Accepted Manuscripts

James H. Page, Paul Hield and Paul G. Tucker
J. Turbomach   doi: 10.1115/1.4040030
The effect of inlet distortion from curved intake ducts on jet engine fan stability is an important consideration for next generation passenger aircraft such as the boundary layer ingestion (BLI) ‘Silent Aircraft’. Highly complex inlet flows which occur can significantly affect fan stability. Future aircraft designs are likely to feature more severe inlet distortion, pressing the need to understand the important factors influencing design. This paper presents the findings from a large CFD investigation into which aspects of inlet distortion cause the most significant reductions in stall margin and, therefore, which flow patterns should be targeted by mitigating technology. The study considers the following aspects of distortion commonly observed in intakes: steady vortical distortion due to secondary flow, unsteady vortical distortion due to vortex shedding and mixing, static pressure distortion due to curved streamlines, and low momentum endwall flow due to thickened boundary layers or separation. Unsteady CFD was used to determine the stall points of a multipassage transonic rotor geometry with each of the inlet distortion patterns applied. Interesting new evidence is provided which suggests that low momentum flow in the tip region, rather than distortion in the main body of the flow, leads to damaging instability.
Jason Town, Douglas Straub, James Black, Karen A. Thole and Tom I-P. Shih
J. Turbomach   doi: 10.1115/1.4039942
Effective internal and external cooling of airfoils is key to maintaining component life for efficient gas turbines. Cooling designs have spanned the range from simple internal convective channels to more advanced double-walls with shaped film-cooling holes. This paper describes the development of an internal and external cooling concept for a state-of-the-art cooled turbine blade. These cooling concepts are based on a review of literature and patents, as well as, interactions with academic and industry turbine cooling experts. The cooling configuration selected and described in this paper is referred to as the "baseline" design, since this design will simultaneously be tested with other more advanced blade cooling designs in a rotating turbine test facility using a "rainbow turbine wheel" configuration. For the baseline design, the leading edge is cooled by internal jet impingement and showerhead film cooling. The mid-chord region of the blade contains a three-pass serpentine passage with internal discrete V-shaped trip strips to enhance the internal heat transfer coefficient. The film cooling along the mid-chord of the blade uses multiple rows of shaped diffusion holes. The trailing edge is internally cooled using jet impingement and externally film cooled through partitioned cuts on the pressure side of the blade.
TOPICS: Rotors, Turbines, Blades, Cooling, Turbochargers, Design, Film cooling, Chords (Trusses), Gas turbines, Turbine blades, Pressure, Diffusion (Physics), Patents, Strips, Test facilities, Wheels, Heat transfer coefficients, Airfoils
Yanfeng Zhang, Shuzhen Hu, Ali Mahallati, Xue Feng Zhang and Edwards Vlasic
J. Turbomach   doi: 10.1115/1.4039936
The present work, a continuation of a series of investigations on the aerodynamics of aggressive inter-turbine ducts (ITD), is aimed at providing detailed understanding of the flow physics and loss mechanisms in four different ITD geometries. A systematic experimental and computational study was carried out by varying duct outlet-to-inlet area ratios and mean rise angles while keeping the duct length-to-inlet height ratio, Reynolds number and inlet swirl constant in all four geometries. The flow structures within the ITDs were found to be dominated by the boundary layer separation and counter-rotating vortices in both the casing and hub regions. The duct mean rise angle determined the severity of adverse pressure gradient in the casing's first bend whereas the duct area ratio mainly governed the second bend's static pressure rise. The combination of upstream wake flow and the first bend's adverse pressure gradient caused the boundary layer to separate and intensify the strength of counter-rotating vortices. At high mean rise angle, the separation became stronger at the casing's first bend and moved farther upstream. At high area ratios, a two-dimensional separation appeared on the casing and resulted in increased loss. Pressure loss penalties increased significantly with increasing duct mean rise angle and area ratio.
