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Accepted Manuscripts

BASIC VIEW  |  EXPANDED VIEW
research-article  
Nafiz H. K. Chowdhury, Hootan Zirakzadeh and Je-Chin Han
J. Turbomach   doi: 10.1115/1.4036302
The growing trend to achieve a higher Turbine Inlet Temperature (TIT) in the modern gas turbine industry requires a more efficient and advanced cooling system design. Therefore, a complete study of heat transfer is necessary to predict the thermal loadings on the gas turbine vanes and blades. In the current work, a predictive model for the gas turbine blade cooling analysis has been developed. The model is capable of calculating the distribution of coolant mass flow rate and metal temperatures of a turbine blade using the mass and energy balance equations at given external and internal boundary conditions. Initially, the performance of the model is validated by demonstrating its capability to predict the temperature distributions for a NASA E3 blade. The model is capable of predicting the temperature distributions with reasonable accuracy, especially on the suction side. Later, this paper documents the overall analysis for the same blade profile but at different boundary conditions to demonstrate the flexibility of the model for other cases. Additionally, guidelines are provided to obtain external HTC distributions for highly turbulent mainstream.
TOPICS: Cooling, Turbochargers, Blades, Gas turbines, Boundary-value problems, Temperature distribution, Temperature, Heat transfer, Turbine blades, Design, Metals, Cooling systems, Turbulence, Suction, Energy budget (Physics), Coolants, NASA, Turbines, Flow (Dynamics)
research-article  
Weihong Li, Li Yang, Xueying Li, Jing Ren and Hongde Jiang
J. Turbomach   doi: 10.1115/1.4036297
This study comprehensively illustrates the effect of Reynolds number, hole spacing, jet-to-target distance and target plate thickness on the conjugate heat transfer performance of an impinging jet array. Test model is constructed with a relatively high conductivity material so that the Biot number of the models match engine condition. Highly resolved temperature distributions on the target plate are obtained utilizing steady liquid crystal over a range of Reynolds numbers varying between 5,000 and 27,5000. Effect of streamwise and spanwise jet-to-jet spacing (X/D, Y/D: 4-8), jet-to-target plate distance (Z/D: 0.75-3) and target plate thickness (t/D: 0.75-2.75) are employed composing a test matrix of 108 different geometries. Measured data are utilized as boundary conditions to conduct finite element simulation. Local and averaged non-dimensional temperature and averaged temperature uniformity of target plate “hot side” are obtained. Optimum hole spacing arrangements, impingement distance and target plate thickness are pointed out to minimize hot side temperature, the amount of cooling air and maximize the temperature uniformity. Also included are 2D predictions with different convective boundary conditions, i.e. row-averaged and local heat transfer coefficients, to estimate the accuracy of temperature prediction in comparison with the conjugate results.
TOPICS: Heat transfer, Cooling, Reynolds number, Turbochargers, Optimization, Temperature, Boundary-value problems, Temperature uniformity, Heat transfer coefficients, Temperature distribution, Thermal conductivity, Finite element analysis, Electrical conductivity, Simulation, Liquid crystals, Engines
research-article  
Graham Pullan and John J. Adamczyk
J. Turbomach   doi: 10.1115/1.4036296
A class of problems in turbomachinery is characterised by unsteady interactions at low reduced frequencies. These interactions are often the result of perturbations with length-scale on the order of the machine circumference and examples include axial compressors operating with inlet distortion, fans with downstream pylons, and turbine rotors downstream of mid-frame struts. Typically, this unsteadiness is accompanied by higher frequency fluctuations caused by perturbations with a length-scale on the order of a blade pitch. Conventional numerical analysis of this class of problem requires computations with a time step governed by the high frequency content but a greatly reduced run time could be achieved if the time step were dictated solely by the low reduced frequency, long length-scale, interaction of interest. In this paper, a filtering mixing plane technique is proposed that removes unwanted short length-scale perturbations at the interfaces between blade rows. This approach gives the user control over the amount of mixing that occurs at these interfaces with the limits being fully mixed-out to pitchwise uniformity (conventional mixing plane) or no mixing (conventional sliding plane). By choosing to retain only enough harmonics to resolve the low reduced frequency interaction of interest, an order of magnitude reduction in run time can be achieved.
