bluuebunny/crossref_metadata_2025 huggingface dataset:
Effect of Variations in Shroud Geometry on Single Phase Flow Over a Shrouded Single Spiral Gear
Authors: Rapley Steve , Eastwick Carol , Simmons Kathy
Published: August 2009

In some aero-engine applications a spiral bevel gear is mounted in a bearing chamber. A gear can be separated from the surrounding chamber by the use of a shroud (a close fitting cover) to reduce the parasitic windage power loss due to the drag on the gear caused by the surrounding fluid. In this paper a single-phase computational study of the effect of variations in shroud geometry on the single phase flow over a shrouded single spiral bevel gear is presented, with a baseline case compared to available experimental data. The aim of the work is to identify good working practice in designing shrouds. The work is a continuation of work reported in ASME Turbo 2007 [1]. In the parametric study three variables have been identified as controlling the shroud behaviour and these are the clearances at shroud inlet, outlet and between shroud and teeth. Clearances at these three locations have been varied between less than 1mm (smaller than would typically be found in an aeroengine) and 4mm (wider than generally found). Varying the shroud geometry is shown to affect the flow structure, causing transience and different flow paths/recirculations within the flow. This affects the windage power loss experienced by the gear. The effects of different shroud geometries on the flow is analysed and presented within the paper. For a given rotational speed, changing the shroud geometry is seen to cause a variation in torque levels of up to 50% (comparing best and worst cases). The possibility of directional bias in the flow structure and torque levels is investigated. Static pressure profiles on the shroud for the baseline case are compared with experimental data and reasonable agreement is shown.
Efficient FEA/CFD Thermal Coupling for Engineering Applications
Authors: Sun Zixiang , Chew John W. , Hills Nicholas J. , Volkov Konstantin N. , Barnes Christopher J.
Published: August 2009

An efficient finite element analysis/computational fluid dynamics (FEA/CFD) thermal coupling technique has been developed and demonstrated. The thermal coupling is achieved by an iterative procedure between FEA and CFD calculations. Communication between FEA and CFD calculations ensures continuity of temperature and heat flux. In the procedure, the FEA simulation is treated as unsteady for a given transient cycle. To speed up the thermal coupling, steady CFD calculations are employed, considering that fluid flow time-scales are much shorter than those for the solid heat conduction and therefore the influence of unsteadiness in fluid regions is negligible. To facilitate the thermal coupling, the procedure is designed to allow a set of CFD models to be defined at key time points/intervals in the transient cycle and to be invoked during the coupling process at specified time points. To further enhance computational efficiency, a “frozen flow” or “energy equation only” coupling option was also developed, where only the energy equation is solved while the flow is frozen in CFD simulation during the thermal coupling process for specified time intervals. This option has proven very useful in practice, as the flow is found to be unaffected by the thermal boundary conditions over certain time intervals. The FEA solver employed is an in-house code, and the coupling has been implemented for two different CFD solvers; a commercial code and an in-house code. Test cases include an industrial low pressure (LP) turbine and a high pressure (HP) compressor, with CFD modeling of the LP turbine disc cavity and the HP compressor drive cone cavity flows, respectively. Good agreement of wall temperatures with the industrial rig test data was observed. It is shown that the coupled solutions can be obtained in sufficiently short turn-around times (typically within a week) for use in design.
Nonlinear Numerical Prediction of Gas Foil Bearings Stability and Unbalanced Response
Authors: Le Lez Se´bastien , Arghir Mihai¨ , Frene Jean
Published: August 2009

One of the main interests of gas foil bearings lies in their superior rotordynamic characteristics compared to conventional bearings. A numerical investigation on the stability limit and on the unbalanced response of foil bearings is presented in this paper. The main difficulty in modeling the dynamic behavior of such bearings comes from the dry friction that occurs within the foil structure. Indeed, dry friction is highly non linear and is strongly influenced by the dynamic amplitude of the pressure field. To deal with these non linearities, a structural dynamic model has been developed in a previous work. This model considers the entire corrugated foil and the interactions between the bumps by describing the foil bearing structure as a multiple degrees of freedom system. It allows the determination of the dynamic friction forces at the top and at the bottom of the bumps by simple integration of ordinary differential equations. The dynamic displacements of the entire corrugated sheet are then easily obtained at each time step. The coupling between this structural model and a gas bearing prediction code is presented in this paper and allows performing full non linear analyses of a complete foil bearing. The bearing stability is the first investigated problem. The results show that the structural deflection enhances the stability of compliant surface bearings compared to rigid ones. Moreover, when friction is introduced, a new level of stability is reached revealing the importance of this dissipation mechanism. The second investigated problem is the unbalanced response of foil bearings. The shaft trajectories depict a non linear jump in the response of both rigid and foil bearings when the value of the unbalance increases. Again, it is evidenced that the foil bearing can support higher mass unbalance before this undesirable step occurs.
Assessment of Models for Liquid Jet Breakup
Authors: Hutcheson Paul S. , Chew John W. , Thorpe Rex B. , Young Colin
Published: August 2009

