bluuebunny/crossref_metadata_2025 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.
Efficient Computation of Thermoacoustic Modes in Annular Combustion Chambers Based on Bloch-Wave Theory
Authors: Mensah Georg A. , Moeck Jonas P.
Published: August 2015

Most annular combustors feature a discrete rotational symmetry so that the full configuration can be obtained by copying one burner–flame segment a certain number of times around the circumference. A thermoacoustic model based on the Helmholtz equation then admits special solutions of the so-called Bloch type that can be obtained by considering one segment only. We show that a significant reduction in computational effort for the determination of thermoacoustic modes can be achieved by exploiting this concept. The framework is applicable even in complex cases including a non-homogeneous temperature field and a frequency-dependent, spatially distributed flame response. A parametric study on a three-dimensional combustion chamber model is conducted using both the full scale chamber simulation and a one-segment model with the appropriate Bloch-type boundary conditions. The results for both computations are compared in terms of mode frequencies and growth rates as well as the corresponding mode shapes. This comparison demonstrates the benefits of the Bloch-wave based analysis. It is further shown that even the effect of circumferential asymmetries can be assessed based on computations of one burner–flame segment only by resorting to spectral perturbation theory.
Development of an Interlaminar Toughness Test for Ceramic Matrix Composites
Authors: Mansour Rabih , Maillet Emmanuel , Morscher Gregory N.
Published: August 2015

Interlaminar fracture properties in fiber-reinforced ceramic matrix composites have been found to be a limiting feature in certain applications, a limitation that intensifies at elevated temperatures. Few if any fibers bridge the cracks in the interlaminar direction, resulting in a weaker toughness and a higher probability of failure in this direction. The main challenge in evaluating interlaminar fracture toughness is the difficulty to detect the crack initiation and to measure the crack length directly at high temperature, primarily because of the limited visual accessibility of the area of interest. A DCB test-method using a ceramic wedge is developed. The method uses electrical resistance to monitor crack growth for melt-infiltrated woven SiC/SiC composites at room temperature, with the goal of doing the same at high temperature. In situ optical measurements of crack growth in addition to micro-CT post-inspection were used to validate and calibrate the resistance method. The estimated crack length was in excellent agreement with the measured crack length in the three considered specimen geometries. Preliminary estimates of Mode I energy release rate were also provided.
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.
Effects of Averaging the Heat-Transfer Coefficient on the Predicted Material Temperature Distribution
Authors: Shih T. I-P. , Lee C.-S. , Bryden K. M.
Published: August 2015

The heat-transfer coefficient (HTC) in internal-coolant passages can vary appreciably about a heat-transfer enhancement feature such as a pin fin, a rib, and a concavity because of stagnation regions and wakes about the enhancement feature. However, the computed or measured HTC is often averaged spatially in the spanwise direction or over some region when used in the design of cooling strategies. Since the variation in the HTC could be a factor of eight or more about an enhancement feature, it is of interest to understand the effects of averaging the HTC on the predicted temperature distribution in the solid subjected to the heating and cooling. In this computational study, a flat plate of thickness H (1 mm) and length L = 20H is heated on one side by either a constant heat flux (68 W/cm2) or a constant HTC (1,167.2 W/m2-K) and a constant hot-gas temperature (1,482 °C). On the cooled side, the free stream or bulk temperature is kept constant (400 °C) and the average HTC (1,442.5 W/m2-K) is kept constant as well. This average HTC on the cooled side is the average of a higher HTC (hH) and a lower HTC (hL). Two types of changes from hH to hL are considered — abrupt (or step) and gradual. When the HTC changes abruptly, hH is imposed over LH, and hL is imposed over LL=L–LH. When the HTC changes gradually from hH to hL, hH is imposed from from x = 0 to LH/2, and hL is imposed from x = 3LH/2 to L with a smooth variation in the HTC to connect hH and hL. Results obtained show that when the averaged HTC is used, the maximum temperature in the plate is 900 °C on the heated side of the plate. However, if the variation in the HTC is accounted for, then the maximum temperature in the plate could be as high as 1.363 times the maximum temperature predicted by assuming an averaged HTC. Also, for the range of parameters studied, the difference in the maximum and minimum temperature in the plate can increase by a factor of 16, which strongly affects thermal stress.
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.
Assessment of Scale Adaptive Simulation Model for Honeywell Combustor
Authors: Mahapatra Debabrata , Basani Jaydeep , Rida Samir
Published: August 2015