TOPICS: Turbochargers, Filtration, Turbomachinery, Blades, Computation, Compressors, Fluctuations (Physics), Fans, Numerical analysis, Rotors, Turbines, Machinery
research-article  
Shane E. Haydt, Stephen Lynch and Scott D. Lewis
J. Turbomach   doi: 10.1115/1.4036199
Shaped film cooling holes are used extensively in gas turbines to reduce component temperatures. These holes generally consist of a metering section through the material and a diffuser to spread coolant over the surface. These two hole features are created separately using electrical discharge machining, and occasionally an offset can occur between the meter and diffuser due to misalignment. The current study examines the potential impact of this manufacturing defect to the film cooling effectiveness for a well-characterized shaped hole known as the 7-7-7 hole. Five meter-diffuser offset directions and two offset sizes were examined, both computationally and experimentally. Adiabatic effectiveness measurements were obtained at a density ratio of 1.2 and blowing ratios ranging from 0.5 to 3. The detriment in cooling relative to the baseline 7-7-7 hole was worst when the diffuser was shifted upstream (aft meter-diffuser offset), and least when the diffuser was shifted downstream (fore meter-diffuser offset). At some blowing ratios and offset sizes, the fore meter-diffuser offset resulted in slightly higher adiabatic effectiveness than the baseline hole, due to a reduction in the high-momentum region of the coolant jet caused by a separation region created inside the hole by the fore meter-diffuser offset. Steady RANS predictions did not accurately capture the levels of adiabatic effectiveness or the trend in the offsets, but it did predict the fore offset's improved performance.
TOPICS: Turbochargers, Diffusers, Film cooling, Coolants, Gas turbines, Electrical discharge machining, Reynolds-averaged Navier–Stokes equations, Density, Momentum, Temperature, Cooling, Separation (Technology), Manufacturing
research-article  
Kyle Chavez, Thomas N. Slavens and David G. Bogard
J. Turbomach   doi: 10.1115/1.4036200
Understanding the sensitivity of film cooling to a range of inlet conditions is necessary when verifying the robustness of a cooling design. In order to do this, adiabatic and overall effectiveness data were measured at various blowing ratios on an airfoil with nine shaped hole rows. This was performed in a low-speed linear cascade at two incidence angles, an inlet Reynolds number of 120000, a turbulence intensity of 5%, and a density ratio of 1.23. The test section was first adjusted so that the airfoil pressure distribution and stagnation line matched a CFD model. Infrared thermography was then used to measure effectiveness levels. This process was then repeated for the second incidence angle. Airfoil inlet pressures for each blowing ratio were matched between incidence angles. The stagnation line position changed the laterally averaged adiabatic effectiveness by as much as 0.2 near the showerhead. The effect persisted strongly 35 hole diameters downstream of the stagnation row but was visible over the whole viewable area of 160 diameters. The showerhead was then investigated in detail and it was found that the stagnation line dramatically increased the near-hole adiabatic and overall effectiveness levels when positioned within the breakout footprint of the stagnation hole row. This is the first study to present measured effectiveness values over both the pressure- and suction-sides of a fully-cooled airfoil for appreciably off-nominal incidence angles as well as examine adiabatic and overall effectiveness levels for a conical stagnation row of holes.
