Accepted Manuscripts

Hamid Hazby, Michael Casey and Luboš Březina
J. Turbomach   doi: 10.1115/1.4043786
The impact of the hub and shroud leakage flows on the compressor efficiency has been investigated for four compressor stages with flow coefficients of 0.017, 0.0265, 0.063 and 0.118 belonging to a family of centrifugal compressor stages, designed for process compressor applications. Very good agreement was observed between the measured and predicted performance when the detailed geometrical features were included in the calculations. The CFD calculations indicated that addition of leakage cavities and leakage flows resulted in about 3% drop in stage polytropic efficiency for the highest flow coefficient stage. The detrimental effect of leakages increased to about 8% for the lowest flow coefficient stage investigated here. The increase in the compressor work input due to the disc windage and the leakage recirculation was estimated from the CFD calculations and compared with values obtained using 1D methods, showing a very good agreement between the two. The impact of parasitic losses on compressor efficiency has been investigated and the contribution of various loss sources to the stage efficiency is discussed.
TOPICS: Compressors, Turbochargers, Leakage flows, Leakage, Flow (Dynamics), Computational fluid dynamics, Disks, Cavities
A. Duncan Walker, Bharat Koli and Peter A. Beecroft
J. Turbomach   doi: 10.1115/1.4043781
The drive for greater efficiencies means that aero engine overall pressure ratios are continuing to rise, and this also causes cycle temperatures to increase. A solution to managing this is to employ a cooled cooling air system. In such a system a purge flow into the main gas path, downstream of the compressor, will be required to prevent hot gas ingestion into rotor drive cavity. It is important to understand, and mitigate, the impact such a flow may have on the compressor outlet guide vanes (OGV), pre-diffuser and the downstream combustion system aerodynamics. Initial CFD predictions demonstrated the potential of the purge flow to modify the incidence and negatively affect the OGV such that the secondary flows increased. An experimental assessment using a fully annular, isothermal test facility confirmed the CFD results and demonstrated that the degradation in the combustor inlet flow resulted in an increased combustion system loss of up to 13%. To mitigate these effects CFD was used to examine the effect of varying the purge flow swirl fraction to better align the flow at the hub of the outlet guide vane. With a swirl fraction of 0.65 (x rotor speed) the secondary flows were reduced below that of the datum case (no purge flow). Experimental assessment agreed with the CFD but the measured effects had a smaller magnitude. However, differences in system loss were still measured with savings of around 10% for the feed to the turbine cooling ports.
TOPICS: Flow (Dynamics), Aerodynamics, Compressors, Turbochargers, High pressure (Physics), Diffusers, Combustion systems, Computational fluid dynamics, Rotors, Cooling, Guide vanes, Pressure, Temperature, Combustion chambers, Gates (Closures), Turbines, Cavities, Cycles, Test facilities, Aircraft engines
Riccardo Rubini, Roberto Maffulli and Tony Arts
J. Turbomach   doi: 10.1115/1.4043782
The study of the boundary layer transition plays a fundamental role in the field of turbomachinery. The main reason is the strong influence of the transition on the flow field local parameters, such as skin friction and heat transfer, this variation is reflected on the global ones such as efficiency and heat load of the blade row.Turbulent transition models are nowadays commonly used tools in both CFD research and design practice. It is then of particular interest to understand if they are able to predict the effect of temperature on bypass transition and, in case of positive answer, the reasons of their behaviour.This becomes even more interesting as the effect of the flow aero-thermal coupling becomes prominent in the analysis of such phenomena and is typically not accounted for in the validation of turbulence models. In this work we focus our attention on two state of the art transition model that use two radically different approaches to describe transition.To isolate the effects of the temperature ratio on the transition the simulations have been performed keeping the same values of Reynolds and Mach numbers and changing the value of the wall to freestream Temperature Ratio (TR).The results of the two transition models have been compared between them as well as with experimental results. They show that both models are sensitive to TR, though a locally based (rather than correlation based) approach for transition modelling should be favoured.
TOPICS: Flow (Dynamics), Mach number, Heat, Temperature, Heat transfer, Turbulence, Simulation, Stress, Turbochargers, Skin friction (Fluid dynamics), Boundary layers, Computational fluid dynamics, Engineering design processes, Engineering simulation, Modeling, Blades, Turbomachinery, Wall temperature
Maxime Fiore, Nicolas Gourdain, Jean-François Boussuge and Eric Lippinois
J. Turbomach   doi: 10.1115/1.4043660
This paper investigates the flow in a low-speed linear cascade rig with upstream hub cavity at a Reynolds number commonly observed in modern low pressure turbine stages by the use of numerical simulation. Numerical predictions are validated by comparing against experimental data available. Three different purge mass flow rates are tested using three different rim seal geometries. Numerical simulations are performed using a Large Eddy Simulation (LES) solver on structured grids. An investigation of the different mechanisms associated to turbine flow including cavity and purge air is intended through this simplified configuration. The underlying mechanisms of loss are tracked using an entropy generation formulation. Once described for a baseline case, the influence of purge flow and rim seal geometry on flow mechanisms and loss generation are described with the emphasis to obtain design parameters for losses reduction in the turbine. The study quantifies loss generation due to boundary layer on wetted surfaces and secondary vortices developing in the passage. The analysis shows different paths by which purge flow and rim seal geometry can change loss generation including a modification of the shear layer between purge and main stream, interaction with secondary vortices and a modification of the flow behavior close to the hub compared to a smooth configuration. The study shows the influence of purge flow rate and swirl on the strengthening of secondary vortices in the passage and the ability of axial overlapping rim seal geometries to delay the development of secondary vortices compared to simple axial gaps.