The complex flow phenomena inside a gas turbine combustor demands alternative simulation methods to the Reynolds Averaged Navier-Stokes (RANS) model, where a portion of turbulence scales is resolved inside the flow domain. Large Eddy Simulation (LES) is the most-widely acknowledged method for its attractive feature of resolving large turbulent structures down to the grid limit for the entire flow domain. However, for practical industrial problems where the Reynolds number is high and the flow domain is large, the grid resolution for LES becomes excessively high making it computationally very expensive. Scale Adaptive Simulation (SAS), on the other hand, adjusts to the resolved structures in an Unsteady RANS (URANS) simulation resulting in LES-like behavior in unsteady regions of the flow field. At the same time, it provides RANS capabilities in the stable flow regions. It allows a larger time step than LES resulting in the possibility of computation time advantage with LES-like solution fidelity. In the current paper, the SAS model is compared to the LES model for a Honeywell combustor using the commercial CFD code ANSYS FLUENT. Several time-steps are considered for SAS simulations. Results show that SAS is promising in terms of predicting combustor performance parameters like LES, but with a substantially reduced turn-around time.
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.
Effects of Turbulence Intensity and Scale on Turbine Blade Heat Transfer
Authors: Boyle Robert J. , Ameri Ali A.
Published: August 2015

The effects of turbulence intensity and length scale on turbine blade heat transfer and aerodynamic losses are investigated. The importance of freestream turbulence on heat transfer increases with Reynolds number and turbulence intensity, and future turbine blade Reynolds numbers are expected to be higher than in current engines. Even when film cooling is used, accurate knowledge of baseline heat transfer distributions are needed. Heat flux reductions due to film cooling depend on the ratio of film cooled-to-solid blade heat transfer coefficients. Comparisons are made between published experimental data and published correlations for leading edge heat transfer. Stagnation region heat transfer rates of vanes and blades of high pressure turbines can be nearly double those predicted when predictions neglect freestream turbulence effects. Correlations which included the scale of turbulence gave better agreement with data. Two-dimensional Navier-Stokes analysis were done for several existing test cases where measures of the turbulence scale are available. The test cases had significant regions where the flow was not fully turbulent. Freestream turbulence increases laminar heat transfer, but has little influence on turbulent heat transfer. The Navier-Stokes analysis included a model for the effects of high freestream turbulence on laminar or transitioning boundary layers. Comparisons were made with vane and rotor blade data, as well as with high Reynolds number test data that simulated the favorable pressure gradient regions seen in the forward portions of turbine blades. Predictions of surface heat transfer showed the appropriate trends in heat transfer with turbulence intensity and turbulence scale. However, the absolute level of agreement indicated that further verification of approaches to predicting turbulence intensity and scale effects is needed. Significant increases in losses were calculated for vane and rotor blade geometries as inlet turbulence increased.
Numerical Study of Flow Through 3x3 Multiswirler Arrays With Co and Counter Swirler Arrangements
Authors: Rojatkar Prachi , Jog Milind A. , Jeng San-Mou
Published: August 2015

A numerical study of turbulent flow through 3×3 multi swirler arrangement has been performed using the realizable k-ε turbulence model on a grid with about 19 million points. All co and alternate co/counter swirler configurations comprised of radial-radial swirler with counter rotating vanes are analyzed. The offset distances of swirler exit from the base wall of confinement of 0.02D and 0.31D are considered where D is the diameter of swirler exit. For both arrangements, a strong jet is issued as the flow exits individual swirl cup. Recirculation is observed at the walls and between each swirl cup along with the formation of central toroidal recirculation zone (CTRZ) at each individual swirler. It is observed that all co swirling arrangement has a stronger more compact individual CTRZ. On the other hand alternate co and counter arrangement produces more swirler-to-swirler interactions. When the offset between swirler exit and base wall of confinement is increased to 0.31D, longer but more compact CTRZ are formed at each swirler cup. The velocity gradient for 0.31D offset case is also higher than that of 0.02D. These differences in the flow field indicate better combustion performance, fuel breakup and flame anchoring for the higher offset case.
Predicting Flameholding for Hydrogen and Natural Gas Flames at Gas Turbine Premixer Conditions
Authors: Sullivan-Lewis Elliot , McDonell Vincent
Published: August 2015

