bluuebunny/crossref_metadata_2025 huggingface dataset:
Further Investigation of the Influence of Real-World Blade Profile Variation on the Aerodynamic Performance of Transonic Nozzle Guide Vanes
Authors: Ilott J. , Asghar A. , Allan W. D. E. , Woodason R.
Published: May 2012

This paper addresses the issue of aerodynamic consequences of variations in airfoil profile. An analysis of new and repaired airfoils was used to synthesize profiles representative of specific repair types. Five variations of a reference new low pressure turbine vane were obtained by changing the characteristic parameters of trailing-edge tweaking and laminate-repair methods used to refurbish turbine vanes. Flow visualization of shock structure and total pressure measurements were made by experimentation in a cascade rig and by calculations through Computational Fluid Dynamics (CFD). The performance of the modified profiles was compared with that of the reference new vane. The total pressure losses increased when the profile was bent at the trailing edge towards the pressure side. The losses for synthesized laminate repair profiles increased with an increase in the thickness of laminate repair. The numerical results were used to supplement experimental results in cases where the experimental conditions were not representative of typical design operating conditions.
Large Eddy Simulation of Flow and Convective Heat Transfer in a Gas Turbine Can Combustor With Synthetic Inlet Turbulence
Authors: Patil Sunil , Tafti Danesh
Published: May 2012

Large eddy simulations of swirling flow and the associated convective heat transfer in a gas turbine can combustor under cold flow conditions for Reynolds numbers of 50,000 and 80,000 with characteristic Swirl number of 0.7 are carried out. A precursor Reynolds Averaged Navier-Stokes (RANS) simulation is used to provide the inlet boundary conditions to the large-eddy simulation (LES) computational domain, which includes only the can combustor. A stochastic procedure based on the classical view of the turbulence as superposition of the coherent structures is used to simulate the turbulence at the inlet plane of the computational domain using the mean flow velocity and Reynolds stress data from the precursor RANS simulation. To further reduce the overall computational resource requirement and the total computational time, the near wall region is modeled using zonal two layer model. A novel formulation in generalized co-ordinate system is used for solution of effective tangential velocity and temperature in the inner layer virtual mesh. LES predictions are compared with the experimental data of Patil et al. [1] for the local heat transfer distribution on the combustor liner wall obtained using robust infrared thermography technique. The heat transfer coefficient distribution on the liner wall predicted from LES is in good agreement with experimental values. The location and the magnitude of the peak heat transfer are predicted in very close agreement with the experiments.
Time-Dependent Deposition Characteristics of Fine Coal Flyash in a Laboratory Gas Turbine Environment
Authors: Laycock Robert G. , Fletcher Thomas H.
Published: May 2012

Time-dependent deposition characteristics of fine coal flyash were measured in the Turbine Accelerated Deposition Facility (TADF) at Brigham Young University. Two samples of subbituminous coal fly ash, with mass mean diameters of 3 and 13 μm, were entrained in a hot gas flow with a gas temperature of 1250°C and Mach number of 0.25. A nickel base super alloy metal coupon approximately 0.3 cm thick was held in a hot particle-laden gas stream to simulate deposition in a gas turbine. Tests were conducted with deposition times of 20, 40, and 60 minutes. Capture efficiencies and surface roughness characteristics (e.g., Ra) were obtained at different times. Capture efficiency increased exponentially with time while Ra increased linearly with time. The increased deposition with time caused the surface temperature of the deposit to increase. The increased surface temperature caused more softening, increasing the propensity for impacting particles to stick to the surface. These data are important for improving models of deposition in turbines from syngas flows.
Optimization of a 3-Stage Booster: Part 1—The Axisymmetric Multi-Disciplinary Optimization Approach to Compressor Design
Authors: Park Kevin , Turner Mark G. , Siddappaji Kiran , Dey Soumitr , Merchant Ali
Published: May 2012

In this first part of a two part paper, an axisymmetric multi-disciplinary optimization approach for compressors is presented and applied to the design of a three stage booster. The booster has been chosen because its optimization gets little attention in the literature, it has low rotational speed and high curvature, and is also a component with only a few stages to test the capabilities of the approach. Optimization of compressors using a meanline approach have been done in the past, but a mean-line code cannot easily deal with complex curvature effects that are accentuated in a booster. An axisymmetric flow solver with a coupled boundary layer and compressor loss models is used for the aerodynamics, and an axisymmetric disk analysis code is used to generate weight-optimum disks for every rotor. The process is driven by the DAKOTA optimization package available from Sandia Labs. A genetic optimizer is used to create the Pareto front for a multi-objective function that includes efficiency, weight, length and number of airfoils. Following the genetic algorithm, a gradient based algorithm is also used. The design space is specified using physical parameters that completely define the multistage compressor. A booster made of titanium is presented in addition to two design studies. One design study explores using carbon-carbon composites and another design study explores restricting the last stage stator to 10 blades to understand if an integrated strut concept is feasible. Several optimum results along the Pareto front are described, and they are not intuitive. The optimizer has found solutions that have very high reactions in the last stage. The near-wall streamlines at the edge of the boundary layer are used as the resulting flowpath for the design. The benefit of the high stage reaction is to keep the rotor at a high tip radius, and have high turning in the following stator with very low diffusion as it matches to a lower radius high pressure compressor. The optimization process is fast enough to replace a meanline approach and explores a large design space to create a novel design.
Hydroprocessed Renewable Jet Fuel Evaluation, Performance, and Emissions in a T-63 Turbine Engine
Authors: Klingshirn Christopher D. , DeWitt Matthew J. , Striebich Rich , Anneken David , Shafer Linda , Corporan Edwin , Wagner Matt , Brigalli Dean
Published: May 2012