For many gas turbine architectures a failure modes and effects analysis identifies a potential mode in which failure of an oil transfer pipe could result in oil leakage into the secondary air system. Such an event would result in a complex two-phase interacting flow. The atomisation and transport of the oil within the air system is of interest, but is difficult to predict. Available data for the droplet size resulting from jet breakup in crossflow are limited. A dimensional analysis shows jet breakup in a crossflow to involve many factors. The atomisation process has been shown experimentally to include many physical processes and is still not completely understood. Currently, the most practical method of modelling these breakup processes in sprays is by using a CFD package with a set of sub-models within an Euler-Lagrangian (discrete-droplet) approach. The strengths and weaknesses of each of these sub-models cannot reasonably be tested when used in combination with other approximations to model a spray in crossflow. The purpose of this study was to assess various submodels for liquid breakup with a series of simple test cases.
Heat Transfer Measurements for a Film Cooled Turbine Vane Cascade
Authors: Thurman Douglas R. , Poinsatte Philip E. , Heidmann James D.
Published: August 2009

Experimental heat transfer and pressure measurements were obtained on a large scale film cooled turbine vane cascade. The objective was to investigate heat transfer on a commercial high pressure first stage turbine vane at near engine Mach and Reynolds number conditions. Additionally blowing ratios and coolant density were also matched. Numerical computations were made with the Glenn-HT code of the same geometry and compared with the experimental results. A transient thermochromic liquid crystal technique was used to obtain steady state heat transfer data on the midspan geometry of an instrumented vane with 12 rows of circular and shaped film cooling holes. A mixture of SF6 and Argon gases was used for film coolant to match the coolant-to-gas density ratio of a real engine. The exit Mach number and Reynolds number were 0.725 and 2.7 million respectively. Trends from the experimental heat transfer data matched well with the computational prediction, particularly for the film cooled case.
Validation and Calibration of Modern CFD RANS Codes for the Prediction of Transonic Axial-Centrifugal Compressors
Authors: Mansour Mahmoud L. , Holbrook Greg , Gunaraj John , Qizar Mohamed , Mangat Vikram
Published: August 2009

This paper summarizes the results of evaluating and validating three 3D viscous “RANS” Reynolds-Averaged Navier-Stokes codes for a Honeywell modern design of axial-centrifugal compressor test case with high quality test data. The compressor consists of a single stage axial followed by a single stage centrifugal designs. The axial stage consists of a transonic axial rotor and stator. The centrifugal stage consists of a centrifugal impeller, a vane-island radial diffuser, a 90-degrees bend and tandem deswirl vanes. The compressor was designed and rig tested by Honeywell in the early ‘00s. Detailed measurements of total pressure and total temperature were made at the compressor inlet, exit, and between the stages, including an X-Y survey at the axial stage exit, with sufficient resolution to provide a good estimate of the performance of each stage of this compressor. The validation and calibration was performed by running simulations for the entire compressor as tested in the rig. The RANS CFD programs used in this study are the NASA Glenn Research Center’s average-passage program APNASA, the Numeca’s Fine Turbo program and the ANSYS/CFX. The APNASA program employs only the standard K-ε turbulence model while the Numeca’s Fine Turbo and ANSYS/CFX employ various turbulence closures that are based on the standard two-equation K-ε turbulence model and the Wilcox K-ω model. The paper presents a comparison between the three CFD codes’ predictions and the measured test data for the overall compressor performance and the performance of the different compressor stages. These results shed some light on the ability of these codes to predict the overall compressor performance and the performance of the individual components. Sensitivity of the two codes predictability for mesh resolution and turbulence modeling variations are also presented.
Mesh Generation for Conjugate Heat Transfer Analysis of a Cooled High Pressure Turbine Stage
Authors: Goormans-Francke Colinda , Carabin Guy , Hirsch Charles
Published: August 2009

The presented work demonstrates the feasibility of quasi-automatic structured mesh generation for all details in the complex cooling system of an industrial high pressure turbine stage, as required by advanced Conjugate Heat Transfer (CHT) simulations. The grid generation software has been adapted in order to quasi-automatically mesh typical cooling configurations such as cooling passages, basins, inserts, solid bodies, cooling holes, slots, and rib turbulators. A multi-domain structured mesh with about 154 million grid points and 12,316 blocks has been generated for the turbine stage. It includes 1,000 cooling holes, over 250 rib turbulators and 150 pin fins for the turbine stage. In order to verify the CFD response to the grid properties, simulations were performed as a first step on the coarse grid level (of 21.8 million grid points) using the 3D flow solver package FINE™/Turbo. The conductivity equation was solved for the solid part of the computational domain using the same temporal discretization scheme as for the flow solver. Parallel, coupled fluid/solid calculations using the k-ε turbulence model were performed on three different configurations: nozzle guide vane alone, rotor-blade alone, and full stage. These results show the feasibility of this approach to mesh generation for use in CHT modeling of the complex configuration of cooled turbine stages.
Design Tuning of High Aspect Ratio Shrouded Turbine Blades
Authors: Arkhipov Alexander N. , Pipopulo Andrey V. , Putchkov Igor V.
Published: August 2009

