bluuebunny/crossref_metadata_2025 huggingface dataset:
Radial Lift-Based Cascade for Bi-Directional Wave Energy Air-Turbines
Authors: Moisel Christoph , Carolus Thomas H.
Published: September 2014

Over the last decades various turbine designs have been proposed capable of handling a bi-directional flow in an oscillating water column (OWC) ocean wave power plant. The radial bi-directional lift-based turbine proposed by Kentfield in the early 1980ies is one candidate for power take-off in OWC’s. Key element of this air-turbine is a radial cascade of non-staggered blades made of symmetrical airfoils, with their mean lines geometrically mapped on the rotor’s circumference. The objective of this study is to understand and quantify the performance behavior of an isolated cascade of unstaggered radial blades under bi-directional flow by means of a numerical experiment.

3D boundary layer interaction in the corners where the blades are attached to back and front plate degrade cascade performance. But even with side wall effects neglected the through flow direction (centrifugal or centripetal) has a substantial impact. A variation of the radial cascade parameter ‘chord length/radius ratio’ revealed that the performance characteristics can be improved considerably in terms of stall margin by choosing large values of the chord length/radius ratio. This effect is more pronounced for the centrifugal but also existing for centripetal through flow direction. As one increases the chord length/radius ratio the mean flow, the cascade operates in, is modified from plane to curved.

Chordwise pressure and friction distributions on the blade surfaces did not disclose a plausible mechanism. However, one hypothesis to explain to the most favorable stall behavior of the more curved cascades is a large stable vortex detected at the blade suction that generates a new ‘blade’ contour. This allows the flow to be attached on the blade suction side at large angles of attack and hence increases lift.
Effect of Rotation on Heat Transfer and Flow Field Inside Leading Edge Cooling Passage Using Impinging Jets
Authors: Paramanandam Karthikeyan , Narayanan Sridharan , Jayamurugan Chandiran , Srinivasan Balamurugan
Published: September 2014

In this paper, the effect of rotation on impingement cooling on the internal surface (profile) of the leading edge region in a turbine rotor blade is investigated using Computational Fluid Dynamics (CFD) simulations. The flow domain is obtained by stretching the middle cross section of the blade. The simulations are performed for 3 different leading edge profiles to increase the heat transfer rate in the cooling flow passage in stationary domain. In all the profiles, the nozzle position and Mach number of cooling fluid is kept constant at E/D = 4.32 and 0.4 respectively. In the profile 1, the suction side profile is modified to facilitate vortex shift so that it may reduce the crossflow effect which will enhance the Nuavg. But Nuavg reduced by 1.9% when compared to base case. In the profile 2, the coolant flow passage is smoothened at the apex to reduce dead zones and to enhance spreading of the jet. In this case, a 3.48% increase in Nuavg is obtained. Based on the analysis of velocity contours in the profile 2, the leading edge profile is further modified (profile 3). This resulted in 5.37% increase in the Nuavg. The effect of rotation on the Nuavg for different Mach numbers (M) in the profile 3 and base case are studied. When compared to stationary domain Nuavg is reduced for base case and profile 3. In rotating domain profile 3 shows improvement for different M when compared with base case. For M = 0.2, the Nuavg is increased by 2.7%, for M = 0.4, Nuavg is increased by 3.98% and for M = 0.6, it is negligible.
Influence of Degree of Reaction on Turbine Performance for Pulsating Flow Conditions
Authors: Roclawski Harald , Gugau Marc , Langecker Florian , Böhle Martin
Published: September 2014

This paper presents a study on the influence of the degree of reaction (DoR) on turbine performance under highly pulsating inflow. A reference test turbine wheel is designed and scaled to three different wheel diameters while an identical flow capacity of all three turbines is provided by adjusting the volute size. Hence, the three turbines differ by their DoR, inertia and efficiency characteristic. The investigation is done completely numerically using highly validated models. Naturally, the pulsating flow character of a 4-cylinder gasoline engine requires unsteady CFD. In addition steady-state turbine maps were calculated beforehand as a reference base.

The results of the steady state calculation show that for the combination of the bigger turbine wheel with the smaller turbine volute the peak efficiency is smaller but is shifted towards higher pressure ratios respectively to lower blade speed ratios. This is fundamentally beneficial for turbines in automotive turbochargers for gasoline engines characterized by highly pulsating flow conditions, in particular at lower engine speeds. For the transient flow calculations with pulsating turbine inflow, the hysteresis loop and the turbine power generation was investigated. It is shown that the smallest volute compared to the biggest one causes a more contracted hysteresis loop combined with increased power output within one pulse cycle.

