bluuebunny/crossref_metadata_2025 huggingface dataset:
Turbine Blade Entropy Generation Rate: Part I — The Boundary Layer Defined
Authors: Davies M. R. D. , O’Donnell F. K. , Niven A. J.
Published: August 2014

The profile loss of a gas turbine blade is normally associated with the entropy increase due to the boundary layer phenomena of viscous shear, Reynolds stress generation and heat transfer. To establish the relative contributions of laminar, transitional and turbulent adiabatic boundary layer flow, to the overall entropy generation (as described in part two of this paper), detailed hot film and hot wire measurements have been made over the suction surface of a turbine blade mounted within a subsonic linear cascade. At a Reynolds number of 185 × 103, a natural transition region was found between 53 and 70% suction surface length, followed by a slowly relaxing turbulent boundary layer. The wall shear stress distribution indicated a peak in the leading edge region, dropping to a minimum value prior to transition, with a sharp rise over the transitional length before decreasing within the turbulent portion. The measurements were compared with predictions obtained from a commercial computational fluid dynamics code, which utilised the renormalisation group theory (RNG) turbulence model.
Turbine Blade Entropy Generation Rate: Part II — The Measured Loss
Authors: O’Donnell F. K. , Davies M. R. D.
Published: August 2014

Using detailed boundary layer velocity measurements the profile loss, expressed in terms of local entropy generation rate, is evaluated at discrete locations along the suction surface of a turbine blade in a subsonic linear cascade at a chord Reynolds number of 1.8 × 103 under adiabatic test conditions. The distribution of loss through the entire boundary layer is thus established with particular attention given to the comparison of the relative contributions from the laminar, transitional and turbulent regions. It is found that 75% of the entropy generation occurs in the laminar region of the blade, with transition being one of the key features of the overall loss distribution. Traditional correlation methods are considered and shown to give accurate results when compared to the experimental measurements within both the laminar and turbulent regions, the application of such correlations is however dependant upon knowledge of the onset and extent of transition. Finally it is demonstrated that an existing method for the evaluation of local entropy generation rate from measurements of wall shear stress in laminar flow, may be adapted for use in turbulent flow and hence the possibility is presented for the measurement of loss from surface mounted sensors.
Bypass Transition Modelling: A New Method Which Accounts for Free-Stream Turbulence Intensity and Length Scale
Authors: Roach Paul E. , Brierley David H.
Published: August 2014

The publication of the present authors’ boundary layer transition data in 1992 (now widely known as the ERCOFTAC test case T3) has led to a spate of new experimental and modelling efforts aimed at improving our understanding of this problem. This paper describes a new method of determining boundary layer transition with zero mean pressure gradient. The approach examines the development of a laminar boundary layer to the start of transition, accounting for the influences of free-stream turbulence and test surface geometry. It is presented as a “proof of concept”, requiring a significant amount of work before it can be considered as a practically applicable model for transition prediction. The method is based upon one first put forward by G.I. Taylor in the 1930’s, and accounts for the action of local, instantaneous pressure gradients on the developing laminar boundary layer. These pressure gradients are related to the intensity and length scale of turbulence in the free-stream using Taylor’s simple isotropic model.

The findings demonstrate the need to account for the separate influences of free-stream turbulence intensity and length scale when considering the transition process. Although the length scale has less of an effect than the intensity, its influence is, nevertheless, significant and must not be overlooked. This fact goes a long way towards explaining the large scatter to be found in simple correlations which involve only the turbulence intensity. Intriguingly, it is demonstrated that it is the free-stream turbulence at the leading edge of the test surface which is important, not that found locally outside the boundary layer.

The additional influence of leading edge geometry is also shown to play a major role in fixing the point at which transition begins. It is suggested that the leading edge geometry will distort the incident turbulent eddies, modifying the effective “free-stream” turbulence properties. Consequently, it is shown that the scale of the eddies relative to the leading edge thickness is a further important parameter, and helps bring together a large number of test cases.
Total Pressure Losses in Rotor Systems With Radial Inflow
Authors: Brillert D. , Lieser D. , Reichert A. W. , Simon H.
Published: August 2014

Gas turbines with a splined-disc rotor design allow the compressor bleed pressure to be adapted precisely to the requirements of rotor cooling air systems in which the cooling air is routed through the spaced between the rotating discs.

Calculation of such flows is extremely difficult; particularly so if the flow is directed radially inward. In such cases the circumferential component of the absolute velocity can be very high and can thus lead to pronounced total pressure losses.

The paper gives a brief description of the flow phenomena, and details in the calculation methods cited in the literature.

Navier–Stokes calculations were carried out for the flow through a model test bed engine. The results are compared with experimental data.

A simple calculation model is discussed and its result compared with test data.