TOPICS: Turbochargers, Turbines, Airfoils, Pressure, Cooling, Turbulence, Suction, Reynolds number, Thermography, Cascades (Fluid dynamics), Computational fluid dynamics, Design, Film cooling, Robustness, Density
research-article  
Anna M. Young, Teng Cao, Ivor J. Day and John P. Longley
J. Turbomach   doi: 10.1115/1.4036201
In this paper, experiments and numerical modelling are used to quantify the effects of clearance and eccentricity on compressor performance and to examine the influence of each on flow distribution and stall margin. A change in the size of the tip-clearance gap influences the pressure rise and the stall margin of a compressor. Eccentricity of the tip-clearance gap then further exacerbates the negative effects of increasing tip-clearance. There are few studies dealing with the combined effect of clearance and eccentricity. There is also little guidance for engine designers, who have traditionally used rules of thumb to quantify these effects. One such rule states the stall margin of an eccentric machine to be equal to that of a concentric machine with uniform clearance equal to the maximum eccentric clearance. In this paper, this rule of thumb is checked using experimental data and found to be overly pessimistic. In addition, eccentric clearance causes a variation in axial velocity around the circumference of the compressor. The current study uses a three dimensional model which demonstrates the importance of radial flow gradients in capturing this redistribution. The circumferential variation in axial velocity is also examined in terms of the local stability of the flow. The large clearance sector of the annulus is found to operate beyond its equivalent axisymmetric stall limit, which means that the small clearance sector of the annulus must be stabilising the large clearance sector. An improved rule of thumb dealing with the effects of eccentricity is presented.
TOPICS: Compressors, Turbochargers, Accounting, Clearances (Engineering), Flow (Dynamics), Machinery, Annulus, Radial flow, Three-dimensional models, Engines, Pressure, Stability, Modeling
research-article  
Christopher J. Clark, Graham Pullan, Eric M. Curtis and Frederic Goenaga
J. Turbomach   doi: 10.1115/1.4036190
Low aspect ratio vanes, often the result of overall engine architecture constraints, create strong secondary ?ows and high endwall loss. In this paper, a splitter concept is demonstrated that reduces secondary ?ow strength and improves stage performance. The total secondary kinetic energy of the secondary ?ow vortices is reduced when the number of passages is increased and, for a given number of vanes, when the inlet endwall boundary layer is evenly distributed between the passages. Viscous computations show that, for this to be achieved in a splitter con?guration, the pressure-side leg of the low aspect ratio vane horseshoe vortex, must enter the adjacent passage (and not “jump” in front of the splitter leading edge). For a target turbine application, four vane designs were produced using a multi-objective optimization approach. These designs represent: current practice for a low aspect ratio vane; a design exempt from thickness constraints; and two designs incorporating splitter vanes. Each geometry is tested experimentally, as a sector, within a low-speed turbine stage. The vane designs with splitters geometries were found to reduce the measured secondary kinetic energy, by up to 85%, to a value similar to the design exempt from thickness constraints. The resulting ?ow?eld was also more uniform in both the circumferential and radial directions. One splitter design was selected for a full annulus test where a mixed-out loss reduction, compared to the current practice design, of 15.3% was measured and the stage ef?ciency increased by 0.88%.
TOPICS: Pressure, Engines, Kinetic energy, Turbochargers, Boundary layers, Design, Turbines, Vortices, Annulus, Computation, Flow control, Geometry, Pareto optimization
research-article  
Hubert Mathias Diefenthal, Piotr Luczynski, Christian Rakut, Manfred Wirsum and Tom Heuer
J. Turbomach   doi: 10.1115/1.4036104
In turbomachinery design the accurate prediction of the life cycle is one of the most challenging issues. Traditionally, life cycle calculations for radial turbine wheels of turbochargers focus on mechanical loads such as centrifugal and vibration forces. Due to the increase in exhaust gas temperatures in the last years, thermomechanical fatigue in the turbine wheel came more into focus. In order to account for the thermally induced stresses in the turbine wheel as a part of the standard design process, a fast method is required for predicting metal temperatures. In order to develop a suitable method, the mechanisms have to be understood that cause the thermal stresses. Thus, in a first step a detailed analysis of the temperature fields is conducted in the present paper. Extensive numerical simulations of a thermal shock process are carried out and validated by experimental data from a test rig. Based on the results the main heat transfer mechanisms are identified, that are crucial for the critical thermal stresses in transient operation. It is shown that these critical stresses mainly depend on local 3D flow structures. With this understanding, a fast method to calculate the transient temperatures in a radial turbine was developed. It is based on a standard method for transient fluid/solid heat transfer. This standard method was modified in order to achieve a sufficient accuracy in the calculation of the investigated heat transfer processes. The results show a good agreement with experimental data and with the results of the extensive numerical calculations.