In order to include the influence of moment of inertia, the turbines with varying DoR but same flow capacity were analytically compared with a 1D code simulating engine load step operation. Thus, the paper shows the effect of turbine DoR on both, steady-state turbine performance under pulsating inflow and the capability for optimum engine load step operation.
Time-Linearized Forced Response Analysis of a Counter Rotating Fan: Part I — Theoretical Concept of a Fully Time-Linear Forced Response Analysis
Authors: Blocher Michael , Gómez Fernández Io Eunice
Published: September 2014

From 2010 to 2013, several institutes of the German Aerospace Center investigated a counter rotating fan arrangement called CRISP2. While having efficiency advantages over rotor-stator arrangements, counter rotating fans might be subject to increased aerodynamic excitation as well. It is well known that rotor-stator interaction (in this case rotor-rotor-interaction) can result in aerodynamic excitation of blades leading to high cycle fatigue. Therefore, forced response calculations and an estimation of the static and dynamic loads during operation are essential prior to production and rig tests.

At the Institute of Aeroelasticity, loosely coupled approaches for forced response calculations have been employed for a number of years (Vasanthakumar [1]).

Due to a limited time frame for the forced response simulation, it was decided to use this loosely coupled approach for the forced response simulations in CRISP2 as well and to reduce its time consumption even more by using a time-linear approach for the forcing calculations. The theoretical concepts and their implementation into a forced response simulation are highlighted. Special consideration is given to the linearized unsteady simulations with gust boundary conditions that have been implemented into the DLR in-house flow solver TRACE [2].

The advantages and drawbacks of a fully time-linear loosely coupled method compared to more accurate and complicated approaches are discussed.
A Numerical Study of Inlet Geometry for a Low Inertia Mixed Flow Turbocharger Turbine
Authors: Leonard Thomas M. , Spence Stephen , Early Juliana , Filsinger Dietmar
Published: September 2014

Mixed flow turbines can offer improvements over typical radial turbines used in automotive turbochargers, with regards to transient performance and low velocity ratio efficiency. Turbine rotor mass dominates the rotating inertia of the turbocharger, and any reductions of mass in the outer radii of the wheel, including the rotor back-disk, can significantly reduce this inertia and improve the acceleration of the assembly. Off-design, low velocity ratio conditions are typified by highly tangential flow at the rotor inlet and a non-zero inlet blade angle is preferred for such operating conditions. This is achievable in a Mixed Flow Turbine without increasing bending stresses within the rotor blade, which is beneficial in high speed and high inlet temperature turbine design. A range of mixed flow turbine rotors was designed with varying cone angle and inlet blade angle and each was assessed at a number of operating points. These rotors were based on an existing radial flow turbine, and both the hub and shroud contours and exducer geometry were maintained. The inertia of each rotor was also considered. The results indicated that there was a trade-off between efficiency and inertia for the rotors and certain designs may be beneficial for the transient performance of downsized, turbocharged engines.
Concurrent Blade Aerodynamic-Aeroelastic Design Optimization With Re-Scaled Response Surface
Authors: Valero O. , He L. , Li Y. S.
Published: September 2014

Given the ever increasing demands on turbomachinery performance, various advanced blade shape optimizations have been actively developed and applied in modern blading designs. Multidisciplinary and concurrent optimizations have attracted considerable attention, offering the advantage of disciplinary interactions being included more simultaneously in a design process.

This paper presents the development of a multidisciplinary optimization algorithm for the concurrent blade aerodynamic and aeromechanic shape optimization of realistic 3D turbine stages. A non-gradient algorithm is enhanced by a new re-scaled response surface (RSM) model. This meta-model is able to rescale the design space and redefine the response surface during a blade shape optimization process, leading to a much enhanced convergence compared to a standard RSM approach. The optimization algorithm is developed in conjunction with an efficient nonlinear harmonic phase solution method solving the unsteady flow equations in the frequency domain, combined with a finite element analysis (FEA) to extract the structural dynamic characteristics of the blades.

The effectiveness of the concurrent method is examined for an optimized design of a realistic LP turbine stage. The optimization goals are the maximization of the isentropic stage efficiency and aeroelastic flutter stability (aero-damping). Two sets of cases are considered. In the first set, the shaping is applied only to stator blades, while for the second set, both stator and rotor blades are shaped. The concurrent cases are compared with their single-disciplinary counterparts. For both sets of the cases, the advantages of the concurrent treatment are clearly demonstrated.
Flow Characteristics of Contra- and Co-Rotating Swirler Arrangements of an Industrial Combustor
Authors: Wu X. , Norster E. R. , Xie Gang
Published: September 2014