TOPICS: Flow (Dynamics), Turbochargers, Cascades (Fluid dynamics), Cavities, Vortices, Turbines, Geometry, Computer simulation, Reynolds number, Entropy, Shear (Mechanics), Boundary layers, Design, Large eddy simulation, Pressure, Delays
Dusan Perovic, Cesare A. Hall and Ewan Gunn
J. Turbomach   doi: 10.1115/1.4043644
Jet engines with Boundary Layer Ingestion (BLI) could offer significant reductions in aircraft fuel burn compared with podded turbofans. However, the engine fans must run continuously with severe inlet distortion, which is known to reduce stability. In this paper an experimental study has been completed on a low-speed rig fan operating with a BLI-type inlet distortion. Unsteady casing static pressure measurements have been made at multiple locations during stall events. Steady state, full-annulus area traverses have also been performed at rotor inlet and exit at a near-stall operating point. The reduction in stability caused by BLI is found to be small. It is found that with BLI the fan can operate stably despite the presence of localised regions where the rotor operating point lies beyond the stability boundary measured in clean flow. With the BLI-type distortion applied, the measured rotor incidence varies around the annulus due to non-uniform upstream swirl and velocity. The measured amplitude of unsteady casing pressure fluctuations just prior to stall is found to correlate with the circumferential variation of rotor incidence, suggesting that rotor incidence is a key variable affecting the creation and growth of flow disturbances. In regions of high incidence, disturbances resembling local flow separations are initiated. However, in regions of low or negative incidence, any disturbances decay rapidly. Full rotating stall with BLI occurs when high incidence regions are widespread enough to sustain disturbances which can propagate around the entire annulus.
TOPICS: Turbochargers, Boundary layers, Stall inception, Rotors, Annulus, Stability, Flow (Dynamics), Pressure measurement, Fuels, Engines, Fans, Fluctuations (Physics), Flow separation, Jet engines, Aircraft, Steady state, Turbofans, Pressure
Tianrui Sun, Paul Petrie-Repar, Damian M. Vogt and Anping Hou
J. Turbomach   doi: 10.1115/1.4043407
Blade flutter in the last stage is an important design consideration for the manufacturers of steam turbines. Therefore, the accurate prediction method for blade flutter is critical. Since the majority of aerodynamic work contributing to flutter is done near the blade tip, resolving the tip leakage flow can increase the accuracy of flutter predictions. Previous research has shown that the induced vortices in the tip region can have a significant influence on the flow field near the tip. The structure of induced vortices due to the tip leakage vortex cannot be resolved by URANS simulations because of the high dissipation in turbulence models. To the best of author's knowledge, the influence of induced vortices on flutter characteristics has not been investigated. In this paper, the results of DES and URANS flutter simulations of a realistic-scale last stage steam turbine are presented and the influence of induced vortices on the flutter stability is investigated. Significant differences for the predicted aerodynamic work coefficient distribution on the blade surface, especially on the rear half of the blade suction side near the tip are observed. At the least stable IBPA, the induced vortices show a destabilization effect on blade aeroelastic stabilities. The motion of induced vortices is dependent on the blade amplitude and hence the aerodynamic damping is also dependent on the blade vibration amplitude. In conclusion, the induced vortices can influence the predicted flutter characteristics of the steam turbine test case.
TOPICS: Stability, Eddies (Fluid dynamics), Blades, Steam turbines, Simulation, Turbochargers, Vortices, Flutter (Aerodynamics), Damping, Design, Vibration, Suction, Flow (Dynamics), Turbulence, Energy dissipation, Leakage flows, Leakage
Alexander Hergt, Sebastian Grund, Joachim Klinner, Wolfgang Steinert, Manfred Beversdorff and Ulrich Siller
J. Turbomach   doi: 10.1115/1.4043280
The development of axial compressors has already reached a high level. Therefore an enlargement of the design space by means of new or advanced aerodynamic methods is necessary in order to achieve further enhancements. The tandem arrangement of profiles in a transonic compressor blade row is such a method. It is necessary to address the design aspects a bit more in detail in order to efficiently apply this blading concept to turbomachinery. Therefore, in the current study the known design aspects of tandem blading in compressors will be summed up under consideration of the aerodynamic effects and construction characteristics. Based on this knowledge, a new transonic compressor tandem cascade (DLR TTC) with an inflow Mach number of 0.9 is designed using modern numerical methods and a multi objective optimization process. The aim of the ensuing experimental and numerical investigation is to answer the question, whether the tandem cascade resulting from the modern design process fulfils the described design aspects and delivers the requested performance and efficiency criteria. The experiments are performed at the Transonic Cascade Wind Tunnel of DLR in Cologne. The experimental results show an increase in losses and a reduction of the cascade deflection by about 2 degrees at the ADP compared to design concept. Due to the extremely high loading the cascade performance at off-design conditions are not achieved. The reasons for the discrepancies are discussed in the paper. Nevertheless, the experimental and numerical results allow a good understanding of the aerodynamic effects.
