Accepted Manuscripts

Roland Brachmanski and Reinhard Niehuis
J. Turbomach   doi: 10.1115/1.4036436
The present measurements for each low pressure turbine profile were conducted at midspan under a range of Reynolds- and exit Mach numbers. The exit Mach number was varied in a range covering low subsonic up to values where a transonic flow regime on the suction side of the blade could be expected. The variation of the exit Mach number was also used to create different locations of the maximum Mach number and to evaluate the resulting total pressure losses. This work focuses on two profiles with a diffusion factor between 0.18 ≤ DF ≤ 0.22. The integral total pressure losses were evaluated by wake traverses downstream of the profile. Numerical studies were also conducted to investigate further the influence of a reduced turbulence intensity on the boundary layer of the suction side of design B. The results show that the optimum of the integral total pressure losses are significantly dependent on the Reynolds number. Therefore a correlation between the maximum Mach number on the suction side and the integral total pressure losses has been successfully established. It also results in an equivalent change of the total pressure losses, which has been predicted by the trend line. However, the trend lines, which are based on the data of the integral total pressure losses of an attached boundary layer, are not able to predict the integral total pressure loss or the location of the maximum Mach number on the suction side of the blade since an open separation bubble occurs.
TOPICS: Pressure, Mach number, Diffusion (Physics), Turbochargers, Turbines, Suction, Blades, Boundary layers, Design, Transonic flow, Reynolds number, Wakes, Bubbles, Separation (Technology), Turbulence
G. Barigozzi, H. Abdeh, A. Perdichizzi, M. Henze and J. Krueckels
J. Turbomach   doi: 10.1115/1.4036437
In the present paper, the influence of the presence of an inlet flow non uniformity on the aerodynamic and thermal performance of a film cooled linear nozzle vane cascade is fully assessed. Tests have been carried out with platform cooling, with coolant ejected through a slot located upstream of the leading edge. Cooling air is also ejected through a row of cylindrical holes located upstream of the slot, simulating a combustor cooling system. An obstruction was installed upstream of the cascade at variable tangential and axial position to generate a flow non uniformity. The cascade was tested at a high inlet turbulence intensity level (Tu1 = 9%) and at a constant inlet Mach number of 0.12 and nominal cooling condition. Aero-thermal characterization of vane platform was obtained through 5-hole probe and end wall adiabatic film cooling effectiveness measurements. Results show a relevant negative impact of inlet flow non uniformity on the cooled cascade aerodynamic and thermal performance. Higher film cooling effectiveness and lower aerodynamic losses are obtained when the inlet flow non uniformity is located at mid pitch, while lower effectiveness and higher losses are obtained when it is aligned to the vane leading edge. Moving the non uniformity axially or changing its blockage only marginally influences the platform thermal protection.
TOPICS: Flow (Dynamics), Turbochargers, Cascades (Fluid dynamics), Nozzles, Film cooling, Cooling, Cooling systems, Turbulence, Coolants, Combustion chambers, Mach number, Probes
Martin Elfert, Anton Weber, David Wittrock, Andreas Peters, Christian Voss and Eberhard Nicke
J. Turbomach   doi: 10.1115/1.4036357
Outgoing from a well-proven radial compressor design which has been extensively being tested in the past known as SRV4 impeller (Krain impeller), an optimization has been performed using the AutoOpti tool developed at DLR. The optimization package AutoOpti was applied to the redesign and optimization of a radial compressor stage with a vaneless diffusor. The optimization was performed for the SRV4 compressor geometry without fillets using a coarse structured mesh in combination with wall functions. In order to filter out the improvements of the new SRV5 radial compressor design, two work packages were conducted: The first one was the rig test to evaluate the classical performance map and the efficiency chart achieved by the new compressor design. The efficiencies realized in the performance chart were enhanced by nearly 1.5 %. A 5 % higher maximum mass flow rate was measured in agreement with the RANS simulations during the design process. The second work package comprises the CFD analysis. The numerical investigations were conducted with the exact geometries of both, the baseline SRV4 as well as the optimized SRV5 impeller including the exact fillet geometries. To enhance the prediction accuracy of pressure ratio and impeller efficiency the geometries were discretized by high resolution meshes of approximately 5 million cells allowing a low-Reynolds approach in order to get high quality results. The comparison of the numerical predictions and the experimental results shows a very good agreement and confirms the improvement of the compressor performance using the optimization tool AutoOpti.
TOPICS: Turbochargers, Optimization, Compressor impellers, Compressors, Design, Impellers, Filters, Geometry, Reynolds-averaged Navier–Stokes equations, Vaneless diffusers, Engineering simulation, Resolution (Optics), Computational fluid dynamics, Simulation, Pressure, Flow (Dynamics)
Kathryn Evans and John P. Longley
J. Turbomach   doi: 10.1115/1.4036341
The effect of stator clocking has been experimentally and computationally investigated using a low-speed, two-stage, Low-Pressure Turbine which was specifically designed to maximise the clocking potential by aligning the Stator 1 wake segments with the Stator 2 leading edge along the span. It was verified that the wake segments are aligned to within 10% of stator pitch across the span. The measured clocking effect on the work extraction is 0.12% and on efficiency is 0.08%. Although the effect of clocking is small, it is repeatable, periodic across four stator pitches and consistent between independent measurements. Furthermore, factors to consider for a reliable clocking investigation are discussed. The measurements revealed that the majority of the clocking effect on the work extraction occurs in Stage 2 and it originates at Stator 2 exit. This indicates that the flow is being processed differently within Stator 2. There is also an effect on the Stage 1 work. In each blade row the measured clocking effect on the lost work is similar across the span. The computations with meshed cavities do not capture any clocking effects in Stage 1. This indicates that an unsteady viscid phenomenon within Rotor 1 is not captured by the fully turbulent calculation e.g. unsteady transition. However, the computations do capture the measured clocking effect on the Stage 2 work extraction. It is hypothesised that the clocking effect on Stator 2 flow turning is dominated by a steady, inviscid process.