The model predicts the flow pattern more accurately than the Navier-Stokes calculations, and this paper shows that the simple model can be improved further.
Flow and Heat Transfer in an Industrial Rotor-Stator Rim Sealing Cavity
Authors: Mirzamoghadam Alexander V. , Xiao Zhenhua
Published: August 2014

Flow and heat transfer in the row-1 upstream rotor-stator disc cavity of a large 3600-rpm industrial gas turbine was investigated using an integrated approach. A 2D axisymmetric transient thermal analysis using aero engine-based correlations was performed to predict the steady state metal temperatures and hot running seal clearances at ISO rated power condition. The cooling mass flow and the flow pattern assumption for the thermal model were obtained from the steady state 2D axisymmetric CFD study. The CFD model with wall heat transfer was validated using cavity steady state air temperatures and static pressures measured at inlet to the labyrinth seal and four cavity radial positions in an engine test which included the mean annulus static pressure at hub radius. The predicted wall temperature distribution from the matched thermal model was used in the CFD model by incorporating wall temperature curve-fit polynomial functions. Results indicate that although the high rim seal effectiveness prevents ingestion from entering the cavity, the disc pumping flow draws air from within the cavity to satisfy entrainment leading to an inflow along the stator. The supplied cooling flow exceeds the minimum sealing flow predicted from both the rotational Reynolds number-based correlation and the annulus Reynolds number correlation. However, the minimum disc pumping flow was found to be based on a modified entrainment expression with a turbulent flow parameter of 0.08. The predicted coefficient of discharge (Cd) of the industrial labyrinth seal from CFD was confirmed by modifying the carry-over effect of a correlation reported recently in the literature. Moreover, the relative effects of seal windage and heat transfer were obtained and it was found that contrary to what was expected, the universal windage correlation was more applicable than the aero engine-based labyrinth seal windage correlation. The CFD predicted disc heat flux profile showed reasonably good agreement with the free disc calculated heat flux. The irregular cavity shape and high rotational Reynolds number (in the order of 7×107) leads to entrance effects that produce a thicker turbulent boundary layer profile compared to that predicted by the 1/7 power velocity profile assumption.
Influence of a Honeycomb Facing on the Heat Transfer in a Stepped Labyrinth Seal
Authors: Willenborg K. , Schramm V. , Kim S. , Wittig S.
Published: August 2014

The influence of a honeycomb facing on the heat transfer of a stepped labyrinth seal with geometry typical for modern jet engines was investigated. Heat transfer measurements were obtained for both a smooth stator and a stator lined with a honeycomb structure. In addition, an LDV system was used with the scaled up geometry to obtain a high local resolution of the velocity distribution in the seal. The experiments covered a wide range of pressure ratios and gap widths, typical for engine operating conditions. Local heat transfer coefficients were calculated from the measured wall and gas temperatures using a finite element code. By averaging the local values, mean heat transfer coefficients were determined and correlations for the global Nusselt numbers were derived for the stator and the rotor. The LDV results showed strong geometrical effects of the honeycomb structure on the development of the flow fields for the honeycomb seal. The distribution of the local heat transfer coefficients are compatible with to the flow features identified by the LDV results and reveal a significantly reduced heat transfer with the honeycomb facing compared to the smooth facing.
Large-Scale Testing to Validate the Influence of External Crossflow on the Discharge Coefficients of Film Cooling Holes
Authors: Rowbury D. A. , Oldfield M. L. G. , Lock G. D.
Published: August 2014

This paper discusses large-scale, low-speed experiments that explain unexpected flow-interaction phenomena witnessed during annular cascade studies into the influence of external crossflow on film cooling hole discharge coefficients. More specifically, the experiments throw light on the crossover phenomenon, where the presence of the external crossflow can, under certain circumstances, increase the discharge coefficient. This is contrary to most situations, where the external flow results in a decrease in discharge coefficient. The large-scale testing reported helps to explain this phenomena through an increased understanding of the interaction between the emerging coolant jet and the freestream flow.

The crossover phenomenon came to light during an investigation into the influence of external crossflow on the discharge coefficients of nozzle guide vane film cooling holes. These experiments were performed in the Cold Heat Transfer Tunnel (CHTT), an annular blowdown cascade of film cooled vanes which models the three-dimensional external flow patterns found in modern aero-engines. (Rowbury et al., 1997 and 1998).

The variation in static pressure around the exit of film cooling holes under different flow conditions was investigated in the large-scale tests. The study centred on three holes whose geometries were based on those found in the leading edge region of the CHTT vanes, as the crossover phenomenon was witnessed for these rows during the initial testing. The experiments were carried out in a low speed wind tunnel, with the tunnel freestream flow velocity set to match the freestream Reynolds number (as based on the local radius of curvature) and the ‘coolant’ flow velocity set to replicate the engine coolant-to-freestream momentum flux ratio.