Due to potential beneficial environmental impacts and increased supply availability, alternative fuels derived from renewable resources are evolving on the forefront as unconventional substitutes for fossil fuel. Focus is being given to the evaluation and certification of Hydroprocessed Renewable Jet (HRJ), a fuel produced from animal fat and/or plant oils (triglycerides) by hydroprocessing, as the next potential synthetic aviation fuel. Extensive efforts have recently been performed at the Air Force Research Laboratory (AFRL) at Wright Patterson Air Force Base (WPAFB) to evaluate the potential of two HRJ fuels produced from camelina and tallow feedstocks. These have included characterization of the fuel chemical and physical fuel characteristics, and Fit-for-Purpose properties (FFP). The present effort describes general combustion performance and emission propensity of a T63-A-700 Allison turbine engine operated on the HRJs and 50/50 (by volume) HRJ/JP-8 fuel blends relative to a specification JP-8. In addition, engine and emission testing with a blend of the tallow-derived HRJ and 16% bio-derived aromatic components was completed. Fundamental engine performance characterization allows for determination of the suitability of potential synthetic fuels while quantitation of gaseous and particulate matter emissions provides an assessment of the potential environmental impact compared to current petroleum-derived fuels. In addition, an extended 150 hour endurance test was performed using a 50/50 blend of tallow-derived HRJ with JP-8 to evaluate the long-term operation of the engine with the synthetic fuel blend. This paper discusses the laboratory testing performed to characterize HRJs and results from the basic engine operability and emissions studies of the alternative fuel blends.
Computation of Aerodynamic Damping for Flutter Analysis of a Transonic Fan
Authors: Vasanthakumar Parthasarathy
Published: May 2012

This paper describes the computational analysis of aerodynamic damping for prediction of flutter characteristics of a transonic fan stage that consists of a highly loaded rotor along with a tandem stator. Three dimensional, linearized Navier-Stokes flow solver TRACE is used to numerically analyse the flutter stability of the fan. The linear flow solver enables the modeling of a single blade passage to simulate the desired inter-blade phase angle. The unsteady aerodynamic load on a vibrating blade is obtained by solving the unsteady Navier-Stokes equations on a dynamically deforming grid and the energy exchange method is used to calculate the aerodynamic damping. The calculation of aerodynamic damping for the prediction of flutter characteristics of the fan rotor is carried out with and without considering the influence of the disk. The blade mode shapes from finite element modal analysis are obtained accordingly and the flutter calculations are carried out for three blade vibration modes at the design speed and at part speeds for all possible inter-blade phase angles. Two operating points, one on the working line and the other near stall are investigated at every rotational speed. Different aspects that affect the aerodynamic damping behaviour like part speed operation, variation in unsteady blade surface pressure fluctuation between operating points on the working line and at near stall and the corresponding variation in aerodynamic work, inter-blade phase angle etc., are described. This analysis primarily focuses on the variations in aerodynamic damping of the fan with and without the influence of the disk. In addition, influence and effect of shock wave on the aerodynamic damping is also discussed.
A One-Dimensional Investigation of the Effect of an Active Control Turbocharger on Internal Combustion Engine Performance
Authors: Pesiridis Apostolos , Dubois Benjamin , Martinez-Botas Ricardo F.
Published: May 2012

The present paper discusses the impact of a new type of turbocharger, namely, the Active Control Turbocharger (ACT). The aim of this work was to prove the advantage of this type of turbocharger over the current state-of-the-art: the Variable Geometry Turbocharger (VGT). This was achieved by carrying out a comparison between two commercial Diesel engine models (through the use of a commercial engine simulation software), which belong to the same family: one 10 litre engine equipped with VGT (originally) was consecutively compared to the same model of engine modified for ACT operation and through the integration of the ACT into the 81 version of the same engine in order to demonstrate the ACT’s downsizing capability. The study has been carried out for speeds between 800 and 2000 rpm, and a fuel-air ratio range of between 0.017 and 0.057. The results showed that the actuation of the turbine in ACT mode (through the sinusoidal regulation of the turbine inlet area with each incoming exhaust gas pressure pulse) increases greatly the energy available at the turbine inlet. This leads to an increase of the boost pressure at the intake of the engine by an average 30%. The specific fuel consumption was found to be similar throughout engine operating range with a penalty of up to 10% for the ACT engine of the same size (10 litre). A comparison was then carried out between the 10 litre VGT engine and the 8 litre ACT engine. The 8 litre has been found to produce up to 37% more torque and horse power under 1400 rpm and obtained very similar performance to the 10 litre VGT engine at higher speeds. At constant power output between the 8 and 10 litre engines, it has been found that the fuel consumption was decreased by a maximum of 9% when using the 8 litre engine. The results of the present study were encouraging with respect to the potential of ACT to downsize the internal combustion engine.
The Effects of Conjugate Heat Transfer on the Thermal Field Above a Film Cooled Wall
Authors: Dees Jason E. , Bogard David G. , Ledezma Gustavo A. , Laskowski Gregory M.
Published: May 2012

