Accepted Manuscripts

Kuen-Bae Lee, Mark J. Wilson and Mehdi Vahdati
J. Turbomach   doi: 10.1115/1.4039051
It was found that the standard Spalart-Allmaras (SA) model implemented in the CFD solver used in this work predicts premature stall, which is in line with the observation of other researchers who use the SA model. Therefore, to improve the prediction of the stall boundary, the standard SA model was modified by scaling the source term in the model based on the local pressure gradient and the velocity helicity of the flow. Furthermore, a generalized wall function valid for non-zero wall pressure gradient was implemented to improve the accuracy of boundary conditions at the solid wall. This work aims to produce a turbulence model which can be used to model the flows near the stall boundary for the transonic fan blades on relatively coarse grids of around 600k points per passage. Initially, two fan blades with different design and operating speeds were used to optimize the new parameters in the modified turbulence model. The optimization was based on improving the correlation between measured and numerical radial profiles of the pressure ratio. Thereafter, steady computations were performed for two other blades (by using the same parameters) and the predictions were compared with the experimental data for all the four fan blades. Numerical results showed a significant improvement over those obtained with the original SA model, when compared against the measured data. Finally, for completeness it was decided to test the performance of the modified model by comparing the result with measured data for a simple canonical case.
TOPICS: Flow (Dynamics), Turbulence, Computer simulation, Turbochargers, Blades, Pressure gradient, Pressure, Boundary-value problems, Computation, Computational fluid dynamics, Design, Optimization
Kuen-Bae Lee, John Dodds, Mark J. Wilson and Mehdi Vahdati
J. Turbomach   doi: 10.1115/1.4039052
This paper investigates the flow near the stall boundary for a low-speed/low-pressure ratio fan blade. Three-dimensional, RANS computations are performed for a modern low speed fan rig for which extensive measured data are available. Simulations are conducted at 80% corrected speed, for which the measured constant speed characteristic contains a part with positive slope. It is shown in this paper that by using an unsteady whole assembly approach, it is possible to predict the flow for all the points on the measured constant speed characteristic (including those on the positive slope part), which is not achievable by using a single passage strategy as it would result in premature 'numerical stall'. The results of the computations reveal that for the operating points on the positive slope part of the characteristic, the flow structure becomes asymmetric and hence requires a whole assembly numerical model. The type of asymmetry which appears at lower flow coefficients is similar to the multi cell, part span rotating stall, which can occur on the front stages of the core compressors at stable operating conditions. The numerical results showed a good correlation with the measured data in terms of stall characteristics.
TOPICS: Flow (Dynamics), Computer simulation, Turbochargers, Computation, Manufacturing, Simulation, Compressors, Reynolds-averaged Navier–Stokes equations, Engineering simulation, Blades, Pressure
Yunfeng Fu, Fu Chen, Huaping Liu and YanPing Song
J. Turbomach   doi: 10.1115/1.4039049
In this paper, the effect of a novel honeycomb tip on suppressing tip leakage flow has been experimentally and numerically studied. The research focuses on the mechanisms of honeycomb tip on suppressing tip leakage flow and affecting the secondary flow, as well as the influences of different clearance heights on leakage flow characteristics. In addition, two kinds of local honeycomb tip structures are proposed to explore the positive effect on suppressing leakage flow. Honeycomb tip rolls up a number of small vortices and forms radial jets in honeycomb cavities, increasing the flow resistance in the clearance and reducing the velocity of leakage flow. As a result, honeycomb tip not only reduces the leakage flow effectively, but also has positive effect on reducing the losses by reducing the strength of leakage vortex. Compare to the flat tip cascade, the relative leakage flow reduces from 3.05% to 2.73%, and the loss at exit section is decreased by 10.63%. With the increase of the gap height, the tip leakage flow and loss have variations of direct proportion with it, but the growth rates in the honeycomb tip cascade is smaller. Consider the abradable property of the honeycomb seal, a smaller gap height is allowed with honeycomb tip, which means honeycomb tip can perform better. Two various local honeycomb tip structures has also been discussed. It shows that local raised honeycomb tip has better suppressing leakage flow effect than honeycomb tip, while local concave honeycomb tip has no more effect.