It was found that the apparent enhancement of film cooling hole discharge coefficients with external crossflow was caused by a reduction in the static pressure around the hole exit, associated with the local acceleration of the freestream around the emerging coolant jet. When these measured static pressures (rather than the freestream static pressure) were used to calculate the discharge coefficient, the crossover effect was absent. The improved understanding of the crossover phenomenon and coolant-to-freestream interactions that has been gained will be valuable in aiding the formulation of predictive discharge coefficient schemes.
Influence of Disturbances on Transients of a Gas Turbine Ship Propulsion System — Preliminary Investigations
Authors: Dzida Marek , Domachowski Zygfryd
Published: August 2014

A gas turbine ship propulsion control system transients have been investigated. On the basis of a mathematical model composed of blocks modelling a two-shaft gas turbine, a gear (mechanical or electric), and a coupling shaft, some preliminary simulations have been carried out. Ship propeller shaft angular velocity, fuel flow rate, and gas turbine combustion chamber outlet temperature response to the ship propeller shaft angular velocity set point, and fuel flow rate, changes have been analyzed. Influences of limiters in the controller action on analyzed transients have been compared.
Cyclic Oxidation of a NiCoCrAlY Coating Applied by HVOF and LPPS
Authors: Bernstein Henry L. , Greaves Wayne L. , van Esch Hans , Talley Peggy L.
Published: August 2014

A comparison of a NiCoCrAlY coating applied by the Diamond Jet Hybrid 2600 system to the same coating applied by LPPS are presented for cyclic oxidation. The results show that the HVOF version, HICoat H11, has similar cyclic oxidation resistance to the LPPS coating at 980°C (1800°F) and 1095°C (2000°F). Microstructurally, the HVOF and LPPS coatings were similar at 980°C (1800°F), but the HVOF coating was superior at 1095°C (2000°F).
Effect of Crystal Orientation on Fatigue Failure of Single Crystal Nickel Base Turbine Blade Superalloys
Authors: Arakere Nagaraj K. , Swanson Gregory
Published: August 2014

High Cycle Fatigue (HCF) induced failures in aircraft gas turbine and rocket engine turbopump blades is a pervasive problem. Single crystal nickel turbine blades are being utilized in rocket engine turbopumps and jet engines throughout industry because of their superior creep, stress rupture, melt resistance and thermomechanical fatigue capabilities over polycrystalline alloys. Currently the most widely used single crystal turbine blade superalloys are PWA 1480/1493, PWA 1484, RENE’ N-5 and CMSX-4. These alloys play an important role in commercial, military and space propulsion systems. Single crystal materials have highly orthotropic properties making the position of the crystal lattice relative to the part geometry a significant factor in the overall analysis. The failure modes of single crystal turbine blades are complicated to predict due to the material orthotropy and variations in crystal orientations. Fatigue life estimation of single crystal turbine blades represents an important aspect of durability assessment. It is therefore of practical interest to develop effective fatigue failure criteria for single crystal nickel alloys and to investigate the effects of variation of primary and secondary crystal orientation on fatigue life.

A fatigue failure criterion based on the maximum shear stress amplitude [Δτmax] on the 24 octahedral and 6 cube slip systems, is presented for single crystal nickel superalloys (FCC crystal). This criterion reduces the scatter in uniaxial LCF test data considerably for PWA 1493 at 1200F in air. Additionally, single crystal turbine blades used in the alternate advanced high-pressure fuel turbopump (AHPFTP/AT) are modeled using a large-scale 3D finite element (FE) model. This FE model is capable of accounting for material orthotrophy and variation in primary and secondary crystal orientation. Effects of variation in crystal orientation on blade stress response are studied based on 297 FE model runs. Fatigue lives at critical points in the blade are computed using FE stress results and the failure criterion developed. Stress analysis results in the blade attachment region are also presented. Results presented demonstrates that control of secondary and primary crystallographic orientation has the potential to significantly increase a component’s resistance to fatigue crack growth without adding additional weight or cost.
Adaptive Strategies for Manufacturing and Repair of Blades and Blisks
Authors: Bremer Claus
Published: August 2014

This paper presents Reverse Engineering based methods for the adaptive manufacturing and repair of blades and blisks. This technology compensates for part-to-part variation as well as inaccurate clamping position and maintains manufacturing tolerances for the actual parts within a minimal value. The systems are integrated in standard machine tools as well as in the existing CMM, CAD, CAM and CAQ environment. All solutions utilize robust and precise 3D scanning systems and application specific data processing software. The presented solutions fit into the sphere of workshop. The main premises are the quality of the machining results and the simplicity of operation. The flexibility of the adaptive approach is demonstrated by different existing and future applications: reprofiling of straight blade tips, reprofiling of blades with specially shaped final domains, automatic blisk repair and adaptive belt grinding.
An Innovative Device for Passive Control of Surge in Industrial Compression Systems
Authors: Arnulfi Gianmario L. , Giannattasio Pietro , Micheli Diego , Pinamonti Piero
Published: August 2014

