bluuebunny/crossref_metadata_2025_split huggingface dataset:
Numerical and Experimental Investigations of Clocking Effect in a Two-Stage Compressor With πC = 3.7
Authors: Mileshin Victor , Druzhinin Yaroslav , Stepanov Alexander , Savin Nikolay 
Published: August 2015

In recent years, a number of studies in Russia and abroad was completed with the aim of decreasing pressure fluctuations and losses in blade cascades by controlling the unsteady interactions of blade rows (known as “clocking effect”) [1–4]. Tests of individual stages demonstrated that the clocking effect is responsible for 1.5–2.0% in efficiency and 50% in pressure fluctuations [5]. This paper presents the results of experimental and theoretical studies of the clocking effect on gas-dynamic characteristics of a high-loaded two-stage compressor simulating the first two stages of HPC for an advanced engine. The compressor is designed with the help of up-to-date 1D, 2D and 3D direct and inverse problem solutions and distinguished by high aerodynamic loads of stages with πk=3.7 total pressure ratio, 17% stall margin and 88% adiabatic efficiency at Ncor=88% rotational speed that was demonstrated experimentally [6].
The compressor was tested at CIAM’s C-3 test facility in the assembly with d=0.5, 0.75, 1.0-mm tip clearance in both rotors (relative clearance in first stage 4.6·10−3; 6.9·10−3; 9.2·10−3 and relative clearance in second stage 9.1·10−3; 13.7·10−3; 18.3·10−3). When tested, clocking effects were checked up for separate and simultaneous changes in clocking positions of stator and rotor blade rows. Indications of a blade tip-timing system and pressure pulsation sensors were used as experimental data. Earlier, it was shown that physics of the rotor clocking is a wake interaction which modifies the behavior of a boundary layer in Rotor 2 blades. This work studies the mechanism of rotor clocking in combination with changes in angular position of Rotor 2 blades due to interactions with Rotor 1 wakes. Tests showed that changes in the clocking position of the rotor with a multiple number of Rotor 1 and Rotor 2 blades affected the static position of Rotor 2 blades causing re-position of the blades depending on the rotor clocking-position. To confirm this result, 3D unsteady aerodynamic calculation was completed with the help of NUMECA software package simulating one of the test points. This work presents the calculated and experimental data showing that vortex wakes from Rotor 1 blades extend downstream, reach Rotor 2 and cause a variable aerodynamic load and a variable blade pitch.
Two-Phase Computational Modelling of a Spiral Bevel Gear Using a Eulerian Multiphase Model
Authors: Al Baydu C. , Simmons Kathy , Morvan Hervé P. 
Published: August 2015

The efficiency of power transmission systems is increasingly targeted with a view to reducing parasitic losses and improving specific fuel consumption (SFC). One of the effects associated with such parasitic losses is gear windage power loss and this mechanism can be a significant contributor to overall heat-to-oil within large civil aeroengines. The University of Nottingham Technology Centre in Gas Turbine Transmission Systems has been conducting experimental and computational research into spiral bevel gear windage applicable to an aeroengine internal gearbox (IGB).
The two-phase flows related to gear lubrication, shrouding and scavenging are complex. Good understanding of such flows can be used to balance lubrication needs with need to minimise oil volumes and parasitic losses. Previous computational investigations have primarily employed discrete phase modelling (DPM) to predict oil behaviour under the shroud [1, 2]. In this paper modelling capability has been investigated and extended through application of FLUENT’s Eulerian multiphase model. In addition, DPM modelling linked to FLUENT’s Lagrangian film model has been conducted. A control volume with periodic symmetry comprising a single tooth passage of the bevel gear has been modelled to keep the computational cost down.The results from both models are compared to each other and to available experimental visual data.
Both models are found to perform acceptably with the Eulerian multiphase model yielding results closer to those observed experimentally. The use of DPM with a Eulerian film model is suggested for future work and extension to a full 360° model is recommended.
Velocity Measurements in the Vicinity of the Exit and Inlet of a Spiral Bevel Gear Shroud
Authors: Simmons Kathy , Johnson Graham 
Published: August 2015