A large height to chord ratio, or aspect ratio, of industrial gas turbine blades, especially for the last rows, requires an increase of the blade natural frequencies above the 2nd Engine Order (EO). Generally this is possible only by utilizing a shrouded blade design. During the design process there are also several other mechanical factors, such as LCF, HCF, creep, shroud coupling, and aero efficiency, which have to be taken into consideration. It is shown that mathematical models of different complexity levels (1D / 3D models) can give comparable results for many important parameters if the boundary conditions for the 1D model are corrected by using the results of 3D model analysis. These parameters include quantities such as shroud contact pressure and coupling forces, the rotor speed at which full coupling is achieved, the first three blade natural frequencies, blade profile untwist moment and angle, and von Mises elastic stresses in critical sections. Thus a combination of 1D and 3D models can achieve a reduction of the time required for design of a shrouded turbine blade. The type of shroud design, the initial shroud clearance, and the contact surface angle are the important parameters for a shrouded turbine blade. The influence of these factors on dynamic response and lifetime prediction is non-linear and not always obvious, so a good blade design usually requires searching for an optimal combination of different design, technological and operational factors. Significant design parameters for blade frequency tuning are the radial distributions of cross-sectional geometrical characteristics and especially the radial distribution of the stagger angle. The latter one allows the designer to bring together or to move apart the first and the second vibration modes. More suitable parameter combinations are determined by the usage of mathematical optimization methods. The method presented and the corresponding techniques were used during design of shrouded blades for a few new turbines, where the length of the blade airfoil has exceeded 800mm.
Proposal of a Novel Multi-Functional Energy System for Cogeneration of Coke, Hydrogen and Power
Authors: Jin Hongguang , Sun Shien , Han Wei , Gao Lin
Published: August 2009

This paper proposes a novel multi-functional energy system (MES), which co-generates coke, hydrogen, and power, through the use of coal and coke oven gas (COG). In this system, a new type of coke oven, firing coal instead of COG as heating resource for coking, is adopted. In this method, the coal is substituted for COG and can be considered as transforming it into coal gas indirectly only by burning instead of gasification. The COG rich in H2 is sent to a pressure swing adsorption (PSA) unit to separate about 80% of hydrogen firstly, and then the PSA purge gas is fed to a combined cycle as fuel. The new system combines the chemical processes and power generation system, along with the integration of chemical conversion and thermal energy utilization. In this manner, both chemical energy of fuel and thermal energy can be used more effectively. With the same inputs of fuel and the same output of coking heat, the new system can product about 65% more hydrogen than that of individual systems. As a result, the thermal efficiency of the new system is about 70%, and the exergy efficiency is about 66%. Compared with conventional systems, the primary energy saving ratio can reach as high as 12.5%. Based on the graphical exergy analyses, we disclose that the integration of synthetic utilization of COG and coal plays a significant role in decreasing the exergy destruction of the MES system. The promising results obtained may lead to a clean coal technology that will utilize COG and coal more efficiently and economically.
Damping Ratio Estimation Techniques for Rotordynamic Stability Measurements
Authors: Cloud C. Hunter , Maslen Eric H. , Barrett Lloyd E.
Published: August 2009

Rotor stability is most commonly estimated using methods derived from a simple, single degree of freedom (SDOF) system. When the modes of more complex systems, such as rotors, are closely spaced, we demonstrate that such methods can yield very poor estimates of the modal stability (damping ratio). Multiple output backward autoregression (MOBAR) is proposed as an alternative approach and is demonstrated to yield reasonably accurate estimates of modal damping even when modes are closely spaced. The performance of the MOBAR approach is then examined on an experimental rotor in tilt-pad bearings, demonstrating good performance in a realistic measurement setting.
Aero-Thermal Investigation of a Rib-Roughened Trailing Edge Channel With Crossing-Jets: Part I—Flow Field Analysis
Authors: Armellini Alessandro , Coletti Filippo , Arts Tony , Scholtes Christophe
Published: August 2009

The present contribution addresses the aero-thermal experimental and computational study of a trapezoidal cross-section model simulating a trailing edge cooling cavity with one rib-roughened wall. The flow is fed through tilted slots on one side wall and exits through straight slots on the opposite side wall. The flow field aerodynamics is investigated in part I of the paper. The reference Reynolds number is defined at the entrance of the test section and set at 67500 for all the experiments. A qualitative flow model is deduced from surface-streamline flow visualizations. Two-dimensional Particle Image Velocimetry measurements are performed in several planes around mid-span of the channel and recombined to visualize and quantify three-dimensional flow features. The jets issued from the tilted slots are characterized and the jet-rib interaction is analyzed. Attention is drawn to the motion of the flow deflected by the rib-roughened wall and impinging on the opposite smooth wall. The experimental results are compared with the numerical predictions obtained from the finite volume, RANS solver CEDRE.
Tip Clearance Flow Visualization of a Turbine Blade Cascade With Active and Passive Flow Control
Authors: Van Ness Daniel K. , Corke Thomas C. , Morris Scott C.
Published: August 2009