The present paper reports a numerical-experimental study on the dynamic behaviour of a compression system based on a multi-stage centrifugal blower and fitted with an innovative device for the dynamic suppression of surge instability. The control device is of passive type and is based on the aeroelastic coupling of the basic compression system with a hydraulic oscillator. The controlled system is modelled at first by using a non-linear lumped parameter approach. The simulated system dynamics within a wide range of operating conditions allows a parametric analysis to be performed and the optimal values of the control parameters to be singled out. Such optimal values are then used to design the hydraulic oscillator, which results in a technically feasible and very simple configuration. Finally, experimental tests are carried out on the compression plant with and without the passive control device, which demonstrate the effectiveness of the proposed control system in suppressing surge instabilities, at least within the limits predicted by the numerical simulation.
A Novel Technique for the Measurement of Blade Damping Using Piezoelectric Actuators
Authors: Jeffers T. R. , Kielb J. J. , Abhari R. S.
Published: August 2014

Reducing damage to turbomachinery blades due to high cycle resonant excitation is an important engineering challenge facing the engine design community. To minimize peak resonant stresses, designers attempt to reduce the forcing function and increase structural and aerodynamic damping. For higher frequency modes in particular, understanding of blade damping has been a challenge, due in part to a lack of detailed experimental measurements. The focus of this paper is on the description and validation of a novel experimental technique for the measurement of damping in turbomachinery blades.

The technique utilizes a continuously powered piezoelectric actuator embedded within the structure of the blade to excite a range of structural modes, while monitoring the structural response by the use of strain gage transducers. A novel data reduction technique is developed in order to 1) minimize the contamination of the strain gage output from the piezoelectric high voltage electrical field and to 2) efficiently obtain an accurate measure of the critical damping ratio and resonant frequency. The range of applicability and practical limitations of the technique is explored through a detailed uncertainty analysis.

Three different case studies were investigated to show the application of the technique including: a cantilever plate, broach block mounted turbine blades, and rotating turbine blades mounted in a rotor disk. Two different blade geometries were analyzed using full-scale engine hardware in a high speed rotating blade-disk assembly. Future papers will present more detailed results obtained from numerous spin tests that have been completed to date using the technique presented here.
A Comparison of Two Finite Element Reduction Techniques for Mistuned Bladed-Disks
Authors: Moyroud François , Fransson Torsten , Jacquet-Richardet Georges
Published: August 2014

The high performance bladed-disks used in today’s turbomachines must meet strict standards in terms of aeroelastic stability and resonant response level. One structural characteristic that can significantly impact on both these area is that of bladed-disk mistuning. To predict the effects of mistuning, computationally efficient methods are necessary to make it feasible, especially in an industrial environment, to perform free vibration and forced response analyses of full assembly finite element models. Due to the size of typical finite element models of industrial bladed-disks, efficient reduction techniques must be used to systematically produce reduced order models. The objective of this paper is to compare two prevalent reduction methods on representative test rotors, including a modern design industrial shrouded bladed-disk, in terms of accuracy (for frequencies and mode shapes), reduction order, computational efficiency, sensitivity to inter-sector elastic coupling, and ability to capture the phenomenon of mode localization. The first reduction technique employs a modal reduction approach with a modal basis consisting of mode shapes of the tuned bladed-disk which can be obtained from a classical cyclic symmetric modal analysis. The second reduction technique is based on a Craig and Bampton substructuring and reduction approach. The results show a perfect agreement between the two reduced order models and the non-reduced finite element model. It is found that the phenomena of mode localization is equally well predicted by the two reduction models. In terms of computational cost, reductions from 1 to 2 orders of magnitude are obtained for the industrial bladed-disk, with the modal reduction method being the most computationally efficient approach.
Limit Cycle due to Interlocked Shroud Friction: Instrumental to the High-Bypass Turbofan
Authors: Klompas Nicholas
Published: August 2014

The limit cycle induced by friction at interlocked shrouds of a bladed disk flexing in the mode of one nodal diameter is defined and validated by results of manufacturer’s tests designed to diagnose breakage of a fan blade outer panel. Despite indications in reported rotor failures, this mode, first demonstrated by Wilfred Campbell in the twenties, is not yet recognized in research or design. The bases of possible excessive vibratory stresses and limitations of the unproven but widespread assumption of beneficial damping due to slippage at interlocked shrouds are advanced for this mode. The action of friction in limiting vibratory amplitudes is shown as instrumental to development of the high-bypass turbofan. Questions remaining after official investigations are exposed to show the need to redirect research into analyses of potential destabilizing mechanisms and their combined actions in the whole engine.
A Resonance Tracking Algorithm for the Prediction of Turbine Forced Response With Friction Dampers
Authors: Bréard C. , Green J. S. , Vahdati M. , Imregun M.
Published: August 2014