TOPICS: Pressure, Turbochargers, Turbines, Stators, Wakes, Computation, Rotors, Flow (Dynamics), Turbulence, Flow turning, Blades, Cavities
Kathryn Evans and John P. Longley
J. Turbomach   doi: 10.1115/1.4036342
It is common to assume that the performance of low-speed turbines depend only on the flow coefficient and Reynolds number. However, when the turbine has an atmospheric inlet and uses unconditioned air, variations in ambient pressure, temperature and humidity are introduced. Whilst it is still possible to maintain the required values for the flow coefficient and Reynolds number, the ambient variations affect additional non-dimensional quantities which are related to the blade speed and gas properties. Generally, these additional non-dimensional quantities are uncontrolled and affect the turbine performance. In addition, thermal effects, which are exacerbated by the use of plastic blades, can cause changes in the blade row seal clearance and these also affect the performance. Therefore to obtain measurements with greater accuracy and repeatability, the changes in the uncontrolled non-dimensional quantities must be accounted. This paper contains four parts. Firstly, it is described how suitable data acquisition parameters can be determined to eliminate short time scale facility unsteadiness within the measurements. Secondly, by the analysis of models, the most appropriate forms for the additional non-dimensional quantities are obtained. Since the variations in the uncontrolled non-dimensional quantities affect repeatability the size of the effect on the turbine performance is quantified. Thirdly, a best-fit accounting methodology is described which reduces the effects of the uncontrolled non-dimensional quantities on turbine performance provided sufficient directly related measurements are available. Finally, the observations are generalised to high-speed turbomachines.
TOPICS: Turbines, Turbochargers, Accounting, Blades, Flow (Dynamics), Reynolds number, Pressure, Temperature, Turbomachinery, Data acquisition, Temperature effects, Clearances (Engineering)
Christian Peeren and Konrad Vogeler
J. Turbomach   doi: 10.1115/1.4036343
This paper focus on the geometrical modification of the unsteady pressure to reduce aerodynamic flutter excitation. Flutter occurs when the blade structure is absorbing energy from the surrounding fluid, which leads to hazardous amplification of vibrations. The unsteady pressure is caused by the motion of the vibrating blades and is responsible for local stability. Especially for free-standing blades, where most exciting aerodynamic work transfer is found at the upper tip sections, a geometrical redesign is expected to beneficially influence stability. Two approaches are pursued in this work. This first approach is based on flow physics considerations and analytical models. The unsteady pressure field is decomposed into three physical mechanisms or effects and each effect investigated. One of these effects is the the contract-and expansion of the channelled regions of the blades for which an analytical model is derived for the limiting case of a reduced frequency of zero. Another effect discussed are the excitation by shocks and gasdynamic effects in transonic flows. The second approach is used to validate the conclusions made in the theoretical part by numerical optimizing the geometry of a representative turbine blade. As optimization targets the aerodynamic damping and loss behaviour are used. The steady blade loading is fixed by constraining the outlet flow angle and velocity. Selected optimized designs are picked and compared with each other in terms of local excitation, aerodynamics and robustness with respect to the boundary conditions. Based on these observations some recommendations are made for an improved turbine design.
TOPICS: Pressure, Flutter (Aerodynamics), Turbochargers, Blades, Excitation, Stability, Flow (Dynamics), Aerodynamics, Fluids, Damping, Design, Optimization, Turbines, Vibration, Physics, Turbine blades, Shock (Mechanics), Boundary-value problems, Geometry, Robustness, Transonic flow
Karsten Knobloch, Lars Neuhaus, Friedrich Bake, Paolo Gaetani and Giacomo Persico
J. Turbomach   doi: 10.1115/1.4036344
The noise originating from the core of an aero-engine is usually difficult to quantify and the knowledge about its generation and propagation is less advanced than that for other engine components. In order to overcome the difficulties associated with dynamic measurements in the crowded core region, dedicated experiments have been set up in order to investigate the processes associated with the generation of noise in the combustor, its propagation through the turbine and the interaction of these two components, which may produce additional - so-called indirect combustion - noise. In the current work, a transonic turbine stage installed at the Laboratorio di Fluidodinamica delle Macchine of the Politechnico di Milano was exposed to acoustic, entropic, and vortical disturbances. The incoming and outgoing sound fields were analyzed in detail by two large arrays of microphones. The mean flow field and the disturbances were carefully mapped by several aerodynamic and thermal probes. The results include transmission and reflection characteristics of the turbine stage, latter one was found to be much lower than usually assumed. The modal decomposition of the acoustic field in the upstream and downstream section show beside the expected rotor-stator interaction modes additional modes. At the frequency of entropy or respectively vorticity excitation, a significant increase of the overall sound power level was observed.
TOPICS: Turbochargers, High pressure (Physics), Turbines, Noise (Sound), Acoustics, Engines, Reflection, Entropy, Combustion chambers, Flow (Dynamics), Combustion, Vorticity, Rotors, Microphones, Probes, Stators, Aircraft engines, Excitation
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)
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
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
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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