Common gas turbine heat transfer analysis methods rely on the assumption that the driving temperature for heat transfer to a film cooled wall can be approximated by the adiabatic wall temperature. This assumption implies that the gas temperature above a film cooled adiabatic wall is representative of the overlying gas temperature on a film cooled conducting wall. This assumption has never been evaluated experimentally. In order for the adiabatic wall temperature as driving temperature for heat transfer assumption to be valid, the developing thermal boundary layer that exists above a conducting wall must not significantly affect the overriding gas temperature. In this paper, thermal fields above conducting and adiabatic walls of identical geometry and at the same experimental conditions were measured. These measurements allow for a direct comparison of the thermal fields above each wall in order to determine the validity of the adiabatic wall temperature as driving temperature for heat transfer assumption. In cases where the film cooling jet was detached, a very clear effect of the developing thermal boundary layer on the gas temperature above the wall was measured. In this case, the temperatures above the wall were clearly not well represented by the adiabatic wall temperature. For cases where the film cooling jet remained attached, differences in the thermal fields above the adiabatic and conducting wall were small, indicating a very thin thermal boundary layer existed beneath the coolant jet.
Mu Analysis for Turbomachinery Stall Flutter
Authors: Forhad Md. M. I. , Vishwakarma Puneet , Xu Yunjun
Published: May 2012

Flutter is an aeroelastic instability phenomenon that can result either in serious damage or complete destruction of a gas turbine blade structure. To assure a reliable and safe operation, potential for blade flutter must be eliminated from the turbo-machinery stages. In this paper, the robustness of an axial compressor blade design is studied with respect to parametric uncertainties through the Mu analysis. The analytical description of the nominal model used is based on matching a two dimensional incompressible flow field across the flexible rotor and the rigid stator. The aerodynamic load on the blade is derived via the control volume analysis. For use in the Mu analysis, first the model originally described by a set of partial differential equations is reduced to ordinary differential equations by the Fourier series based collocation method. After that, the nominal model is obtained by linearizing the achieved non-linear ordinary differential equations. The uncertainties coming from the modeling assumptions, model reduction, and linearization approximations, as well as imperfectly known parameters and coefficients are all modeled as parametric uncertainties through the Monte Carlo simulation. As compared with other robustness analysis tools, such as Hinf, the Mu analysis is less conservative and can handle both structured and unstructured perturbations. Simulation results show that the procedure described in this paper can be effective in studying the flutter stability margin and can be used to guide the gas turbine blade design.
Measurement of Volatile Particulate Matter Emissions From Aircraft Engines Using a Simulated Plume Aging System
Authors: Peck Jay , Timko Michael T. , Yu Zhenhong , Wong Hsi-Wu , Herndon Scott C. , Yelvington Paul E. , Miake-Lye Richard C. , Wey Changlie , Winstead Edward L. , Ziemba Luke D. , Anderson Bruce E.
Published: May 2012

Aircraft exhaust contains nonvolatile (soot) particulate matter (PM), trace gas pollutants, and volatile PM precursor material. Nonvolatile soot particles are predominantly present at the engine exit plane, but volatile PM precursors form new particles or add mass to the existing ones as the exhaust is diluted and cooled. Accurately characterizing the volatile PM mass, number, and size distribution is challenging due to this evolving nature and the impact of local ambient conditions on the gas-to-particle conversion processes. To accurately and consistently measure the aircraft PM emissions, a dilution and aging sampling system that can condense volatile precursors to particle phase to simulate atmospheric evolution of aircraft engine exhaust has been developed. In this paper, field demonstration of its operation is described. The dilution/aging probe system was tested using both a combustor rig and on-wing CFM56-7 engines. During the combustor rig testing at NASA Glenn Research Center, the dilution/aging probe supported formation of both nucleation/growth mode particles and soot coatings. The results showed that by increasing residence time, the nucleation particles become larger in size, increase in total mass, and decrease in number. During the on-wing CFM56-7 engine testing at Chicago Midway Airport, the dilution/aging probe was able to form soot coatings as well as nucleation mode particles, unlike conventional 1-m probe engine measurements. The number concentration of nucleation particles depended on sample fraction and relative humidity of the dilution air. The performance of the instrument is analyzed and explained using computational microphysics simulations.
Investigation of Transient CFD Methods Applied to a Transonic Compressor Stage
Authors: Blumenthal Rob , Hutchinson Brad , Zori Laith
Published: May 2012