This paper presents an iterative method for determining the resonant speed shift when non-linear friction dampers are included in turbine blade roots. Such a need arises when conducting response calculations for turbine blades where the unsteady aerodynamic excitation must be computed at the exact resonant speed of interest. The inclusion of friction dampers is known to raise the resonant frequencies by up to 20% from the standard assembly frequencies. The iterative procedure uses a viscous, time-accurate flow representation for determining the aerodynamic forcing, a look-up table for evaluating the aerodynamic boundary conditions at any speed, and a time-domain friction damping module for resonance tracking. The methodology was applied to an HP turbine rotor test case where the resonances of interest were due to the 1T and 2F blade modes under 40 engine-order excitation. The forced response computations were conducted using a multi-stage approach in order to avoid errors associated with “linking” single stage computations since the spacing between the two bladerows was relatively small. Three friction damper elements were used for each rotor blade. To improve the computational efficiency, the number of rotor blades was decreased by 2 to 90 in order to obtain a stator/rotor blade ratio of 4/9. However, the blade geometry was skewed in order to match the capacity (mass flow rate) of the components and the condition being analysed. Frequency shifts of 3.2% and 20.0% were predicted for the 1T/40EO and 2F/40EO resonances in about 3 iterations. The predicted frequency shifts and the dynamic behaviour of the friction dampers were found to be within the expected range. Furthermore, the measured and predicted blade vibration amplitudes showed a good agreement, indicating that the methodology can be applied to industrial problems.
An Integrated Time-Domain Aeroelasticity Model for the Prediction of Fan Forced Response Due to Inlet Distortion
Authors: Bréard C. , Vahdati M. , Sayma A. I. , Imregun M.
Published: August 2014

The forced response of a low aspect-ratio transonic fan due to different inlet distortions was predicted using an integrated time-domain aeroelasticity model. A time-accurate, non-linear viscous, unsteady flow representation was coupled to a linear modal model obtained from a standard finite element formulation. The predictions were checked against the results obtained from a previous experimental programme known as “Augmented Damping of Low-aspect-ratio Fans” (ADLARF). Unsteady blade surface pressures, due to inlet distortions created by screens mounted in the intake inlet duct, were measured along a streamline at 85% blade span. Three resonant conditions, namely 1F/3EO, 1T&2F /8EO and 2S/8EO, were considered. Both the amplitude and the phase of the unsteady pressure fluctuations were predicted with and without the blade flexibility. The actual blade displacements and the amount of aerodynamic damping were also computed for the former case. A whole-assembly mesh with about 2,000,000 points was used in some of the computations. Although there were some uncertainties about the aerodynamic boundary conditions, the overall agreement between the experimental and predicted results was found to be reasonably good. The inclusion of the blade motion was shown to have an effect on the unsteady pressure distribution, especially for the 2F/1T case. It was concluded that a full representation of the blade forced response phenomenon should include this feature.
Interaction Effects in a Transonic Turbine Stage
Authors: Barter J. W. , Vitt P. H. , Chen J. P.
Published: August 2014

A 3D, viscous, time-accurate code has been used to predict the time-dependent flowfield in a transonic turbine stage. Two analytical techniques are used to understand the unsteady physics. One technique takes into account interaction effects associated with reflected waves bouncing between blade rows while the other neglects them. Both techniques model the exact blade counts using phase-lag boundary conditions. The analytical techniques are validated by comparing to unsteady pressure measurements which have been made on the vane and blade surfaces at midspan. The analytical results are then used to understand the importance of interaction effects when the blade rows are close-coupled and when they are more widely spaced. The results show that interaction effects must be taken into account in order to accurately predict the unsteady loading on the upstream blade row. However, for the downstream blade row, interaction effects are second order and do not routinely need to be taken into account in the design process.
The Impact of Viscous Effects on the Aerodynamic Damping of Vibrating Transonic Compressor Blades — A Numerical Study
Authors: Grüber Björn , Carstens Volker
Published: August 2014

A parametric study which investigates the influence of viscous effects on the damping behaviour of vibrating compressor cascades is presented here.

To demonstrate the dependence of unsteady aerodynamic forces on the flow viscosity, a computational study was performed for a transonic compressor cascade of which the blades underwent tuned pitching oscillations while the flow conditions extended from fully subsonic to highly transonic flow. Additionally, the reduced frequency and Reynolds number were varied. In order to check the linear behavior of the aerodynamic forces, all calculations were carried out for three different oscillation amplitudes. Comparisons with inviscid Euler results helped identify the influence of viscous effects.

The computations were performed with a Navier-Stokes code, the basic features of which are the use of an AUSM upwind scheme, an implicit time integration, and the implementation of the Baldwin-Lomax turbulence model. In order to demonstrate the possibility of this code to correctly predict the unsteady behavior of strong shock-boundary layer interactions, the experiment of Yamamoto and Tanida on a self-induced shock oscillation due to shock-boundary layer interaction was calculated. A significant improvement in the prediction of the shock amplitude was achieved by a slight modification of the Baldwin Lomax turbulence model.

