Showing papers on "Turbine blade published in 1989"

Heat transfer in rotating passages with smooth walls and radial outward flow

Abstract: Experiments were conducted to determine the effects of rotation on heat transfer in turbine blade internal coolant passages. The experiments were conducted with a smooth wall, large-scale heat transfer model. The objective was to obtain the heat transfer data base required to develop heat transfer correlations and to assess computational fluid dynamic techniques for rotating coolant passages. An analysis of the governing equations showed that four parameters influence the heat transfer in rotating passages (coolant density ratio, Rossby number, Reynolds number, and radius ratio). These four parameters were varied over ranges that exceed the ranges of current open literature results, but that are typical of current and advanced gas turbine engine operating conditions. Rotation affected the heat transfer coefficients differently for different locations in the coolant passage. For example, heat transfer at some locations increased with rotation, but decreased and then increased again at other locations. Heat transfer coefficients varied by as much as a factor of five between the leading and trailing surfaces for the same test condition and streamwise location. Comparisons with previous results are presented.

High temperature coatings for gas turbine blades: A review

Abstract: High temperature coatings are formed to protect many engineering components from environmental degradation. The operating temperature and the nature of the corrodant dictate the choice of the coating. The coatings can be formed by different methods, but invariably the properties are dependent on the coating process. The life of the coating depends upon the mode of degradation in service. It is also influenced by the nature of the substrate. The need for coatings is well illustrated by their wide use in gas turbines. This review covers the development of the oxidation- and hot-corrosion-resistant coatings for turbine blades. The rationale behind such protection, starting from the nickel aluminide to future thermal barrier oxide coatings, is outlined. It is followed by a discussion of current trends in post-coating modifications to improve the properties of the coatings. The principles underlying the development of these coatings should serve as a useful guide in the choice of coatings for other high temperature applications.

Geared counterrotating turbine/fan propulsion system

Abstract: High by-pass fan jet engine includes counterrotating turbine blade sets for receiving hot combustion gases from a core engine portion and driving the fan blades through a planetary gear-type reduction gear assembly. Lightweight, highspeed, concentric, counterrotating shafts transmit power from both turbine blade sets to the fan via the reduction gear assembly. A pair of planetary gear assemblies with fixed fing gears and a rotatable common planetary gear carrier is used to drive a single set of fan blades via the carrier, and a pair of planetary gear assemblies with a fixed common planetary gear carrier and rotatable ring gears is used to drive two or more counterrotating fan blade sets via the ring gears. A low pressure "booster" compressor feeding the core engine portion can be directly driven from one of the two high speed shafts.

State of the art on high-temperature corrosion-resistant coatings

Abstract: A high-temperature protective coating must meet several criteria: provide adequate environmental resistance, be chemically and mechanically compatible with the substrate, be applicable. Comprehensive reviews on high-temperature coatings have appeared regularly since the early 1970s. Our purpose is not to recapitulate the material covered therein but rather to focus on recent trends, and point out some research perspectives. Historically the development of high-temperature protective coatings has been linked with the evolution of demanding applications such as super-alloy components in gas turbine engines; the searches for better performance (higher inlet temperatures, longer lifetimes, etc.) and for cost-saving solutions (use of contaminated low-grade fuels) have been the main incentives for developing the different coatings now available in production: simple and “modified” diffusion coatings, overlay coatings, thermal barriers. In recent years research and development activity has been concentrating on the following points. 1. (a) Degradation mechanisms in high-temperature corrosion of metallic coatings; basic studies on the growth mechanisms of oxide scales, for example are still required, in particular to understand the role of addition elements such as platnum and palladium. 2. (b) Alternative techniques for depositing MCrAlY coatings; electrolytic codeposition and electrophoresis, for example, have been developed at the laboratory stage and these permit the deposition of MCrAlY coatings with claimed economic and technical advantages over processes already in production. 3. (c) Thermal barrier coatings; ceramic coatings have been applied to sheet metal combustor components for about 15 years; only recently have they been used in the turbine section. Two challenges remain though to exploit these coatings on turbine blades: improve their reliability and, in the case of stationary gas turbines, their hot corrosion resistance. Both structural and mechanical approaches are required to determine, in particular, the role of microstructure, microcracking, porosity, residual stresses, oxidation of the bond layer in the degradation mechanisms of these coatings. 4. (d) Mechanical properties of coated systems; the intrinsic mechanical properties of coating materials are still poorly described and the lack of information hinders the adequate modelling of the behaviour of coating-substrate composite systems. In parallel, an increasing activity is noted concerning the design and development of high-temperature coatings for protecting materials other than superalloys, for instance ceramic composites and titanium-based alloys.