Understanding unsteady flow phenomena in compressor stages often requires the use of time-accurate CFD simulations. Due to the inherent differences in blade pitch between adjacent blade rows, the flow conditions at any given instant in adjacent blade rows differ. Simplified computation of the stage represented by a single blade in each row and simple periodic boundary conditions is therefore not possible. Depending on the blade counts, it may be necessary to model the entire annulus of the stage; however, this requires considerable computational time and memory resources. Several methods for modeling the transient flow in turbomachinery stages which require a minimal number of blade passages per row, and therefore reduced computational demands, have been presented in the literature. Recently, some of these methods have become available in commercial CFD solvers. This paper provides a brief description of the methods used, and how they are applied to a transonic compressor stage. The methods are evaluated and compared in terms of computational efficiency and storage requirements, and comparison is made to steady stage simulations. Comparisons to overall performance data and two-dimensional LDV measurements are used to assess the predictive capabilities of the methods. Computed flow features are examined, and compared with reported measurements.
Can-Annular Combustion Chamber Surface Temperature Measurements and Damage Signatures at Operationally Representative Conditions
Authors: Bouchard D. , Pucher G. , Allan W. D. E.
Published: May 2012

An experimental program which investigated the surface temperature distribution of a contemporary gas turbine combustion liner is presented. An array of 65 embedded surface mounted thermocouples was installed on a Rolls Royce/Allison T56 combustion liner and exposed to combustion conditions in the Combustion Chamber Sector Rig (CCSR) at the Royal Military College of Canada. The CCSR was operated at two test points to simulate idle and cruise modes of operation. Corresponding exhaust temperature measurements were taken in the test combustion chamber exhaust plane with a sweeping thermocouple rake. These efforts were the latest in a multi-year program to investigate the impact of service wear related geometric deformations of combustion liners and damaged/fouled fuel nozzles on the exit temperature profile from typical combustion chambers. It has been previously ascertained that real-world geometric anomalies in the T56 combustion chambers, particularly in the transitional zone, can modify the exhaust temperature profile to a sufficient degree so as to risk hot section damage due to excessive heat exposure. The collection and analysis of surface temperature data represents a useful extension of the knowledge base of the T56 combustion system within the context of the overall program and is paramount to upcoming numerical modelling efforts aimed at assessing hot section damage risks.
Low-Frequency Thermoacoustic Instability Mitigation Using Adaptive-Passive Acoustic Radiators
Authors: Kashani Reza , Monfort Jeff
Published: May 2012

A commonly used technique for mitigating thermoacoustic instability in an enclosed combustion environment is removing more acoustic energy from the combustor, at the frequency corresponding to the acoustic mode(s) of the combustor which are sympathetic to such instability. This approach is based on adding tuned acoustic damping to the combustion environment. By incorporating in-situ adjustability into acoustic damping devices, they can change their mechanical attributes, e.g., mass and/or stiffness, and adapt themselves in a semi-active manner to the varying instability frequency. Adaptive-passive thermoacoustic mitigation solutions have less weight penalty than the alternative active solutions mainly because the adaptation is done in a semi-active way, at slow pace, with a small and less power-hungry actuation mechanisms. Moreover, the flexibility they offer make them highly desirable for land and marine instability mitigation applications. In this work, semi-active adjustment of a novel tuned acoustic damper, namely an acoustic radiator, is explored. The paper describes the inner working of a semi-active (adaptive-passive) acoustic radiator and the relevant control schemes to adapt them to the instability frequency on hand. The damping effectiveness of the proposed damper, is demonstrated experimentally. It should be mentioned that the semi-active control strategies developed for acoustic radiators can also be used, with minor modifications, for semi-active control of other acoustic damping mechanisms such as Helmholtz resonators and quarter-wave tubes.
Swirl-Enhanced Internal Cooling of Turbine Airfoils: Part II—90 Degree Flow Entry
Authors: Segura Del , Acharya Sumanta
Published: May 2012