The secondary flow in the tip clearance region of a stationary linear low pressure turbine blade cascade was studied using two types of surface flow visualization and documented using wake pressure measurements in order to identify the potential means and impact of flow control to reduce losses associated with the tip clearance flow. An evaporating fluid mixture was used for flow visualization on the casing surface of the tip clearance. An oil ink-dot tracing method was used on the blade tip. These measurements illustrate the important features of the near-casing flow physics, including the size and chordwise extent of the blade tip separation bubble, separation lines on the casing, the flow direction on the blade tip and casing, the size and exit trajectory of the tip leakage and passage vortices, as well as the total pressure loss and secondary velocity vectors downstream of the blade. The flow was visualized in this way for a plain, flat tip, a tip mounted plasma actuator, and a partial suction side squealer tip. Both flow control devices were observed to affect the flow in the clearance. The plasma actuator was shown to improve the total pressure loss in the tip leakage vortex by as much as 9% from the loss over the plain tip blade. The tests were performed over a Reynolds numbers range between 5.3 × 104 and 1.04 × 105 at a fixed tip clearance of 2% of axial chord.
Preliminary Design Optimization of Impingement Cooled Turbine Airfoils
Authors: Bergholz Robert F.
Published: August 2009

Generally, the objective of the turbine cooling design process is to minimize the amount of cooling flow required to achieve component life goals, including creep rupture, low-cycle (thermal-mechanical) fatigue, and material oxidation. High-pressure turbine airfoils in particular are subject to large variations in external gas temperatures and heat transfer coefficients, and high aero-mechanical loads. The efficient distribution of internal and film cooling flows, the management of internal coolant temperatures, and the optimal arrangement of cooling features to control thermal gradients are major factors in controlling thermal stress. The objective of this paper is to describe a process for conceptual and preliminary turbine airfoil cooling design based on the assembly of primitive cooling elements to define an overall “realizable” 3D airfoil cooling structure. This approach allows the evaluation of multiple cooling configurations, and potentially an improved prediction of cooling flows, early in the design process. This conceptual design method can be used to quickly generate models for fabrication and testing in a rapid prototyping laboratory. The design process is outlined for the simplified case of impingement cooling in a prototypical airfoil shape to illustrate some of the key design parameters and procedures. A future paper will address improvements in automation and more complex geometries. The paper also presents predictions based on a parametric impingement cooling model for the local surface heat transfer coefficient distribution. The model is easily incorporated into the conceptual design method. It is particularly useful for determining axial and radial thermal gradients the airfoil wall induced by sparse impingement arrays, for which the use of average heat transfer coefficients give unsatisfactory results. The model was calibrated using data from past published literature and more recently obtained experimental data from the GE Global Research Center.
Gas Turbine Electric Start System (GT-ESS) Ship Integration
Authors: Quin˜ones Martin , Deshmukh Saurabh
Published: August 2009

The U.S. Navy has incorporated a realm of new technologies into ship warfare in the last few years. This availability has led to advances in ship machinery systems that have enabled the Navy to improve and sustain its mission capabilities. Since 2003, the Navy has worked with manufacturers to develop, test and evaluate a gas turbine electric starter system (GT-ESS) capable of meeting all starting requirements of the General Electric LM2500 marine gas turbine. This engine is used on U.S. Navy Surface Combatants for vessel propulsion. The GT-ESS has also been used to start a newly acquired naval engine from Rolls Royce, the MT30. The GT-ESS has endured extensive test and evaluation at the Land Based Engineering Site (LBES) of the Naval Surface Warfare Center, Carderock Division in Philadelphia, PA. Throughout this time the system has been redesigned as well to comply with more restrictive shipboard requirements. This effort had a direct impact on the packaging of the system. Additional effort has been directed to machinery integration. Full integration of the GT-ESS onto the LM2500 engine entails sharing a common lubrication and cooling loop. This is a challenge for the engine since its lubrication is critical for its service life. It poses a challenge for the GT-ESS since its design must guarantee that the lubrication system will remain free of contamination. The GT-ESS also poses integration challenges for a ship installation. The ideal location of the GT-ESS in the ship is adjacent to the propulsion engine within the machinery room space. Switchboards that provide power to the system are located several decks above the machinery space. Additionally, dedicated power supplies are required to handle the power requirements of the GT-ESS. Thus this power demand calls for an assessment of the power loads onboard ship. This new technology also has an impact on other ship systems. The GT-ESS driver is water cooled and designed to feed from the chill water system of the ship. Its electric motor lubrication has ties with another naval technology (digital fuel control). In turn these two are tied to the lube oil system and conditioning assembly (LOSCA) of the engine. Thus issues of oil pressure, temperature, backpressure, suction, and starvation are addressed and assessed. This paper explores the impact of ship machinery design for ship installation and its impact on other machinery systems as well as ship’s power. It targets proper hardware packaging as a means to achieve a well balanced design for ship application.
Nano-Technology Coatings for Erosion Protection of Turbine Components
Authors: Swaminathan V. P. , Wei Ronghua , Gandy David W.
Published: August 2009

