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

2021 ◽  
Author(s):  
Dimitra Tsakmakidou ◽  
Ian Mariah ◽  
A Duncan Walker ◽  
Chris Hall ◽  
Harry Simpson
Keyword(s):  
Boundary Layer ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Cumulative Effect ◽  
Poor Quality ◽  
Secondary Flows ◽  
Flow Structures ◽  
Momentum Exchange ◽  
Fuel Burn ◽  
Length Reduction

Abstract The need to reduce fuel-burn and emissions, is pushing turbofan engines towards geared architectures with higher bypass ratios and small ultra-high-pressure ratio cores. However, this increases the radial offset between compressor spools leading to a more challenging design for compressor transition ducts. For the duct connecting the fan to the engine core this is further complicated by poor-quality flow generated at the fan hub which is characterised by low total pressure and large rotating secondary flow structures. This paper presents an experimental evaluation of a new rotor designed to produce these larger flow structures and examines their effect on the performance of an engine sector stators (ESS) and compressor transition duct. Aerodynamic data were collected via five-hole probes, for time-averaged pressures and velocities and phase-locked hot-wire anemometry to capture the rotating secondary flows. The data showed that larger structures promoted mixing through the ESS increasing momentum exchange between the core and boundary layer flows. Measurements within the duct showed a continued reduction in the hub boundary layer suggesting the duct had moved further from separation. Consequently, an aggressive duct with 12.5% length reduction was designed and tested and measurements confirmed the duct remained fully attached. Total pressure loss was slightly increased over the ESS, but this was offset by reduced loss in the duct due to improved flow quality. Overall, this length reduction represents a significant cumulative effect in reduced fuel-burn and emissions over the life of an engine.

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.

Aerodynamic Design and Analysis of a High Pressure Ratio Aspirated Compressor Stage

2000 ◽  
Author(s):  
Ali A. Merchant ◽  
Mark Drela ◽  
Jack L. Kerrebrock ◽  
John J. Adamczyk ◽  
Mark Celestina
Keyword(s):  
Boundary Layer ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Adverse Pressure Gradient ◽  
Compressor Stage ◽  
High Pressure Ratio ◽  
Total Pressure Ratio ◽  
Isentropic Efficiency

The pressure ratio of axial compressor stages can be significantly increased by controlling the development of blade and endwall boundary layers in regions of adverse pressure gradient by means of boundary layer suction. This concept is validated and demonstrated through the design and analysis of a unique aspirated compressor stage which achieves a total pressure ratio of 3.5 at a tip speed of 1500 ft/s. The aspirated stage was designed using an axisymmetric through-flow code coupled with a quasi three-dimensional cascade plane code with inverse design capability. Validation of the completed design was carried out with three-dimensional Navier-Stokes calculations. Spanwise slots were used on the rotor and stator suction surfaces to bleed the boundary layer with a total suction requirement of 4% of the inlet mass flow. Additional bleed of 3% was also required on the hub and shroud near shock impingement locations. A three-dimensional viscous evaluation of the design showed good agreement with the quasi three-dimensional design intent, except in the endwall regions. The three-dimensional viscous analysis predicted a mass averaged total pressure ratio of 3.7 at an isentropic efficiency of 93% for the rotor, and a mass averaged total pressure ratio of 3.4 at an isentropic efficiency of 86% for the stage.

Investigation of the Dihedral Angle Effect on the Boundary Layer Development Using Special-Shaped Expansion Pipes

2018 ◽  
Author(s):  
Teng Fei ◽  
Lucheng Ji ◽  
Weilin Yi
Keyword(s):  
Boundary Layer ◽  
Boundary Layers ◽  
Dihedral Angle ◽  
Adverse Pressure Gradient ◽  
Secondary Flows ◽  
Flow Structures ◽  
Wall Surface ◽  
Geometric Structures ◽  
Corner Region ◽  
End Wall

The corners between the blades and end walls are common geometric structures in turbomachinery, where boundary layers on the blade and end wall surface interact with each other. This boundary layer interaction enlarges the region of low momentum fluid which leads to the boundary layers grow thicker at the corner region. Then the corner separation is likely to occur, and even worse by the adverse pressure gradient along the streamwise as well as secondary flows along the pitchwise. The key issue to design the geometric structures of the corner region is to control the dihedral angle between the blade and end wall surface. However, from the current published literature, few researchers have studied the influence of dihedral angle on the flow structures at the corner region in detail. In this paper, a series of expansion pipes with different cross sections which represent different dihedral angles are simulated. Then, some useful conclusions about how the dihedral angle affects the flow structures at the corner region are drawn. Moreover, a new method to predict the boundary layer thickness at the corner region is introduced, and the predicted results are in good agreement with simulation results.