TOPICS: Turbochargers, Cascades (Fluid dynamics), Honeycomb structures, Turbines, Leakage flows, Vortices, Clearances (Engineering), Flow (Dynamics), Jets, Cavities, Leakage
Christoph Brandstetter, Maximilian Juengst and Heinz-Peter Schiffer
J. Turbomach   doi: 10.1115/1.4039053
The phenomena prior to rotating stall were investigated in a high-speed compressor test rig using optical and pneumatic measurement techniques. A number of throttling procedures was performed at transonic and subsonic speedlines with the aim to detect the unsteady effects initiating rotating stall or large amplitude blade vibrations. At transonic speed radial vortices traveling around the circumference were detected in the upstream part of the rotor using phase-locked PIV measurements above 92% span and unsteady wall pressure measurements. When these radial vortices impinge on a blade leading edge, they cause a forward spill of fluid around the leading edge. The effects are accompanied by a large-scale vortex breakdown in the blade passage leading to immense blockage in the endwall region. At subsonic speeds, the observed flow phenomena are similar but differ in intensity and structure. During the throttling procedure, blade vibration amplitudes were monitored using strain gauges and blade tip timing instrumentation. Non-synchronous blade vibrations in the first torsional eigenmode were measured as the rotor approached stall. Using the different types of instrumentation, it was possible to align the aerodynamic flow features with blade vibration levels. The results show a clear correlation between the occurrence of radial vortices and blade vibrations.
TOPICS: Compressors, Turbochargers, Vortices, Blades, Vibration, Instrumentation, Rotors, Strain gages, Aerodynamic flow, Flow (Dynamics), Fluids, Pressure measurement
Tiziano Ghisu and Shahrokh Shahpar
J. Turbomach   doi: 10.1115/1.4038982
Uncertainty Quantification is an increasingly important area of research. As components and systems become more efficient and optimized, the impact of uncertain parameters. It is fundamental to consider the impact of these uncertainties as early as possible during the design process, with the aim of producing more robust designs (less sensitive to the presence of uncertainties). The cost of UQ with high-fidelity simulations becomes therefore of fundamental importance. This work makes use of Least Squares Approximations in the context of appropriately selected Polynomial Chaos bases. An efficient technique based on QR column pivoting has been employed to reduce the number of evaluations required to construct the approximation, demonstrating the superiority of the method with respect to full tensor and sparse grid quadratures. Orthonormal polynomials used for the PC expansion are calculated numerically based on the given uncertainty distribution, making the approach optimal for any type of input uncertainty. The approach is used to quantify the variability in the performance of two large bypass-ratio jet engine fans in the presence of shape uncertainty due to possible manufacturing processes. The impacts of shape uncertainty on the two geometries are compared, and sensitivities to the location of the blade shape variability are extracted. The mechanisms at the origin of the change in performance are analyzed in detail, as well as the differences between the two configurations. These results provide important information both for controling the manufacturing process, and for designing blades that are less sensitive to the presence of manufacturing uncertainties.
TOPICS: Turbochargers, Fans, Aircraft engines, Uncertainty quantification, Uncertainty, Shapes, Manufacturing, Design, Polynomials, Blades, Chaos, Jet engines, Least squares approximations, Engineering simulation, Approximation, Simulation, Tensors
Ioanna Aslanidou and Budimir Rosic
J. Turbomach   doi: 10.1115/1.4038907
In gas turbines with can combustors the trailing edge of the combustor transition duct wall is found upstream of every second vane. This paper presents an experimental and numerical investigation of the effect of the combustor wall trailing edge on the aerothermal performance of the nozzle guide vane. In the measurements carried out in a high speed experimental facility, the wake of this wall is shown to increase the aerodynamic loss of the vane. On the other hand, the wall alters secondary flow structures and has a protective effect on the heat transfer in the leading edge-endwall junction, a critical region for component life. The different clocking positions of the vane relative to the combustor wall are tested experimentally and are shown to alter the aerothermal field. The experimental methods and processing techniques adopted in this work are used to highlight the differences between the different cases studied.