TOPICS: Temperature, Turbochargers, Transients (Dynamics), Turbines, Wheels, Stress, Heat transfer, Thermal stresses, Design, Cycles, Exhaust systems, Thermal shock, Turbomachinery, Vibration, Flow (Dynamics), Fatigue, Fluids, Metals, Computer simulation
research-article  
Francesco Papa, Umesh Madanan and Richard Goldstein
J. Turbomach   doi: 10.1115/1.4036106
Measurements of the mass/heat transfer coefficients on the blade and endwall surfaces of a linear turbine cascade are compared to numerical predictions using the standard Shear Stress Transport (SST) closure and the SST model in combination with the Re?-? transition model. Experiments were carried out in a wind tunnel test section composed of five large-scale turbine blades, using the naphthalene sublimation technique. Two cases were tested, with exit Reynolds number of 600,000 and inlet turbulence values of 0.2% and 4% respectively. The main secondary flow features, consisting of the horseshoe vortex system, the passage vortex and the corner vortices are identified and their influence on heat/mass transfer is analyzed. Numerical simulations were carried out to match the conditions of the experiments. Results show that large improvements are obtained with the introduction of the Re?-? transition model. In particular, excellent agreement with the experiments is found, for the whole spanwise extension of the blade, on the pressure surface. On the suction surface, performance is very good in the highly three-dimensional region close to the endwall, but some weaknesses appear in predicting the location of transition in the two dimensional region. On the endwall surface, the SST model in combination with the transition model produces satisfactory results, greatly improved compared to the standard SST model.
TOPICS: Heat, Mass transfer, Turbochargers, Modeling, Turbines, Cascades (Fluid dynamics), Vortices, Blades, Wind tunnels, Heat transfer coefficients, Shear stress, Corners (Structural elements), Turbine blades, Pressure, Flow (Dynamics), Turbulence, Suction, Computer simulation, Reynolds number
research-article  
Tien Dat Phan, Patrick Springer and Robert Liebich
J. Turbomach   doi: 10.1115/1.4036107
In order to prevent critical effects due to pulsed detonation propulsion, e.g. incidence fluctuations, an elastomer-piezo-adaptive stator blade with a deformable front part is developed. Numerical investigations with respect to the interaction of fluid and structure including the piezoelectric properties and the hyperelastic material behavior of an elastomer membrane are conducted in order to investigate a concept of the elastomer-piezo-adaptive blade for developing the best suitable concept for subsequent experiments with a stator cascade in a wind tunnel. Results of numerical investigations of the structure-dynamic and fluid mechanical behavior of the elastomer-piezo-adaptive blade by using a novel Fluid-Structure-Piezoelectric-Elastomer-Interaction-Simulation (FSPEI-Simulation) show that the latent danger of a laminar flow separation at the leading edge at incidence fluctuations can be prevented by using an adaptive blade. Therefore, the potential of the concept of the elastomer-piezo-adaptive blade for active flow control is verified. Furthermore, it is essential to consider the interactions between fluid and structure of the transient FSPEI-Simulations, since not only the deformation of the adaptive blade affects the flow around the blade, the flow has a significant effect on the dynamic behavior of the adaptive blade, as well.
TOPICS: Fluids, Simulation, Turbochargers, Elastomers, Blades, Flow control, Unsteady flow, Stators, Fluctuations (Physics), Flow (Dynamics), Deformation, Separation (Technology), Explosions, Propulsion, Cascades (Fluid dynamics), Transients (Dynamics), Laminar flow, Fluid mechanics, Membranes, Wind tunnels
research-article  
Jeremy Nickol, Randall M. Mathison, Michael Dunn, Jong Liu and Malak Malak
J. Turbomach   doi: 10.1115/1.4036109
Cooling flow behavior is investigated within multiple serpentine passages with turbulators on the leading and trailing walls of an axial gas turbine blade operating at design-corrected conditions with accurate external flow conditions. Pressure and temperature measurements at midspan within the passages are obtained using miniature butt-welded thermocouples and miniature Kulite pressure transducers. These measurements are used with airfoil surface pressure distributions from a full CFD simulation as boundary conditions for a model that provides quantitative values of film-cooling blowing ratio for each film cooling hole on the blade. The model accounts for the continuously changing cross-sectional area and shape of the channels, frictional pressure loss, convective heat transfer from the solid portion of the blade, massflow reduction as coolant bleeds out through film-cooling or impingement holes, compressibility effects, and the effects of blade rotation. The results provide detailed coolant ejection information for a film-cooled turbine airfoil rotating at design-corrected conditions, and also accounts for the variable freestream conditions on the airfoil. While these values are commonly known for simpler experimental geometries, they have previously either been unknown or estimated crudely for full-stage experiments of this nature. The quantified cooling parameters provide a bridge for better comparison with the wealth of film-cooling work already reported for simplified geometries. The calculation also shows the significant range in blowing ratio that arises among a single row of cooling holes associated with one passage due to significant changes in both coolant and local freestream massfluxes.
