Accepted Manuscripts

Johan Dahlqvist and Jens Fridh
J. Turbomach   doi: 10.1115/1.4038468
The aspect of hub cavity purge has been investigated in a high-pressure axial low-reaction turbine stage. A full-scale cold-flow experimental rig featuring a rotating stage was used in the investigation, quantifying main annulus flow field impact with respect to purge flow rate as it was injected upstream of the rotor. Five operating speeds were investigated of which three with respect to purge flow, namely a high loading design case, and two high speed points encompassing the peak efficiency. At each of these operating speeds, the amount of purge flow was varied from 0% to 2%. The prominent effect due to purge is seen in the efficiency, showing a linear sensitivity to purge of 1.3%-points for every 1% of added purge flow for the investigated speeds. While spatial average values of flow angle and Mach number are essentially unaffected by purge injection, important spanwise variations are observed and highlighted. The secondary flow structure is strengthened in the hub region, leading to a generally increased over-turning and lowered flow velocity. Meanwhile, the added volume flow through the rotor leads to higher outlet flow velocities visible at higher span, with associated decreased turning. A radial efficiency distribution is utilized, showing negative impact through span heights from 15% to 70%. Pitchwise variation of investigated flow parameters is significantly influenced by purge flow, making this a parameter to include for instance when evaluating benefits of stator clocking positions.
TOPICS: Flow (Dynamics), Turbochargers, Turbines, Cavities, Rotors, Stators, Annulus, High pressure (Physics), Design, Mach number
Emmanuel Laroche, Matthieu Fenot, Eva Dorignac, Jean-Jacques Vuillerme, Laurent-emmanuel Brizzi and Juan-Carlos Larroya
J. Turbomach   doi: 10.1115/1.4038411
The present study aims at characterizing the flow field and heat transfer for a schematic but realistic vane cooling scheme. Experimentally, both velocity and heat transfer measurements are conducted to provide a detailed database of the investigated configuration. From a numerical point of view, the configuration is investigated using isotropic as well as anisotropic Reynolds-Averaged Navier-Stokes (RANS) turbulence models. An hybrid RANS/LES technique is also considered to evaluate potential unsteady effects. Both experimental and numerical results show a very complex 3D flow. Air is not evenly distributed between the different injections, mainly because of a large recirculation flow. Due to the strong flow deviation at the hole inlet, the velocity distribution and the turbulence characteristics at the hole exit are far from fully developed profiles. The comparison between PIV measurements and numerical results shows a reasonable agreement. However, coming to heat transfer, all RANS models exhibit a major overestimation compared to IR thermography measurements. The Billard-Laurence model does not bring any improvement compared to a classical k-? SST model. The hybrid RANS/LES simulation provides the best heat transfer estimation, exhibiting potential unsteady effects ignored by RANS models. Those conclusions are different from the ones usually obtained for a single fully developed impinging jet.
TOPICS: Flow (Dynamics), Heat transfer, Turbochargers, Turbines, Reynolds-averaged Navier–Stokes equations, Turbulence, Thermography, Simulation, Cooling, Databases, Anisotropy
Philip Bear, J. Mitch Wolff, Andreas Gross, Christopher Marks and Rolf Sondergaard
J. Turbomach   doi: 10.1115/1.4038413
Improvements in turbine design methods have resulted in the development of blade profiles with both high lift and good Reynolds lapse characteristics. An increase in aerodynamic loading of blades in the low pressure turbine section of aircraft gas turbine engines has the potential to reduce engine weight or increase power extraction. Increased blade loading means larger pressure gradients and increased secondary losses near the endwall. The present study analyzes the secondary flow field of the front-loaded low-pressure turbine blade designated L2F with and without blade profile contouring at the junction of the blade and endwall. Stereoscopic particle image velocimetry data and total pressure loss data are used to describe the secondary flow field. The flow is analyzed in terms of total pressure loss, vorticity, Q-Criterion, turbulent kinetic energy and turbulence production. The flow description is then expanded upon using an Implicit Large Eddy Simulation of the flow field. The RANS momentum equations contain terms with pressure derivatives. These equations can be rearranged to form an equation for the change in total pressure along a streamline as a function of velocity only. A comparison of the total pressure transport calculated from the velocity components and the total pressure loss is presented and discussed. Peak values of total pressure transport those of total pressure loss through and downstream of the passage suggesting that total pressure transport is a useful tool for localizing and predicting loss origins and loss development using velocity data which can be obtained non-intrusively.