TOPICS: Aerodynamics, Turbochargers, Turbines, Ducts, Separation (Technology), Pressure, Flow (Dynamics), Boundary layers, Pressure gradient, Vortices, Physics, Wakes, Reynolds number
Daniel Feseker, Mats Kinell and Matthias Neef
J. Turbomach   doi: 10.1115/1.4039842
The ability to understand and predict the pressure losses of orifices is important in order to improve the air flow within the secondary air system. This experimental study investigates the behavior of the discharge coefficient for circular orifices with inlet cross flow which is a common flow case in gas turbines. Examples of this are at the inlet of a film cooling hole or the feeding of air to a blade through an orifice in a rotor disc. Measurements were conducted for a total number of 38 orifices, covering a wide range of length-to-diameter ratios, including short and long orifices with varying inlet geometries. Up to five different chamfer-to-diameter and radius-to-diameter ratios were tested per orifice length. Furthermore, the static pressure ratio across the orifice was varied between 1.05 and 1.6 for all examined orifices. The results of this comprehensive investigation demonstrate the beneficial influence of rounded inlet geometries and the ability to decrease pressure losses, which is especially true for higher cross flow ratios where the reduction of the pressure loss in comparison to sharp edged holes can be as high as 54%. With some exceptions, the chamfered orifices show a similar behavior as the rounded ones but with generally lower discharge coefficients. Nevertheless, a chamfered inlet yields lower pressure losses than a sharp edged inlet. The obtained experimental data was used to develop two correlations for the discharge coefficient as a function of geometrical as well as flow properties.
TOPICS: Turbochargers, Pressure, Orifices, Cross-flow, Discharge coefficient, Flow (Dynamics), Air flow, Film cooling, Gas turbines, Rotors, Disks, Blades
Jichao Li, Juan Du, Zhiyuan Li and Feng Lin
J. Turbomach   doi: 10.1115/1.4039806
Self-recirculating injection, which bleeds air from the downstream duct of the last blade row and injects the air as a wall jet upstream of the first rotor blade row, is experimentally studied after well-design of its structure in single- and three-stage axial flow compressor respectively. The external injection and outlet bleed air are selected for comparison. Results show that the self-recirculating injection can improve the stall margin by 13.67% and 13% on the premise of no efficiency penalty respectively in single- and three-stage compressor for only 0.7% and 4.2% of the total injected momentum ratio re-circulated near stall. It is best among all the three cases if comprehensively considering the impact on pressure rise and efficiency. The details of flow field are captured by using a collection of pressure transducers on the casing with circumferential and chord-wise spatial resolution. The detailed comparative analysis of the endwall flow indicates that the self-recirculating injection can postpone the occurrence of stalling in the proposed compressor through delaying the forward movement of the interface between the tip leakage flow (TLF) and main stream flow, weakening the unsteadiness of TLF, and sharply declining the circumferentially propagating speed induced by TLF that triggers the spike-type stall inception. Finally, the stall control concept on the stage that first generates stall inception using self-recirculating injection is proposed. This study may be helpful to guide the design of self-recirculating injection in actual application.
TOPICS: Stability, Turbochargers, Axial flow, Flow (Dynamics), Compressors, Blades, Design, Stall inception, Leakage flows, Rotors, Ducts, Pressure transducers, Resolution (Optics), Chords (Trusses), Pressure, Momentum
John Leggett, Stephan Priebe, Aamir Shabbir, Vittorio Michelassi, Richard Sandberg and Ed S Richardson
J. Turbomach   doi: 10.1115/1.4039807
Axial compressors may be operated under off-design incidences due to variable operating conditions. Therefore, a successful design requires accurate performance and stability limits predictions under a wide operating range. Designers generally rely both on correlations and on RANS, the accuracy of the latter often being questioned. The present study investigates profile losses in an axial compressor linear cascade using both RANS and wall-resolved Large Eddy Simulation (LES), and compares with measurements. The analysis concentrates on "loss buckets", local separation bubbles and boundary layer transition with high levels of free stream turbulence, as encountered in real compressor environment without and with periodic incoming wakes. The work extends previous research with the intention of furthering our understanding of prediction tools and improving our quantification of the physical processes involved in loss generation. The results show that while RANS predicts overall profile losses with good accuracy, the relative importance of the different loss mechanisms does not match with LES, especially at off-design conditions. This implies that a RANS based optimisation of a compressor profile under a wide incidence range may require a thorough LES verification at off-design incidence.
TOPICS: Compressors, Turbochargers, Cascades (Fluid dynamics), Design, Reynolds-averaged Navier–Stokes equations, Large eddy simulation, Optimization, Stability, Separation (Technology), Turbulence, Wakes, Bubbles, Boundary layers
Dajan Mimic, Bastian Drechsel and Florian Herbst
J. Turbomach   doi: 10.1115/1.4039821
Exhaust diffusers significantly enhance the available power output and efficiency of gas and steam turbines by allowing for lower turbine exit pressures. The residual dynamic pressure of the turbine outflow is converted into static pressure, which is referred to as pressure recovery. Since total pressure losses as well as construction costs increase drastically with diffuser length, it is more than favourable to design shorter diffusers with rather steep opening angles. However, those designs are more susceptible to boundary layer separation. In this paper, the stabilising properties of tip leakage vortices generated in the last rotor row and their effect on the boundary layer characteristics are examined. Based on analytical considerations, for the first time a correlation between the pressure recovery of the diffuser and integral rotor parameters of the last stage, namely the loading coefficient, flow coefficient and reduced frequency, is established. Both, experimental data and scale resolving simulations, carried out with the SST-SAS method, show excellent agreement with the correlation. Blade tip vortex strength predominantly depends on the amount of work performed in the rotor, which in turn is described by the non-dimensional loading coefficient. The flow coefficient influences mainly the orientation of the vortex, which affects the interaction between vortex and boundary layer. The induced velocity field accelerates the boundary layer, essentially reducing the thickness of the separated layer or even locally preventing separation.