Lean premixed gas turbines are one of the most common methods for stationary power generation. By creating a homogeneous mixture of fuel and air upstream of the combustion chamber, temperature variations are reduced within the combustor, which reduces emissions of nitrogen oxides. However, by premixing fuel and air, a potentially flammable mixture is established in a part of the engine not designed to contain a flame. If the flame propagates upstream from the combustor (flashback), significant engine damage can result. While significant effort has been put into developing flashback resistant combustors, these combustors are only capable of preventing flashback during steady operation of the engine. Transient events (e.g. auto ignition within the premixer, pressure spikes during ignition) can trigger flashback that cannot be prevented with even the best combustor design. In these cases, preventing engine damage requires designing premixers that will not allow a flame to be sustained. Experimental studies were conducted to determine under what conditions premixed flames of hydrogen and natural gas can be anchored in a simulated gas turbine premixer. Tests have been conducted at pressures up to 9 atm, temperatures up to 750 K, and free stream velocities between 20 and 100 m/s. Flames were anchored in the wakes of features typical of premixer passageways, including cylinders, steps and airfoils. The results of this study have been used to develop tools that predict when a flame remains anchored to a particular feature.
Compliant Gas Foil Bearings and Analysis Tools
Authors: Branagan Michael , Griffin David , Goyne Christopher , Untaroiu Alexandrina
Published: August 2015

Compliant gas foil bearings are composed of two layers of thin metallic foil and a thin film of gas to support the journal. The bottom foil creates an elastic structure which supports the top foil. This support structure can take a variety of shapes that range from a series of bumps around the circumference to a series of overlapping leaves. The top foil and the rotation of the rotor create a wedge of air that supports the rotor. The complaint foil structure deforms in response to the pressure developed within the gas film. These bearings have several advantages over conventional fluid film bearings. These advantages include reduced weight due to the elimination of the oil system, stable operation at higher speeds and temperatures, low power loss at high speeds and long life with little maintenance. Some disadvantages of gas foil bearings are low load capacities at low speed and modest stiffness and damping values. Due to these properties, compliant gas foil bearings are commonly used in specialized applications such as compressors for aircraft pressurization, engines for turboshaft propulsion, air cycle machines, turboexpanders, and small micro turbines. The ability to predict the behavior of these bearings and design them to meet the needs of the application is invaluable to the design process. This behavior can include things such as bearing stiffness, damping, and load capacity. Currently most foil bearing analysis tools involve some sort of coupling between hydrodynamics and structural analyses. These analysis tools can often have convergence issues and can require the use of empirically derived characteristics. This paper reviews the current status of the compliant gas foil bearings research, focusing mainly on the journal bump-type gas foil bearings and the development of the analysis tools for these bearings. This paper contributes to the field by making recommendations of the future developments of the analytical tools of journal bump-type gas foil bearings.
Structural and Rotordynamic Force Coefficients of a Shimmed Bump Foil Bearing: An Assessment of a Simple Engineering Practice
Authors: San Andrés Luis , Norsworthy Joshua
Published: August 2015

High speed rotors supported on bump-type foil bearings (BFBs) often suffer from large sub synchronous whirl motions. Mechanically preloading BFBs through shimming is a common, low cost practice that shows improvements in rotordynamic stability. However, there is absence of empirical information related to the force coefficients (structural and rotordynamic) of shimmed BFBs. This paper details a concerted study towards assessing the effect of shimming on a first generation BFB (L=38.1 mm, D =36.5 mm). Three metal shims, 120° apart, are glued to the inner surface of the bearing cartridge and facing the underside of the bump foil strip. The shim sets are of identical thickness, either 30 μm or 50 μm. Static load tests show that shimming produces nonlinear static load vs. deflection curves leading to a larger structural stiffness than for the bearing without shims. Torque measurements during shaft acceleration also demonstrate a shimmed BFB has a larger friction coefficient. For a static load of 14.3 kPa, dynamic loads with a frequency sweep from 250 Hz to 450 Hz are exerted on the BFB, without and with shims, to estimate its rotordynamic force coefficients while operating at ∼50 krpm (833 Hz). Similar measurements are conducted without shaft rotation. Results are presented for the original BFB (without shims) and the two shimmed BFB configurations. The direct stiffnesses of the BFB, shimmed or not, increase with excitation frequency thus evidencing a mild hardening effect. The BFB stiffness and damping coefficients decrease slightly for operation with rotor speed as opposed to the coefficients when the shaft is stationary. For frequencies above 300 Hz, the direct damping coefficients of the BFB with 50 μm thick shims are ∼ 30% larger than the coefficients of the original bearing. The bearing structural loss factor, a measure of its ability to dissipate mechanical energy, is derived from the direct stiffness and damping coefficients. The BFB with 50 μm thick shims has a 25% larger loss factor — average from test data collected at 300 Hz to 400 Hz — than the original BFB. Further measurements of rotor motions while the shaft accelerates to ∼50 krpm demonstrate the shimmed BFB (thickest shim set) effectively removes sub synchronous whirl motions amplitudes that were conspicuous when operating with the original bearing.
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.
Forced Response Analysis of High-Mode Vibrations for Mistuned Bladed Discs With Effective Reduced-Order Models
Authors: Duan Yongliang , Zang Chaoping , Petrov E. P.
Published: August 2015