An important result of the presented compressor cascade investigations is that viscous effects may cause a significant change in the aerodynamic damping. This behaviour is demonstrated by two cases in which an Euler calculation predicts a damped oscillation whereas a Navier-Stokes computation leads to an excited vibration. It was found that the reason for these contrary results are shock-boundary-layer interactions which dramatically change the aerodynamic damping.
Multilevel Optimization of Rotor Bearing System With Dynamic Behavior Constraints
Authors: Shiau T. N. , Lee H. J. , Tsai Y. J.
Published: August 2014

The main purpose of this study is to investigate the multilevel optimization of rotor-bearing systems. The design variables include the shaft inner radius, the bearings stiffness, and the axial position of the bearings and disks. The design objectives are minimization of the shaft weight and transmitted force to the bearings. Constraints are placed on the critical speeds, the maximum shaft bending stress, and the maximum amplitude of the steady state response.

In the multilevel optimization, three levels are considered and each level includes a single objective and/or multiobjective with various design variables and constraints. For each level, the method of feasible direction (MFD) is used. In addition, the weighting method (WM) is used for multiobjective optimization. The dynamic analysis is carried out using the generalized polynomial expansion method.

The results show that the shaft weight and transmitted forces can be simultaneously reduced with the multilevel technique and are better than those obtained using the multiobjective optimization technique with only a single level.
The Effect of Viscous and Hysteretic Damping on Rotor Stability
Authors: Cerminaro Anthony M. , Nelson Frederick C.
Published: August 2014

A rotordynamic finite element program has been modified to include internal viscous and hysteretic damping. A numerical study has been done on a rotor system with variable external damping to predict the effect of various amounts of internal viscous and hysteretic damping on rotor stability. Three cases are considered: (1) external viscous damping with internal viscous damping; (2) external hysteretic damping with internal hysteretic damping; and (3) external viscous damping with internal hysteretic damping.
Performance of a Foil-Magnetic Hybrid Bearing
Authors: Swanson Erik E. , Heshmat Hooshang , Walton James
Published: August 2014

To meet the advanced bearing needs of modern turbomachinery, a hybrid foil-magnetic hybrid bearing system was designed, fabricated and tested in a test rig designed to simulate the rotor dynamics of a small gas turbine engine (31 kN to 53 kN thrust class). This oil-free bearing system combines the excellent low and zero-speed capabilities of the magnetic bearing with the high load capacity and high speed performance of the compliant foil bearing. An experimental program is described which documents the capabilities of the bearing system for sharing load during operation at up to 30,000 RPM and the foil bearing component’s ability to function as a back-up in case of magnetic bearing failure. At an operating speed of 22,000 RPM, loads exceeding 5300 N were carried by the system. This load sharing could be manipulated by an especially designed electronic control algorithm. In all tests, rotor excursions were small and stable. During deliberately staged magnetic bearing malfunctions, the foil bearing proved capable of supporting the rotor during continued operation at full load and speed, as well as allowing a safe rotor coast-down. The hybrid system tripled the load capacity of the magnetic bearing alone and can offer a significant reduction in total bearing weight compared to a comparable magnetic bearing.
Comparative Studies on Short and Long Length-Scale Stall Cell Propagating in an Axial Compressor Rotor
Authors: Inoue M. , Kuroumaru M. , Tanino T. , Yoshida S. , Furukawa M.
Published: August 2014

In a low-speed compressor test rig at Kyushu University, multiple short length-scale stall cells appeared under a mild stall condition and turned into a long length-scale cell under a deep stall condition. Then, for the both types of stall cell, the pressure distribution on the casing wall and the velocity distributions upstream and downstream of the rotor have been measured by high response pressure transducers and a slanted hot-wire, respectively. The time-dependent ensemble averages of these distributions have been obtained phase-locking to both of the rotor and the stall cell rotation by using a so-called ‘double phase-locked averaging technique’ developed by the authors.

Structure of the two stall cells are compared with each other: The short length-scale stall cell is characterized by a concentrated vortex spanning from the casing wall ahead of the rotor to the blade suction surface. In the long length-scale stall cell, the separation vortices go upstream irregularly when blade separation develops in the front half of the cell, and reenter the rotor on the hub side in the rear half of it.

