A Numerical Investigation Into Secondary Flows at the Inlet to a Low-Pressure Compressor Transition Duct

2020 ◽  
Author(s):  
Dimitra Tsakmakidou ◽  
A. Duncan Walker ◽  
Chris Hall
Keyword(s):  
Boundary Layer ◽  
Layer Thickness ◽  
Gas Turbines ◽  
Boundary Layer Thickness ◽  
Pressure Ratio ◽  
Low Pressure ◽  
Secondary Flows ◽  
Test Facility ◽  
Rotor Loss ◽  
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 larger 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. This paper presents a numerical design study aimed at modifying an existing fully annular, low-speed test facility to produce larger, more representative rotor loss cores. The CFD domain comprises a single stage axial compressor and transition duct representative of the low-pressure spool in a high bypass ratio turbofan. The rotating and stationary frames were coupled using a sliding mesh and a uRANS approach with Reynolds Stress closure for the turbulence modelling. The methodology was first validated using existing experimental data before examining a number of parameters such as inlet boundary layer thickness, inlet swirl, tip gap, blade solidity, thickness, lean. It was found that a combination of an increased inlet hub boundary layer and a reduction in solidity sufficiently increased the hub loading and encouraged the development of significantly large loss cores. Preliminary CFD results also showed that the increased rotor loss cores promoted better mixing through the outlet guide which subsequently reduced hub boundary layer thickness and secondary flows. Despite incurring a larger total pressure loss this suggested that the transition duct moved further from stall and could therefore potentially be made shorter.

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.

scholarly journals The Influence of Fan Root Flow on the Aerodynamic of a Low-Pressure Compressor Transition Duct

2019 ◽  
Vol 142 (1) ◽  
Author(s):  
A. Duncan Walker ◽  
Ian Mariah ◽  
Dimitra Tsakmakidou ◽  
Hiren Vadhvana ◽  
Chris Hall
Keyword(s):  
Pressure Ratio ◽  
Axial Compressor ◽  
Pressure Profile ◽  
Low Pressure ◽  
Secondary Flows ◽  
Test Facility ◽  
Navier Stokes ◽  
High Pressure Ratio ◽  
Fuel Burn ◽  
Bypass Ratio

Abstract To reduce fuel-burn and emissions, there is a drive toward higher bypass ratio and smaller high-pressure ratio core engines. This makes the design of the ducts connecting compressor spools more challenging as the higher radius change increases aerodynamic loading. This is exacerbated at inlet to the engine core by fan root flow which is characterized by a hub-low-pressure profile and large secondary flow structures. Additionally, shorter, lighter nacelles mean that the intake may not provide a uniform inlet flow when the aircraft is at an angle of attack or subject to cross winds. Such inlet distortion can further degrade the flow entering the engine. A combination of experiments and computational fluid dynamics (CFD) has been used to examine the effects on the aerodynamics of an engine section splitter (ESS) and transition duct designed to feed the low-pressure spool of a high bypass ratio turbofan. A test facility incorporating a 1½ stage axial compressor was used to compare system performance for a flat rotor exit profile to one with a hub deficient flow. Validated Reynolds averaged Navier–Stokes (RANS) CFD was then used to further investigate the effects of increased inlet boundary layer thickness and bulk swirl distortion at rotor inlet. These changes were seen to have a surprisingly small effect on the flow at the duct exit. However, increased secondary flows were observed which degraded the performance of the ESS and significantly increased loss. Nevertheless, the enhanced mixing delayed separation in the duct suggesting that overall the design was reasonably robust albeit with increased system loss.

Experimental Investigation of Secondary Flows and Length Reduction for a Low-Pressure Compressor Transition Duct

2021 ◽  
Author(s):  
D. Tsakmakidou ◽  
I. Mariah ◽  
A. D. Walker ◽  
C. Hall ◽  
H. Simpson
Keyword(s):  
Boundary Layer ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Low Pressure ◽  
Secondary Flows ◽  
Test Facility ◽  
Momentum Exchange ◽  
Fuel Burn ◽  
Operational Life ◽  
Length Reduction