TOPICS: Pressure, Turbochargers, Cascades (Fluid dynamics), Turbines, Blades, Flow (Dynamics), Turbulence, Engines, Kinetic energy, Turbine blades, Particulate matter, Momentum, Junctions, Pressure gradient, Reynolds-averaged Navier–Stokes equations, Large eddy simulation, Vorticity, Design methodology, Gas turbines, Aircraft, Weight (Mass)
Anirban Garai, Laslo T Diosady, Scott M Murman and Nateri Madavan
J. Turbomach   doi: 10.1115/1.4038403
ABSTRACT The application of a new computational capability for accurate and efficient high-fidelity scale-resolving simulations of turbomachinery is presented. The focus is on the prediction of heat transfer and boundary layer characteristics with comparisons to the experiments of Arts et al. for an uncooled, transonic, linear high-pressure turbine (HPT) inlet guide vane cascade that includes the effects of elevated inflow turbulence. The computational capability is based on an entropy-stable, discontinuous-Galerkin spectral-element approach that extends to arbitrarily high orders of spatial and temporal accuracy. The suction side of the vane undergoes natural transition for the clean inflow case, while bypass transition mechanisms are observed in the presence of elevated inflow turbulence. The airfoil suction-side boundary layer turbulence characteristics during the transition process thus differ significantly between the two cases. Traditional simulations based on the Reynolds-averaged Navier Stokes (RANS) fail to predict these transition characteristics. The heat transfer characteristics for the simulations with clean inflow agree well with the experimental data, while the heat transfer characteristics for the bypass transition cases agree well with the experiment when higher inflow turbulence levels are prescribed. The differences between the clean and inflow turbulence cases are also highlighted through a detailed examination of the characteristics of the transitional and turbulent flow fields.
TOPICS: Simulation, Turbochargers, Cascades (Fluid dynamics), High pressure (Physics), Engineering simulation, Turbines, Galerkin method, Inflow, Turbulence, Heat transfer, Suction, Entropy, Boundary layers, Airfoils, Inlet guide vanes, Reynolds-averaged Navier–Stokes equations, Turbomachinery, Boundary layer turbulence
Daniel Espinal, Hong-Sik Im and Gecheng Zha
J. Turbomach   doi: 10.1115/1.4038337
A high speed 1-1/2 axial compressor stage is simulated in this paper using an Unsteady Reynolds-Averaged Navier-Stokes (URANS) solver for a full-annulus configuration to capture its non-synchronous vibration (NSV) flow excitation with rigid blades. The predicted dominant frequencies using the blade tip response signals are not harmonic to the engine order, which is the NSV excitation. The simulation is based on a rotor blade with a 1.1% tip-chord clearance. Comparison with the previous 1/7 annulus simulations show that the time-shifted phase-lag BCs used in the 1/7 annulus are accurate. For most of the blades, the NSV excitation frequency is 6.2% lower than the measurement in the rig test, although some blades displayed slightly different NSV excitation frequencies. The simulation confirms that the NSV is a full annulus phenomenon. The instability of the circumferential traveling vortices in the vicinity of the rotor tip due to the strong interaction of incoming flow is the main cause of the NSV excitation. This instability is present in all blades of the rotor annulus. For circumferentially averaged parameters like total pressure ratio, NSV is observed to have an effect on the radial profile, particularly at radial locations above 70% span. A design with a lower loading of the upper blade span and a higher loading of the mid blade spans is recommended to mitigate or remove NSV.