Thermal barrier coatings for gas turbine use

Abstract: The use of thermal barrier coatings on high-pressure turbine components can improve gas turbine efficiency through reduction of cooling airflow. However, the risk involved in reducing cooling airflow requires a highly reliable thermal barrier coating. This increased reliability will be achieved through several complementary approaches; material and process development, life prediction method development and engine service experience. The two processes available for deposition of thermal barrier coatings (plasma spray and physical vapor deposition) are compared, and the advantages and disadvantages of each discussed as they apply to gas turbine components. The results of bond coat material development which has increased the thermal cycle life of plasma spray thermal barrier coatings are presented. Improvements were achieved by two methods: (1) the use of creep-resistant bond coat compositions and (2) overaluminiding of the bond coat. Results of engine testing of thermal barrier coatings in an environment that produces hot corrosion are also presented.

Cooling of turbine blades

Abstract: In order to facilitate efficient use of high pressure cooling air in a turbine rotor blade its aerofoil has a triple-pass convoluted cooling air duct in its leading edge region, and a triple-pass convoluted cooling air duct in its mid-chord region, both of them being fed from a common high pressure inlet in the root of the blade. The trailing edge region has a single-pass duct fed by low pressure cooling air from an inlet located just under the blade platform. The shroud of the blade also has cooling air passages and these are fed from the ends of respective ones of the ducts. Air from the ducts also exits through rows of film cooling holes to film cool the concave flank of the aerofoil and through a row of holes in its trailing edge to remove heat from the thinner metal section in that area.

Combined air-hydrogen turbo-rocket power plant

Abstract: A combined air-hydrogen turbo-rocket engine is disclosed having a simplified construction in which the hydrogen driven turbine is formed integrally with the rotor wheel of the axial air compressor stages. The rotor stages are located downstream of a stator vane structure and are driven by gaseous hydrogen passing across the turbine blades. The hydrogen is subsequently injected into an air duct surrounding the axial air compressor and defining an airflow path having an air inlet. The hydrogen-air mixture is ignited and the burned gases are expanded through a converging-diverging exhaust nozzle.

Life modeling of thermal barrier coatings for aircraft gas turbine engines

Abstract: Thermal barrier coating life models developed under the NASA Lewis Research Center's Hot Section Technology (HOST) Program are summarized. An initial laboratory model and three design-capable models are discussed. Current understanding of coating failure mechanisms are also summarized. The materials and structural aspects of thermal barrier coatings have been successfully integrated under the HOST program to produce models which may now or in the near future be used in design. Efforts on this program continue at Pratt and Whitney Aircraft where their model is being extended to the life prediction of physical vapor deposited thermal barrier coatings.

Apparatus and method for optimizing the air inlet temperature of gas turbines

Abstract: A heat exchanger is disclosed for heating air entering a combustion gas turbine to increase the power output of the turbine when the turbine operates in a cold environment. The heat exchanger may be used also as a cooler to cool air entering the turbine to increase the power output of the turbine when the turbine operates in a hot environment.

Turbine blade fatigue monitor

Abstract: An on-line vibratory fatigue monitor measures displacement of an object such as the blade of a turbine rotor to generate a displacement signal and calculates accumulated fatigue in the object in dependence upon the displacement signal. The displacement of the object may be detected by passive proximity probes which generate signals indicative of changes in magnetic flux or capacitance between the sensor and the turbine blade. An analog/digital converter converts the signals from the probes to digital signals indicating displacement of the turbine blades. The digital signals are processed, preferably by taking a Hilbert transform, to detect the amplitude envelope and instantaneous frequency of the displacement. The amplitude and frequency are combined with steady state stress on the object to determine fatigue usage which is accumulated by constantly monitoring vibratory displacement and changes in the steady state stress.