Solid particle erosion (SPE) and liquid droplet erosion (LDE) cause severe damage to turbine components and lead to premature failures, business loss and repair costs to power plant owners and operators. Under a program funded by the Electric Power Research Institute (EPRI), TurboMet International (TMET) and Southwest Research Institute (SWRI) have developed hard erosion resistant nano-coatings and conducted evaluation tests. These coatings are targeted for application in steam and gas turbines to mitigate the adverse effects of SPE and LPE on rotating blades and stationary vanes. Based on a thorough study of the available information, most promising coatings such as nano-structured titanium silicon carbo-nitride (TiSiCN), titanium nitride (TiN) and multilayered nano coatings were selected. State-of-the-art nano-technology coating facilities at SwRI were used to develop the coatings. Plasma enhanced magnetron sputtering (PEMS) method was used to apply these coatings on various substrates. Ti-6Al-4V, 12Cr, 17-4PH, and Custom 450 stainless steel substrates were selected based on the current alloys used in gas turbine compressors and steam turbine blades and vanes. Coatings with up to 30 micron thickness have been deposited on small test coupons. Initial screening tests on coated coupons by solid particle erosion testing indicate that these coatings have excellent erosion resistance by a factor of 20 over the bare substrate. Properties of the coating such as modulus, hardness, microstructural conditions including the interface, and bond strength were determined. Tests are in progress to determine the effects of coatings on the tensile and high-cycle fatigue strengths of these alloys.
Losses in High-Speed Permanent Magnet Machines Used in Microturbine Applications
Authors: Huynh Co , Zheng Liping , Acharya Dipjyoti
Published: August 2009

High speed permanent magnet (PM) machines are used in microturbine applications due to their compactness, robust construction and high efficiency characteristics. These machines are integrated with the turbines and rotate at same speeds. This paper discusses in details losses in high speed PM machines. A typical PM machine designed for microturbine application is presented with its detailed loss calculations. Various loss verification methods are also discussed.
Numerical and Experimental Studies on Separated Boundary Layers Over Ultra-High Lift Low-Pressure Turbine Cascade Airfoils With Variable Solidity: Effects of Free-Stream Turbulence
Authors: Funazaki Ken-Ichi , Yamada Kazutoyo , Chiba Yasuhiro , Tanaka Nozomi
Published: August 2009

This paper deals with LES investigation, along with measurements, on the interaction between inlet freestream turbulence and boundary layers with separation bubble over ultra-high lift low-pressure turbine airfoils. The cross section of the test airfoils is typical for highly-loaded LP turbines for civil aeroengines. The solidity of the cascade can be reduced by increasing the airfoil pitch by at least 25%, while maintaining the throat in the blade-to-blade passage. Reynolds number examined is 57,000, based on chord length and averaged exit velocity. Free-stream turbulence is about 0.85% (no grid condition) and 2.1% (with grid condition). Hot-wire probe measurements of the boundary layer are carried out to obtain time-averaged and time-resolved characteristics of the boundary layers under the influence of the freestream turbulence. A newly developed probe positioning tool, which is installed downstream of the cascade with minimal blockage, enables precise probe positioning along lines normal to the airfoil surface. Numerical analysis based on high-resolution LES (Large-Eddy Simulation) is executed to enhance the understanding of the flow field around the Ultra-High Lift and High Lift LP turbine airfoils. Emphasis is placed on the relationship of inherent instability of the shear layer of the separation bubble and the free-stream turbulence. Standard Smagorinsky model is employed for subgrid scale modeling. The flow solver used is an in-house code that was originally developed by one of the authors as FVM (Finite Volume Method)-based fully implicit and time-accurate Reynolds-Averaged Navier-Stokes code. Homogeneous isotropic turbulence created with SNGR (Stochastic Noise Generation and Radiation) method using von Karman-Pao turbulent energy spectrum is applied in the present study for the emulation of inlet turbulence.
Probabilistic Analysis on Turbine Disk Under LCF-Creep
Authors: Hu Dianyin , Wang Rongqiao
Published: August 2009