TOPICS: Compressors, Simulation, Turbochargers, Vibration, Annulus, Blades, Excitation, Rotors, Flow (Dynamics), Engines, Arches, Clearances (Engineering), Chords (Trusses), Design, Pressure, Vortices, Signals
Daniel Möller, Maximilian Juengst, Heinz-Peter Schiffer, Thomas Giersch and Frank Heinichen
J. Turbomach   doi: 10.1115/1.4038316
Rotor blade vibrations observed in the Darmstadt transonic compressor rig are investigated in this paper. The vibrations are non-synchronous and occur in the near stall operating re- gion. Rotor tip flow fluctuations traveling near the leading edge against the direction of rotation (in the rotor relative frame of reference) with about 50% blade tip speed are found to be the reason for the occurrence of the vibrations. The investigations show, that the blockage at the rotor tip is an important factor for the aeroelastic stability of the compressor in the near stall region. It is found, that by application of a recirculating tip in- jection casing treatment, the aeroelastic stability increases as a result of reduced blockage in the rotor tip region.
TOPICS: Compressors, Turbochargers, Rotors, Vibration, Blades, Stability, Flow (Dynamics), Fluctuations (Physics), Rotation
Feng Wang, Mauro Carnevale, Luca di Mare and Simon Gallimore
J. Turbomach   doi: 10.1115/1.4038317
Computational Fluid Dynamics (CFD) has been widely used for compressor design, yet the prediction of performance and stage matching for multi-stage, high-speed machines remain challenging. This paper presents the authors' effort to improve the reliability of CFD in multistage compressor simulations. The endwall geometry features are meshed with minimal approximations. Turbulence models with linear and non-linear eddy viscosity models are assessed. The non-linear eddy viscosity model predicts a higher production of turbulent kinetic energy in the passages, especially close to the endwall region. This results in a more accurate prediction of the choked mass flow and the shape of total pressure profiles close to the hub. The non-linear viscosity model generally shows an improvement on its linear counterparts based on the comparisons with the rig data. For geometrical details, truncated fillet leads to thicker boundary layer on the fillet and reduced mass flow and efficiency. Shroud cavities are found to be essential to predict the right blockage and the flow details close to the hub. At the part speed the computations without the shroud cavities fail to predict the major flow features in the passage and this leads to inaccurate predictions of massflow and shapes of the compressor characteristic. The paper demonstrates that an accurate representation of the endwall geometry and an effective turbulence model, together with a good quality and sufficiently refined grid result in a credible prediction of compressor matching and performance with steady state mixing planes.
TOPICS: Compressors, Simulation, Turbochargers, Design, Flow (Dynamics), Turbulence, Viscosity, Computational fluid dynamics, Eddies (Fluid dynamics), Cavities, Geometry, Shapes, Steady state, Computation, Approximation, Boundary layers, Machinery, Pressure, Kinetic energy, Reliability
Martin Lange, Matthias Rolfes, Ronald Mailach and Henner Schrapp
J. Turbomach   doi: 10.1115/1.4038319
Since the early work on axial compressors the penalties due to radial clearances between blades and side walls are known and an ongoing focus of research work. The periodic unsteadiness of the tip clearance vortex, due to its interaction with the stator wakes, has only rarely been addressed in research papers so far. The current work presents experimental and numerical results from a four stage low speed research compressor modeling a state of the art compressor design. Time-resolved experimental measurements have been carried out at three different rotor tip clearances (gap to tip chord: 1.5%, 2.2%, 3.7%) to cover the third rotor's casing static pressure and exit flow field. These results are compared with either steady simulations using different turbulence models or harmonic RANS calculations to discuss the periodical unsteady tip clearance vortex development at different clearance heights. The prediction of the local tip leakage flow is clearly improved by the EARSM turbulence model compared to the standard SST model. The harmonic RANS calculations (using the SST model) improve the prediction of time-averaged pressure rise and are used to analyze the rotor stator interaction in detail. The interaction of the rotor tip flow field with the passing stator wakes cause a segmentation of the tip clearance vortex and result in a sinusoidal variation in blockage downstream the rotor row.