TOPICS: Coolants, Turbochargers, High pressure (Physics), Blades, Gas turbines, Film cooling, Airfoils, Pressure, Cooling, Flow (Dynamics), Design, Rotation, Compressibility, Boundary-value problems, Shapes, Thermocouples, Bridges (Structures), Temperature measurement, Pressure transducers, Simulation, Computational fluid dynamics, Convection, Turbines
research-article  
Bai-Tao AN, Jian-Jun LIU and Si-Jing ZHOU
J. Turbomach   doi: 10.1115/1.4036007
This paper presents an experimental investigation of the rectangular diffusion hole. The effects of rectangular aspect ratio and lateral diffusion angle on film cooling effectiveness were studied at a low-speed flat plate experimental facility. The pressure sensitive paint measurement technique was employed to determine the adiabatic effectiveness. The experiments were performed at a density ratio of DR=1.38 and a mainstream turbulence intensity of Tu=3.5%. The blowing ratio was varied from M=0.5 to M=2.5. Three aspect ratios and three lateral diffusion angles were chosen to match the semi-circle and straight line sidewall shape of the rectangular cross section. A comparative investigation was performed among a typical fan-shaped hole and ten rectangular diffusion holes. The experimental results exhibited diversified film distribution patterns of the rectangular diffusion hole, including single-, bi-, and tri-peak patterns. The overall cooling effectiveness increased with the increase of rectangular aspect ratio. The improved magnitude was amplified as blowing ratio increased. The holes with semi-circle sidewall were shown to be more suitable for high blowing ratio conditions. The maximum increase of cooling effectiveness was over 70% compared to the fan-shaped hole. The reduction of the lateral diffusion angle affected the film distribution pattern significantly, thereby influenced the cooling effectiveness. To obtain a fixed coverage ratio of film hole row, the rectangular diffusion hole with a larger cross-sectional aspect ratio and a slightly smaller lateral diffusion angle is a preferred scheme.
TOPICS: Diffusion (Physics), Turbochargers, Film cooling, Cooling, Turbulence, Density, Pressure, Flat plates, Shapes
research-article  
Robin Prenter, Ali Ameri and Jeffrey P. Bons
J. Turbomach   doi: 10.1115/1.4036008
Ash particle deposition in a high-pressure turbine stage was numerically investigated using steady (RANS) and unsteady (URANS) methods. An inlet temperature profile consisting of Gaussian non-uniformities (hot streaks) was imposed on the vanes, with vane cooling simulated using a constant vane wall temperature. The steady case utilized a mixing plane at the vane-rotor interface, while a sliding mesh was used for the unsteady case. Corrected speed and mass flow were matched to an experiment involving the same geometry, so that the flow solution could be validated against measurements. Particles ranging from 1 to 65 µm were introduced into the vane domain, and tracked using an Eulerian-Lagrangian tracking model. A novel particle rebound and deposition model was employed to determine particles' stick/bounce behavior upon impact with a surface. Predicted impact and capture distributions for different diameters were compared between the steady and unsteady methods, highlighting effects from the circumferential averaging of the mixing plane. The mixing plane simulation was found to over predict impact and capture efficiencies compared with the unsteady calculation, as well as over predict particle temperature upon impact with the blade surface. Blade impact efficiencies increased with higher Stokes numbers in both simulations, with multiple rebounds occurring on the pressure surface in the mixing plane case, and on the suction surface in the unsteady case.
