Updates / published 19.06.24
Papers being presented at Turbomachinery Technical Conference & Exposition 2024
Dissipation in Skewed Boundary Layers
Transitional Flows Within Cooled HP Turbines
Data-driven Radial Compressor Design Space Mapping
Design and Analysis of a Liftfan for eVTOL Aircraft
Nature of Transonic Compressor Flow: Imortance of 3D Athroat/Ainlet
Reverse thrust aerodynamics of variable pitch fans with inlet distortion
A rapid test facility for transonic compressors and fans
Effect of blade damage on low pressure ratio fan windmill aerodynamics
Dissipation in Skewed Boundary Layers
Robert Peacock, Graham Pullan, Masha Folk
GT2024-125855
Session: 32-03 Boundary Layer & Transition
Date: Monday, June 24, 1:30 PM – 3:30 PM
Abstract: The entropy generation within a turbulent, collateral boundary layer is well understood and is characterised by a dissipation coefficient, 𝐶𝑑 . However, it is common for the transverse pressure gradients in turbomachines to create highly skewed boundary layers, where the velocity varies in direction as well as magnitude. A combined experimental and high-fidelity computational approach is used to quantify the effect of skew on the dissipation coefficient for the first time. At a nominal condition of 14 degrees of skew and 𝑅𝑒𝜃 of 1000, the increase in dissipation coefficient is 20% as determined from direct numerical simulation, and 28% from experimental measurements, relative to the collateral boundary layer. Experimental data over a range of skew angles and 𝑅𝑒𝜃 values show that 𝐶𝑑 increases approximately linearly with skew so that, at a skew of 25°, loss is 70% greater than in the collateral boundary layer. The implications for loss estimation are examined by evaluating boundary layer loss, in the Harrison turbine cascade, with and without the influence of skew on 𝐶𝑑. By accounting for the skew in the boundary layer, a new proposed model has been used to calculate the loss coefficient in the Harrison cascade to within 4% of the experimentally measured value.
Transitional Flows Within Cooled HP Turbines
Dr C.Clark, Dr N.Atkins, Dr M.Stokes, Dr R.Vazquez
GT2024-124918
Session: 32-03 Boundary Layer & Transition
Date: Monday, June 24, 1:30 PM – 3:30 PM
Abstract: The transitional landscape within the HP turbine features multiple routes to turbulent flow. Free stream turbulence, surface roughness and film cooling ejections all provide bypass transition mechanisms. In reality the mechanisms are often at play simultaneously. Current understanding of this transitional landscape is often simplified, either with mechanisms assumed independent or even dominant and always active. Experimental data has been taken in a variable density transonic linear cascade featuring representative HP blade sections. Testing has been undertaken with variable inlet turbulence, surface roughness and film cooling configurations. Sweeps of Reynolds number enable the onset of bypass transition to be observed as both loss measurements and using IR interferometry to infer boundary layer states and features. The data is used to show the independent effects of each transition route as well as demonstrations of their superposition under certain circumstances. The transitional landscape within the HP turbine features multiple routes to turbulent flow. Free stream turbulence, surface roughness and film cooling ejections all provide bypass transition mechanisms. In reality the mechanisms are often at play simultaneously. Current understanding of this transitional landscape is often simplified, either with mechanisms assumed independent or even dominant and always active. Experimental data has been taken in a variable density transonic linear cascade featuring representative HP blade sections. Testing has been undertaken with variable inlet turbulence, surface roughness and film cooling configurations. Sweeps of Reynolds number enable the onset of bypass transition to be observed as both loss measurements and using IR interferometry to infer boundary layer states and features. The data is used to show the independent effects of each transition route as well as demonstrations of their superposition under certain circumstances.
