
Pressure Distribution ◽  
Film Cooling ◽  
Gas Turbines ◽  
Guide Vane ◽  
Leading Edge ◽  
Suction Side ◽  
Test Facility ◽  
Endwall Contouring

An increased demand for lower emission of stationary gas turbines as well as civil aircraft engines has led to new, low emission combustor designs with less liner cooling and a flattened temperature profile at the outlet. As a consequence, the heat load on the endwall of the first nozzle guide vane is increased. The secondary flow field dominates the endwall heat transfer, which also contributes to aerodynamic losses. A promising approach to reduce these losses is non-axisymmetric endwall contouring. The effects of non-axisymmetric endwall contouring on heat transfer and film cooling are yet to be investigated. Therefore, a new cascade test rig has been set up in order to investigate endwall heat transfer and film cooling on both a flat and a non-axisymmetric contoured endwall. Aerodynamic measurements that have been made prior to the upcoming heat transfer investigation are shown. Periodicity and detailed vane Mach number distributions ranging from 0 to 50% span together with the static pressure distribution on the endwall give detailed information about the aerodynamic behavior and influence of the endwall contouring. The aerodynamic study is backed by an oil paint study, which reveals qualitative information on the effect of the contouring on the endwall flow field. Results show that the contouring has a pronounced effect on vane and endwall pressure distribution and on the endwall flow field. The local increase and decrease of velocity and the reduced blade loading towards the endwall is the expected behavior of the 3d contouring. So are the results of the oil paint visualization, which show a strong change of flow field in the leading edge region as well as that the contouring delays the horse shoe vortex hitting the suction side.

scholarly journals The Development of Supersonic Axial Compressor Boost Stages for Small Gas Turbines

1969 ◽  
Author(s):  
C. H. Muller ◽  
A. Sabatiuk
Keyword(s):  
Gas Turbines ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Performance Goals ◽  
Ample Evidence ◽  
Rotor Type ◽  
Highly Loaded ◽  
Percent Efficiency

The axial supersonic compressors of the “shock-in-rotor” type are under development for application to small gas turbines. A passage flow approach and passage criteria were used to design and develop the airfoils for the highly loaded rotor and stator blading of these 4 lb/sec machines. The overall stage performance values demonstrated to date are 2.06:1 pressure ratio at 78 percent adiabatic efficiency and 2.56:1 at 74.4 percent efficiency. The loss data and static pressure rise measured for the rotors and exit stators provide ample evidence that the higher design performance goals can be met.

A New Test Facility for Investigating the Interaction Between Swirl Flow and Wall Cooling Films in Combustors

2009 ◽  
Author(s):  
B. Wurm ◽  
A. Schulz ◽  
H.-J. Bauer
Keyword(s):  
Gas Turbines ◽  
Static Pressure ◽  
Thermal Analyses ◽  
Thermal Studies ◽  
Swirl Flow ◽  
Operating Conditions ◽  
Test Facility ◽  
Wall Region ◽  
Test Rig ◽  
Combustor Liner

Swirl stabilization of flames is typically used in combustors of aero engines and gas turbines for power generation. In the near wall region of the combustor liner, the swirling flow interacts in a very particular way with wall cooling films. This interaction and its effect on the local wall cooling performance gave reason to design and commission a new atmospheric test rig for detailed aerodynamic and thermal studies. The new test rig includes three burners in a planar arrangement. Special emphasis was placed on the simulation of realistic operating conditions as Reynolds number and temperature ratio. The liner cooling and the formation of a starter cooling film can be independently controlled. The rectangular flow channel is equipped with large windows to allow for laser optical diagnostics like PIV and 3-component LDA. The thermal analyses are based on highly resolved temperature mappings of the cooled surface utilizing infrared thermography. First experimental results are presented in terms of static pressure distributions on the combustor liner and PIV contour plots of the swirl flow. The static pressure pattern corresponds well to the up wash and downwash regions of the swirl flow.

scholarly journals The Influence of Blade Leaning on the Performance of an Integrated OGV-Diffuser System

1991 ◽  
Author(s):  
J. F. Carrotte ◽  
S. J. Stevens ◽  
A. P. Wray
Keyword(s):  
Static Pressure ◽  
Guide Vane ◽  
Integrated Design ◽  
Outer Wall ◽  
Flow Conditions ◽  
Trailing Edge ◽  
Outlet Angle ◽  
Lean Angle ◽  
Wall Flow ◽  
Inlet Conditions

An experimental investigation has been carried out to study the performance of an integrated design of compressor outlet guide vane and combustor pre-diffuser system. The trailing edge of each OGV was located within the outwardly canted diffuser by a distance equal to 27% of the diffuser axial length. In order to obtain representative inlet conditions a rotor, providing a fully sheared velocity profile and an air outlet angle of approximately 40°, was located upstream of the OGVs. Compared with the measured performance when the trailing edge of each OGV was located at diffuser inlet, a small increase in total pressure loss and a corresponding decrease in static pressure recovery was observed for the shortened system. This change in performance reflected a deterioration in the flow conditions along the outer wall, with transitory stalling of the flow being observed at diffuser exit. By leaning the blades in a circumferential direction through angles of 10° and 15° the outer wall flow conditions could be progressively improved, although at the largest angle tested stalling of the flow occurred at the hub of each OGV. However, at a lean angle of 10° the performance, in terms of loss and flow stability, could be virtually restored to the levels obtained when the trailing edge of each OGV was located radially at diffuser inlet.