TOPICS: Turbochargers, Combustion chambers, Nozzle guide vanes, Experimental methods, Gas turbines, Ducts, Junctions, Wakes, Flow (Dynamics), Heat transfer
Jekyoung Lee, Seong Kuk Cho and Jeong Ik Lee
J. Turbomach   doi: 10.1115/1.4038879
From the efforts of many researchers and engineers related to the S-CO2 Brayton cycle technology development, the S-CO2 Brayton cycle is now considered as one of the key power technologies for the future. Since the S-CO2 Brayton cycle has advantages in economics due to high efficiency and compactness of the system, various industries have been trying to develop baseline technology on the design and analysis of the S-CO2 Brayton cycle components. According to the previous researches on the S-CO2 Brayton cycle component technology, the treatment of a thermodynamic property near the critical point of CO2 is one of the main concerns since conventional design and analysis methodologies cannot be used for the near critical point region. Among many thermodynamic properties, the stagnation to static condition conversion process is important since the flow in a compressor is at high flow velocity. In this paper, the impact of various stagnation to static conversion methods on the S-CO2 compressor design near the critical point will be evaluated. From the evaluation, the limitation of a certain stagnation to static conversion method will be discussed to provide a guideline for the future S-CO2 compressor designers.
TOPICS: Compressors, Turbochargers, Design, Approximation, Supercritical carbon dioxide, Brayton cycle, Flow (Dynamics), Engineers, Carbon dioxide, Technology development, Economics
Alexander Hehn, Moritz Mosdzien, Daniel Grates and Peter Jeschke
J. Turbomach   doi: 10.1115/1.4038908
A transonic centrifugal compressor was aerodynamically optimized by means of a numerical optimization process. The ob- jectives were to increase the isentropic efficiency and to reduce the acoustic signature by decreasing the amplitude of pre shock pressure waves at the inlet of the compressor. The optimization was performed at three operating points on the 100% speed line in order to maintain choke mass flow and surge margin. At the de- sign point, the specific work input was kept equal. The baseline impeller was designed by using ruled surfaces due to require- ments for flank milling. To investigate the benefits of arbitrary blade surfaces, the restrictions of ruled surfaces were abolished and fully 3D blade profiles allowed. In total therefore, 45 pa- rameters were varied during the optimization. The combined ge- ometric and aerodynamic analysis reveals that a forward swept leading edge and a concave suction side at the tip of the leading edge are effective design features for reducing the shock strength. Beyond that the blade shape of the optimized compressor creates a favorable impeller outlet flow, which is the main reason why the performance of the vaneless diffuser improves. In total a gain of 1.4%-points in isentropic total-to-static efficiency, evaluated by CFD at the exit plane of the vaneless diffuser, is achieved.
TOPICS: Compressors, Turbochargers, Optimization, Blades, Vaneless diffusers, Shock (Mechanics), Impellers, Flow (Dynamics), Acoustics, Suction, Computational fluid dynamics, Design, Waves, Pressure, Milling, Shapes, Surges
Xinqian Zheng, Zhenzhong Sun, Tomoki Kawakubo and Hideaki Tamaki
J. Turbomach   doi: 10.1115/1.4038875
The non-uniformity of the flow field induced by the non-axisymmetric volute significantly exacerbates the stability of a turbocharger centrifugal compressor. In this paper, non-axisymmetric vaned diffuser as an instance of non-axisymmetric flow control method is investigated with both three-dimensional CFD and experiment. The numerical study firstly focuses on the relationship between the flow field and the static pressure distortion, and numerical results indicate that the positive static pressure gradient along rotating direction facilitates flow separations and makes the flow field non-uniform. Non-axisymmetric flow control method with variable stagger and solidity of the vaned diffuser is developed to suppress the flow separation, and it suggests to narrow the flow passages where flow separation happens or to close the diffuser vanes upstream of the flow separations. The steady CFD also presents the flow field of the investigated turbocharger centrifugal compressor with volute, and flow separation appears in flow passages near the volute tongue. Under the guidance of the non-axisymmetric flow control method, several non-axisymmetric vaned diffusers are designed, numerical analysis shows they uniform the flow field. Finally, experiment validation shows the non-axisymmetric vaned diffuser extends the stable flow range of the compressor by 26% relatively compared with axisymmetric vaned diffuser.