The impact of manufacturing variations on the aerothermal performance of high-pressure turbine blade shrouds
Bram Hulhoven, John D. Coull, Dougal Jackson and Nicholas R. Atkins
GT2024-125688
Session: 13-07 Cooling Methods and Radiation Effects
Date: Monday, June 24, 4:00 PM – 5:30 PM
Abstract: High pressure turbine blade (HPTB) shrouds suffer manufacturing variations in both platform alignment and inter-platform gap width. Compared to hub endwalls, the aerothermal effects of shroud platform steps and gaps has had little attention, which introduces uncertainty in the sentencing of such manufacturing variations. This paper presents a shroud step sentencing correlation developed using a quasi-2D (Q2D) model of a shroud endwall step. The use of a Q2D model follows from the study of a 3D steady RANS simulation matrix of engine-representative platform steps and gap widths, based on a sample of scanned HPTB castings and finished parts. This study showed that the aftchord shroud step flow is Q2D and resembles canonical step flow with enhanced heat transfer at the reattachment point. The shroud step sentencing correlation is tested on the platform steps in the simulation matrix giving prediction errors below 20% for the vast majority of cases. Finally, the correlation is tuned using experimental data to mitigate the uncertainty associated with RANS simulations of separated flows.
Data-driven Radial Compressor Design Space Mapping
James Brind
GT2024-123250
Session: 37-02 Radial Turbomachinery Optimization
Date: Wednesday, June 26, 1:30 PM – 3:30 PM
Abstract: Estimates of turbomachinery performance trends inform system-level compromises during preliminary design. Existing empirical correlations for efficiency use limited experimental data, while analytical loss models require calibration to yield predictive results. From a set of 3708 radial compressor computations, this paper maps efficiency as a function of mean-line aerodynamics, and determines the governing loss mechanisms. An open-source turbomachinery design code creates annulus and blade geometry, then runs a Reynolds-average Navier–Stokes simulation for compressors sampled from the mean-line design space. Polynomial surface fits yield a continuous eight-dimensional representation of the design space for analysis, predicting efficiency with a root-mean-square error of 1.2\% points. The results show a balance between surface dissipation in boundary layers and mixing losses due to casing separations sets optimum values for inlet Mach number, hub-to-tip ratio, de Haller number, and backsweep angle. Surface dissipation drives the effect of flow coefficient, with high surface areas at low values, and high velocities at high values. Compact compressor designs are achieved by increasing inlet Mach number, reducing hub-to-tip ratio, and minimising the radial loading coefficient — all of which reduce efficiency approaching design space boundaries. An interactive web-based tool makes the results available to practising engineers, demonstrating large ensembles of automated designs and simulations as a higher-fidelity replacement for legacy empirical correlations in preliminary design.
Design and Analysis of a Liftfan for eVTOL Aircraft
Rory Hine, Dominic Cousins, Leo Maden, Samuel Walker, Nick Atkins, Samuel Grimshaw, James Taylor
GT2024-123904
Session: 01-12 Propellers and Open Rotors
Date: Wednesday, June 26, 4:00 PM – 5:30 PM
Abstract: Ducted liftfans can increase the hovering efficiency of eVTOL aircraft but are not as common as open propellers. Their increased complexity poses three challenges which are addressed in this paper, with the ultimate goal of reducing overall fan length and mass without suffering increased losses due to separations from stronger adverse pressure gradients. First, instead of the single row of a propeller, the liftfan consists of a complete stage which must be designed as a whole. The preliminary and 3D design variables of the rotor row and splittered diffuser stator row are optimised together using 3D CFD to maximise the hovering figure of merit, a non-dimensional measure of power consumption. The resulting design is then validated by experimental tests of a prototype liftfan. Second, unlike an open propeller, the liftfan is enclosed by a complete nacelle. To be viable it must simultaneously perform effectively during forward flight of the aircraft, as well as in pure hover. A low-order model is developed to investigate the maximum cruise range for different fan designs with varying area ratio (𝜎) and cruise pitch angles (𝛼). The resulting design is validated using full annulus 3D CFD and experimental wind tunnel tests. Third, as the electric motor in a liftfan is mounted inside the hub, fan-driven active cooling is necessary to prevent overheating. The stagnation pressure losses incurred by cooling airflow must be minimised without impeding heat transfer. Low-order models are developed to predict motor heat transfer and loss and to guide the design of a mixed flow cooling fan. The addition of forward sweep and adoption of a high blade count are found to significantly reduce cooling fan loss. This is confirmed with 3D CFD calculations of the cooling fan and experimental test of the prototype.