In many aeroengines the accessory power offtake is achieved using a spiral bevel gear set running off one of the main shafts. The crown and bevel gears are housed in an internal gearbox. Over the past few years the Nottingham University Technology Centre (UTC) in Gas Turbine Transmission Systems has researched flow near spiral bevel gears both computationally and experimentally using a purpose-built test rig. In the current investigation the rig was configured with a Trent crown gear and slightly modified shroud covering the full 360° of the gear. No external containment chamber was fitted and all testing was conducted single-phase (air only) at 5,000 rpm. Laser Doppler Anemometry (LDA) was used to obtain the three components of flow velocity at a shroud exit slot and at shroud inlet. A 2D system was utilised and thus two measurements were required at each point to give the 3 velocity components.
The LDA technique enabled detailed mapping of flow features over the chosen regions, which included areas very near the shroud surfaces. Data was obtained over two measurement regions:
1) a volume mapping the air “jet” exiting the shroud exit slot at top dead centre (TDC) and
2) an area capturing the flow structures local to the shroud inlet.
Combined the results form an excellent set of high quality, detailed, 3-component flow data for direct use in validating CFD models and/or to define CFD boundary conditions.
At the shroud exit slot the maximum velocity measured was 46.2 m/s with the jet velocity dispersing over the measurement volume such that by 26 mm from slot plane the maximum velocity was less than 20 m/s. The jet angle was found to be only 16° off perpendicular azimuthally and 22° down from perpendicular. Data from the top 5 slots shows good similarity indicating the detailed data for the TDC slot is probably applicable to all slots.
Air entering the shroud comes down the shroud face and up the rotating end face of the gear shaft. The azimuthal velocity component at shroud inlet was around 20 m/s; this is of the order of 50% of the maximum linear shaft surface speed. Within 3 mm of the rotating gear face the azimuthal velocity is less than 1 m/s. Detailed measurements were obtained only at one angular location but sufficient additional measurements were obtained to determine that for the purposes of CFD validation the results can be considered representative.
A Computational Investigation Into the Effects of Inlet Swirl on a Shrouded Spiral Bevel Gear
Authors: Turner A. , Morvan H. P. , Simmons K. 
Published: August 2015

This paper presents a computational investigation into the effects of inlet swirl on the fluid flows computed for a spiral bevel gear and shroud pairing. This is with a specific focus on aeroengine internal gearboxes where the power offtake gear is located in a chamber where a highly swirled environment exists.
Previous work for a single rotating gear and stationary shroud [1] has been performed in isolation and with idealised boundary conditions at both the inlet and outlet of the system. This effectively decouples the shrouded gear subsystem from the rest of the gearbox. In the present study a parametric investigation has been conducted varying the amount of inlet swirl and the effect this has on the mass flowrate of air “pumped” through the gear. The paper presents data showing that compared to no inlet swirl, a higher mass flowrate is induced when there is swirl, with up to 20% higher mass flowrate occurring when the tangential velocity at shroud inlet is 50% of the axial velocity component. For swirl above 50% the mass flowrate drops back somewhat remaining higher than for the no-swirl case for values investigated. Gear windage power loss is a function of mass flowrate and consequently higher windage losses occur for higher mass flowrates.
In an aeroengine flow through the gear exits through slots in the shroud and this paper shows that the swirl velocity component at the shroud exit holes is relatively insensitive to the velocity components entering the gear/shroud system for a given geometry.
Further to this, the effect of the shroud outlet geometry on the flow leaving the back of the gear has been investigated and quantified. It is shown that accurate geometric representation is required as the outlet geometry has a significant effect on the computed mass flow through the gear-shroud system and consequently on the computed windage power loss.
Independent Effects of Surface and Gas Temperature on Coal Flyash Deposition in Gas Turbines at Temperatures Up to 1400°C
Authors: Laycock Robert , Fletcher Thomas H. 
Published: August 2015