TOPICS: Simulation, Turbochargers, High pressure (Physics), Turbines, Particulate matter, Flow (Dynamics), Blades, Geometry, Reynolds-averaged Navier–Stokes equations, Temperature profiles, Wall temperature, Pressure, Temperature, Cooling, Suction, Rotors
research-article  
Jianhui Qi, Thomas Reddell, Kan Qin, Kamel Hooman and Ingo Jahn
J. Turbomach   doi: 10.1115/1.4035920
Supercritical CO2 cycles are considered a promising technology for next generation concentrated solar thermal, waste heat recovery and nuclear applications. Particularly at small scale, where radial inflow turbines can be employed, using sCO2 results in both system advantages and simplifications of the turbine design, leading to improved performance and cost reductions. This paper aims to provide new insight towards the design of radial turbines for operation with sCO2 in the 100~200kW range. The quasi one dimensional mean line design code TOPGEN is enhanced to explore and map the radial turbine design space. This mapping process over a state space defined by Head and Flow coefficients allows the selection of an optimum turbine design, while balancing performance and geometrical constraints.By considering three operating points with varying power levels and rotor speeds the effect of these on feasible design space and performance is explored. This provides new insight towards the key geometric features and operational constraints that limit the design space as well as scaling effects. Finally review of the loss break-down of the designs elucidates the importance of the respective loss mechanisms. Similarly it allows the identification of a design directions that lead to improved performance. Overall this work has shown that turbine design with efficiencies in the range 78~82% are possible in this power range and provides insight into the design space that allows the selection of optimum designs.
TOPICS: Turbochargers, Design, Turbines, Supercritical carbon dioxide, Concentrating solar power, Cycles, Inflow, Rotors, Flow (Dynamics), Heat recovery
research-article  
Jeffrey P. Bons, Robin Prenter and Steven Whitaker
J. Turbomach   doi: 10.1115/1.4035921
A new model is proposed for predicting particle rebound and deposition in environments relevant for gas turbine engines. The model includes the following physical phenomena: elastic deformation, plastic deformation, adhesion, and shear removal. It also incorporates material property sensitivity to temperature and tangential-normal rebound velocity cross-dependencies observed in experiments. The model is well-suited for incorporation in CFD simulations of complex gas turbine flows due to its algebraic (explicit) formulation. Model predictions are compared to coefficient of restitution data available in the open literature as well as deposition results from two different high temperature turbine deposition facilities. While the model comparisons with experiments are in many cases promising, several key aspects of particle deposition remain elusive. The simple phenomenological nature of the model allows for parametric dependencies to be evaluated in a straightforward manner. It is hoped that this feature of the model will aid in identifying and resolving the remaining stubborn holdouts that prevent a universal model for particle deposition.
TOPICS: Physics, Turbochargers, Particulate matter, Turbomachinery, Gas turbines, Deformation, Temperature, Adhesion, Flow (Dynamics), Shear (Mechanics), Materials properties, Computational fluid dynamics, Engineering simulation, Turbines, Simulation, High temperature, Algebra
research-article  
Charles E. Seeley, Christian Wakelam, Xuefeng Zhang, Douglas Hofer and Wei-Min Ren
J. Turbomach   doi: 10.1115/1.4035840
Flutter is a self-excited and self-sustained aero-elastic instability, caused by the positive feedback between structural vibration and aerodynamic forces. A two-passage linear turbine cascade was designed, built and tested to better understand the phenomena and collect data to validate numerical models. The cascade featured a center airfoil that had its pitch axis as a degree of freedom to enable coupling between the air flow and mechanical response in a controlled manner. The airfoil was designed to be excited about its pitch axis using an electromagnetic actuation system over a range of frequencies and amplitudes. The excitation force was measured with load cells and the airfoil motion was measured with accelerometers. Extraordinary effort was taken to minimize the mechanical damping so that the damping effects of the airflow over the airfoil, that were of primary interest, would be observable. Assembling the cascade required specialized alignment procedures due to the tight clearances and large motion. The aerodynamic damping effects were determined by observing changes in the mechanical frequency response of the system. Detail aero and mechanical measurements were conducted within a wide range of flow conditions. Experimental results indicate interesting changes in aerodynamic damping over a range of Mach numbers from 0.4 to 1.2. The aero damping was also found to be independent of displacement amplitude within the tested range, giving credence to linear numerical approaches.