This paper establishes a structural model of the turbine disk of a certain aero-engine subjected to loads due to centrifugal force and duration at elevated temperature, and calculates the plastic and creep deformation with finite element (FE) analysis software. During creep analysis, the Norton constitutive equation was taken into account. After the temperature field distribution of the turbine disk had been obtained, structural static analysis results showed that the maximum equivalent stress and strain appeared at the area near to the disk center, which was taken as the critical point. The Manson-Coffin formula was used to predict low cycle fatigue (LCF) life based on the calculations derived from cyclic plastic-creep analysis, and creep life was obtained by applying the Larson-Miller equation. Lognormal distributions of LCF life and creep life were used as pointed out in the literature, which mainly considered the uncertainties in material parameters. Taking the lives and loads as random variables including LCF life, creep life, fatigue load and hold time, a probabilistic analysis on the turbine disk was performed under LCF-Creep, in which a failure function was proposed based on linear cumulative damage rule (LCDR). The probabilistic method of response surface (RS) was applied to fit the regression model of the LCF-Creep life with a quadratic approximation function including cross-terms. The Monte Carlo Simulation sampling technique was employed to carry out probabilistic analysis on the turbine disk life, resulting in the conclusion that the life of turbine disk under LCF-Creep follows a lognormal distribution. Furthermore, the effect of different random variables on the disk life was investigated through sensitivity analysis in order to increase the component’s life and improve its reliability. However, studies on the distributions of loads and damage theory of the component under LCF-Creep still need to be discussed in future study.
Numerical Investigation of a Stagnation Point Reverse Flow Combustor
Authors: Cui Yufeng , Lu Xuan , Xu Gang , Chen Jianli , Nie Chaoqun , Huang Weiguang
Published: August 2009

Flameless combustion is characterized by ultra-low NOx emissions, high combustion efficiency and very stable flame, which is able to operate stably at very lean fuel-air mixtures without problems of combustion oscillation and flashback. Stagnation Point Reverse Flow (SPRF) combustors, as an important application of flameless combustion, have been experimentally studied by various optical diagnostic techniques. In this paper, Eddy Dissipation Concept (EDC) model with detailed chemical reaction mechanisms of natural gas GRI 2.11 is used to investigate the flame characteristics of a SPRF combustor operating at premixed mode with various mass flow rates and equivalence ratios of CH4 and air mixtures. The numerical results indicate that as the fuel and air mixture injection velocities increase, there are no distinct changes in the jet penetrations. However, flame temperatures and NOx emissions decrease, CO emissions increase, and OH is distributed in wider area and more evenly. For turbulent flow, intense reactions take place in the shear layer and the stagnation zone and they gradually shift to the combustor outlet as the jet velocities increase. As the equivalence ratios increase from 0.5 to 1, the NOx emissions always increase, although they are very low when equivalence ratio is below 0.7. However, the CO emissions decrease firstly, reaching the minimum value at equivalence ratio of 0.58, and then increase. The numerical results are compared with experimental data and it is verified that EDC model can capture the important flow field characteristics and flame structure and is appropriate for modeling SPRF combustor.
Influence of Blade Fillets on the Performance of a 15 Stage Gas Turbine Compressor
Authors: Ku¨geler E. , Nu¨rnberger D. , Weber A. , Engel K.
Published: August 2009

In the modern process of the aerodynamic design of multistage compressors and turbines for jet engines as well as for stationary gas turbines, 3D-CFD plays a key role. Before building the first test rig several designs have been investigated using numerical simulations. To understand the characteristics of the individual components it is necessary to simulate their behavior in a multistage simulation and investigate for example, the single stage maps of the compressor in order to understand how the load is divided between the different parts of the compressor during throttling. Increasing computing resources allow ever more details to be incorporated in a 3D simulation. In former times only single blade rows were investigated with a high resolution of the boundary layers, whereas in multistage configurations wall functions were state of the art. Today we are able to apply Low Reynolds resolution even for multistage configurations, so the designer is required to include more and more geometrical details into the simulation. One important such feature is the fillets of rotor and stator blades. Fillets reduce the flow deflection at the endwalls and therefore the loading of the downstream blade rows. This effect is accumulated in a multistage simulation. In this paper a 15-stage compressor with additional inlet and outlet guide vane designed for a stationary gas turbine was investigated with a modern CFD tool by using a real gas approach for two speedlines. Two simulations were done: first a clean configuration with tip and hub clearances but without blade fillets; in the second simulation all rotor blades and the cantilevered stator blades were additionally modeled with fillets. The comparison of the overall global values with measurement data shows a better performance of the simulation with fillets, especially by throttling the compressor. A deeper look into the compressor shows different loads for a considerable number of single stages. The analysis of the steady multistage simulations shows that the numerical stability is reached in different regions of the machine.
Measurement Error Influence on Gas Turbine Operability for Condition-Based Maintenance and Reliability/Availability Improvement
Authors: Fadlun Ever Avriel , Michelizzi Ilaria , De Iaco Marco
Published: August 2009