TOPICS: Compressors, Turbochargers, Diffusers, Stability, Flow (Dynamics), Flow separation, Flow control, Computational fluid dynamics, Numerical analysis, Pressure, Pressure gradient
Fangyuan Lou, John C. Fabian and Nicole L. Key
J. Turbomach   doi: 10.1115/1.4038876
This paper introduces a new approach for the preliminary design and aero-thermal analysis of centrifugal impellers using a relative diffusion effectiveness parameter. The relative diffusion effectiveness is defined as the ratio of the achieved diffusion to the maximum available diffusion in an impeller. It represents the quality of the relative diffusion process in an impeller. This parameter is used to evaluate impeller performance by correlating the relative diffusion effectiveness with the impeller isentropic efficiency using the experimental data acquired on a single stage centrifugal compressor. By including slip, which is appropriate considering it is an inviscid effect that should be included in the determination of maximum available diffusion in the impeller, a linear correlation between impeller efficiency and relative diffusion effectiveness resulted for all operating conditions.Additionally, a new method for impeller preliminary design was introduced using the relative diffusion effectiveness parameter, in which the optimal design is selected to maximize relative diffusion effectiveness. While traditional preliminary design methods are based on empirical loss models or empirical knowledge for selection of diffusion factor in the impeller, the new method does not require any such models, and it also provides an analytical approach for the selection of diffusion factor that gives optimal impeller performance. Validation of the method was performed using three classic impeller designs available in the open literature, and very good agreement was achieved.
TOPICS: Impellers, Turbochargers, Design, Diffusion (Physics), Compressors, Diffusion processes, Design methodology
Harald Schoenenborn
J. Turbomach   doi: 10.1115/1.4038868
The aeroelastic prediction of blade forcing is still a very important topic in turbomachinery design. Usually, the wake from an upstream airfoil and the potential field from a downstream airfoil are considered as the main disturbances. In recent years, it became evident that in addition to those two mechanisms Tyler-Sofrin modes may have a significant impact on blade forcing. In Schrape et al. it was found that in multi-row configurations not only the next, but also the next but one blade row is very important as it may create a large circumferential forcing variation fixed in the rotating frame of reference. In the present paper a study of these effects is performed on the basis of a quasi 3D multi-row and multi-passage compressor configuration. A harmonic balancing code is used for various setups and the results are compared to full-annulus unsteady calculations. It is shown that the effect of the circumferentially different blade excitation is mainly contributed by the Tyler-Sofrin modes and not to blade-to-blade variation in the steady flowfield. The influence of various clocking positions, coupling schemes and number of harmonics onto the forcing is investigated. It is shown that along a compressor speedline the blade-to-blade forcing variation changes significantly. In addition, multi-row flutter calculations show the influence of the upstream and downstream blade row onto aerodynamic damping. The effect of these forcing variations onto random mistuning effects is investigated in the second part of the paper.
TOPICS: Compressors, Turbochargers, Excitation, Blades, Airfoils, Turbomachinery, Wakes, Flutter (Aerodynamics), Damping, Design, Annulus
Johann Gross, Malte Krack and Harald Schoenenborn
J. Turbomach   doi: 10.1115/1.4038869
The prediction of aerodynamic blade forcing is a very important topic in turbomachinery design. Usually, the wake from the upstream blade row and the potential field from the downstream blade row are considered as the main causes for excitation, give rise to dynamic forcing of the blades. In addition, so-called Tyler-Sofrin modes, which refer to the acoustic interaction with blade rows further up- or downstream, may have a significant impact on blade forcing. In particular, they lead to considerable blade-to-blade variations of the aerodynamic loading. In part 1 a study of these effects is performed on the basis of a quasi 3D compressor configuration.Part 2 of the paper proposes a method to analyze the interaction of the aerodynamic forcing asymmetries with the already well-studied effects of random mistuning stemming from blade-to-blade variations of structural properties. Based on a finite element model of a sector, the equations governing the dynamic behavior of the entire bladed disk can be efficiently derived using substructuring techniques. The disk substructure is assumed as cyclically symmetric, while the blades exhibit structural mistuning and linear aeroelastic coupling. In order to avoid the costly multi-stage analysis, the variation of the aerodynamic loading is treated as an epistemic uncertainty, leading to a stochastic description of the annular force pattern. The effects of structural mistuning and stochastic aerodynamic forcing are first studied separately and then in a combined manner for a blisk of a research compressor without and with aeroelastic coupling.