TOPICS: Compressors, Clearances (Engineering), Turbochargers, Vortices, Rotors, Stators, Reynolds-averaged Navier–Stokes equations, Wakes, Pressure, Flow (Dynamics), Turbulence, Simulation, Chords (Trusses), Design, Engineering simulation, Modeling, Blades, Image segmentation, Leakage flows
Quentin Rendu, Yannick Rozenberg, Stephane Aubert and Pascal Ferrand
J. Turbomach   doi: 10.1115/1.4038279
In order to predict oscillating loads on a structure, time-linearized methods are fast enough to be routinely used in design and optimization steps of a turbomachine stage. In this work, frequency-domain time-linearized Navier-Stokes computations are proposed to predict the unsteady separated flow generated by an oscillating bump in a transonic nozzle. We also investigate the interaction of backward travelling pressure waves and moving surface on the unsteady behavior of a strong shock-wave with separated boundary-layer. This case is representative of transonic stall flutter of a compressor blade submitted to downstream stator potential effects. The influence of frequency is first investigated on a generic oscillating bump to identify the most unstable configuration. Introducing backward travelling pressure waves, we then show that the aeroelastic stability of the system depends on the phase-shift between the waves source and the bump motion. Finally, we propose to actively control the instability by generating backward travelling pressure waves at prescribed amplitude, frequency and phase.
TOPICS: Acoustics, Turbochargers, Nozzles, Flutter (Aerodynamics), Waves, Pressure, Stability, Flow (Dynamics), Phase shift, Compressors, Shock waves, Stress, Boundary layers, Design, Optimization, Blades, Computation, Stators, Turbomachinery
Myeonggeun Choi, David R.H. Gillespie and Leo Lewis
J. Turbomach   doi: 10.1115/1.4038280
Thermal closure of the engine casing is widely used to minimize undesirable blade tip leakage flows thus improving jet engine performance. This may be achieved using an impingement cooling scheme on the external casing wall, provided by manifolds attached to the outside of the engine. The assembly tolerance of these components leads to variation in the standoff distance between the manifold and the casing and its effects on casing contraction must be understood to allow build tolerance to be specified. For cooling arrangements with promising performance, the variation in closure with standoff distance of z/d = 1 - 6 were investigated through a mixture of extensive numerical modelling and experimental validation. A cooling manifold, typical of that adopted by several engine companies, incorporating three different arrays of short cooling holes (chosen from previous study by Choi et al. (2016)) and thermal control dummy flanges were considered. Typical contractions of 0.5 - 2.2mm are achieved from the 0.02 - 0.35kg/s of the current casing cooling flows. The variation in heat transfer coefficient observed with standoff distance is much lower for the sparse array investigated compared to a previous designs employing arrays typical of blade cooling configurations. The reason for this is explained through interrogation of the local flow field and resultant heat transfer coefficient. This implies acceptable control of the circumferential uniformity of case cooling can be achieved with relatively large assembly tolerance of the manifold relative to the casing.
TOPICS: Turbochargers, Clearances (Engineering), Impingement cooling, Blades, Manifolds, Cooling, Engines, Manufacturing, Flow (Dynamics), Heat transfer coefficients, Leakage flows, Flanges, Jet engines, Modeling
Pedro M. Milani, Julia Ling, Gonzalo Saez-Mischlich, Julien Bodart and John Eaton
J. Turbomach   doi: 10.1115/1.4038275
In film cooling flows, it is important to know the temperature distribution resulting from the interaction between a hot main flow and a cooler jet. However, current Reynolds-averaged Navier-Stokes (RANS) models yield poor temperature predictions. A novel approach for RANS modeling of the turbulent heat flux is proposed, in which the simple gradient diffusion hypothesis (GDH) is assumed and a machine learning algorithm is used to infer an improved turbulent diffusivity field. This approach is implemented using three distinct data sets: two are used to train the model and the third is used for validation. The results show that the proposed method produces significant improvement compared to the common RANS closure, especially in the prediction of film cooling effectiveness.