Nowadays the major oil & gas companies are forced to move to a condition-based maintenance in order to extend mean time between maintenance (MTBM) and increase productivity, predicting gas turbine behavior and preventing unexpected failures. Gas turbine operability is influenced by many technological limitations, such as high temperature exposure limit of hot components, requested fuel composition flexibility, combustion chamber pulsations, pollutant emissions, rotor vibrations, compressor surge limit etc. In order to prevent exceeding these limitations certain significant measurable variables are monitored and controlled during machine operation; the instrumentation accuracy and stability have a strong impact on gas turbine reliability, availability and control. The aim of this paper is to demonstrate how the accuracy and stability of the instrumentation installed on a gas turbine has a direct influence on its operability. Monte Carlo simulations have been carried out to assess each instrument deviation effect and the level of impact. The second phase is oriented towards establishing an improvement process for a safe and reliable machine operation. Among the above-mentioned critical limits the main focus of this analysis is the firing temperature estimation and its control, which is performed by measuring gas turbine exhaust temperature and compressor discharge pressure. This study has highlighted how an error on compressor discharge pressure measurement during unit operation may provoke gas turbine over-firing or under-firing, respectively causing fast life reduction of hot components and the limitation of maximum delivered power. Furthermore, an error on compressor discharge pressure measurement would strongly decrease firing temperature estimation confidence, which is the main parameter for components residual life assessment to target a condition-based maintenance approach. In this paper the corrective actions to increase the accuracy of firing temperature estimation and control will be described. Based on the results of this study, reliability and availability improvements will also be presented. Using the present work as a guideline, similar analyses can be carried out for all other critical aspects that can be related to measurable variables.
Analysis of Noise Generated by Low Solidity Cascade Diffuser in a Centrifugal Blower
Authors: Sakaguchi Daisaku , Ishida Masahiro , Ueki Hironobu , Hayami Hiroshi , Senoo Yasutoshi
Published: August 2009

This paper deals with the effect of the blade leading edge location (RLSD) of a low solidity cascade diffuser (LSD) on noise and diffuser performance in a centrifugal blower. The noise of the LSD was measured and analyzed comparing with that of vaneless diffuser (VLD) in view points of overall noise, discrete frequency noise and broadband noise. The numerical flow analysis was conducted in the impeller and the diffuser by using a Navier-Stokes solver. The noise of the VLD varied little in a wide flow rate range, on the other hand, that of the LSD increased remarkably in the small flow rate by about 7 dB. The noise of the LSD did not increase near the design flow and was almost equal to that of the VLD. It was found that the increase in noise due to LSD is dependent mainly on the broadband noise between 600∼1000Hz, which was closely correlated to the lift force of the LSD blade. The two kinds of discrete frequency noise appeared due to an interaction between the rotating impeller and the LSD blade and another interaction between the rotating impeller blades and the reverse flow toward the impeller exit, but their influence on the overall noise were relatively small. By shifting the LSD blade leading edge location downstream from RLSD = 1.1 to 1.2, the noise was reduced by about 3 dB at the maximum without deterioration of the diffuser performance. The maximum lift coefficient of the LSD blade was achieved as high as 1.5 at the high attack angle of 17 degrees even in the case of RLSD = 1.2, resulting in improvement of the diffuser performance by about 40% and in reduction of the unstable flow range by about 11%.
Distortion Compensation by Shape Modification of Complex Turbine Geometries in the Presence of High Temperature Gradients
Authors: McElhaney John
Published: August 2009

It is often the case that the turbine casings have rather complex shapes, and are loaded by three dimensional, high temperature gradients. This is especially so with e.g. steam turbine casings’ requiring shapes such as spirals for steam admission, spherical bowls for steam extraction, and the usual rectangular shaped flanges for bolting the casing halves together. The combination of shape complexity in the presence of high temperature gradients creates significant distortion at the inside casing walls along the steam path. This leads to flow losses between the blade tips and labyrinth seals which results in a reduction in turbine efficiency. The aim in this paper is to present methods for modifying standard design casings’ that lead to significant reduction in two primary forms of distortion. The first is based on metal blanket designs that implement a variable added thickness to the casing outside casing wall. The shape and thickness of the metal blankets are shown to be very naturally scaled to the distortion of the turbine itself. The second method involves trimming and cutting modifications to the bolting flanges which in fact leads to a reduction of material. The methods presented are thought to be generally applicable to any casing that distorts due to temperature gradients. To the authors best knowledge, the approach implementing metal blankets is unique and has not previously been covered in the literature. In fact, there seems to be little, if any literature related to the topic. The reduction in distortion of a modified standard design casing and a brief summary of the subsequent improvement in steady state efficiency is described. A 420 MW, intermediate pressure steam turbine inner casing with 60 bar, 600° C steam at the inlet, and 6 bar, 300° C at the exhaust is presented as an example.
An Analytical Approach to Support High Cycle Fatigue Validation for Turbocharger Turbine Stages
Authors: Heuer Tom , Gugau Marc , Klein Achim , Anschel Paul
Published: August 2009