Heat transfer measurements and analysis have been performed on a uniquely designed multi-channel passage consisting of a slot shaped channel with a 3:1 aspect ratio with coolant-feed tubes located adjacent to the main slot shaped channel. Small round jets connect the outer feed passages to the main channels at a 15 degree angle relative to the main channel flow direction and at a position tangent to the floor/roof of the main channel. Flow entering the multi-channel passages is directed into the main channel through orifices that reduce the pressure in the main channel, thereby enabling positive pressure differences between the feed and the main channel and allowing high velocity flow through the jets. The flow enters the main channel via a 90-degree turn through the orifice. The resulting flow through the side jets and main channel causes high shear flow along the roof and floor of the channel where the jet flow enters the main channel, swirl motion as the high velocity side jet flow enters the main channel flow at an angle relative to the main flow direction, and high turbulence regions as the lower velocity main channel flow tumbles when coming in contact with the high velocity jets issuing from the side channels. The heat transfer characteristics were compared to the slot channel with a 90 degree inlet with no additional heat transfer enhancements. Four different jet configurations are presented along with three different orifice diameters. While a single channel passage with flow exiting freely is not a design typically found in a turbine airfoil, the benefits of this unique concept can be a basis for further studies with geometries more typical of a production airfoil. The results yield average normalized Nusselt numbers enhancement for the entire main channel as high as 10.7, when compared to a smooth slot channel without heat transfer enhancements. Pressure losses, mainly due to the orifices, were high but the overall performance shows significant improvements when compared to other heat transfer enhancement methods in turbine airfoil mid-span regions.
Simultaneous Noise and Performance Measurements for High Speed Jet Noise Reduction Technologies: Part II—Near Field Array Calibration
Authors: Long Dean , Martens Steve
Published: May 2012

Part I of this paper describes a methodology for assessing the far field jet noise from high speed exhaust nozzles using a microphone array in the near field of the exhaust plume. The near field noise measurement is mathematically propagated producing an estimate of the noise level at the new location. Outward propagation produces an estimate of the far field noise. Propagation toward the jet axis produces the source distribution. Part II described here provides a direct validation of this process using a generic CD nozzle in a facility where both the near field and the far field are measured simultaneously. Comparison of these data sets show good agreement over the typical operating range for this type of nozzle. The far field noise is characterized by two independent processes: Shock cell noise radiating in the forward quadrant is produced when the nozzle is operated at non-ideally expanded conditions. Mach wave radiation propagates into the aft quadrant when the exhaust temperature is elevated. Subsequent tests in an acoustically treated nozzle thrust stand demonstrate the value of the near field array allowing immediate feedback on the noise/performance tradeoff for high speed jet noise reduction technologies.
Quantifying Optical Tip-Timing Probe Error With Laboratory Apparatus
Authors: Hockaday Bruce D.
Published: May 2012

Detection of airfoil time of arrival with optical probes has been evolving since the 1980s. Time of arrival data are used to infer airfoil stresses caused by vibration through a sequence of manipulations. The data conversion begins by converting arrival time to blade position, so blade deflection can be determined from the expected non-vibrating position. Various methods are used in the industry to convert deflection data to frequency, amplitude, and stress, which is beyond the scope of this paper. Regardless of the analytical approach used, producing accurate stress information relies on the precise detection and measurement of time of arrival, which equates to blade position. Recent improvements have been made in time of arrival system accuracy by running faster clocks to increase temporal resolution of the measurement. Greater timing resolution, afforded by clock speed, will have diminishing returns when probe and blade-tip interactions begin producing dominant errors. In the case of optical probes, the blade-tip needs to be treated as a curved reflector in the optical system that is capable of introducing dynamic errors. In engine operation the blade-tip moves axially under the probe from untwist, static deflection, and vibration, causing the light to reflect from different parts of the blade-tip. This relative movement between the probe and blade-tip cause the arrival time to change dynamically. Neglecting the dynamic arrival errors caused by the blade-tip’s optical properties will result in blade deflection-errors that propagate into the stress information. This paper presents a laboratory study that quantifies time of arrival errors due to optical interaction with tip radii. The study reports measured arrival position error as a function of location and optical signal power levels. The work is presented in terms of arrival position, producing information that is independent of rotational speed, and vibratory mode.
Probabilistic High-Cycle Fretting Fatigue Assessment of Gas Turbine Engine Components
Authors: Chan Kwai S. , Enright Michael P. , Golden Patrick J. , Naboulsi Samir , Chandra Ramesh , Pentz Alan C.
Published: May 2012

High-cycle fatigue (HCF) is arguably one of the costliest sources of in-service damage in military aircraft engines. HCF of turbine blades and disks can pose a significant engine risk because fatigue failure can result from resonant vibratory stresses sustained over a relatively short time. A common approach to mitigate HCF risk is to avoid dangerous resonant vibration modes (first bending and torsion modes, etc.) and instabilities (flutter and rotating stall) in the operating range. However, it might be impossible to avoid all the resonance for all flight conditions. In this paper, a methodology is presented to assess the influences of HCF loading on the fracture risk of gas turbine engine components subjected to fretting fatigue. The methodology is based on an integration of a global finite element analysis of the disk-blade assembly, numerical solution of the singular integral equations using the CAPRI (Contact Analysis for Profiles of Random Indenters) and Worst Case Fret methods, and risk assessment using the DARWIN (Design Assessment of Reliability with Inspection) probabilistic fracture mechanics code. The methodology is illustrated for an actual military engine disk under real life loading conditions.
Robust Lean Blowout Sensing for Turbine Engine Combustors
Authors: Bompelly Ravi K. , Seitzman Jerry M.
Published: May 2012