An Advanced Single-Stage Turbine Facility for Investigating Nonaxisymmetric Contoured Endwalls in the Presence of Purge Flow

2019 ◽  
Vol 141 (12) ◽  
Author(s):  
Robin R. Jones ◽  
Oliver J. Pountney ◽  
Bjorn L. Cleton ◽  
Liam E. Wood ◽  
B. Deneys J. Schreiner ◽  
...  
Keyword(s):  
Gas Turbines ◽  
Static Pressure ◽  
Optical Measurement ◽  
Single Stage ◽  
Test Facility ◽  
Pressure Distributions ◽  
Mainstream Flow ◽  
Planar Laser ◽  
Flow Interactions ◽  
Purge Flow

Abstract In modern gas turbines, endwall contouring (EWC) is employed to modify the static pressure field downstream of the vanes and minimize the growth of secondary flow structures developed in the blade passage. Purge flow (or egress) from the upstream rim-seal interferes with the mainstream flow, adding to the loss generated in the rotor. Despite this, EWC is typically designed without consideration of mainstream–egress interactions. The performance gains offered by EWC can be reduced, or in the limit eliminated, when purge air is considered. In addition, EWC can result in a reduction in sealing effectiveness across the rim seal. Consequently, industry is pursuing a combined design approach that encompasses the rim-seal, seal-clearance profile, and EWC on the rotor endwall. This paper presents the design of and preliminary results from a new single-stage axial turbine facility developed to investigate the fundamental fluid dynamics of egress–mainstream flow interactions. To the authors' knowledge, this is the only test facility in the world capable of investigating the interaction effects between cavity flows, rim seals, and EWC. The design of optical measurement capabilities for future studies, employing volumetric velocimetry (VV) and planar laser-induced fluorescence (PLIF), is also presented. The fluid-dynamically scaled rig operates at benign pressures and temperatures suited to these techniques and is modular. The facility enables expedient interchange of EWC (integrated into the rotor bling), blade-fillet and rim-seal geometries. The measurements presented in this paper include: gas concentration effectiveness and swirl measurements on the stator wall and in the wheel-space core; pressure distributions around the nozzle guide vanes (NGV) at three different spanwise locations; pitchwise static pressure distributions downstream of the NGV at four axial locations on the stator platform.

Investigation of Axial Spacing and Effect of Interface Location in a 1.5 Stage Transonic Axial Compressor

2015 ◽  
Author(s):  
Jaeho Choi ◽  
Taewoo Choi
Keyword(s):  
Shock Waves ◽  
Gas Turbines ◽  
Guide Vane ◽  
Axial Compressor ◽  
Inlet Guide Vane ◽  
3 Dimensional ◽  
Fluid Analysis ◽  
Interface Location ◽  
Transonic Axial Compressor ◽  
Transonic Rotor

An axial compressor that is used for gas turbines has to meet axial length constraints as well as performance requirements. Therefore, the axial spacing between the stator and the rotor is one of the important design variables in the compressor design. Thus, a 3-dimensional fluid analysis is usually conducted to predict the performance of the compressor and to determine axial spacing in the preliminary design phase. The length of the inlet computational domain, which is influenced by shock waves from the transonic rotor, affects the compressor performance in the computational fluid analysis. This paper deals with a transonic axial compressor stage which consists of an inlet guide vane, a rotor and a stator which are designed as part of a high pressure compressor in a gas turbine. The purpose of this paper is to investigate the axial spacing between an IGV and a rotor, and effects of interface location in the spacing on the performance prediction of the axial compressor with the 3-dimensional fluid analysis. A multi-stage transonic axial compressor with an inlet guide vane was designed. Three dimensional Reynolds-averaged Navier-Stokes equation was used for the numerical analysis and the k-ω SST turbulence model was used to analyze the fluid flow in the transonic axial compressor. Numerical analyses were carried out at three interface locations in the axial spacing between the IGV and the rotor for two axial spacings in order to investigate the effect of interface location limiting the upstream flow of shock waves by the transonic rotor. The results show that the flow characteristics upstream of the rotor depend on the interface location. The detailed results on the performance and the flow fields are shown and discussed in the paper.

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