As the pursuit of improved aerodynamic performance on turbochargers continues to push the boundaries of mechanical design, the risk of high cycle fatigue (HCF) failures of turbine wheels is elevated and drives the need for improved methods of analytical predication. Turbine wheel HCF is caused by the asymmetrical aerodynamic loading associated with the application of variable geometry turbines and compromised inlet/outlet geometries that accommodate packaging turbochargers in engine compartments. Historically, the turbocharger industry has focused on blade pass as the primary cause of HCF. BorgWarner presents an analytical technique to limit the risk of HCF by calculating all critical orders that intersect blade natural frequencies and quantify the relative energy of each forced excitation frequency. The geometry of a turbine volute mainly determines the inflow angle into the turbine wheel. Thus, for the thermodynamically optimized turbine a variable inflow angle dependent on the engine operating point is desirable. The Variable Turbine Geometry (VTG) concept applies adjustable turbine inlet guide vanes to approach the ideal velocity triangle. One of the inevitable disadvantages associated with either fixed or variable turbine nozzle vanes is the generation of wakes and pressure fluctuations upstream of the turbine wheel inducer. The resulting circumferentially non-uniform flow conditions apply a transient load on the rotating turbine wheel. Due to complexity of the VTG design including non-uniformly spaced vanes and struts, the excitation sources and resulting excitation orders are not readily apparent. The analytical method described in this paper applies a transient 3D Computational Fluid Dynamics (CFD) model of the rotor-stator interaction to calculate the time-dependent pressure fluctuations experienced by the turbine wheel blade. This data is used to extract the forced excitation function at the turbine wheel. Finite Element (FE) analysis is applied to determine the mean and dynamic stress. In case of dynamic stress, system vibration modes and the influence of local harmonic excitation on blade stress amplitudes is analyzed. The fixed-speed FE results are scaled for the effect of flow and speed by use of empirical data from strain-gauge measurement. Hence, computational methods combined with experience from experimental measurements are used to determine critical rotational speeds for a given turbocharger geometry. This method allows the analyst to predict the highest energy excitation orders and reduce the risk of turbine wheel fatigue damage. Since the durability of a turbine wheel cannot yet satisfactorily be quantified by the described computational method, the analysis results are used as rotational speed input in subsequent durability tests in order to reduce the necessary amount of testing resources.
Condition Based Maintenance of Gas Turbines Using Simulation Data and Artificial Neural Network: A Demonstration of Feasibility
Authors: Fast Magnus , Assadi Mohsen , De Sudipta
Published: August 2009

Gas turbine maintenance is crucial due to high cost for the replacement of its components and associated loss of power during shutdown period. Conventional scheduled maintenance, based on equivalent operating hours, is not the best alternative as it can require unnecessary shut downs. Condition based maintenance is an attractive alternative as it decreases unnecessary shut downs and has other advantages for both the manufacturers and the plant owners. However, this has shown to be a complex/difficult task. A number of methods and approaches have been presented to develop condition monitoring tools during the past decade. Condition monitoring tools can e.g. be developed by means of training artificial neural networks (ANN) with historical operational data. Such tools can be used for online gas turbine performance prediction where input data from the plant is fed directly to the trained ANN models. The predicted outputs from the models are compared with corresponding measurements and possible deviations are evaluated. With this method both recoverable degradation, caused by fouling, and irrecoverable degradation, caused by wear, can be detected and hence both compressor wash and overhaul periods optimized. However, non-availability of operational data at the beginning of the gas turbine operation may cause problems for the development of ANN based condition monitoring tools. Simulation data, on the other hand, may be generated by using a manufacturer’s engine design program. This data can be used for training artificial neural networks to overcome the problem of non-availability of operational data. ANN models trained with simulation data could be used to monitor the engine from the very beginning of its operation. A demonstration case using a Siemens gas turbine has been shown for this proposed method by comparing two ANN models, one trained with operational data and the other with simulation data. For the comparison an arbitrary section of operational data was used to produce predictions from both models, whereupon these were plotted with corresponding measured data. The comparison shows that the trends are very similar but the parameter values for the measured and the simulated data are shifted by a constant. Using this knowledge, one can provide an ANN based engine monitoring tool that could be adjusted to a certain engine using engine performance test data. The study shows promising results and motivates further investigations in this field.
A Study on Multi Fault Diagnostics of Smart Unmanned Aerial Vehicle Propulsion System Using Data Sorting and Neural Networks
Authors: Kong Changduk , Kho Seonghee , Ki Jayoung , Lee Changho
Published: August 2009

The types and severities of most engine faults are so complex that it is not easy for a conventional model based fault diagnosis approach like the GPA (Gas Path Analysis) method be used to monitor all engine fault conditions. This study therefore discusses on the newly proposed diagnostic algorithm for isolating and effectively identifying the faulted components of the smart UAV propulsion system, that has been developed by KARI (Korea Aerospace Research Institute) based on the fuzzy logic and the neural network algorithms. The diagnosis procedure of the proposed diagnostic system has the following steps. First obtaining database of fault patterns through performance simulation, followed by training the database using the FFBP networks. The third step involve analyzing the trend of the measured parameters due to fault patterns, linked to this is the fourth step that involve isolating the faulty components using fuzzy logic, and finally the magnitudes of the detected faults are obtained by the trained neural networks. The analysis showed that the detected faults had almost same degradation values to those of the implanted fault pattern, confirming that the proposed diagnostic system can be used to effectively detect the engine faults.