The unsteady aerodynamic force and torsional moment acting on the blade tip section have been evaluated from the time-dependent ensemble averages of the casing wall pressure distribution. The force fluctuation due to the short length-scale cells is somewhat smaller than that for the long length-scale cell. The blade suffers two peaks of the force during a period of the short length-scale cells passing through it. The moment fluctuation for the short length-scale cells is considerably larger than that for the long length-scale cell.
Analysis of Unsteady Blade Row Interaction Using Nonlinear Harmonic Approach
Authors: Chen T. , Vasanthakumar P. , He L.
Published: August 2014

An efficient non-linear harmonic methodology has been developed for predicting unsteady blade row interaction effects in multistage axial flow compressors. Flow variables are decomposed into time averaged variables and unsteady perturbations, resulting in time averaged equations with deterministic stress terms depending on the unsteady perturbation. The non-linear interaction between the time averaged flow field and the unsteady perturbations are included by a simultaneous pseudotime integration approach, leading to a strongly coupled solution. The stator/rotor interface treatment follows a flux averaged characteristic based mixing plane approach and includes the deterministic stress terms due to upstream running potential disturbances and downstream running wakes, resulting in the continuous nature of all parameters across the interface. The basic computational methodology is applied to the three-dimensional Navier-Stokes equations and validated against several cases. Results show that this method is much more efficient than the non-linear time-marching methods while still modeling the nonlinear unsteady blade row interaction effects.
Investigation of the Unsteady Rotor Flow Field in a Single HP Turbine Stage
Authors: Kost Friedrich , Hummel Frank , Tiedemann Maik
Published: August 2014

Within a European project a high-pressure turbine stage was investigated at DLR, Göttingen. The investigations consisted primarily of experiments carried out in the windtunnel for Rotating Cascades (RGG), but some numerical work was also performed.

Detailed measurements were carried out at mid section of a turbine rotor using a Laser-2-Focus device which served as a velocimeter measuring 2D-velocity vectors and turbulence quantities and as a tool to determine the concentration of coolant ejected at the trailing edge of the stator blades. The measurement of coolant concentration downstream of the stator and inside the rotor provided a detailed picture of the stator wake development and its interaction with the moving rotor. Axial measurement locations reached from the stator exit through the rotor to a downstream measurement plane. Measurement results are presented as instantaneous flow values. Unsteady flow vectors and turbulence intensities could be related at 16 time instants representing one rotor blade passsing period to the wake development made visible by the coolant concentration. The measured unsteady flow vectors and unsteady pressures, measured with semi-conductor pressure transducers, are compared with results from a numerical calculation using the Navier-Stokes code “TRACE-U” which allows the computation of the unsteady flow field. The measured steady and unsteady flow quantities served to validate the Navier-Stokes code. A comparison of the wake entropy trajectories outside the blade boundary layers and at the wall gives an impression of the lag between the arrival time of the wake in the freestream near the blade surface and the time the boundary layer quantities at the blade surface itself are affected.
Numerical and Experimental Investigation of Trailing Edge Vortex Shedding Downstream of a Linear Turbine Cascade
Authors: Gehrer A. , Lang H. , Mayrhofer N. , Woisetschläger J.
Published: August 2014

In this study, the evolution of the unsteady trailing edge vortex street downstream a linear turbine cascade is experimentally and computationally investigated. In a transonic cascade test stand, Laser Doppler velocimeter (LDV) measurements were acquired in several axial planes downstream of the blade trailing edge. In addition, direct detection of density changes near the trailing edge provide information about the frequency of a vortex shedding cycle.

A two-dimensional upwind-biased Navier-Stokes solver has then been used to perform a series of steady and unsteady cascade simulations, allowing an in-depth study into the mechanisms of the trailing edge vortex shedding.

The numerical results are compared with the experimental data to test the quality of the numerical simulations.
Investigation of the Unsteady Rotor Aerodynamics in a Transonic Turbine Stage
Authors: Dénos R. , Arts T. , Paniagua G. , Michelassi V. , Martelli F.
Published: August 2014

The paper focuses on the unsteady pressure field measured around the rotor mid-span profile of the VKI Brite transonic turbine stage. The understanding of the complex unsteady flow field is supported by a quasi-3D unsteady Navier-Stokes computation using a k-? turbulence model and a modified version of the Abu-Ghannam and Shaw correlation for the onset of transition.

The agreement between computational and experimental results is satisfactory. They both reveal the dominance of the vane-shock in the interaction. For this reason, it is difficult to identify the influence of vane-wake ingestion in the rotor passage from the experimental data. However, the computations allow to draw some useful conclusions in this respect.

The effect of the variation of the rotational speed, the stator-rotor spacing and the stator trailing edge coolant flow ejection is investigated and the unsteady blade force pattern is analyzed.
Unsteady Transport Mechanisms in an Axial Turbine
Authors: Casciaro Claudia , Treiber Martin , Sell Michael
Published: August 2014

A numerical analysis using a commercial unsteady Navier-Stokes solver has been performed on a pin/blade configuration, in order to assess the efficacy of a commercial code in calculating time-periodic interactions and to gain a better understanding of the unsteady flow physics in axial turbines.

Two cases have been investigated, with the pin positioned at 25% and 50% of true chord ahead of the leading edge. Both configurations have been computed both 2D and 3D.

The 2D case was used to examine the influence of numerical parameters, such as mesh, time and space discretisation. The 3D case allowed insight into the complete flow field including the wake influence on the secondary flow and mixing processes of the blade row.