Lean blowout (LBO) proximity detection is demonstrated in a Jet-A fueled, single-element Lean Direct Injection (LDI) combustor operating at conditions intended to simulate supersonic cruise (2–4 atm, ∼700 K inlet air). Detection is based on identifying LBO precursors in the optical radiation produced by the combustion process. The precursors result from short duration partial flame extinction and re-ignition events. Thresholding of the low-pass filtered optical signal provides robust precursor identification in the presence of the moderate levels of combustion dynamics exhibited by the combustor. Acoustic-based precursor sensing was also attempted, but the precursors could not be reliably identified in the presence of the combustion noise and dynamics; likely causes are discussed. The average event occurrence rate in the optical signal provides a reliable measure of the combustor’s proximity to LBO, monotonically increasing as the LBO limit is approached. The precursor events occur intermittently, and the occurrence rate statistics are reasonably modeled with a normal probability distribution. A probabilistic approach is developed to estimate the fastest fuel reduction transients that could be controlled based on the need to detect some events before the blowout conditions are reached. The results from the LDI combustor are also compared to tests in a premixed, swirl- and dump-stabilized gas-fueled combustor with a center body. High-speed flame images of the precursor events indicate the existence of different static stability modes in the premixed combustor, which are absent in the LDI combustor. The absence of mode-switching in the LDI combustor is suggested to be the reason why it produces much shorter but more frequent precursor events compared to the premixed combustor.
An Improved Core Reaction Mechanism for Saturated C0–C4 Fuels
Authors: Naik Chitralkumar V. , Puduppakkam Karthik V. , Meeks Ellen
Published: May 2012

Accurate chemistry models are required to predict the combustion behavior of different fuels, such as synthetic gaseous fuels and liquid jet fuels. A detailed reaction mechanism contains chemistry for all the molecular components in the fuel or its surrogates. Validation studies that compare model predictions with the data from fundamental combustion experiments under well defined conditions. Such fundamental experiments are least affected by the effect of transport on chemistry. Therefore they are the most reliable means for determining a reaction mechanism’s predictive capabilities. Following extensive validation studies and analysis of detailed reaction mechanisms for a wide range of hydrocarbon components reported in our previously published work [1–5], we identified some common issues in the predictive nature of the mechanisms that are associated with inadequacies of the core (C0–C4) mechanism. For example predictions of laminar flame speeds and autoignition delay times for several fuels were inaccurate beyond the level of uncertainty in the data. This core mechanism is shared by all of the mechanisms for the larger hydrocarbon components. Unlike the reaction paths for larger hydrocarbon fuels, however, reaction paths for the core chemistry do not follow prescribed reaction rate-rules. In this work, we revisit our core reaction mechanism for saturated C0–C4 fuels, with the goal of improving predictions for the widest range of fundamental experiments as possible. To evaluate and validate the mechanism improvements, we performed a broad set of simulations of fundamental experiments. These experiments include measurements of ignition delay, flame speed and extinction strain rate, as well as species composition in stirred reactors, flames and flow reactors. The range of conditions covers low to high temperatures, very lean to very rich fuel-air ratios, and low to high pressures. Our core reaction mechanism contains thermochemical parameters derived from a wide variety of sources, including experimental measurements, ab initio calculations, estimation methods and systematic optimization studies. Each technique has its uncertainties and potential inaccuracies. Using a systematic approach that includes sensitivity analysis, reaction-path analysis, consideration of recent literature studies, and an attention to data consistency, we have identified key updates required for the core mechanism. These updates resulted in accurate predictions for various saturated fuels when compared to the data over a broad range of conditions. All reaction rate constants and species thermodynamics and transport parameters remain within known uncertainties and within physically reasonable bounds. Unlike most mechanisms in the literature, the mechanism developed in this work is self-consistent and contains chemistry of all saturated C0–C4 fuels.
Experimental Simulation of Contaminant Deposition on a Film Cooled Turbine Vane Pressure Side With a Trench
Authors: Albert Jason E. , Bogard David G.
Published: May 2012

An important issue in the use of coal- or biomass-derived synthetic gaseous (syngas) fuels is the deposition of contaminants on film cooled turbine surfaces, which alter cooling and aerodynamic performance and increase material degradation. The current study applied a new experimental technique that simulated the key physical aspects of contaminant deposition on a film cooled turbine vane. The depositing contaminants were modeled in a wind tunnel facility with a spray of molten wax droplets of a size range that matched the Stokes number of the contaminant particles in engine conditions. Most experiments were performed using a vane model with a thermal conductivity selected such that the model had the same Biot number of an actual engine airfoil, resulting in a cooler surface temperature. Some experiments were performed using an approximately adiabatic model for comparison. The film cooling design consisted of three rows of showerhead cooling at the leading edge and one row of body film cooling holes on the pressure side. Two designs of pressure side body film cooling holes were considered: a standard design of straight, cylindrical holes and an advanced design of “trenched” cooling holes in which the hole exits were situated in a recessed, transverse trench. The results showed thin deposits formed in the trench, with the thickest deposits on its downstream wall between coolant jets. Adiabatic film effectiveness levels were essentially unchanged by the presence of deposits for either film configuration. Deposit formation was strongly influenced by the model surface temperature with cooler surfaces inhibiting deposition. There was evidence of a threshold surface temperature above which deposits became significantly thicker.
Characterization of Pre- and Post-Service Grain Boundary Phases in a Cast Austenitic Steel
Authors: Unocic Raymond R. , Unocic Kinga A. , Pint Bruce A. , Lipschutz Mark D.
Published: May 2012