The Influence of Turbine Clearance Gap Leakage on Passage Velocity and Heat Transfer Near Blade Tips: Part II—Source Flow Effects on Blade Suction Sides

Abstract: A study has been conducted to investigate influences of tip leakage flow on heat transfer and flow development along the pressure side of a gas turbine blade. An analysis of the sink character of the flow situation indicates that high velocities and accelerations are generated very near the gap, and an apparatus was specifically designed to model the phenomena and to permit resolution of the expected localized near-gap heat transfer enhancement

Investigation of coatings at high temperature for use in turbomachinery

Abstract: Aircraft gas turbines, steam turbines and coal utilization turbines operating under particulate flows are exposed to erosion and performance deterioration. Almost all advanced aircraft turbines today are manufactured with a protective coating on the blades. The particulates in the steam turbines are produced in the boiler from exfoliated oxide scales and carried in the steam to the turbine. In coal utilization turbines, the particulates are by-products of the combustion process. There are many problem areas which need future research to be able to produce better and more durable high-temperature coatings. However, this paper will present only some of the work done at the University of Cincinnati on erosion testing at high temperatures and velocities for different materials and coatings. The testings have been performed with a special high-temperature erosion wind tunnel which simulates the aerodynamic conditions on the blades.

Turbofan engine having a counterrotating partially geared fan drive turbine

Abstract: A high bypass ratio turbofan engine having a fan section, a booster compressor disposed aft of the fan section relative to the flow of combustion gases through the engine, and a core section disposed aft of the booster compressor. A low pressure counterrotating turbine, disposed aft of the core section, is used for driving the fan section and the booster compressor. The counterrotating turbine includes at least one set of rotating turbine blades and at least one set of oppositely rotating counterrotating turbine blades. A twin spool shaft is provided for coupling the turbine blades to the booster compressor and for coupling the counterrotating turbine blades to the fan section. A reduction gear is disposed in the drive shaft for coupling the turbine blades to the fan section and for reducing the rotational speed of the turbine output power to match the rotational speed of the fan section thereby splitting the usable work of the turbine blades between the fan section and the booster.

Nonsynchronous turbine blade vibration monitoring system

Abstract: A nonsynchronous turbine blade vibration monitoring system is comprised of a maximum of two sensors circumferentially mounted along the turbine blade row for detecting the actual arrival times of the turbine blades at the sensors. Expected arrival times of the turbine blades at the sensors are determined and then compared to the actual arrival times in order to obtain turbine blade deflection data. Harmonic analysis is then performed on the blade deflection data to determine the level of vibration at each non-integral harmonic.

Dynamic modelling and robust control of a wind energy conversion system

Abstract: The application of wind energy conversion systems for the production of electrical energy requires a cheap and reliable operation. Especially at high wind velocities fluctuations from the wind field result in large mechanical loads of the wind turbine. Also fluctuations in the grid voltage may yield large dynamic excitations. In order to realize a long lifetime and a reliable operation active control systems are necessary. The main goal of the study described in this thesis is to develop an approach for the design of a high performance control system for a wind turbine with variable speed. The wind turbine system under investigation has a three-bladed rotor which is connected to the generator by a transmission. The electrical conversion system consists of a synchronous generator with a rectifier, direct current transmission and an inverter. The manipulable inputs are the pitch angle of the turbine blades, the field voltage of the generator and the delay angle of the rectifier. Both the generator speed and the direct current are being measured. The control design problem at full load is to minimize fluctuations in speed and current while reducing the mechanical (fatigue) loads. The feedback system should realize this without excessive use of the input variables and must also be simple to implement. I In order to be able to design a high performance control system a high quality dynamic model is required. Much attention has been given to the. modelling of the electrical conversion system. The switching of the thyristors of the rectifier bridge results in periodic behaviour at a high frequency. In order to design a control system an averaged model has been derived through the application of Floquet theory for periodic systems. The properties of the aerodynamic transfer and of the drive train only have been approximately modelled. Deviations of these nominal models from the real system are accounted for using norm-bounded uncertainty models. Using the nominal model and the uncertainty models the control system design has been carried out. The control design problem can suitably be handled by the Linear Quadratic design method. However, instead of using the standard solution with observers, in this study the optimization theory has been applied with respect to a predefined structure of the feedback law. In this approach the order and structure of the controller can be selected as part of the problem formulation. The application to the wind turbine system shows that a high performance control system can be obtained using a relatively simple, low order multivariable feedback law. The use of frequency weighting effectively reduces the role of mechanical parasitic dynamics. Application of the multi-model principle in combination with LQ optimization theory provides a way to synthesize controllers which are robust for large (aerodynamic) changes in operating conditions. A quantitative robustness analysis shows how the design parameters of the feedback law can be adapted in order to enhance the robustness of the controller. The approach taken, involving extensive modelling combined with uncertainty models and with the use of optimization theory and robustness analysis, has been shown to result in a high performance control system. Its main characteristic is the integrated approach of the control problem, with combined control action via the mechanics and the electrical conversion system. It is recommended to apply this integrated approach also to other types of wind turbine systems.