TOPICS: Turbochargers, Turbine blades, Flutter (Aerodynamics), Damping, Experimental characterization, Airfoils, Cascades (Fluid dynamics), Air flow, Stress, Degrees of freedom, Turbines, Vibration, Displacement, Feedback, Frequency response, Fluid-dynamic forces, Flow (Dynamics), Mach number, Aerodynamics, Computer simulation, Accelerometers, Excitation
research-article  
Holger Werschnik, Jonathan Hilgert, Manuel Wilhelm, Martin Bruschewski and Heinz-Peter Schiffer
J. Turbomach   doi: 10.1115/1.4035832
At the Large Scale Turbine Rig (LSTR) at Technische Universiẗat Darmstadt the aerothermal interaction of combustor exit flow conditions on the subsequent turbine stage is examined. The rig resembles a high pressure turbine and is scaled to low Mach numbers. A baseline configuration with axial inflow and a swirling inflow representative for a lean combustor is modeled by swirl generators, whose clocking position towards the NGV leading edge can be varied. A staggered double-row of cylindrical film cooling holes on the endwall is examined. The effect of swirling inflow on heat transfer and film cooling effectiveness is studied, while the coolant mass flux rate is varied. Nusselt numbers are calculated using infrared thermography and the auxiliary wall method. Boundary layer, turbulence and five-hole probe measurements as well as numerical simulations complement the examination. The results for swirling inflow show a decrease of film cooling effectiveness of up to 40 % and an increase of Nusselt numbers of 10-25% in comparison to the baseline case for low coolant mass flux rates. For higher coolant injection, the heat transfer is on a similar level as the baseline. The differences vary depending on the clocking position. The turbulence intensity is increased to 30 % for swirling inflow.
TOPICS: Heat transfer, Turbochargers, Combustion chambers, Turbines, Film cooling, Inflow, Swirling flow, Coolants, Nozzle guide vanes, Boundary layer turbulence, Generators, Probes, High pressure (Physics), Flow (Dynamics), Mach number, Turbulence, Computer simulation, Thermography
research-article  
Greg Natsui, Zachary Little, Jayanta S. Kapat and Jason E. Dees
J. Turbomach   doi: 10.1115/1.4035520
Adiabatic film cooling effectiveness measurements are obtained using pressure-sensitive paint (PSP) on a flat film cooled surface. The effects of blowing ratio and hole spacing are investigated for four multi-row arrays comprised of 8 rows containing 52 holes of 3.8 mm diameter with 20º inclination angles and hole length-to-diameter ratio of 11.2. The four arrays investigated have two different hole-to-hole spacings composed of cylindrical and diffuser holes. For the first case, lateral and streamwise pitches are 7.5 times the diameter. For the second case, pitch-to-diameter ratio is 14 in lateral direction and 10 in the streamwise direction. The holes are in a staggered arrangement. Adiabatic effectiveness measurements are taken for a blowing ratio range of 0.3 to 1.2 and a density ratio of 1.5, with CO2 injected as the coolant. Local effectiveness, laterally averaged effectiveness, boundary layer thickness, momentum thickness, turbulence intensity and turbulence length scale are presented. For the cylindrical holes, at the first row of injection, the film jets are still attached at a blowing ratio of 0.3. By a blowing ratio of 0.5, the jet is observed to lift off, and then impinge back onto the test surface. At a blowing ratio of 1.2, the jets lift off, but reattach much further downstream, spreading the coolant further along the test surface.
TOPICS: Turbochargers, Film cooling, Jets, Turbulence, Coolants, Density, Pressure, Momentum, Diffusers, Boundary layers, Carbon dioxide
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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