TOPICS: Machinery, Turbulence, Turbochargers, Flow (Dynamics), Film cooling, Reynolds-averaged Navier–Stokes equations, Temperature distribution, Trains, Heat flux, Algorithms, Modeling, Temperature, Diffusion (Physics)
David Cerantola and Michael Birk
J. Turbomach   doi: 10.1115/1.4038277
Effusion cooling was a popular technology integrated into the design of gas turbine combustor liners. A staggering amount of research was completed that quantified performance with respect to operating conditions and cooling hole geometric properties; however, most of these investigations did not address the influence of the manufacturing process on the hole shape. This study completed an adiabatic wall numerical analysis using the realizable k-epsilon turbulence model of a laser-drilled hole that had a nozzled profile with an area ratio of 0.24 and five additional cylindrical, nozzled, diffusing, and filleted holes that yielded the same hole mass flow rate at representative engine conditions. The traditional methods for quantifying blowing ratio yielded the same value for all holes that was not useful considering the substantial differences in film cooling performance. It was proposed to define hole mass flux based on the outlet y-cross sectional area projected onto the inclination angle plane. This gave blowing ratios that correlated to better and worse cooling performance for the diffusing and nozzled holes respectively. The diffusing hole delivered the best film cooling due to having the lowest effluent velocity and greatest amount of in-hole turbulent production, which coincided with the worst discharge coefficient
TOPICS: Cooling, Turbochargers, Turbulence, Film cooling, Flow (Dynamics), Engines, Manufacturing, Lasers, Combustion chambers, Design, Gas turbines, Numerical analysis, Discharge coefficient, Shapes
Chao-Cheng Shiau, Andrew F Chen, Je-Chin Han, Salam Azad and Ching-Pang Lee
J. Turbomach   doi: 10.1115/1.4038278
Turbine vane endwalls are highly susceptible to intensive heat load due to their large exposed area and complex flow field especially for the first-stage of the vane. Therefore, a suitable film cooling design that properly distributes the given amount of coolant is critical to keep the vane endwall from failure at the same time to maintain a good balance between manufacturing cost, performance, and durability. This work is focused on film cooling effectiveness evaluation on full-scale heavy duty turbine vane endwall and the performance comparison with different film cooling pattern designs in the literature. The area of interest of this study is on the inner endwall (hub) of turbine vane. Tests were performed in a three-vane annular sector cascade under the mainstream Reynolds number 350,000; the related inlet Mach number is 0.09 and the freestream turbulence intensity is 12%. Two variables, coolant-to-mainstream mass flow ratios (MFR = 2%, 3%, 4%) and density ratios (DR = 1.0, 1.5) are investigated. The conduction-error free Pressure-sensitive paint (PSP) technique is utilized to evaluate the local flow behavior as well as the film cooling performance. The presented results are expected to provide the gas turbine engine designer a direct comparison between two film-hole configurations on a full-scale vane endwall under the same amount of coolant usage.
TOPICS: Turbochargers, Turbines, Film cooling, Coolants, Flow (Dynamics), Mach number, Heat, Turbulence, Manufacturing, Reynolds number, Heat conduction, Stress, Errors, Failure, Density, Pressure, Cascades (Fluid dynamics), Design, Durability, Gas turbines
Justin Varty, Loren W. Soma, Forrest Ames and Dr. Sumanta Acharya
J. Turbomach   doi: 10.1115/1.4038281
Secondary flows in vane passages sweep off the endwall and onto the suction surface having an immediate impact on heat transfer. The present paper documents the impact of secondary flows on suction surface heat transfer acquired over a range of turbulence levels (0.7% through 17.4%) and a range of exit chord Reynolds numbers (500,000 through 2,000,000). The vane design includes an aft loaded suction surface and a large leading edge diameter. The heat transfer visualization for the heated endwall condition shows no initial high heat transfer level near the edge of heating on the vane. The heat transfer level in the region affected by the secondary flows is largely uniform, except for a notable depression in an affected region believed due to an upwash region generated above the separation line of the passage vortex, likely in conjunction with the counter rotating suction leg of the horseshoe vortex. The extent and definition of the secondary flow affected region on the suction surface is clearly evident at lower Reynolds numbers and lower turbulence levels when the suction surface flow is largely laminar. The heat transfer in the plateau region has a magnitude similar to a turbulent boundary layer. However, the location and extent of this secondary flow affected region is less perceptible at higher turbulence levels where transitional or turbulent flow is present. Also, aggressive mixing at higher turbulence levels serves to smooth out discernable differences in the heat transfer due to the secondary flows.
TOPICS: Turbulence, Suction, Reynolds number, Turbochargers, Flow (Dynamics), Heat transfer, Boundary-value problems, Vortices, Boundary layer turbulence, Separation (Technology), Chords (Trusses), Design, Visualization, Heating
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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