TOPICS: Pressure, Turbochargers, Diffusers, Design, Boundary layers, Rotors, Vortices, Turbines, Flow (Dynamics), Separation (Technology), Simulation, Construction, Engineering simulation, Wake turbulence, Blades, Exhaust systems, Steam turbines, Leakage, Outflow
Ruzbeh Hadavandi, Fabrizio Fontaneto and Julien Desset
J. Turbomach   doi: 10.1115/1.4039727
CFD is nowadays extensively used for turbomachinery design and performance prediction. Nevertheless, compressors numerical simulations still fail in correctly predicting the stall inception and the post-stall behavior. Several authors address such a lack of accuracy to the incomplete definition of the boundary conditions and of the turbulence parameters at the inlet of the numerical domain. The aim of the present paper is to contribute to the development of compressors CFD by providing a complete set of input data for numerical simulations. A complete characterization has been carried out for a state-of-art 1.5 stage highly loaded low pressure compressor for which previous CFD analyses have failed to predict its behavior. The experimental campaign has been carried out in the R4 facility at the Von Karman Institute for Fluid Dynamics. The test item has been tested in different operative conditions for two different speed lines (90% and 96% of the design speed) and for two different Reynolds numbers. Stable and unstable operative conditions have been investigated along with the stalling behavior, its inception and the stall-cell flow field. Discrete hot-wire traverses have been performed in order to characterize the span-wise velocity field and the turbulence characteristics.
TOPICS: Pressure, Simulation, Turbochargers, Computational fluid dynamics, Engineering simulation, Compressors, Reynolds number, Design, Turbulence, Computer simulation, Fluid dynamics, Flow (Dynamics), Wire, Boundary-value problems, Stall inception, Turbomachinery
Jeff Defoe, Majed Etemadi and David Hall
J. Turbomach   doi: 10.1115/1.4039433
Applications such as boundary-layer-ingesting fans, and compressors in turboprop engines require continuous operation with distorted inflow. A low-speed axial fan with incompressible flow is studied in this paper. The objectives are to (1) identify the physical mechanisms which govern the fan response to inflow distortions and (2) determine how fan performance scales as the type and severity of inlet distortion varies at the design flow coefficient. A distributed source term approach to modeling the rotor and stator blade rows is used in numerical simulations in this paper. The model does not include viscous losses so that changes in diffusion factor are the primary focus. Distortions in stagnation pressure and temperature as well as swirl are considered. The key findings are that unless sharp pitchwise gradients in the diffusion response, strong radial flows, or very large distortion magnitudes are present, the response of the blade rows for strong distortions can be predicted by scaling up the response to a weaker distortion. In addition, the response to distortions which are composed of non-uniformities in several inlet quantities can be predicted by summing up the responses to the constituent distortions.
TOPICS: Pressure, Flow (Dynamics), Temperature, Diffusion (Physics), Computer simulation, Engines, Compressors, Turbochargers, Design, Fans, Modeling, Rotors, Blades, Radial flow, Stators, Inflow
Dr. John Clark, Joseph Beck, Alex Kaszynski, Angela Still and Ron Ho Ni
J. Turbomach   doi: 10.1115/1.4039361
This effort focuses on the comparison of unsteadiness due to as-measured turbine blades in a transonic turbine to that obtained with blueprint geometries via computational fluid dynamics (CFD). The nominal turbine CFD grid data defined for analysis of the blueprint blade was geometrically modified to reflect as-manufactured turbine blades using an established mesh metamorphosis algorithm. The approach uses a modified neural network to iteratively update the source mesh to the target mesh. The approach avoids the tedious manual approach of regenerating the CFD grid and does not rely on geometry obtained from Coordinate Measurement Machine (CMM) sections, but rather a point cloud representing the entirety of the turbine blade. Surface pressure traces and the discrete Fourier transforms thereof from numerical predictions of as-measured geometries are then compared both to blueprint predictions and to experimental measurements. The importance of incorporating as-measured geometries in analyses to explain deviations between numerical predictions of blueprint geometries and experimental results is readily apparent. Further analysis of every casting produced in the creation of the test turbine yields variations that one can expect in both aero-performance and unsteady loading as a consequence of manufacturing tolerances. Finally, the use of measured airfoil geometries to reduce the unsteady load on a target blade in a region of interest is successfully demonstrated.