Abstract The need to reduce fuel-burn and CO2 emissions, is pushing turbofan engines towards geared architectures with very high bypass-ratios and small ultra-high-pressure ratio core engines. However, this increases the radial offset between compressor spools and leads to a more challenging design for the compressor transition ducts. To minimise weight, these ducts must achieve the radial turning in as short a length, but this leads to strong curvature induced pressure gradients, increased aerodynamic loading and likelihood of flow separation. For the duct connecting the low-pressure fan to the engine core this is further complicated by the poor-quality flow generated at the fan hub which is characterised by low total pressure and large rotating secondary flow structures. In a previous paper the authors numerically designed modifications to an existing test facility such that the rotor would produce these large structures. The current paper presents an experimental evaluation of the new rotor design and examines the effect of the increased loss cores on the performance of a set of engine sector stators (ESS) or outlet guide vanes (OGV) and an engine representative compressor transition duct. Aerodynamic data were collected via miniature five-hole probes, for the time-averaged pressure and velocity field, and phase-locked hot-wire anemometry to capture the rotating secondary flows. Analysis of the experimental data showed that these structures promoted mixing through the ESS increasing the momentum exchange between the core and boundary layer flows. Measurements within the duct showed a continued reduction in the hub-wall boundary layer suggesting that the duct has been moved further from separation. Consequently, a more aggressive duct with 12.5% length reduction was designed and tested with the data confirming that the more aggressive duct remained fully attached. Total pressure loss data suggested a slight increase in loss over the vane row but that was offset by a reduced loss in the duct due to improved flow quality and reduced length. Overall, the 12.5% length reduction represents a significant cumulative effect in terms of reduced fuel burn and CO2 over the operational life of an engine.

Freestream Flow Effects on Film Effectiveness and Heat Transfer Coefficient Augmentation for Compound Angle Shaped Holes

2017 ◽  
Author(s):  
Joshua B. Anderson ◽  
John W. McClintic ◽  
David G. Bogard ◽  
Thomas E. Dyson ◽  
Zachary Webster
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Layer Thickness ◽  
Film Cooling ◽  
Gas Turbines ◽  
Boundary Layer Thickness ◽  
Adiabatic Effectiveness ◽  
Compound Angle

The use of compound-angled shaped film cooling holes in gas turbines provides a method for cooling regions of extreme curvature on turbine blades or vanes. These configurations have received surprisingly little attention in the film cooling literature. In this study, a row of laid-back fanshaped holes based on an open-literature design, were oriented at a 45-degree compound angle to the approaching freestream flow. In this study, the influence of the approach flow boundary layer thickness and character were experimentally investigated. A trip wire and turbulence generator were used to vary the boundary layer thickness and freestream conditions from a thin laminar boundary layer flow to a fully turbulent boundary layer and freestream at the hole breakout location. Steady-state adiabatic effectiveness and heat transfer coefficient augmentation were measured using high-resolution IR thermography, which allowed the use of an elevated density ratio of DR = 1.20. The results show adiabatic effectiveness was generally lower than for axially-oriented holes of the same geometry, and that boundary layer thickness was an important parameter in predicting effectiveness of the holes. Heat transfer coefficient augmentation was highly dependent on the freestream turbulence levels as well as boundary layer thickness, and significant spatial variations were observed.

Secondary Losses in a Large Deflection Annular Turbine Cascade: Effect of the Entry Boundary Layer Thickness

1986 ◽  
Author(s):  
M. Govardhan ◽  
N. Venkatrayulu ◽  
D. Prithvi Raj
Keyword(s):  
Boundary Layer ◽  
Layer Thickness ◽  
Boundary Layer Thickness ◽  
Three Dimensional ◽  
Secondary Flows ◽  
Impulse Turbine ◽  
Flow Measurements ◽  
Local Loss ◽  
Trailing Edges ◽  
Annular Cascade

The paper presents the results of three dimensional flow measurements behind the trailing edges of an impulse turbine blade row of 120° deflection in an annular cascade. The entry boundary layer thickness was systematically varied on the hub and casing walls separately and its effect on secondary flows and losses is investigated. With the increase of entry boundary layer thickness, it has been found that (i) the contours of local loss coefficient show that the magnitude of the hub loss core increased, (ii) the loss cores near the hub and casing wall are convected away from the walls, (iii) the spanwise variation of the pitchwise averaged losses indicate that the position of large loss peak near the hub wall remains the same, but the magnitude of the loss increases, (iv) the exit static pressure increases and the exit velocity in general decreases, (v) the degree of underturning of flow increases and (vi) the net secondary losses do not change appreciably.

scholarly journals Wind resource assessment in heterogeneous terrain

2017 ◽  
Vol 375 (2091) ◽  
pp. 20160109 ◽  
Author(s):  
C. Vanderwel ◽  
M. Placidi ◽  
B. Ganapathisubramani
Keyword(s):  
Boundary Layer ◽  
Layer Thickness ◽  
Boundary Layer Thickness ◽  
Particle Image Velocimetry Data ◽  
Secondary Flows ◽  
Heterogeneous Terrain ◽  
Power Extraction ◽  
Link Type ◽  
First Case ◽  
Wind Resource