This paper describes the investigation of the flow characteristics of two double radial inflow swirlers configured for use in a gas turbine combustor. The only difference between the two swirlers is in the contra- and co-rotating flow of air in the inner nozzle arrangement. The isothermal vortex flow field created by the double swirlers has been examined using numerical Model 1. The model also includes a cylinder reaction zone downstream of the swirler. The comparison of flow characteristics is carried out by examination of the spatial resolution of three mean velocity components. The contra- and co-rotating configurations show some discrepancy in terms of total loss factor and mass split ratio between the two swirlers. The comparison of flow fields also indicate that there is almost no remaining swirl further downstream in the contra-rotating configuration, while a significant amount of remaining swirl exists for the co-rotating option. The development of Model 1 to include a typical dilution zone and transition duct leads to numerical Model 2, which was used to investigate the impact on downstream mixing with the dilution air and the emerging temperature distribution at the transition duct exit. Comparing the temperature field for both configurations, the dilution effectiveness increases significantly with dilution jet penetration depth and reduces with spread along the circumferential direction. These effects lead to the central hot core persisting along the transition duct to the combustor outlet for the co-rotating option due to the combination of initial cross flow and a strong swirl, resulting in a considerable difference in the predicted outlet temperature distribution factors (OTDF) of 10.8% and 17.7% for the contra- and co-rotating arrangements, respectively.
Experimental and Numerical Investigation of the Unsteady Flow Field in a Vaned Diffuser of a High-Speed Centrifugal Compressor
Authors: Vogel Klemens , Abhari Reza S. , Zemp Armin
Published: September 2014

Vaned diffusers in centrifugal compressor stages are used to achieve higher stage pressure ratios, higher stage efficiencies and more compact designs. The interaction of the stationary diffuser with the impeller can lead to resonant vibration with potentially devastating effects. This paper presents unsteady diffuser vane surface pressure measurements using in-house developed, flush mounted, fast response piezo-resistive pressure transducers. The unsteady pressures were recorded for 9 operating conditions, covering a wide range of the compressor map. Experimental work was complemented by 3D unsteady CFD simulations using ANSYS CFX V12.1 to detail the unsteady diffuser aerodynamics. Pressure fluctuations of up to 34.4% of the inlet pressure were found. High pressure variations are present all along the vane and are not restricted to the leading edge region. Frequency analysis of the measured vane surface pressures show that reduced impeller loading and the corresponding reduction of tip leakage fluid changes the characteristics of the fluctuations from a main blade count to a total blade count. The unsteady pressure fluctuations in the diffuser originate from three distinct locations. The impact of the jet wake flow leaving the impeller results in high variation close to the leading edge. It was observed that CFD results overpredicted the amplitude of the pressure fluctuation on average by 62%.
Reviewing the Implicit Mixing Plane Approach: Theoretical and Applied Cases
Authors: Hanimann Lucian , Casartelli Ernesto , Mangani Luca
Published: September 2014

The CFD assisted design of modern single- and multi-stage turbomachines is usually performed with the mixing plane approach in order to assess the components matching. While the mixing-plane state-of-the-art is based on a boundary-condition based approach, hereafter called explicit, the authors presented last year a novel, fully implicit method, which shows considerable advantages compared to the explicit one. In the present paper the quality and advantages of the novel approach compared to the state-of-the-art will be shown through a variety of detailed examples. The main issues discussed are the built-in ability to reduce incoming disturbances and to manage backflow at the interface due to the implicit formulation. With selected cases it is shown that no special care has to be taken to avoid reflections at the interface also for inviscid transonic or fully-supersonic cases. Moreover, detailed results for a high-pressure centrifugal compressor are presented, showing that the proposed approach is able to capture both, the global behavior as well as local flow-features over the complete speed-line, while the explicit approach partially fails on the same test case.
Large Eddy Simulation of a High Pressure Turbine Stage: Effects of Sub-Grid Scale Modeling and Mesh Resolution
Authors: Papadogiannis Dimitrios , Duchaine Florent , Sicot Frédéric , Gicquel Laurent , Wang Gaofeng , Moreau Stéphane
Published: September 2014

The use of Computational Fluid Dynamics (CFD) tools for integrated simulations of gas turbine components has emerged as a promising way to predict undesired component interactions thereby giving access to potentially better engine designs and higher efficiency. In this context, the ever-increasing computational power available worldwide makes it possible to envision integrated massively parallel combustion chamber-turbomachinery simulations based on Large-Eddy Simulations (LES). While LES have proven their superiority for combustor simulations, few studies have employed this approach in complete turbomachinery stages. The main reason for this is the known weaknesses of near wall flow modeling in CFD. Two approaches exist: the wall-modeled LES, where wall flow physics is modeled by a law-of-the-wall, and the wall-resolved LES where all the relevant near wall physics is to be captured by the grid leading to massive computational cost increases. This work investigates the sensitivity of wall-modeled LES of a high-pressure turbine stage. The code employed, called TurboAVBP, is an in-house LES code capable of handling turbomachinery configurations. This is possible through an LES-compatible approach with the rotor/stator interface treated based on an overset moving grids method. It is designed to avoid any interference with the numerical scheme, allow the proper representation of turbulent structures crossing it and run on massively parallel platforms. The simulations focus on the engine-representative MT1 transonic high-pressure turbine, tested by QinetiQ. To control the computational cost, the configuration employed is composed of 1 scaled stator section and 2 rotors. The main issues investigated are the effect of mesh resolution and the effect of sub-grid scale models in conjunction with wall modeling. The pressure profiles across the stator and rotor blades are in good agreement with the experimental data for all cases. Radial profiles at the rotor exit (in the near and far field) show improvement over RANS predictions. Unsteady flow features, inherently present in LES, are, however, found to be affected by the modeling parameters as evidenced by the obtained shock strengths and structures or turbulence content of the different simulations.
Life Fraction Hardening Applied to a Modified Theta Projection Creep Model for a Nickel-Based Super-Alloy
Authors: Day W. David , Gordon Ali P.
Published: September 2014