This paper is focused on the analysis of effects of mistuning on the forced response of gas-turbine bladed discs vibrating in the frequency ranges corresponding to higher modes. For high modes the blade aerofoils are deformed during vibrations and the blade mode shapes differ significantly from beam mode shapes. A model reduction technique is developed for the computationally efficient and accurate analysis of forced response for bladed discs vibrating in high frequency ranges. High-fidelity finite element models of a tuned bladed disc sector are used to provide primary information about dynamic properties of a bladed disc and the blade mistuning is modelled by specially defined mistuning matrices. The forced response displacement and stress amplitude levels are studied for high frequency ranges. The effects of different types of mistuning are examined and the existence of high amplifications of mistuned forced response levels is shown for high-mode vibrations: in some cases, the resonance peak response of a tuned structure can be lower than out-of-resonance amplitudes of its mistuned counterpart.
A Comparison of Techniques for Particle Rebound Measurement in Gas Turbine Applications
Authors: Bons Jeffrey P. , Blunt Rory , Whitaker Steven
Published: August 2015

The rebound characteristics of 100–500μm quartz particles from an aluminum surface were imaged using the particle shadow velocimetry (PSV) technique. Particle trajectory data were acquired over a range of impact velocity (30–90 m/s) and impact angle (20°–90°) typical for gas turbine applications. The data were then analyzed to obtain coefficients of restitution (CoR) using four different techniques: (1) individual particle rebound velocity divided by the same particle’s inbound velocity (2) individual particle rebound velocity divided by inbound velocity taken from the mean of the inbound distribution of velocities from all particles (3) rebound velocity distribution divided by inbound velocity distribution related using distribution statistics and (4) the same process as (3) with additional precision provided by the correlation coefficient between the two distributions. It was found that the mean and standard deviation of the CoR prediction showed strong dependence on the standard deviation of the inbound velocity distribution. The two methods that employed statistical algorithms to account for the distribution shape [methods (3) and (4)] actually overpredicted mean CoR by up to 6% and CoR standard deviation by up to 100% relative to method (1). The error between the methods is shown to be a strong (and linear) function of correlation coefficient, which is typically 0.2–0.6 for experimental CoR data. Non-Gaussianity of the distributions only accounts for up to 1% of the error in mean CoR, and this largely from the non-zero skewness of the inbound velocity distribution. Particle rebound data acquired using field average techniques that do not provide an estimate of correlation coefficient are most accurately evaluated using method (2). Method (3) can be used with confidence if the standard deviation of the inbound velocity distribution is less than 10% of the mean velocity, or if a linear correction based on an assumed correlation coefficient is applied.
An Experimental and Computational Study on the Aerodynamic and Acoustic Characteristics of Case Fans
Authors: Wang Zhenyu , Hu Hui
Published: August 2015

We report the progress made in our recent study to develop an ultra-quiet axial fan for computer cooling applications. By using a commercially-available cooling fan as the baseline, a number of acoustically tailored modifications are implemented in order to reduce the noise level of the cooling fan, which includes optimizing the rotating blades and guide vanes according to axial fan design theory, adding an intake cone in the front of the hub to guide the airflow into the axial fan smoothly, and reducing the tip clearance to lower the noise generation due to tip vortex structures. A comparison study is conducted to measure the sound pressure level (SPL) of the new-design axial fan in an anechoic chamber, in comparison to that of the baseline fan, in order to assess the effects of the modifications on the fan noise reduction. The measurement results of our study reveal that, at the same flow rate, the SPL of the new-design fan would be up to 5 dB lower than that of the baseline fan. In addition to measuring the sound pressure levels (SPLs) of the fans, a digital particle image velocimetry (PIV) system is also used to conduct detailed flow field measurements to reveal the changes of the flow characteristics and unsteady vortex structures associated with the modifications. “Time-averaged” PIV measurements illustrate the ensemble-averaged statistics of the flow quantities such as mean velocity, Reynolds stress, and turbulence kinetic energy (TKE) distributions at the flow field around the axial fan. Moreover, a numerical study of the new-design fan is also performed and which is aimed to investigate the flow details inside of the fan. The validation comparison between numerical results and experimental results agree very well and that gives the strong confidence for the computational flow field in the fan can be regarded as useful complement for experiments. The detailed flow field measurements are correlated with the SPL measurements in order to elucidate underlying physics associated with the fan noise reduction.
Implicit Large Eddy Simulation of a Stalled Low Pressure Turbine Airfoil
Authors: Memory C. L. , Chen J. P. , Bons J. P.
Published: August 2015