Thermal barrier coatings for turbine applications in aero engines

Abstract: The current status of thermal barrier coatings (TBGs) is reveiewed with reference to gas turbine applications. The principal methods of TBC deposition are outlined and failure processes based on thermomechanical stress effect, bond coat oxidation, corrosion of constituents of the ceramic, and eroswn are described. The performance of TBG systems in engine trials is described and correlated wlth the various degradation processes and experimental data.

Observations of directional gamma prime coarsening during engine operation

Abstract: Two alloys with negative mismatch parameters, NASAIR 100 and a modified NASAIR 100 called Alloy 3 were run as turbine blades in an experimental ground based Garret TFE731 engine for up to 200 hr. The directional coarsening of gamma prime (rafting) that developed during engine testing was analyzed and compared to previous research from laboratory tests. The blades were found to be rafted normal to the centrifugal stress axis over much of the span, but near the surfaces, the blades were found to be rafted parallel to the centrifugal stress axis for certain cycles. Representative photomicrographs of the blades and the effects of stress and temperature on raft formation are shown.

Gas turbine engine with turbine tip clearance control device and method of operation

Abstract: The problem of excessive clearances being generated by rub occurring between the tips of turbine blades and the linings on shroud segments which surround the blades is addressed by way of providing a valve 42 of a metal of different coefficient of expansion than the shroud control ring 24. On ground running prior to take off of an associated aircraft, the valve diverts a cooling airflow away from the ring 24, thus causing it to run hot and expand radially. The shrouds 22 are thus moved away from the blade tips 40 and on take off, are sufficiently far away from the tips 42 as to avoid deep penetration of the lining 38 thereby. The higher temperature generated by take off affects the valve 40 and makes it expand rapidly, to open the cooling airflow paths to the ring 24 via the holes 46 and 50.

Turbine blade clearance controller

Abstract: A control system (30) for measuring the clearance ''x'' between the tip (26) of a blade (24) and wall (28) in a turbine engine (8). A series of probes (32, 32' and 32N) located in the wall (28) each have a reference orifice (52) and a sensing orifice (54). Fluid from a source flows through the reference orifices (52) into the turbine engine. This fluid flow is substantially unrestricted and as a result a fluid pressure P?2? is established in a reference chamber (44). Fluid flow from the sensing orifice (54) is restricted by the relationship of the tip (26) of each blade (24) and as a result fluid pressure P2' is created in chamber (48). A multiplexer (62) sequentially provides transducer (92 and 94) with fluid pressure P2 from the reference chamber (44) and P2' from sensing chamber (48). The transducers (92 and 94) convert the pneumatic signals into electrical signals which are supplied to a computer (116). The computer (116) supplies a control member (118) with an operational signal to allow cooling air to flow to housing (10) to maintain a desired tip clearance ''x'' for the blades of the turbine.