This paper presents the application of a life fraction hardening rule to the analytical calculation of creep in hot section components. Accurate prediction of creep is critical to assuring the mechanical integrity of heavy-duty, industrial gas turbine (IGT) hardware. The accuracy of such predictions depend upon both the creep models assumed and how those models are implemented in a finite element solution. A modified theta projection creep model for a nickel-based super alloy was presented in a previous paper as an accurate simulation of creep behavior [1]. Application of such a user defined creep law depends upon definition of a hardening rule in the form of either an explicit or an implicit integration scheme in order to calculate incremental strains during any time increment. Time hardening is the simplest and least computationally intensive of the two most common hardening rules, but does not correctly show the effect of changing stresses or temperatures. Strain hardening may provide the most accurate solution, but the creep models are too complex to invert, which results in highly iterative and computationally intensive solutions. A life fraction hardening rule has been presented in other works [2] as a compromise between time hardening and strain hardening. Life fraction hardening is presented here as a highly efficient and accurate means of calculating incremental creep strain when applied to a modified theta projection creep model. A user creep subroutine was defined using a state variable to represent the strain life fraction at any time. By using the time to tertiary creep as the denominator for the life fraction, no new material constants are needed to relate to creep failure. The start of tertiary creep is effectively considered to be a failure. Additional design insight can be provided through the inclusion of other state variables to calculate temperature margins at current conditions. Material testing with changing stress levels will be used to help validate the technique. A simplified example of the technique is presented in the paper. More accurate creep predictions allow our company to improve the structural integrity of its turbine blades and vanes.
CFD Evaluations of Unconventional Film Cooling Scaling Parameters on a Simulated Turbine Blade Leading Edge
Authors: Rutledge James L. , Polanka Marc D.
Published: September 2014

While it is well understood that certain nondimensional parameters, such as freestream Reynolds number and turbulence intensity, must be matched for proper design of film cooling experiments, uncertainty continues on the ideal method to scale film cooling flow rate. This debate typically surrounds the influence of the coolant to freestream density ratio and whether mass flux ratio or momentum flux ratio properly accounts for the density effects. Unfortunately, density is not the only fluid property to differ between typical wind tunnel experiments and actual turbine conditions. Temperature differences account for the majority of the property differences; however, attempts to match density ratio through the use of alternative gases can exacerbate these property differences.

A computational study was conducted to determine the influence of other fluid properties besides density — namely specific heat, thermal conductivity, and dynamic viscosity. CFD simulations were performed by altering traditional film cooling nondimensional parameters as well as others such as the Reynolds number ratio, Prandtl number ratio, and heat capacity ratio to evaluate their effects on adiabatic effectiveness and heat transfer coefficient. A cylindrical leading edge with a flat afterbody was used to simulate a turbine blade leading edge region. A single coolant hole was located 21.5° from the leading edge, angled 20° to the surface and 90° from the streamwise direction. Results indicated that thermal properties can play a significant role in understanding and matching results in cooling performance. Density effects certainly dominate; however, variations in conductivity and heat capacity can result in 10% or higher changes in the resulting heat flux to the surface when scaling ambient rig configurations to engine representative conditions.
Forced Response Prediction of an Axial Turbine Rotor With Regard to Aerodynamically Mistuned Excitation
Authors: Figaschewsky Felix , Giersch Thomas , Kühhorn Arnold
Published: September 2014

The design of both efficient and reliable turbomachinery blades demands a detailed knowledge of static and dynamic forces during operation.

This paper aims to contribute to the proper identification of dynamic excitation mechanisms acting on an axial turbine rotor, particularly with regard to deviations of the NGV’s nominal geometry due to the use of variable vanes or tolerances in manufacturing.

As variations of the NGV’s geometry disturb the perfectly periodic pattern of the downstream flow features, other spectral components than those correlated with the number of stator vanes are possible to appear. These frequency components may lead to low engine order excitation of fundamental blade modes at high engine speeds. Under these operating conditions the rotor is already highly loaded with centrifugal forces and additional dynamic excitation may cause unacceptable stresses.

Thus aerodynamic mistuning might be a limiting criterion for the design of a highly loaded turbine rotor.