Time-accurate numerical simulations were conducted on the aft-loaded L1A low pressure turbine airfoil at a Reynolds number of 22,000 (based on inlet velocity magnitude and axial chord length). This flow condition produces a non-reattaching laminar separation zone on the airfoil suction surface. The numerical code TURBO is used to simulate this flow field as an Implicit Large Eddy Simulation. Generally good agreement was found when compared to experimental time-averaged and instantaneous flow measurements. The numerical separation zone is slightly larger than that in the experiments, though integrated wake loss values improved from RANS-based simulations. Instantaneous snapshots of the numerical flow field showed the Kelvin Helmholtz instability forming in the separated shear layer and a large-scale vortex shedding pattern at the airfoil trailing edge. These features were observed in the experiments with similar sizes and vorticity levels. Power spectral density analyses revealed a global passage oscillation in the numerics that was not observed experimentally. This oscillation was most likely a primary resonant frequency of the numerical domain.
Testing and Modeling of an Acoustic Instability in Pilot-Operated Pressure Relief Valves
Authors: Allison Timothy C. , Brun Klaus
Published: August 2015

Pressure relief valves are included as an essential element of many compressor piping systems in order to prevent overpressurization and also to minimize the loss of process gas during relief events. Failure of the valve to operate properly can result in excessive quantities of vented gas and/or catastrophic failure of the piping system. Several mechanisms for chatter and instability have been previously identified for spring-loaded relief valves, but pilot-operated relief valves are widely considered to be stable. In this paper, pilot-operated pressure relief valves are shown to be susceptible to a dynamic instability under certain conditions where valve dynamics couple with upstream piping acoustics. This self-exciting instability can cause severe oscillations of the valve piston, damaging the valve seat, preventing resealing and possibly causing damage to attached piping. Two case studies are presented that show damaging unstable oscillations in a field installation and a blowdown rig, and a methodology is presented for modeling the instability by coupling a valve dynamic model with a 1-D transient fluid dynamics simulation code. Modeling results are compared with measured stable and unstable operation in a blowdown rig to show that the modeling approach accurately predicts the observed behaviors.
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.
Unsteady Simulations of a Low NOx LDI Combustor for Environmentally Responsible Aviation Engines
Authors: Drennan Scott A. , Kumar Gaurav , Steinthorsson Erlendur , Mansour Adel
Published: August 2015

A key objective of NASA’s Environmentally Responsible Aviation (ERA) research program is to develop advanced technologies that enable 75% reduction of LTO NOx emissions of N+2 aviation gas turbine engines relative to the CAEP 6 standard. To meet this objective, a new advanced multi-point fuel injector was proposed and tested under the NASA ERA program. The new injector, called the three-zone injector, or 3ZI, uses fifteen spray cups arranged in three zones. Swirling air flows into each cup and fuel is introduced via pressure swirl atomizers within the cup. Multiple design parameters impact the performance of the injector, such as the location of the atomizer within the spray cup, the spray angle and cup-to-cup spacing. To fully understand the benefits and trade-offs of various injector design parameters and to optimize the performance of the injector, detailed CFD simulations are an essential tool. Furthermore, the CFD methodology must allow easy changes in design parameters and guarantee consistent and comparable accuracy from one design iteration to the next. This paper investigates the use of LES in reacting and non-reacting flows and compares against the NOx experimental data for the multi-point atomization strategy of the injector. The CFD simulations employ an automatically generated Cartesian cut-cell meshing approach with mesh refinement applied near complex geometry and spray regions. Adaptive Mesh Refinement (AMR) is used to refine mesh in regions of high gradients in velocity and temperature. The CFD simulations use boundary and operating conditions based on experimental data for air flow and spray atomization obtained from LDV and PDPA characterizations of the spray respectively. The results are extended to reacting flow using a detailed reaction mechanism and predictions of NOx emissions are compared to experimental data. Overall NOx predictions were consistently less than experimental values. However, the NOx prediction trends showed excellent agreement with experimental data across the wide range of equivalence ratios investigated.