Deposition of coal flyash in gas turbines has been studied to support the concept of integrated gasification combined cycle (IGCC). Although particle filters are used in IGCC, small amounts of ash particles less than 5 μm diameter enter the gas turbine. Previous deposition experiments in the literature have been conducted at temperatures up to about 1288°C. However, few tests have been conducted that reveal the independent effects of gas and surface temperature, and most have been conducted at gas temperatures lower than 1400°C.
The independent effects of gas and surface temperature on particle deposition in a gas turbine environment were measured using the Turbine Accelerated Deposition Facility (TADF) at Brigham Young University. Gas temperatures were measured with a type K thermocouple and surface temperatures were measured with two-color pyrometry using the RGB signals from a camera. This facility was modified for testing at temperatures up to 1400°C. Subbituminous coal fly ash, with a mass mean diameter of 4 μm, was entrained in a hot gas flow at a Mach number of 0.25. A nickel base super alloy metal coupon 2.5 cm in diameter was held in this gas stream to simulate deposition in a gas turbine. The gas temperature (and hence particle temperature) governs the softening and viscosity of the particle, while the surface temperature governs the stickiness of the deposit.
Two tests series were therefore conducted. The first series used backside cooling to hold the initial temperature of the deposition surface (Ts,i) constant at 1000°C while varying the gas temperature (Tg) from 1250°C – 1400°C. The second series held Tg constant at 1400°C while varying the initial Ts,i from 1050°C to 1200°C by varying the amount of backside cooling. Capture efficiency and surface roughness were calculated. Capture efficiency increased with increasing Tg. Capture efficiency also initially increased with Ts,i until a certain threshold temperature where capture efficiency began to decrease with increasing Ts,i.
Methane-Oxygen Flame Stability in a Generic Premixed Gas Turbine Swirl Combustor at Varying Thermal Power and Pressure
Authors: Runyon Jon , Marsh Richard , Valera-Medina Agustin , Giles Anthony , Morris Steve , Pugh Daniel , Sevcenco Yura , Bowen Phil 
Published: August 2015

At low thermal power (<5 kW) conditions, nitrogen and carbon dioxide were added as diluents to a premix of methane-oxygen in an atmospheric generic swirl burner. Results indicate that CO2-diluted oxy-methane flames have a wider stability range than N2-diluted flames in terms of overall oxygen concentration in the premix. Bulk flow Reynolds number, augmented by varying the size of the burner exit nozzle, was also found to increase the stability limits of flames diluted with both CO2 and N2, as the increased flow velocity offsets the higher burning velocity of the oxyfuel mixture. A combination of differing transport properties between diluents and the resulting flame chemistry produces a change in the structure of the premixed oxyfuel swirl flame, shown by combustion PIV to affect the observed lean and rich stability limits.
Utilising the results at low thermal power conditions, enhanced-oxygen combustion of a methane-air flame was investigated in a pressurized generic swirl burner operating at higher thermal power (<50 kW) conditions and pressures up to 3 bar absolute. Over a range of increasing thermal powers, it is seen that a relatively small amount of pure oxygen addition can shift the equivalence ratio at which the lean stability limit or rich stability limit are reached compared with the same phenomenon observed for a methane-air flame. Pressurised operation with CO2 dilution up to 15.5 mol% was validated through stability limit and emissions gas analysis, giving further support to the use of exhaust gas recirculation in premixed swirl-stabilized burners for oxyfuel combustion.
Oxide/Oxide Ceramic Matrix Composite (CMC) Exhaust Mixer Development in the NASA Environmentally Responsible Aviation (ERA) Project
Authors: Kiser J. Douglas , Bansal Narottam P. , Szelagowski James , Sokhey Jagdish (Jack) , Heffernan Tab , Clegg Joseph , Pierluissi Anthony , Riedell Jim , Wyen Travis , Atmur Steven , Ursic Joseph 
Published: August 2015