TOPICS: Design, Compressors, Turbochargers, Cascades (Fluid dynamics), Construction, Mach number, Numerical analysis, Blades, Deflection, Pareto optimization, Turbomachinery, Wind tunnels, Inflow
Andrew J. Saul, Peter T. Ireland, John D. Coull, Tsun Holt Wong, Haidong Li and Eduardo Romero
J. Turbomach   doi: 10.1115/1.4043263
The effect of film cooling on a transonic squealer tip has been examined in a high speed linear cascade, which operates at engine realistic Mach and Reynolds numbers. Tests have been performed on two uncooled tip geometries with differing pressure side rim edge radii, and a cooled tip matching one of the uncooled cases.The pressure sensitive paint technique has been used to measure adiabatic film cooling effectiveness on the blade tip at a range of tip gaps and coolant mass flow rates. Complementary tip heat transfer coefficients have been measured using transient infrared thermography, and the effects of the coolant film on the tip heat transfer and engine heat flux examined.The uncooled data show that the tip heat transfer coefficient distribution is governed by the nature of flow reattachments and impingements. The squealer tip can be broken down into three regions, each exhibiting a distinct response to a change in the tip gap, depending on the local behaviour of the overtip leakage flow. Complementary CFD shows that the addition of casing motion causes no change in flow over the pressure side rim.Injected coolant interacts with the overtip leakage flow, which can locally enhance the tip heat transfer coefficient. The film effectiveness is dependent on both the coolant mass flow rate and tip clearance. At increased coolant mass flow, areas of high film effectiveness on the pressure side rim coincide strongly with a net heat flux reduction and in the subsonic tip region with low heat transfer coefficient.
TOPICS: Turbochargers, Heat transfer coefficients, Film cooling, Coolants, Flow (Dynamics), Pressure, Leakage flows, Heat flux, Engines, Reynolds number, Thermography, Heat transfer, Cascades (Fluid dynamics), Transients (Dynamics), Clearances (Engineering), Computational fluid dynamics, Blades
Alexander Hergt, Joachim Klinner, Jens Wellner, Chris Willert, Sebastian Grund, Wolfgang Steinert and Manfred Beversdorff
J. Turbomach   doi: 10.1115/1.4043329
The flow through a transonic compressor cascade shows a very complex structure due to the occurring shock waves. In addition, the interaction of these shock waves with the blade boundary layer inherently leads to a very unsteady flow behaviour. The aim of the current investigation is to quantify this behaviour and its influence on the cascade performance. Therefore, an extensive experimental investigation of a transonic compressor cascade was performed. In this process, the flow phenomena were thoroughly examined for an inflow Mach number of 1.21. The experiments investigate both, the laminar as well as the turbulent shock wave boundary layer interaction within the blade passage and the resulting unsteady behaviour. The experiments show a fluctuation range of the passage shock wave of about 10 percent chord for both cases, which is directly linked with a change of the inflow angle/operating point of the cascade. In addition to the experiments, an unsteady simulation has been carried out in order to capture the unsteady flow behaviour. The results from this simulation show that the fluctuation of the passage shock wave can be reproduced but not in the correct magnitude. Due to this fact, there exist a delta in working range between the shock position where the numerical stall margin is reached and the averaged shock position of experimental cascade stall margin. This is a weak point within the design process of transonic compressor blades, because the working range will not be correct predicted.
TOPICS: Compressors, Turbochargers, Design, Blades, Cascades (Fluid dynamics), Shock waves, Simulation, Flow (Dynamics), Shock (Mechanics), Unsteady flow, Inflow, Boundary layers, Chords (Trusses), Mach number, Turbulence
Ziyou Wu, Dan Zhao and Shai Revzen
J. Turbomach   doi: 10.1115/1.4042696
Conventional wisdom would have it that moving mechanical systems which dissipate energy by Coulomb friction have no relationship between force and average speed. One could argue that the work done by friction is constant per unit of distance traveled, and if propulsion forces exceed friction, the net work is positive and the system accumulates kinetic energy without bound. We present a minimalistic model for legged propulsion with slipping under Coulomb friction, scaled to parameters representative of single kilogram robots and animals. Our model, amenable to exact solutions, exhibits nearly linear (R2>0.98) relationships between actuator force and average speed over its entire range of parameters, and in both motion regimes it supports. This suggests that the interactions inherent in multi-legged locomotion may lead to governing equations more reminiscent of viscous friction than would be immediately obvious.
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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