Flow Measurements in a Fishtail Diffuser With Strong Curvature

1999 ◽  
Vol 121 (2) ◽  
pp. 410-417 ◽  
Author(s):  
M. I. Yaras
Keyword(s):  
Boundary Layer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Static Pressure ◽  
Secondary Flows ◽  
Total Pressure Loss ◽  
Cross Sectional ◽  
Inlet Flow ◽  
Inlet Conditions ◽  
Strong Curvature

The paper presents detailed measurements of the incompressible flow development in a large-scale 90 deg curved diffuser with strong curvature and significant streamwise variation in cross-sectional aspect ratio. The flow path approximates the so-called fishtail diffuser utilized on small gas turbine engines for the transition between the centrifugal impeller and the combustion chamber. Two variations of the inlet flow, differing in boundary layer thickness and turbulence intensity, are considered. Measurements consist of three components of velocity, static pressure and total pressure distributions at several cross-sectional planes throughout the diffusing bend. The development and mutual interaction of multiple pairs of streamwise vortices, redistribution of the streamwise flow under the influence of these vortices, the resultant streamwise variations in mass-averaged total-pressure and static pressure, and the effect of inlet conditions on these aspects of the flow are examined. The strengths of the vortical structures are found to be sensitive to the inlet flow conditions, with the inlet flow comprising a thinner boundary layer and lower turbulence intensity yielding stronger secondary flows. For both inlet conditions a pair of streamwise vortices develop rapidly within the bend, reaching their peak strength at about 30 deg into the bend. The development of a second pair of vortices commences downstream of this location and continues for the remainder of the bend. Little evidence of the first vortex pair remains at the exit of the diffusing bend. The mass-averaged total pressure loss is found to be insensitive to the range of inlet-flow variations considered herein. However, the rate of generation of this loss along the length of the diffusing bend differs between the two test cases. For the case with the thinner inlet boundary layer, stronger secondary flows result in larger distortion of the streamwise velocity field. Consequently, the static pressure recovery is somewhat lower for this test case. The difference between the streamwise distributions of measured and ideal static pressure is found to be primarily due to total pressure loss in the case of the thick inlet boundary layer. For the thin inlet boundary layer case, however, total pressure loss and flow distortion are observed to influence the pressure recovery by comparable amounts.

A-10 Wing Leading Edge Effects on Engine Stability: Part 2 — Identifying Total Pressure Distortion Sources and the Effects on Engine/Airframe Integration

2015 ◽  
Author(s):  
Brian A. Binkley ◽  
Grant T. Patterson ◽  
Jerome C. Jenkins
Keyword(s):  
Total Pressure ◽  
Pressure Ratio ◽  
Leading Edge ◽  
Maintenance Cost ◽  
Flow Structures ◽  
Inlet Distortion ◽  
Test And Evaluation ◽  
Engine Stability ◽  
The Many ◽  
Wing Stall

The A-10 aircraft has fuselage mounted engines with inlets just above the rear of the wing. A deployable slat system is used on the A-10 to delay wing stall directly in front of the engines. Wing stall can lead to high inlet distortion and ultimately engine stall for this aircraft. Many alternate wing leading-edge designs were recently considered for replacement of the slat system to reduce maintenance cost, reduce system complexity and increase system reliability. Fifteen potential wing leading-edge proposals for replacing the A-10 slat system were evaluated through test and analysis. Performance of the wing leading-edge candidates was characterized by the effect of inlet/engine distortion on loss of stability pressure ratio (ΔPRS) for the TF-34 engine fan and compressor. The many protuberances and non-aerodynamic shapes of the A-10 outer mold lines can generate flow structures that cause significant inlet/engine total pressure distortion. Thorough understanding of these flow structures and their impact on inlet/engine distortion was necessary to fully assess the performance of candidate wing leading edge configurations. The paper discusses the integrated test and evaluation tools and methods used to identify the sources of inlet/engine total pressure distortion and the associated impact to engine/airframe integration.