TOPICS: Compressors, Turbochargers, Blades, Disks, Acoustics, Finite element model, Turbomachinery, Uncertainty, Excitation, Wakes, Mechanical properties, Design
Yousef Kanani, Dr. Sumanta Acharya and Forrest Ames
J. Turbomach   doi: 10.1115/1.4038877
Slot film cooling in an accelerating boundary layer with high free-stream turbulence is studied numerically using Large Eddy Simulations (LES). Calculations are done for a symmetrical leading edge geometry with the slot fed by a plenum populated with pin fins. To generate the inflow turbulence, the Synthetic Eddy Method is used by which the turbulence intensity and length scales in each direction can be specified at the inflow. Different levels of turbulence are imposed at the inflow cross-plane. Calculations are done for a Reynolds number of 250,000 and freestream turbulence levels of 0.7%, 3.5%, 7.8% and 13.7% are reported. These conditions correspond to the experimental measurements of Busche and Ames (2014). Numerical results show good agreement with experiment data and show the observed decay of thermal effectiveness with turbulence intensity. The turbulence and non-uniformity exiting the slot play an important role in the cooling effectiveness distributions downstream of the slot. Generation of freestream structures is observed at the leading edge, and the amplification of the corresponding fluctuations downstream is one of the parameters influencing the slot cooling performance. Predictions show the higher growth rate of the thermal boundary layer with increasing turbulence which is a clear indication of the increase in turbulent thermal diffusivity and reduction of the effective turbulence Prandtl number. The self-similar temperature profiles deviate from those measured under low freestream turbulence condition.
TOPICS: Turbulence, Simulation, Turbochargers, Engineering simulation, Film cooling, Inflow, Cooling, Fins, Geometry, Prandtl number, Temperature profiles, Large eddy simulation, Thermal boundary layers, Fluctuations (Physics), Thermal diffusivity, Boundary layers, Eddies (Fluid dynamics), Reynolds number, Symmetry (Physics)
Shane E. Haydt, Stephen Lynch and Scott D. Lewis
J. Turbomach   doi: 10.1115/1.4038871
Shaped film cooling holes are used as a cooling technology in gas turbines to reduce metal temperatures and improve durability, and they generally consist of a small metering section connected to a diffuser that expands in one or more directions. The area ratio of these holes is defined as the area at the exit of the diffuser, divided by the area at the metering section. A larger area ratio increases the diffusion of the coolant momentum, leading to lower average momentum of the coolant jet at the exit of the hole and generally better cooling performance. Cooling holes with larger area ratios are also more tolerant of high blowing ratio conditions, and the increased coolant diffusion typically better prevents jet liftoff from occurring. Higher area ratios have traditionally been accomplished by increasing the expansion angle of the diffuser while keeping the overall length of the hole constant. The present study maintains the diffuser expansion angles and instead increases the length of the diffuser, which results in longer holes. Various area ratios have been examined for two shaped holes: one with forward and lateral expansion angles of 7° (7-7-7 hole) and one with forward and lateral expansion angles of 12° (12-12-12 hole). Each hole shape was tested at numerous blowing ratios to capture trends across various flow rates. Adiabatic effectiveness measurements indicate that for the baseline 7-7-7 hole, a larger area ratio provides higher effectiveness, especially at higher blowing ratios.