Within this paper 2 dimensional CFD-models are used to investigate both, the determination of the wake of a geometric mistuned stator guide vane and the influence of the resulting excitation on the adjacent rotor stage due to aerodynamically mistuned flow. In order to generate a mistuned NGV geometry, variations of pitch and stagger angle are taken into account and a mesh morpher is used to produce computational domains of the mistuned geometry on the basis of a nominal mesh.

Additionally a simplified reconstruction process based on a set of CFD computations will be introduced, being able to reproduce the spectral components of the mistuned wake by specifying a certain geometric mistuning distribution.

The prediction of the resulting modal forces is carried out in time domain and approaches with lower fidelity are investigated with respect to their capability of reproducing the key features of an aerodynamically mistuned excitation mechanism.
Influence of a Novel 3D Leading Edge Geometry on the Aerodynamic Performance of Low Pressure Turbine Blade Cascade Vanes
Authors: Asghar A. , Allan W. D. E. , LaViolette M. , Woodason R.
Published: September 2014

This paper addresses the issue of aerodynamic performance of a novel 3D leading edge modification to a reference low pressure turbine blade. An analysis of tubercles found in nature and used in some engineering applications was employed to synthesize new leading edge geometry. A sinusoidal wave-like geometry characterized by wavelength and amplitude was used to modify the leading edge along the span of a 2D profile, rendering a 3D blade shape. The rationale behind using the sinusoidal leading edge was that they induce streamwise vortices at the leading edge which influence the separation behaviour downstream. Surface pressure and total pressure measurements were made in experiments on a cascade rig. These were complemented with computational fluid dynamics studies where flow visualization was also made from numerical results. The tests were carried out at low Reynolds number of 5.5 × 104 on a well-researched profile representative of conventional low pressure turbine profiles. The performance of the new 3D leading edge geometries was compared against the reference blade revealing a downstream shift in separated flow for the LE tubercle blades; however, total pressure loss reduction was not conclusively substantiated for the blade with leading edge tubercles when compared with the performance of the baseline blade. Factors contributing to the total pressure loss are discussed.
Reverse-Pass Cooling Systems for Improved Performance
Authors: Kirollos Benjamin , Povey Thomas
Published: September 2014

Total heat transfer between a hot and a cold stream of gas across a non-porous conductive wall is greatest when the two streams flow in opposite directions. This counter-current arrangement outperforms the co-current arrangement because the mean driving temperature difference is larger. This simple concept, whilst familiar in the heat exchanger community, has received no discussion in papers concerned with cooling of hot-section gas turbine components (e.g. turbine vanes/blades, combustor liners, afterburners). This is evidenced by the fact that there are numerous operational systems which would be significantly improved by the application of ‘reverse-pass’ cooling. That is, internal coolant flowing substantially in the opposite direction to the mainstream flow. A reverse-pass system differs from a counter-current system in that the cold fluid is also used for film cooling. Such systems can be realised when normal engine design constraints are taken into account.

In this paper, the thermal performance of reverse-pass arrangements is assessed using bespoke 2D numerical conjugate heat transfer models, and compared to baseline forward-pass and adiabatic arrangements. It is shown that for a modularised reverse-pass arrangement implemented in a flat plate, significantly less coolant is required to maintain metal temperatures below a specified limit than for the corresponding forward-pass system. The geometry is applicable to combustor liners and afterburners. Characteristically, reverse-pass systems have the benefit of reducing lateral temperature gradients in the wall.

The concept is demonstrated by modelling the pressure and suction surfaces of a typical nozzle guide vane with both internal and film cooling. For the same cooling mass flow rate, the reverse-pass system is shown to reduce the peak temperature on the suction side and reduce lateral temperature gradients on both suction and pressure sides.

The purpose of this paper is to demonstrate that by introducing concepts familiar in the heat exchanger community, engine hot-section cooling efficiency can be improved whilst respecting conventional manufacturing constraints.
Combustion System Development and Testing for MAN’s New Industrial Gas Turbines MGT 6100 and MGT 6200
Authors: Reiss Frank , Wiers Sven-Hendrik , Orth Ulrich , Aschenbruck Emil , Lauer Martin , El Masalme Jaman
Published: September 2014

This paper describes the development and test results of the low emission combustion system for the new industrial gas turbines in the 6–7 MW class from MAN Diesel & Turbo.

The design of a robust combustion system and the achievement of very low emission targets were the most important design goals of the combustor development.

During the design phase, the analysis of the combustor (i.e. burner design, air distribution, liner cooling design) was supported with different CFD tools.

This advanced Dry Low Emission can combustion system (ACC) consists of 6 cans mounted externally on the gas turbine.

The behavior and performance of a single can sector was tested over a wide load range and with different boundary conditions; first on an atmospheric test rig and later on a high pressure test rig with extensive instrumentation to ensure an efficient test campaign and accurate data.

The atmospheric tests showed a very good performance for all combustor parts and promising results. The high pressure tests demonstrated very stable behavior at all operation modes and very low emissions to satisfy stringent environmental requirements.