Austenitic steel castings are currently being used in components for industrial gas turbine engines. Service experience has indicated a degradation of mechanical properties with extended exposures at elevated temperature. The purpose of this study was to characterize the grain boundary phases that develop during the casting processes as a likely explanation for the observed performance. In order to isolate these precipitates, a variety of electron microscopy characterization techniques were used to characterize their composition after various heat treatments and service exposure. In the baseline, as-cast and annealed condition, a discontinuous network of grain boundary metal carbides was observed. These precipitates coarsened during short-term annealing at 649°C and a denuded zone formed in the adjacent matrix. When the 38,600 h service-exposed material was analyzed, the grain boundaries were highly decorated with a more continuous film of grain boundary carbides as well as voids attributed to creep cavitation. In addition to carbides, acicular AlN precipitates were identified on the grain boundaries of the casting examined after service exposure.
Turbine Hub and Shroud Sealing Flow Loss Mechanisms
Authors: Zlatinov Metodi Blagoev , Tan Choon Sooi , Montgomery Matthew , Islam Tito , Seco-Soley Melissa
Published: May 2012

Purge air is injected through seals in the hub and shroud of axial turbines in order to prevent hot gas ingestion into the inter-stage gaps. An investigation into the losses involved with the injection of purge air has been undertaken, with the objectives of answering where the losses are generated, how they are generated, and what are the most effective ways for reducing them. In order to address these questions, a consistent framework for interpreting entropy generation as a measure of loss is developed for turbomachinery applications with secondary air streams. A procedure for factoring out distinct effects is also presented. These tools, applied to steady computations, elucidate four routes though which change in loss generation is brought about by injection of purge air: a shear layer between purge and main streams, interaction with the passage vortex system that generates radial velocity gradients, changes in wetted loss and tip clearance flow due to an increased degree of reaction, and the potential for reducing tip clearance flow for the case of purge flow injected from the shroud. An emphasis is placed on tracing these effects to specific purge flow characteristics that drive them. The understanding gained provides a rationale for the observed sensitivity of purge flow losses to the design parameters purge air mass fraction and swirl, compared to purge slot axial inclination and gap width. Pre-swirling of purge flow is less effective in mitigating losses in the case of shroud-injection, since there is a tradeoff with the tip clearance flow suppression effect.
3D RANS Prediction of Gas-Side Heat Transfer Coefficients on Turbine Blade and Endwall
Authors: Luo Jiang , Razinsky Eli H. , Moon Hee-Koo
Published: May 2012

This paper presents a study using 3D computational fluid dynamics (CFD) based on Reynolds-averaged Navier-Stokes (RANS) equations to predict turbine gas-side heat transfer coefficients (HTC) on the entire airfoil and endwall. The CFD results at different spanwise sections and endwall have been compared with the flat-plate turbulent boundary layer correlation and with the data in a NASA turbine rotor passage with strong secondary flows, under three different flow conditions. The enhancement effects of secondary flow vortices on the blade surface and endwall heat transfer rate have been examined in detail. Analyses were conducted for the impact of Reynolds number and exit Mach number on heat transfer. The SST, k-ε, V2F, and realizable k-ε turbulence models have been assessed. The classical log-law wall-functions have been found to be comparable to the wall-integration methods, but with much reduced sensitivity to inlet turbulence conditions. The migration of hot gas was simulated with a radial profile of inlet temperature. CFD results for mid-span HTCs of two other airfoils were also compared with test data. Overall results are encouraging and indicate improved HTC and temperature predictions from 3D CFD could help optimize the design of turbine cooling schemes.
Aerodynamic Improvements of Wind-Turbine Airfoil Geometries With the Prescribed Surface Curvature Distribution Blade Design (CIRCLE) Method
Authors: Korakianitis T. , Rezaienia M. A. , Hamakhan I. A. , Avital E. , Williams J. J. R.
Published: May 2012

The prescribed surface curvature distribution blade design (CIRCLE) method can be used for the design of two-dimensional (2D) and three-dimensional (3D) turbomachinery blade rows with continuous curvature and slope of curvature from leading edge (LE) stagnation point to trailing edge (TE) stagnation point and back to the LE stagnation point. This feature results in smooth surface pressure distribution airfoils with inherently good aerodynamic performance. In this paper the CIRCLE blade design method is modified for the design of 2D isolated airfoils. As an illustration of the capabilities of the method, it is applied to the redesign of two representative airfoils used in wind turbine blades: the Eppler 387 airfoil; and the NREL S814 airfoil. Computational fluid dynamic analysis is used to investigate the design point and off-design performance of the original and modified airfoils, and compare with experiments on the original ones. The computed aerodynamic advantages of the modified airfoils are discussed. The surface pressure distributions, drag coefficients, and lift-to-drag coefficients of the original and redesigned airfoils are examined. It is concluded that the method can be used for the design of wind turbine blade geometries of superior aerodynamic performance.
Comparison of the Steady-State and Dynamic Performance of Two Fixed-Geometry Journal Bearings
Authors: Conlon Martin J. , Dadouche Azzedine , Dmochowski Waldemar M.
Published: May 2012