High-resolution particle image velocimetry data obtained in rough-wall boundary layer experiments are re-analysed to examine the influence of surface roughness heterogeneities on wind resource. Two different types of heterogeneities are examined: (i) surfaces with repeating roughness units of the order of the boundary layer thickness (Placidi & Ganapathisubramani. 2015 J. Fluid Mech. 782 , 541–566. ( doi:10.1017/jfm.2015.552 )) and (ii) surfaces with streamwise-aligned elevated strips that mimic adjacent hills and valleys (Vanderwel & Ganapathisubramani. 2015 J. Fluid Mech. 774 , 1–12. ( doi:10.1017/jfm.2015.228 )). For the first case, the data show that the power extraction potential is highly dependent on the surface morphology with a variation of up to 20% in the available wind resource across the different surfaces examined. A strong correlation is shown to exist between the frontal and plan solidities of the rough surfaces and the equivalent wind speed, and hence the wind resource potential. These differences are also found in profiles of and (where U is the streamwise velocity), which act as proxies for thrust and power output. For the second case, the secondary flows that cause low- and high-momentum pathways when the spacing between adjacent hills is beyond a critical value result in significant variations in wind resource availability. Contour maps of and show a large difference in thrust and power potential (over 50%) between hills and valleys (at a fixed vertical height). These variations do not seem to be present when adjacent hills are close to each other (i.e. when the spacing is much less than the boundary layer thickness). The variance in thrust and power also appears to be significant in the presence of secondary flows. Finally, there are substantial differences in the dispersive and turbulent stresses across the terrain, which could lead to variable fatigue life depending on the placement of the turbines within such heterogeneous terrain. Overall, these results indicate the importance of accounting for heterogeneous terrain when siting individual turbines and wind farms. This article is part of the themed issue ‘Wind energy in complex terrains’.

scholarly journals Effects of spanwise spacing on large-scale secondary flows in rough-wall turbulent boundary layers

2015 ◽  
Vol 774 ◽  
Author(s):  
Christina Vanderwel ◽  
Bharathram Ganapathisubramani
Keyword(s):  
Boundary Layer ◽  
Layer Thickness ◽  
Boundary Layers ◽  
Boundary Layer Thickness ◽  
Large Scale ◽  
Turbulent Boundary Layers ◽  
Surface Flow ◽  
Secondary Flows ◽  
Flow Conditions ◽  
Roughness Elements

Large-scale secondary flows can sometimes appear in turbulent boundary layers formed over rough surfaces, creating low- and high-momentum pathways along the surface (Barros & Christensen, J. Fluid Mech., vol. 748, 2014, R1). We investigate experimentally the dependence of these secondary flows on surface/flow conditions by measuring the flows over streamwise strips of roughness with systematically varied spanwise spacing. We find that the large-scale secondary flows are accentuated when the spacing of the roughness elements is roughly proportional to the boundary layer thickness ${\it\delta}$, and do not appear for cases with finer spacing. Cases with coarser spacing also generate ${\it\delta}$-scale secondary flows with tertiary flows in the spaces in between. These results show that the ratio of the spanwise length scale of roughness heterogeneity to the boundary layer thickness is a critical parameter for the occurrence of these secondary motions in turbulent boundary layers over rough walls.

EFFECTS OF BOUNDARY-LAYER THICKNESS ON UNSTEADY FLOW CHARACTERISTICS INSIDE OPEN CAVITIES AT SUBSONIC AND TRANSONIC SPEEDS

2012 ◽  
Vol 19 ◽  
pp. 206-213
Author(s):  
DANG-GUO YANG ◽  
JIAN-QIANG LI ◽  
ZHAO-LIN FAN ◽  
XIN-FU LUO
Keyword(s):  
Boundary Layer ◽  
Unsteady Flow ◽  
Layer Thickness ◽  
Boundary Layer Thickness ◽  
Sound Pressure ◽  
Flow Characteristics ◽  
Open Cavity ◽  
Aerodynamic Noise ◽  
Sustained Oscillation ◽  
Cavity Depth

An experimental study was conducted in a 0.6m by 0.6m wind-tunnel to analyze effects of boundary-layer thickness on unsteady flow characteristics inside a rectangular open cavity at subsonic and transonic speeds. The sound pressure level (SPL) distributions at the centerline of the cavity floor and Sound pressure frequency spectrum (SPFS) characteristics on some measurement positions presented herein was obtained with cavity length-to-depth ratio (L/D) of 8 over Mach numbers (Ma) of 0.6 and 1.2 at a Reynolds numbers (Re) of 1.23 × 107 and 2.02 × 107 per meter under different boundary-layer thickness to cavity-depth ratios (δ/D). The experimental angle of attack, yawing and rolling angles were 0°. The results indicate that decrease in δ/D leads to severe flow separation and unsteady pressure fluctuation, which induces increase in SPL at same measurement points inside the cavity at Ma of 0.6. At Ma of 1.2, decrease in δ/D results in enhancing compressible waves. Generally, decrease in δ/D induces more flow self-sustained oscillation frequencies. It also makes severer aerodynamic noise inside the open cavity.

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