TOPICS: Turbochargers, Film cooling, Diffusers, Cooling, Coolants, Diffusion (Physics), Momentum, Flow (Dynamics), Temperature, Metals, Durability, Gas turbines, Shapes
Jonna Tiainen, Ahti Jaatinen-Värri, Aki Grönman, Teemu Turunen-Saaresti and Jari Backman
J. Turbomach   doi: 10.1115/1.4038872
The estimation of boundary layer losses requires the accurate specification of the free-stream velocity, which is not straightforward in centrifugal compressor blade passages. This challenge stems from the jet-wake flow structure, where the free-stream velocity between the blades cannot be clearly specified. In addition, the relative velocity decreases due to adverse pressure gradient. Therefore, the common assumption of a single free-stream velocity over the blade surface might not be valid in centrifugal compressors.Generally in turbomachinery, the losses in the blade cascade boundary layers are estimated e.g. with different loss coefficients, but they often rely on the assumption of a uniform flow field between the blades. To give guidelines for the estimation of the mentioned losses in highly distorted centrifugal compressor flow fields, this paper discusses the difficulties in the calculation of the boundary layer thickness in the compressor blade passages, compares different free-stream velocity definitions, and demonstrates their effect on estimated boundary layer losses. Additionally, a hybrid method is proposed to overcome the challenges of defining a boundary layer in centrifugal compressors.
TOPICS: Compressors, Turbochargers, Boundary layers, Blades, Flow (Dynamics), Pressure gradient, Turbomachinery, Cascades (Fluid dynamics), Wakes
Chen He, Dakun Sun and Xiaofeng Sun
J. Turbomach   doi: 10.1115/1.4038873
This paper concentrates on the stall inception analysis of transonic compressors with chordwise and axial sweep. A new prediction approach of stall inception is developed based on global stability analysis and immersed boundary theory, which makes it possible to takes both the concrete blade geometry and complicated base flow into consideration. The prediction of stall inception boils down to an eigenvalue problem. Spectral collocation method is adopted to discretize the eigenvalue equations and the eigenvalues are solved by using singular value decomposition method. The developed prediction approach is validated on two different typical transonic compressors, a single stage compressor and an isolated-rotor compressor, which shows a good agreement with the experimental data. The latter is adopted as a baseline rotor for the investigation of chordwise and axial sweep. By adjusting the stacking line of the baseline rotor, a series of swept rotors are modelled and the stall inception behavior of them are predicted by using the developed approach. The comparison of stall inception behaviors between these rotors is presented, and in combination with steady flow analyses, the effects of sweep features on the stall inception in transonic compressors are discussed.
TOPICS: Compressors, Turbochargers, Stall inception, Rotors, Eigenvalues, Flow (Dynamics), Concretes, Boiling, Geometry, Blades, Stability
Liangjun Hu, Harold Sun, Jianwen James Yi, Eric Curtis and Jizhong Zhang
J. Turbomach   doi: 10.1115/1.4038878
Variable geometry turbine (VGT) has been widely applied in internal combustion engines to improve engine transient response and torque at light load. One of the most popular variable geometry turbines is the variable nozzle turbine (VNT), in which the nozzle vanes can be rotated along the pivoting axis and thus the flow passage through the nozzle can be adjusted to match with different engine operating conditions. One disadvantage of the VNT is the turbine efficiency degradation due to the leakage flow in the nozzle endwall clearance, especially at small nozzle open condition. With the purpose to reduce the nozzle leakage flow and to improve turbine stage efficiency, a novel split sliding variable nozzle turbine (SSVNT) has been proposed. In the SSVNT design, the nozzle is divided into two parts: one part is fixed and the other part can moved along the partition surface. When sliding the moving vane to large radius position, the nozzle flow passage opens up and the turbine has high flow capacity. When sliding the moving vane to small radius position, the nozzle flow passage closes down and the turbine has low flow capacity. As the fixed vane doesn't need endwall clearance, the leakage flow through the nozzle can be reduced. Based on calibrated numerical simulation, there is up to 12% turbine stage efficiency improvement with the SSVNT design at small nozzle open condition while maintaining the same performance at large nozzle open condition. The mechanism of efficiency improvement in the SSVNT design has been discussed.