Nature of Transonic Compressor Flow: Importance of 3D Athroat/Ainlet
Demetrios Lefas
GT2024-128748
Session: 31-02 Transonic Flow
Date: Wednesday, June 26, 4:00 PM – 5:30 PM
Abstract:A key problem in transonic compressor and fan design is that although a 3D description of the flow is necessary to correctly capture the shock, accounting for it during the sectional detailed design is difficult because the key driving design parameters are unknown. In this paper, it is shown that for inlet relative Mach numbers between 0.85 to 1.30, the pressure rise across the shock is primarily a function of the 3D streamtube area at the throat At over the inlet area A1. This key finding is based on three key transonic flow features, discussed in detail within this paper, being present together across a wide range of more than 3000 representative transonic compressor and fan designs published online: (https://whittle.digital/). Only the subsonic regime approaching a Mach number of unity is analysed in this paper, whilst the supersonic regime is analysed separately in Part II. Moreover, it is shown that the effect of changes in the blade geometry, or the 3D streamtube height, on the transonic flow field is one of the same and can be explained simply by keeping track of the associated changes in At/A1. Surprisingly, the pre-shock Mach number at a given At/A1 is shown to be insensitive to the details of the blade surface geometry. Only geometric design choices made in the preliminary design phase, such as the maximum thickness and the inlet relative flow angles, are shown to have a second-order effect. These findings suggest, that the purpose of the sectional detailed design phase should be solely to make the desired changes in the real spanwise 3D At/A1. The final part of the paper concerns itself with the level of fidelity necessary when calculating the spanwise 3D At/A1, for it to positively influence design; especially while approaching a Mach number of unity when small changes in At/A1 become increasingly important. A key conclusion is that not resolving the subtle changes in the 3D radial flow within the blade passage at the appropriate level of fidelity, especially at the early throughflow multistage compressor design stage, could potentially mislead the transonic design process. As a result, for the rapid exploration of future compressor designs, this paper advocates utilising the database of more than 3000 transonic designs to generate an initial 3D blade, which is then assessed early in the design process using At/A1 extracted from 3D CFD.
An aerodynamic investigation of a high-pressure turbine using rotor casing static pressure measurements at engine representative conditions with different tip designs, tip gaps and inlet temperature profiles.
Deepanshu Singh, Paul F. Beard, David Cardwell, Vianney Staelens, Pratik Bahulekar, Mark Stokes, Simon Bather, Kam S. Chana
GT2024-122649
Session: 32-13 High Pressure Turbines 2
Date: Wednesday, June 26, 4:00 PM – 5:30 PM
Abstract: A rotating High-Pressure (HP) turbine, in modern aircraft engines, is subjected to high cyclic thermal and mechanical loads. Shroudless turbines experience high heat loads on the rotor tip and casing, which makes them one of the life-limiting components for an engine. If the rotor tip starts to erode, then the tip clearance increases which increases the over-tip leakage flow, resulting in reduced stage efficiency and ultimately, the engine lifetime. Since the 1970s, extensive research has been undertaken to understand the over-tip leakage flow, and develop mitigation strategies such as novel tip designs, to minimize the losses and heat load. However, the majority of the present findings are based on linear cascade or low-speed rotational studies, due to the high cost of the experiments and difficulty in instrumentation at engine conditions. Recent studies (mostly numerical) have shown the importance of testing at engine representative conditions, specifically targeting the transonic tip flow in a high-speed rotational environment, which is essential for an accurate understanding. The Oxford Turbine Research Facility (OTRF) is a high-speed rotating transient test facility that allows unsteady aerodynamics and heat transfer measurements, at engine representative conditions. This paper presents an experimental investigation, of the HP turbine rotor casing aerodynamics, performed in the OTRF. Steady and unsteady casing static pressure measurements were acquired, using pneumatic lines and high-frequency response (~100 kHz) Kulite pressure transducers respectively. The single-stage HP turbine consisted of cooled vanes (featuring film and trailing edge slot cooling), and uncooled rotor blades. Three different tip designs (two squealers and one flat) were tested, at two tip gaps, and with two inlet temperature profiles that included a spatially uniform total temperature profile and an engine representative HP turbine inlet total temperature profile. In parallel, a numerical investigation of the experimental test cases is presented using unsteady Computational Fluid Dynamics (CFD) predictions. The experimental results from this study provide a unique dataset to validate numerical models.