LibertyWorks®, a subsidiary of Rolls-Royce Corporation, first studied CMC (ceramic matrix composite) exhaust mixers for potential weight benefits in 2008. Oxide CMC potentially offered weight reduction, higher temperature capability, and the ability to fabricate complex-shapes for increased mixing and noise suppression.
In 2010, NASA was pursuing the reduction of NOx emissions, fuel burn, and noise from turbine engines in Phase I of the Environmentally Responsible Aviation (ERA) Project (within the Integrated Systems Research Program). ERA subtasks, including those focused on CMC components, were being formulated with the goal of maturing technology from Proof of Concept Validation (Technology Readiness Level 3 (TRL 3)) to System/Subsystem or Prototype Demonstration in a Relevant Environment (TRL 6).
In April 2010, the NASA Glenn Research Center (GRC) and Rolls-Royce (RR) jointly initiated a CMC Exhaust System Validation Program within the ERA Project, teaming on CMC exhaust mixers for subsonic jet engines. The initial objective was to fabricate and characterize the performance of a 0.25 scale low bypass exhaust system that was based on a RR advanced design, with a 16-lobe oxide/oxide CMC mixer and tail cone (center body). Support Services, LLC (Allendale, MI) and COI Ceramics, Inc. (COIC) supported the design of a mixer assembly that consisted of the following oxide/oxide CMC components mounted on separate metallic attachment flanges: a) a lobed mixer and outer fan shrouds, and b) a tail cone. TRL 4 (Component/Subscale Component Validation in a Laboratory Environment) was achieved in a cost-effective manner through subscale rig validation of the aerodynamic and acoustic performance via testing at ASE FluiDyne (Plymouth, MN) and at NASA GRC, respectively. This encouraged the NASA/ RR/COIC team to move to the next phase of component development; full scale CMC mixer design for a RR AE3007 engine. COIC fabricated the full scale CMC mixer, which was vibration tested at GRC under conditions simulating the structural and dynamic environment of a mixer. Air Force Research Laboratory (AFRL, Wright-Patterson Air Force Base (WPAFB)) provided test support by assisting with instrumentation and performing 3D laser vibrometry to identify the mixer mode shapes and modal frequencies over the engine operating range.
Successful vibration testing demonstrated COIC’s new process for fabricating full scale CMC mixers and the durability of the Oxide CMC component at both room and elevated temperatures. A TRL≈5 (Component Validation in a Relevant Environment) was attained and the CMC mixer was cleared for ground testing on a Rolls-Royce AE3007 engine for performance evaluation to achieve TRL 6.
Multi-Angular Flame Measurements and Analysis in a Supersonic Wind Tunnel Using Fiber Based Endoscopes
Authors: Ma Lin , Wickersham Andrew J. , Xu Wenjiang , Peltier Scott J. , Ombrello Timothy M. , Carter Campbell D. 
Published: August 2015

This paper reports new measurements and analysis made in the Research Cell 19 supersonic wind-tunnel facility housed at the Air Force Research Laboratory. The measurements include planar chemiluminescence from multiple angular positions obtained using fiber based endoscopes (FBEs) and the accompanying velocity fields obtained using particle image velocimetry (PIV). The measurements capture the flame dynamics from different angles (e.g., the top and both sides) simultaneously. The analysis of such data by proper orthogonal decomposition (POD) will also be reported.
Non-intrusive and full-field imaging measurements provide a wealth of information for model validation and design optimization of propulsion systems. However, it is challenging to obtain such measurements due to various implementation difficulties such as optical access, thermal management, and equipment cost. This work therefore explores the application of FBEs for non-intrusive imaging measurements in supersonic propulsion systems. The FBEs used in this work are demonstrated to overcome many of the practical difficulties and significantly facilitate the measurements. The FBEs are bendable and have relatively small footprints (compared to high-speed cameras), which facilitates line-of-sight optical access. Also, the FBEs can tolerate higher temperatures than high-speed cameras, ameliorating the thermal management issues. Lastly, the FBEs, after customization, can enable the capture of multiple images (e.g., images of the flowfields at multi-angles) onto the same camera chip, greatly reducing the equipment cost of the measurements.
The multi-angle data sets, enabled by the FBEs as discussed above, were analyzed by POD to extract the dominating flame modes when examined from various angular positions. Similar analysis was performed on the accompanying PIV data to examine the corresponding modes of the flowfields. The POD analysis provides a quantitative measure of the dominating spatial modes of the flame and flow structures and is an effective mathematical tool to extract key physics from large data sets such as the high-speed measurements collected in this study. However, past POD analysis has been limited to data obtained from one orientation only. The availability of data at multiple angles in this study is expected to provide further insights into the flame and flow structures in high-speed propulsion systems.
Performance of a Splittered Transonic Rotor With Several Tip Clearances
Authors: Gannon Anthony J. , Hobson Garth V. , Turner Mark G. , Sanz Wolfgang 
Published: August 2015