The whole operation concept of the combustion system was tested first on the single-can high pressure test bed and later on twin and single shaft gas turbines at MAN’s gas turbine test facility. During the engine tests, the can combustors demonstrated the expected combustion performance under real operation conditions.

All emissions and performance targets were fully achieved. On the single shaft engine, the combustors were running with single digit ppm NOx levels between 50% and 100% load.

The validation phase and further optimization of the gas turbines and the engine components are ongoing.

The highlights of the development process and results of the combustor and engine tests will be presented and discussed within this paper.
Optimization-Aided Forced Response Analysis of a Mistuned Compressor Blisk
Authors: Beirow Bernd , Kühhorn Arnold , Giersch Thomas , Nipkau Jens
Published: September 2014

The forced response of the first rotor of an E3E-type high pressure compressor blisk is analyzed with regard to varying mistuning, varying engine order excitations and the consideration of aeroelastic effects. For that purpose, SNM-based reduced order models are used in which the disk remains unchanged while the Young’s modulus of each blade is used to define experimentally adjusted as well as intentional mistuning patterns. The aerodynamic influence coefficient technique is employed to model aeroelastic interactions. Furthermore, based on optimization analyses and depending on the exciting EO and aerodynamic influences it is searched for the worst as well as the best mistuning distributions with respect to the maximum blade displacement. Genetic algorithms using blade stiffness variations as vector of design variables and the maximum blade displacement as objective function are applied. An allowed limit of the blades’ Young’s modulus standard deviation is formulated as secondary condition. In particular, the question is addressed if and how far the aeroelastic impact, mainly causing aerodynamic damping, combined with mistuning can even yield a reduction of the forced response compared to the ideally tuned blisk. It is shown that the strong dependence of the aerodynamic damping on the inter-blade phase angle is the main driver for a possible response attenuation considering the fundamental blade mode. The results of the optimization analyses are compared to the forced response due to real, experimentally determined frequency mistuning as well as intentional mistuning.
Development of a Fluidic Actuator for Adaptive Flow Control on a Thick Wind Turbine Airfoil
Authors: Niether Sebastian , Bobusch Bernhard , Marten David , Pechlivanoglou Georgios , Nayeri Christian Navid , Paschereit Christian Oliver
Published: September 2014

Wind turbines are exposed to unsteady incident flow conditions such as gusts or tower interference. These cause a change in the blades’ local angle of attack, which often leads to flow separation at the inner rotor sections [1]. Recirculation areas and dynamic stall may occur, which lead to an uneven load distribution along the blade. In this work a fluidic actuator is developed that reduces flow separation. The functional principle is adapted from a fluidic amplifier. High pressure air fed by an external supply flows into the interaction region of the actuator. Two control ports, oriented perpendicular to the inlet, allow for a steering of the actuation flow. One of the control ports is connected to the suction side, the other to the pressure side of the airfoil. Depending on the pressure difference that varies with the angle of attack, the actuation air is directed into one of four outlet channels. These guide the air to different chordwise exit locations on the airfoil’s suction side. The appropriate actuation location adjusts automatically according to the pressure difference between the control ports and therefore incidence. Suction side flow separation is delayed as the boundary layer is enriched with kinetic energy. Experiments were conducted on a DU97-W-300 airfoil [2] at Re = 2.2 · 105. Compared to the baseline, changes in lift with angle of attack were reduced by an order of magnitude. An AeroDyn simulation of a full wind turbine rotor was performed that compares the baseline to a rotor design with adaptive flow control.
Comparative Study of CFD Solver Models for Modeling of Flow Over Wind Turbine Airfoils
Authors: Kirk William T. , Capece V. R. , Pechlivanoglou G. , Nayeri C. N. , Paschereit C. O.
Published: September 2014

This paper presents an evaluation of XFOIL and a commercially available CFD solver to predict the two-dimensional lift and drag coefficients of wind turbine airfoil sections for attached and separated flows. The computational solutions are correlated with the experimental data for the DU 96-W-180 airfoil that has been generated from wind tunnel testing performed at TU Delft and TU Berlin. CFD solutions are presented for turbulent calculations using the Shear-Stress Transport (SST) [1] and Spalart-Allmaras [2] turbulence models and transition calculations using the SST γ-θ model. Transition points from the CFD simulations are compared to the results obtained using XFOIL [3].

The paper culminates in a quantitative analysis identifying the deviations of the XFOIL and CFD solutions from the experimental data. This analysis uses the experimental polars generated by TU Delft and TU Berlin as a baseline for the comparison with the end goal of determining the best computational source of design polars for use in industry.
On-Line Temperature Measurement Inside a Thermal Barrier Sensor Coating During Engine Operation
Authors: Yañez Gonzalez A. , Pilgrim C. C. , Feist J. P. , Sollazzo P. Y. , Beyrau F. , Heyes A. L.
Published: September 2014

Existing thermal barrier coatings (TBC) can be adapted enhancing their functionalities such that they not only protect critical components from hot gases, but also can sense their own material temperature or other physical properties. The self-sensing capability is introduced by embedding optically active rare earth ions into the thermal barrier ceramic. When illuminated by light the material starts to phosphoresce and the phosphorescence can provide in-situ information on temperature, phase changes, corrosion or erosion of the coating subject to the coating design.