scholarly journals A Family of Designs for Aspirated Compressors

1998 ◽  
Author(s):  
Jack L. Kerrebrock ◽  
Mark Drela ◽  
Ali A. Merchant ◽  
Brian J. Schuler
Keyword(s):  
Boundary Layer ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Pressure Rise ◽  
Flow Path ◽  
Fluid Removal ◽  
The Family ◽  
Bypass Ratio ◽  
Total Pressure Ratio ◽  
Rapid Pressure

The performance of compressors can be enhanced by the judicious removal of the viscous boundary layer fluid from the flow path. Removal of the boundary layer fluid just prior to or in a region of rapid pressure rise, either at shock impingement or more generally at the point of rapid pressure rise on the suction surface of the blade, can give significant increases in the diffusion and therefore increase the work done per stage for a given blade speed. It also provides a thermodynamic benefit by removing the high-entropy fluid from the flow path. Design studies have been done using quasi 3-D viscous and 3-D Euler computational tools on a family of fan stages of varying tip speed that lake advantage of such viscous fluid removal. One stage in this family is a low tip speed fan stage designed to produce a pressure ratio of 1.5 at a tip speed of 700 ft/sec. Fan noise reductions resulting from the decrease in tangential Mach number, without sacrificing total pressure ratio, could make this design attractive for the fan of medium-bypass ratio engines. Another stage in the family would produce a total pressure ratio of 2.0 at a tip speed of 1000 ft/sec and could be very attractive as a fan stage on a lower bypass ratio engine or as a first stage of a low speed core compressor. The final stage in the family would achieve a pressure ratio of more than 3.0 at a tip speed of 1500 ft/sec and could be very attractive as a first stage of a core compressor, or as a fan for a military engine. A design for the suction passages to deal with the fluid removal has been completed for an experimental version of the 1.5 pressure ratio design. A tip shroud allows bleeding of the tip surface boundary layer from the rotor, and carries the fluid removed from the blade surfaces through the tip. One of these stages will be tested in the MIT Blowdown Compressor, serving a dual purpose: as a validation of the computational design process and as a test of the concept of aspirated compressors.

Non-Axisymmetric End Wall Profiling in Transonic Compressors—Part II: Design Study of a Transonic Compressor Rotor Using Non-Axisymmetric End Walls—Optimization Strategies and Performance

2009 ◽  
Author(s):  
Steffen Reising ◽  
Heinz-Peter Schiffer
Keyword(s):  
Secondary Flow ◽  
Total Pressure ◽  
Cfd Simulation ◽  
Pressure Ratio ◽  
Cross Flow ◽  
Secondary Flows ◽  
Transonic Compressor ◽  
End Wall ◽  
Total Pressure Ratio ◽  
Isentropic Efficiency

Secondary flows involving cross flow at high stage loading in modern axial compressors contribute significantly to efficiency limits. This paper summarizes an approach to control end wall flow using non-axisymmetric end walls. The challenge is to find the optimal non-axisymmetric end wall shape that results in the largest gain in performance. An automated multi-objective optimizer connected to a 3-D RANS flow solver was used to design the end wall contour. The process chain was applied to the rotor hub end wall of Configuration I of the Darmstadt Transonic Compressor. Several optimization strategies involving different objective functions to be minimized and the corresponding performances were compared. The parameters considered within the optimization process were isentropic stage efficiency, pressure loss in the rotor, throat area and secondary kinetic energy (SKE). A parameter variation was undertaken, leading to the following observations: Strong penalties on SKE at the rotor outlet and moderate penalties on isentropic efficiency, throat area and pressure ratio led to the best design. Isentropic efficiency could be raised by 0.12%, SKE at the rotor exit was reduced while the total pressure ratio of the stage remained constant. Strong penalties on efficiency and pressure ratio, a moderate one on throat area and a small one on SKE at the rotor outlet all led to a smaller increase in efficiency: 0.06%. On the other hand, a slight raise in the total pressure ratio could be achieved. A third optimization, eliminating the restriction on the throat area, was carried out to see which benefit in performance could be achieved without this geometrical restriction. Since the throat areas of all optimized geometries differ slightly from the datum value, an estimation was derived to see the extent to which the end wall profiling and cross section enlargement contribute to the improvements. Finally, a method to display secondary flows in turbomachinery is introduced. A second CFD simulation is used to calculate the primary flow where the hub end wall is defined as an Euler wall to avoid the end wall boundary layer and so eliminate the cause for some of the secondary flow mechanisms. This method clearly shows how the characteristics of secondary flow can be positively influenced by using non-axisymmetric end walls.