TOPICS: Manufacturing, Turbochargers, Turbines, Computational fluid dynamics, Turbine blades, Blades, Fourier transforms, Geometry, Airfoils, Algorithms, Artificial neural networks, Stress, Coordinate measuring machines, Pressure, Machinery, Casting
Stefan Zerobin, Andreas Peters, Sabine Bauinger, Ashwini Bhadravati Ramesh, Michael Steiner, Franz Heitmeir and Emil Goettlich
J. Turbomach   doi: 10.1115/1.4039362
This two-part paper deals with the influence of high-pressure turbine purge flows on the aerodynamic performance of turbine center frames. Measurements were carried out in a product-representative one and a half stage turbine test setup. Four individual purge mass flows differing in flow rate, pressure, and temperature were injected through the hub and tip, forward and aft cavities of the unshrouded high-pressure turbine rotor. Two turbine center frame designs, equipped with non-turning struts, were tested and compared. In this first part of the paper the influence of different purge flow rates is discussed, while in the second part of the paper the impact of the individual hub and tip purge flows on the turbine center frame aerodynamics is investigated. The acquired measurement data illustrate that the interaction of the ejected purge flow with the main flow enhances the secondary flow structures through the turbine center frame duct. Depending on the purge flow rates, the radial migration of purge air onto the strut surfaces directly impacts the loss behavior of the duct. The losses associated with the flow close to the struts and in the strut wakes are highly dependent on the relative position between the high-pressure turbine vane and the strut leading edge, as well as the interaction between vane wake and ejected purge flow. This first-time experimental assessment demonstrates that a reduction in the purge air requirement benefits the engine system performance by lowering the turbine center frame total pressure loss.
TOPICS: Flow (Dynamics), Turbochargers, Turbines, High pressure (Physics), Struts (Engineering), Wakes, Pressure, Ducts, Cavities, Rotors, Aerodynamics, Temperature, Engines
Stefan Zerobin, Christian Aldrian, Andreas Peters, Franz Heitmeir and Emil Goettlich
J. Turbomach   doi: 10.1115/1.4039363
The aerodynamic behavior of turbine center frame ducts under the presence of high-pressure turbine purge flows was experimentally investigated in this two-part paper. While the first part of the paper demonstrated the impact of varying the purge flow rates on the loss behavior of two different turbine center frame designs, this second part concentrates on the influence of individual hub and tip purge flows on the main flow evolution and loss generation mechanisms through the turbine center frame ducts. Therefore, measurements were conducted at six different operating conditions in a one and a half stage turbine test setup, featuring four individual purge flows injected through the hub and tip, forward and aft cavities of the high-pressure turbine rotor. The outcomes of this first-time assessment indicate that a high-pressure turbine purge flow reduction generally benefits turbine center frame performance. Decreasing one of the rotor case purge flow rates leads to an improved duct pressure loss. The purge flows from the rotor aft hub and tip cavities are demonstrated to play a particularly important role for improving the duct aerodynamic behavior. In contrast, the forward rotor hub purge flow actually stabilizes the flow in the turbine center frame duct and reducing this purge flow can penalize turbine center frame performance. These particular high-pressure turbine/turbine center frame interactions should be taken into account whenever high-pressure turbine purge flow reductions are pursued.
TOPICS: Flow (Dynamics), Turbochargers, Turbines, High pressure (Physics), Ducts, Rotors, Cavities, Pressure, Performance
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
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
Technical Brief  
Santosh Patil, Ivana D. Atanasovska and Saravanan Karuppanan
J. Turbomach   doi: 10.1115/1.4030242
The aim of this paper is to provide a new viewpoint of friction factor for contact stress calculations of gears. The idea of friction factor has been coined, for the calculation of contact stresses along the tooth contact for different helical gear pairs. Friction factors were developed by evaluating contact stresses with and without friction for different gear pairs. In this paper, 3D Finite Element Method (FEM) and Lagrange Multiplier algorithm has been used to evaluate the contact stresses. Initially, a spur gear FE model was validated with the theoretical analysis under frictionless condition, which is based on Hertz's contact theory. Then, similar FE models were constructed for 5, 15, 25 and 35 deg. helical gear pairs. The contact stresses of these models were evaluated for different coefficients of friction. These results were employed for the development of friction factor.

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