TOPICS: Turbochargers, Design, Nozzles, Turbines, Flow (Dynamics), Leakage flows, Geometry, Engines, Clearances (Engineering), Torque, Transients (Dynamics), Stress, Computer simulation, Internal combustion engines
Richard Ahlfeld, Francesco Montomoli, Mauro Carnevale and Simone Salvadori
J. Turbomach   doi: 10.1115/1.4038826
Problems in turbomachinery Computational Fluid Dynamics (CFD) are often characterised by non-linear and discontinuous responses. Ensuring the reliability of Uncertainty Quantification (UQ) codes in such conditions, in an autonomous way, is a challenging problem. In this work, we suggest a new approach that combines three state-of-the-art methods: multivariate Padé approximations, Optimal Quadrature Subsampling and Statistical Learning. Its main component is the generalised least squares multivariate Padé -Legendre (PL) approximation. PL approximations are globally fitted rational functions that can accurately describe discontinuous non-linear behaviour. They need fewer model evaluations than local or adaptive methods and do not cause the Gibbs phenomenon like continuous Polynomial Chaos methods. A series of modifications of the Padé algorithm allow us to apply it to arbitrary input points instead of optimal quadrature locations. This property is particularly useful for industrial applications, where a database of CFD runs is already available, but not in optimal parameter locations. One drawback of the PL approximation is that it is non-trivial to ensure reliability. To improve stability we suggest to couple it with Optimal Quadrature Subsampling. Our reasoning is that least squares errors, caused by an ill-conditioned design matrix, are the main source of error. Finally, we use statistical learning methods to check smoothness and convergence. The resulting method is shown to efficiently and correctly fit thousands of partly discontinuous response surfaces for an industrial film cooling and shock interaction problem using only 9 CFD simulations.
TOPICS: Turbochargers, Approximation, Uncertainty quantification, Computational fluid dynamics, Reliability, Errors, Polynomials, Turbomachinery, Film cooling, Stability, Simulation, Shock (Mechanics), Algorithms, Design, Engineering simulation, Chaos, Databases
Yuewen Jiang, Luigi Capone, Peter Ireland and Eduardo Romero
J. Turbomach   doi: 10.1115/1.4038833
An optimal design of film cooling is a key factor in the effort of producing high efficiency gas turbine. Understanding of the fluid dynamics interaction between cooling holes can help engineers to improve overall thermal effectiveness. Modelling and correct prediction is a very complex problem, since the multiple phenomena involved, such as: mixing, turbulence and heat transfer. The present work performs an investigation of different cooling configurations ranging from single hole up to two rows. The main objective is to evaluate the double-rows interaction and the effect on film cooling. Strong nonlinear effects are underlined by different simulations, while varying blowing ratio and geometrical configuration of cooling holes. Meanwhile an initial analysis is performed using flat plate geometry, verification and validation is then extended to realistic stage of high pressure turbine. Multiple cooling holes configurations are embedded on the pressure and suction sides of the single stage. The main outcome is the verification of the thermal effectiveness improvement obtained by cooling jets interaction of multiple rows design. The effects of curvature surface and frame of reference rotation are also evaluated, underlying the differences with standard flat plate test cases.
TOPICS: Cooling, Turbochargers, Flat plates, Design, Film cooling, Engineering simulation, Gas turbines, Modeling, Performance, Turbines, Geometry, High pressure (Physics), Jets, Turbulence, Suction, Engineers, Simulation, Pressure, Rotation, Fluid dynamics, Heat transfer
Alistair John, Ning Qin and Shahrokh Shahpar
J. Turbomach   doi: 10.1115/1.4038834
During engine operation fan casing abradable liners are worn by the blade tip, resulting in the formation of trenches. This paper investigates the influence of these trenches on the fan blade tip aerodynamics. A detailed understanding of the tip flow features for the fan blade under investigation is developed. A parametric model is then used to model trenches in the casing above the blade tip. It is shown that increasing clearance via a trench reduces performance by less than increasing clearance through cropping the blade tip. A response surface method is then used to generate a model that can predict fan efficiency for a given set of clearance and trench parameters. It is shown that the efficiency sensitivity to clearance is greater for cropped tips than trenches, and is biased towards the leading edge for cropped tips, and the trailing edge for trenches.
TOPICS: Aerodynamics, Turbochargers, Blades, Clearances (Engineering), Engines, Response surface methodology, Flow (Dynamics)

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