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.

Low-Speed Annular Cascade Tests of an Ultra-Highly Loaded Turbine With Tip Clearance: Part 1 — Near Design Incidence

1999 ◽  
Author(s):  
A. Yamamoto ◽  
E. Outa
Keyword(s):  
Flow Visualization ◽  
Total Pressure ◽  
Secondary Flows ◽  
Tip Clearance ◽  
Flow Structures ◽  
Turning Angle ◽  
Pressure Losses ◽  
Blade Tip ◽  
Annular Cascade ◽  
Highly Loaded

Annular cascade tests were carried out to study the performance of an ultra-highly loaded turbine cascade (UHLTC) with a design turning angle of 160 deg. The UHLTC is for applications to future high-temperature gas turbine engines. This paper describes details of the secondary flows and the associated total pressure losses of the UHLTC obtained at a test incidence of −2.7 deg. The cascade flows were measured with a small five-hole Pitot probe located at 21 traverse measurement planes upstream, inside and downstream of the UHLTC. From the measurements, detailed flow structures and the loss evolution process were analyzed. Flow visualization tests were also carried out to see more details of the flows on the blade surfaces, on the endwalls and in the blade tip gap. Various flow separations and various small vortices associated with the passage and leakage vortices, such as corner vortices and edge vortices, separation bubbles, and the associated reverse flows, were seen. These were clarified from various flow lines showing separation, attachment/reattachment and division of each flow. The results obtained from the flow visualization were compared with those from the traverse measurements. Large total pressure losses occur inside the cascade passage as well as downstream of the cascade. Various strong passage vortices, strong leakage vortex, strong swirling flows upstream and downstream of the cascade and their associated various flow separations, are the main causes of the loss generation. The coefficient of total pressure loss generated inside the cascade was 0.28 at the test near-design incidence. The actual turning angle of the flow from the cascade inlet and the cascade outlet was 146 deg. Some schematic drawings of the flow structures in the present UHLTC were also given. The basic flow structures did not differ significantly from those seen in the conventional cascades with much smaller turning angles, except for stronger passage vortices, larger internal loss and larger downstream mixing loss due to the very high turning angle of the UHLTC.

Coupled aeropropulsive design optimisation of a boundary-layer ingestion propulsor

2018 ◽  
Vol 123 (1259) ◽  
pp. 121-137 ◽  
Author(s):  
Justin S. Gray ◽  
Joaquim R. R. A. Martins
Keyword(s):  
Boundary Layer ◽  
Power Transmission ◽  
Pressure Ratio ◽  
Relative Size ◽  
Thermodynamic Cycle ◽  
Transmission Efficiency ◽  
Navier Stokes ◽  
Design Optimisation ◽  
Fuel Burn ◽  
Design Variables

AbstractAirframe–propulsion integration concepts that use boundary-layer ingestion (BLI) have the potential to reduce aircraft fuel burn. One concept that has been recently explored is NASA’s STARC-ABL aircraft configuration, which offers the potential for fuel burn reduction by using a turboelectric propulsion system with an aft-mounted electrically driven BLI propulsor. So far, attempts to quantify this potential fuel burn reduction have not considered the full coupling between the aerodynamic and propulsive performance. To address the need for a more careful quantification of the aeropropulsive benefit of the STARC-ABL concept, we run a series of design optimisations based on a fully coupled aeropropulsive model. A 1D thermodynamic cycle analysis is coupled to a Reynolds-averaged Navier–Stokes simulation to model the aft propulsor at a cruise condition and the effects variation in propulsor design on overall performance. A series of design optimisation studies are performed to minimise the required cruise power, assuming different relative sizes of the BLI propulsor. The design variables consist of the fan pressure ratio, static pressure at the fan face, and 311 variables that control the shape of both the nacelle and the fuselage. The power required by the BLI propulsor is compared with a podded configuration. The results show that the BLI configuration offers 6–9% reduction in required power at cruise, depending on assumptions made about the efficiency of power transmission system between the under-wing engines and the aft propulsor. Additionally, the results indicate that the power transmission efficiency directly affects the relative size of the under-wing engines and the aft propulsor. This design optimisation, based on computational fluid dynamics, is shown to be essential to evaluate current BLI concepts and provides a powerful tool for the design of future concepts.

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