Reverse thrust aerodynamics of variable pitch fans with inlet distortion
Kwun Yeung Ma, Cesare A. Hall, Tianhou Wang, Tim S. Williams
GT2024-121728
Session: 01-06 Inlet Distortion and Engine Operability II
Date: Thursday, June 27, 8:00 AM – 10:00 AM
Abstract: Variable pitch fans can improve the operability of low pressure ratio fan systems by re-pitching the rotor blades. If a variable pitch fan can generate sufficient reverse thrust on landing it also eliminates the need for heavy, cascade-type thrust reversers. However, any reverse thrust generation is impacted by high levels of inlet distortion generated as air is drawn into the exhaust nozzle. This paper uses both RANS computations and low-speed rig experiments to explore how a representative inlet distortion from the engine installation affects the aerodynamics and performance of a variable pitch fan operating in reverse thrust mode. The simulations and the experiments both show that the distortion from the engine installation leads to a highly three-dimensional flow field with a large recirculation region within the bypass duct. This is in contrast to the primarily axial flow that is produced for a case with uniform freestream inlet conditions. The distortion substantially redistributes the mass flow, thrust and power in the engine. Streamline tracking combined with a power balance analysis reveals that there is highly radial flow within the fan rotor and almost all of the power from the fan is used to drive the recirculating flow in the bypass duct, which generates high loss and has a high total temperature. The recirculation reduces the net mass flow through the engine to around 5% of a uniform inflow case and the effective reverse thrust is greatly reduced. With uniform inflow the net reverse thrust was found to be 35% of the nominal takeoff thrust. With inlet distortion present this was reduced to 20%. This demonstrates the importance of designing and operating variable pitch fan systems to minimise the reverse thrust inlet distortion.
A rapid test facility for transonic compressors and fans
T Wang, JV Taylor, RJ Miller
GT2024-121459
Session: 31-05 Compressor Design
Date: Thursday, June 27, 1:30 PM – 3:30 PM
Abstract: This paper provides design recommendations for a rapid test facility for transonic compressors and fans, which aims to enable the testing of new geometries in two weeks at a cost of less than $10k. This compares to current industrial test facilities which typically take years and cost around $10M. The approach is important for its potential to unlock creativity allowing designers to ‘play’ at engine-representative conditions. This paper is in two parts. The first part aims to physically enable rapid high-speed testing while preserving the key fluid dynamic mechanisms being studied. The complexity of a facility scales with its power input, driven by safety requirements and complex operations. A turboexpander-type subscale facility is designed to reduce the power by a factor of nine while achieving the full Mach number and a half of the full Reynolds number. The second part aims to shorten the testing cycle time to under two weeks. It analyses the current testing process to identify the time-consuming parts and makes design changes to either eliminate or optimise each process. The final test facility was demonstrated in a trial where an example rotor geometry with a tip Mach number of 1.2 was manufactured, built, and tested. The new geometry was tested in 2 weeks, incurring a marginal cost of less than $7k.
Effect of blade damage on low pressure ratio fan windmill aerodynamics
Sofia I. Medina Cassillas, Alejandro Castillo Pardo, Cesare A. Hall, Ben Mohankumar
GT2024-122107
Session: 31-07 Compressor Off-Design Impacts & Stall Inception
Date: Thursday, June 27, 4:00 PM – 5:30 PM
Abstract: Fan windmill occurs when power to a turbofan is cut during flight and the airflow through the engine causes the fan to freewheel, often following fan damage. The fan rotational speed and drag during windmill determine the loads transmitted to the airframe, and these depend on the fan damage sustained. Idealised patterns of axisymmetric and non-axisymmetric damage have been studied using high-resolution measurements in a flow field representative of a low pressure ratio fan in combination with steady RANS simulations to understand the impact on the windmill flow field. Axisymmetric tip damage decreases the windmill rotational speed by 61% when 25% of the blade span is removed due to a shift in the zero-work radius. The reduced blade span creates an intense tip vortex and redistributes the flow, increasing the axial velocity above the damaged tip. For non-axisymmetric damage, there is still work output at the blade tip, such that removing 25% of half the blades with varying damage pitch-to-chord reduces the fan rotational speed by 9-13% compared to undamaged windmill. Phase-averaged hot-wire measurements show that this additional rotor work is due to radial flow redistribution combined with turning of the flow in the passages above the damaged blades.