The integration of an on-line temperature detection system enables the full potential of TBCs to be realised due to improved accuracy in temperature measurement and early warning of degradation. This in turn will increase fuel efficiency and will reduce CO2 emissions.

This paper reviews the previous implementation of such a measurement system into a Rolls-Royce jet engine using dysprosium doped yttrium-stabilised-zirconia as a single layer and a dual layer sensor coating material. The temperature measurements were carried out on cooled and uncooled components on a combustion chamber liner and on nozzle guide vanes respectively.

The paper investigates the interpretation of those results looking at coating thickness effects and temperature gradients across the TBC. For the study a specialised cyclic thermal gradient burner test rig was operated and instrumented using equivalent instrumentation to that used for the engine test. This unique rig enables the controlled heating of the coatings at different temperature regimes. A long-wavelength pyrometer was employed detecting the surface temperature of the coating in combination with the phosphorescence detector. A correction was applied to compensate for changes in emissivity using two methods. A thermocouple was used continuously measuring the substrate temperature of the sample. Typical gradients across the coating are less than 1K/μm. As the excitation laser penetrates the coating it generates phosphorescence from several locations throughout the coating and hence provides an integrated signal.

The study successfully proved that the temperature indication from the phosphorescence coating remains between the surface and substrate temperature for all operating conditions. This demonstrates the possibility to measure inside the coating closer to the bond coat. The knowledge of the bond coat temperature is relevant to the growth of the thermally grown oxide which is linked to the delamination of the coating and hence determines its life.

Further, the data is related to a one dimensional phosphorescence model determining the penetration depth of the laser and the emission.

Note: a video of the measurement system can be watched under: [http://www.youtube.com/watch?v=T6uXN1__Z7I].
Off-Line Temperature Profiling Utilising Phosphorescent Thermal History Paints and Coatings
Authors: Feist J. P. , Biswas S. Karmaker , Pilgrim C. , Sollazzo P. Y. , Berthier S.
Published: September 2014

Temperature profiling of components in gas turbines is of increasing importance as engineers drive to increase firing temperatures and optimise component’s cooling requirements in order to increase efficiency and lower CO2 emissions. However, on-line temperature measurements and, particularly, temperature profiling are difficult, sometimes impossible, to perform due to inaccessibility of the components. A desirable alternative would be to record the exposure temperature in such a way that it can be determined later, off-line.

The commercially available Thermal Paints are toxic in nature and come with a range of technical disadvantages such as subjective readout and limited durability. This paper proposes a novel alternative measurement technique which the authors call Thermal History Paints and Thermal History Coatings. These can be particularly useful in the design process, but further could provide benefits in the maintenance area where hotspots which occurred during operation can be detected during maintenance intervals when the engine is at ambient temperature.

This novel temperature profiling technique uses optical active ions in a ceramic host material. When these ions are excited by light they start to phosphoresce. The host material undergoes irreversible changes when exposed to elevated temperatures and since these changes are on the atomic level they influence the phosphorescent properties such as the life time decay of the phosphorescence. The changes in phosphorescence can be related to temperature through calibration such that in-situ analysis will return the temperature experienced by the coating. A major benefit of this technique is in the automated interpretation of the coatings. An electronic instrument is used to measure the phosphorescence signal eliminating the need for a specialist interpreter and thus increasing readout speed.

This paper reviews results from temperature measurements made with a water based paint for the temperature range 100°C to 800°C in controlled conditions. Repeatability of the tests and errors will be discussed. Further, some measurements are carried out using an electronic hand-held interrogation device which can scan a component surface and provide a spatial resolution of below 3mm. The instrument enables mobile measurements outside of laboratory conditions.

Further a robust Thermal History Coating is introduced demonstrating the capability of the coating to withstand long term exposures. The coating is based on Thermal Barrier Coating architecture with a high temperature bondcoat and deposited using an air plasma spray process to manufacture a reliable long lasting coating. Such a coating could be employed over the life of the component to provide critical temperature information at regular maintenance intervals for example indicating hot spots on engine parts.
Experimental and Numerical Investigation of the Unsteady Endwall Flow in a Highly Loaded Axial Compressor Stator
Authors: Beselt Christian , Peitsch Dieter , van Rennings Ruben , Thiele Frank , Ehrenfried Klaus
Published: September 2014