The effect of increasing the tip-gap size on the performance of a splittered transonic rotor is presented. Tip clearance has a large influence on the performance and efficiency of compressors and fans during operation. In a gas turbine engine the ratio of tip-gap to blade height or span usually increases in the direction of flow. The front stages usually have a smaller ratio of tip-gap to blade height than the aft core stages. In addition the front stages are usually operating in the transonic regime while the rear stages operate sub-sonically.
In order to be representative of these differing flow regimes the results of a range of tests at varying tip-gaps and speeds from subsonic to transonic are presented. A highly loaded transonic axial splittered rotor is used as the test article in this study. Three experiments with cold tip gaps of 0.53 [mm], 0.76 [mm] and 0.99 [mm] are presented. Each experiment was run at six tip-Mach numbers ranging from Mach 0.72 to Mach 1.2 each over a full speed-line from choked to stalling conditions. Exit temperature and pressure profiles at the rotor exit are presented along with performance maps of pressure ratio and efficiency. Significant differences in performance in terms of pressure ratio, efficiency and operating range due to the tip-gap increase were observed and are presented.
The full mechanical geometry is available upon request to provide an open test case to evaluate simulation codes. This includes the cold-shape, the cavity between the rotor and stationary exit and the mounting bolts and geometry. The manufacturing steps in the preparation of the material and methods used to form the abradable material used over the rotor are presented. In addition the inlet conditions taking into account the effect of relative humidity on gas properties are presented. The complete data set of experimental results is also available electronically.
A Mechanism-Based Approach From Low Cycle Fatigue to Thermomechanical Fatigue Life Prediction
Authors: Wu Xijia , Zhang Zhong 
Published: August 2015

Deformation and damage accumulation occur by fundamental dislocation and diffusion mechanisms. An integrated creep-fatigue theory (ICFT) has been developed, based on the physical strain decomposition rule that recognizes the role of each deformation mechanism and thus relate damage accumulation to its underlying physical mechanism(s). The ICFT formulates the overall damage accumulation as a holistic damage process consisting of nucleation and propagation of surface/subsurface cracks in coalescence with internally distributed damage/ discontinuities. These guiding principles run through both isothermal low cycle fatigue (LCF) and thermomechanical fatigue (TMF) under general conditions.
This paper presents a methodology using mechanism-based constitutive equations to describe the cyclic stress strain curve and the non-linear damage accumulation equation incorporating i) rate-independent plasticity-induced fatigue, ii) intergranular embrittlement, iii) creep and iv) oxidation to predict LCF and TMF lives of ductile cast iron (DCI). The complication of the mechanisms and their interactions in this material provide a good demonstration case for the model, which is in good agreement with the experimental observations.
Design Space of Foil Bearings for Closed Loop Supercritical CO2 Power Cycles Based on Three-Dimensional Thermo-Hydrodynamic Analyses
Authors: Kim Daejong 
Published: August 2015

The closed loop Brayton cycle with super critical CO2 (S-CO2) as an operating fluid is an attractive alternative to conventional power cycles due to very high power density. Foil gas bearings using CO2 is the most promising for small S-CO2 turbomachinery but there are many problems to address; large power loss due to high flow turbulence, lack of design/analysis tool due to non-ideal gas behavior, and lack of load capacity when they are used for large systems.
This paper presents high level design/analysis tool involving three-dimensional thermo-hydrodynamic analyses of radial foil bearings considering real gas effect and flow turbulence inside the film. Simulations are performed for radial foil bearing with 34.9mm in diameter lubricated with CO2 and N2 under various ambient conditions up to above 40 bar gauge pressure. The simulation results using the turbulence model still under-predict the measured data in open literature. However, the error between the prediction and measurements decreases as either speed or ambient pressure increases. In addition, general behavior of substantial increase in power loss with ambient pressure agrees with the measured data. The simulation results indicate the importance of detailed THD analysis of the foil bearings for prediction of power loss under severe turbulent condition.
A conceptual layout of rotor system for 10MWe S-CO2 loop is also presented along with realistic rotor weight and bearing load. A hybrid foil bearings with diameter of 102mm is suggested for gas generator rotor, and its power losses and minimum film thicknesses at various operating conditions are presented.