This paper evaluates two different journal bearings: a cylindrical bore plain journal bearing and a tri-lobe taper land bearing. Each bearing has the same nominal diameter (89mm) and aspect ratio (L/D = 0.7). The shaft rotational speed ranged from 6krpm to 14krpm and the bearing specific load from 700kPa to 2800kPa. The bearings’ steady-state performance is evaluated according to relative bearing and shaft displacement, bearing operating temperature and power loss. A frequency-domain analysis is used to determine bearing rotordynamic coefficients — it treats the bearing as a mass-spring-damper system. Excitation frequencies range between 20Hz and 350Hz. Ultimately, the tri-lobe taper land bearing offers better stability whereas the cylindrical bore plain journal bearing has a lower eccentricity and runs cooler for a given operating condition. The dynamic properties of the two bearings are found to be similar, although the higher stability of the tri-lobe taper land bearing is also reflected in the dynamic coefficients.
Strength Degradation of Oxide/Oxide and SiC/SiC Ceramic Matrix Composites in CMAS and CMAS/Salt Exposures
Authors: Faucett David C. , Choi Sung R.
Published: May 2012

CMAS (Calcium-Magnesium-Aluminosilicate) has shown to induce some deleterious effects on yittria-stabilized-zirconia (YSZ) based thermal barrier coatings (TBCs) of hot section components of aeroengines. The effects were shown to be dependent on the types and operating conditions of engines/components. The work presented here explored how CMAS would affect ceramic matrix composites (CMCs) in terms of strength degradation. Four different, gas-turbine grade CMCs were utilized including two types of MI SiC/SiCs and other two types of oxides/oxides (N720/aluminisilicate and N720/alumina). Test specimens in a simple flexure configuration were CMAS-treated at 1200 °C in air under either isothermal or thermal cycling condition. The effects of CMAS were quantified via residual strengths of treated test specimens. Strength degradation with respect to as-received strengths ranged from 10 to 20% depending on the types of CMCs. It was further observed that significant degradation of strength up to 90% occurred in an oxide/oxide CMC when sodium sulfate was added to CMAS.
Engine Design Strategies to Maximize Ceramic Turbine Life and Reliability
Authors: Vick Michael J. , Jadaan Osama M. , Wereszczak Andrew A. , Choi Sung R. , Heyes Andrew L. , Pullen Keith R.
Published: May 2012

Ceramic turbines have long promised to enable higher fuel efficiencies by accommodating higher temperatures without cooling, yet no engines with ceramic rotors are in production today. Studies cite life, reliability, and cost obstacles, often concluding that further improvements in the materials are required. In this paper, we assume instead that the problems could be circumvented by adjusting the engine design. Detailed analyses are conducted for two key life-limiting processes, water vapor erosion and slow crack growth, seeking engine design strategies for mitigating their effects. We show that highly recuperated engines generate extremely low levels of water vapor erosion, enabling lives exceeding 10,000 hours, without environmental barrier coatings. Recuperated engines are highly efficient at low pressure ratios, making low blade speeds practical. Many ceramic demonstration engines have had design point mean blade speeds near 550 m/s. A CARES/Life analysis of an example rotor designed for about half this value indicates vast improvements in SCG-limited life and reliability. Lower blade speeds also reduce foreign object damage (FOD) particle kinetic energy by a factor of four. In applications requiring very high fuel efficiency that can accept a recuperator, or in short-life simple cycle engines, ceramic turbines are ready for application today.
Effect of Surface Texturing on the Steady-State Properties and Dynamic Coefficients of a Plain Journal Bearing: Experimental Study
Authors: Dadouche Azzedine , Conlon Martin J. , Dmochowski Waldemar , Koszela Waldemar , Galda Lidia , Pawlus Pawel
Published: May 2012

The last decade has seen a significant and increased interest in surface texturing technologies esulting in improving the overall tribological properties of mechanical components. Several techniques could be used to modify the surface topography with laser surface texturing becoming most popular recently. In this investigation, rig experiments have been performed on plain and textured-surface journal bearings with an aspect ratio of 1.1 under a variety of loads and speeds. Percussive burnishing (embossing) was used to create the dimples on the internal surfaces of the test bearings. The dimples have a spherical shape with a diameter of 1 mm and a depth of 60 micrometers. Pit-area ratio was in the range 5–20% of the total bearings surfaces. The effects of surface-texturing and dimple density on the bearing steady-state characteristics and dynamic coefficients are analyzed and discussed in this paper.