The unsteady endwall pressure distribution is investigated by means of highly resolved pressure measurements and numerical simulations in an axial compressor stator exhibiting the phenomenon of rotating instability. The endwall flow has been observed to play an important role in the inception of rotating instability. The statistical properties of the measured pressure fluctuations at the hub endwall of the axial compressor stator are compared for a series of blade loadings. The results allow conclusions about the location of regions being closely connected to the unsteady dynamics related to the rotating instability phenomenon. Furthermore the operating points with detected rotating instability are compared with unsteady numerical results.
Comparison of Laminar Flame Speeds, Extinction Stretch Rates and Vapor Pressures of Jet A-1/HRJ Biojet Fuel Blends
Authors: Munzar Jeffrey D. , Zia Ahmed , Versailles Philippe , Jiménez Rodrigo , Bergthorson Jeffrey M. , Akih-Kumgeh Benjamin
Published: September 2014

An emerging goal within the aviation industry is to replace conventional jet fuel with biologically-derived alternative fuel sources. However, the combustion properties of these potential fuels must be thoroughly characterized before they can be considered as replacements in turbomachinery applications. In this study, seven candidate alternative fuel blends, derived from two biological feedstocks and blended in different quantities with Jet A-1, are considered. For each blend, the laminar flame speed, non-premixed extinction stretch rate, and vapor pressure are experimentally determined and compared to numerical simulations and to Jet A-1 data. Hydrodynamically-stretched flame speeds are determined by applying particle image velocimetry (PIV) to an atmospheric pressure, preheated jet-wall stagnation flame, and the unstretched laminar flame speed is inferred using a direct comparison method in conjunction with a binary jet-fuel surrogate, with results spanning a wide equivalence ratio range. Extinction stretch rates were measured using particle tracking velocimetry (PTV) in a non-premixed counterflow diffusion flame, over a range of fuel mass fractions diluted in nitrogen carrier gas. Finally, the vapor pressure of the seven biojet/Jet A-1 fuel blends was measured using an isoteniscope over a wide temperature range. The results of this study indicate that moderate blends of hydrotreated renewable jet (HRJ) fuel with Jet A-1 have similar combustion properties to conventional jet fuel, highlighting their suitability as drop-in replacements, while higher blend levels of HRJ fuel, regardless of the crop source, lead to definitive changes in the combustion parameters investigated here.
Stability Study of a Bladed Disk in Interaction With a Casing via a Labyrinth Seal
Authors: Andreev F. B. , Thouverez F. , Gouskov A. M. , Blanc L.
Published: September 2014

In this paper a simplified coupled model of a bladed disk and flexible stator developed in a rotating frame is introduced. The interaction between the rotating and stationary parts is adjusted using the gas flow in the straight-through labyrinth seal located on the tip of the blades. This is therefore an “aero elastic contact” problem. An energy-based method is used to express the equations of motion. The computation of eigenvalues provides the Campbell diagram and information on stability. A single-control-volume analysis is performed to predict the destabilizing force exerted by the labyrinth seal.

In terms of strategy a simpler and efficient model with a rigid rotating shaft is compared with models found in the literature before the complete rotor-stator aero elastic contact study is presented. This is followed by a study of the interaction between the bladed disk with a labyrinth seal and the flexible stator.

The model proposed provides a new approach to modelling rotor systems with labyrinth seals. The stator is then modelled as a flexible body and the rotor as a bladed disk. It is shown in the example described that the stability of the system depends on the rigid body movement of the stator and is independent of its nodal diameter modes.
3D Unsteady Multi-Stage CFD Analysis of Combustor-Turbine Hot Streak Migration
Authors: Mollahosseini Khosro , Borns Fred G. , Couey Paul T. , Bonaccorsi Jean-Charles , Demeulenaere Alain
Published: September 2014

With gas temperatures far exceeding the melting point of nickel-base alloys, advanced cooling schemes are essential to meet the desired mission life of turbine airfoils. Naturally, combustion systems produce gas-temperature non-uniformity in the exiting flowfield. Downstream turbine components must be tolerant to the maximum anticipated gas temperatures. On the other hand, excessive use of cooling air reduces engine efficiency and compromises combustor durability. Throughout gas turbine design history it has been the desire of Turbine Aerodynamicists to be able to compute combustor hot streak migration and mixing through multiple turbine airfoil stages. Typically, hot streak migration studies have been performed using (a) mixing-plane models between rotating and stationery domains or (b) unsteady simulations in which the flowpath annulus is represented by a segment containing airfoil counts that are integer multiples in each blade row or (c) Non-Linear Harmonic methods. With the development of highly-parallelized Computational Fluid Dynamic (CFD) codes driving high performance computer clusters simulation of combustor hot streak migration through multiple High Pressure (HP) turbine stages using an unsteady, 360° (full-annulus) model can be achieved. To this end, Honeywell, in collaboration with Numeca Corporation, has performed a study to evaluate the state-of the art for computation of the effect on second-stage HP turbine nozzle metal temperatures of combustor hot streaks migrated through the first-stage of a two-stage HP turbine.