The basic mechanisms of the wake-blade interaction proved, as expected, to be the same for both pin positions. Yet, as the closest pin wake interaction with the blade field was much stronger, its features have helped to identify the respective roles of wake fluid transport and blade potential field for both cases. The latter effect, noticeably strong with the thick leading edge blade form presented in this study, has been often neglected, and this study helps shed new light on this phenomenon.

The code used had been validated in previous work for pin-free steady flow within the same blade row and the new time dependent case has served to confirm the code range and limitations.
A Numerical Study of Secondary Flow in Axial Turbines With Application to Radial Transport of Hot Streaks
Authors: Prasad Dilip , Hendricks Gavin J.
Published: August 2014

The flow field in a low-speed turbine stage with a uniform inlet total pressure is studied numerically. A circular hot streak is superposed on the vane inlet flow. In agreement with previous experimental and numerical work, it is observed that while the streak passes through the vane unaltered, significant radial transport occurs in the rotor. Furthermore, despite the unsteady nature of the flow field, the steady theory of Hawthorne (1974) is found to predict the radial transport velocity well. Making use of this theory, it is shown that the secondary vorticity in the rotor may be attributed to the effects of density stratification, the spatial variation of the vane exit flow angle and the relative eddy. It then follows that the extent of radial transport in the rotor may be influenced by altering the vane exit flow angle distribution. The present study examines one means by which this may be effected, viz., varying the vane twist across the span. It is shown that a “reverse” twist, wherein the flow angle at the vane exit is larger near the tip than it is at mid-span reduces the secondary flow (and consequently, radial transport) in the blade passage. On the other hand, “positive” twist, in which the vane exit flow angle decreases with span is found to markedly worsen the radial transport in the blade. It is to be noted that varying the vane twist is but one method to obtain the desired exit flow angle; possibilities for altering other aspects of the vane geometry also exist.
Characteristics of a Radial Inflow Turbine With Internal Component Measurements: Part I — Data Review
Authors: Japikse David
Published: August 2014

The design and off-design performance of a radial inflow turbine has been carefully measured with both internal component and overall stage data. Observations have been made concerning instrumentation methods, appropriate correlating variables, and resultant component performance. The use of internal measurements changes the conclusions reached concerning component performance. Recommendations for further analysis are given.
Tip Clearance Effects in a Turbine Rotor: Part I — Pressure Field and Loss
Authors: Xiao Xinwen , McCarter Andrew A. , Lakshminarayana Budugur
Published: August 2014

This paper presents an experimental investigation of the effects of the tip clearance flow in an axial turbine rotor. The effects investigated include the distribution and the development of the pressure, the loss, the velocity, and the turbulence fields. These flow fields were measured using the techniques of static pressure taps, rapid response pressure probes, rotating five-hole probes, and Laser Doppler Velocimeter. Part I of this paper covers the loss development through the passage, and the pressure distribution within the passage, on the blade surfaces, on the blade tip, and on the casing wall. Regions with both the lowest pressure and the highest loss indicate the inception and the trace of the tip leakage vortex. The suction effect of the vortex slightly increases the blade loading near the tip clearance region. The relative motion between the turbine blades and the casing wall results in a complicated pressure field in the tip region. The fluid near the casing wall experiences a considerable pressure difference across the tip. The highest total pressure drop and the highest total pressure loss were both observed in the region of the tip leakage vortex, where the loss is nearly twice as high as that near the passage vortex region. However, the passage vortex produces more losses than the tip leakage vortex in total. The development of the loss in turbine rotor is similar to that observed in cascades. Part II of this paper covers the velocity and the turbulence fields.
Using Sweep and Dihedral to Control Three-Dimensional Flow in Transonic Stators of Axial Compressors
Authors: Gümmer V. , Wenger U. , Kau H.-P.
Published: August 2014

The paper describes an advanced 3D blading concept for highly-loaded transonic compressor stators. The concept takes advantage of the aerodynamic effects of sweep and dihedral. To the knowledge of the authors this is the first approach reported in open literature that combines those two basic types of lean in an engine-worthy aerofoil design. The paper makes a contribution to the understanding of the endwall effect of both features with special emphasis put on sweep. The advanced 3D blading concept was applied to an Engine Section Stator (ESS) of an aero-engine fan. In order to demonstrate how three-dimensional flow can be controlled, numerical analysis of the flow structure in a conventional and an advanced stator configuration was performed using a 3D-Navier-Stokes solver. The numerical analysis showed the advanced blade improving the radial loading distribution as well as the three-dimensional endwall boundary layer development. In particular a strong hub corner stall could be largely alleviated. High-speed rig testing of the advanced ESS confirmed the concept and showed good qualitative agreement between measurement and prediction. The work presented was closely linked to the development the BR710 engine on which the advanced ESS is in service today.