Surface Injection Effect on Mass Transfer From a Cylinder in Crossflow: A Simulation of Film Cooling in the Leading Edge Region of a Turbine Blade

Abstract: A naphthalene sublimation technique is used to study the effect of surface injection on the mass (heat) transfer from a circular cylinder in crossflow. Using a heat/mass transfer analogy the results can be used to predict film cooling effects in the leading edge region of a turbine blade. Air injection through one row of circular holes is employed in the stagnation region of the cylinder. Streamwise and spanwise injection inclinations are studied separately, and the effects of blowing rate and injection location relative to the cylinder front stagnation line are investigated. Streamwise injection produces significant mass transfer increases downstream of the injection holes, but a relatively small increase is observed between holes, normal to the injection direction. The mass transfer distribution, measured with spanwise injection through holes located near the cylinder front stagnation line, is extremely sensitive to small changes in the injection hole location relative to stagnation. When the centers of the spanwise injection holes are located 5° or more from the stagnation line, the holes lay entirely on one side of the stagnation line and the injection affects the mass transfer only on that side of the cylinder, approaching the pattern observed with streamwise injection.Copyright © 1989 by ASME

Combination convection and microwave oven having improved microwave energy distribution

Abstract: A cooking oven having both conventional heating as well as microwave heating. The oven muffle includes a metal distribution sheet along a back wall, forming a cavity with the back wall. A bladed turbine fan is included within the cavity, driven by an electric motor. Microwave energy is introduced into the cavity through a waveguide having an exit iris in the cavity. Microwave energy entering the cavity exits through openings in the metallic distribution sheet, as well as past the rotating turbine blade through additional holes in the distribution sheet. The microwave energy exiting the distribution plate provides for a better cooking energy distribution throughout the oven.

Phase-resolved heat-flux measurements on the blade of a full-scale rotating turbine

Abstract: This paper presents detailed phase-resolved heat-flux data obtained on the blade of a Teledyne 702 HP full-stage rotating turbine. A shock tube is used as a short-duration source of heated air, and platinum thin-film gages are used to obtain the heat-flux measurements. Results are presented along the midspan at several locations on the blade suction and pressure surfaces from the stagnation point to near the trailing edge. For these measurements, the turbine was operating at the design flow function and at 100 percent corrected speed. Results are presented for the design vane/blade spacing (0.19 Cs) and at a wide spacing (0.50 Cs). Data are also presented illustrating the phase-resolved blade heat-flux distribution with upstream cold gas injection from discrete holes on the vane surface. The results illustrate that several successive passages can be superimposed upon each other and that a heat-flux pattern can be determined within the passage.

Phase and Time-Resolved Measurements of Unsteady Heat Transfer and Pressure in a Full-Stage Rotating Turbine

Abstract: This paper presents detailed phase-resolved heat-flux data obtained on rotor blades and a comparison of simultaneously obtained time-resolved heat-flux and static pressure data obtained on the stationary shroud of a Garrett TFE 731-2 HP full-stage rotating turbine. A shock tube is used to generate a short-duration source of heated and pressurized air and platinum thin-film gages are used to obtain heat-flux measurements. Blade results are presented at several selected blade locations. Shroud surface pressure and heat-flux time histories are presented for comparable locations relative to the blade position.

Design of airfoils and cascades of airfoils

Abstract: This paper presents a method for generating blade shapes to be used as inputs to the direct- or inverse-bladedesign sequences. This method can generate subsonic or supersonic blades for compressors and turbines, or isolated airfoils, but the discussion in this paper is limited to subsonic-exit turbine blade rows. Gas-turbine blades are usually designed by a mixture of iterative, direct- and inverse-blade-design methods. The iterations are enormously reduced if the initial blade shape has performance characteristics near the desirable ones. The desirable performance characteristics are presented, and ways to manipulate the input striving for such characteristics (minimizing design iterations) are discussed. The overspeed regions near the leading edge are avoided. The performance curves of three blades generated by this method are included. It is concluded that blade performance is extremely sensitive to small changes in the shape of the surfaces and to changes or discontinuities in the derivative of the curvature of the blade surfaces. The order of magnitude of these changes is the same or lower than the order of magnitude of the thickness distribution of the boundary layers.

Method of fabricating an air cooled turbine blade

Abstract: A turbine blade is formed from an airfoil blade 10 having V-grooves 12 in the outer surface and indentations 26 in the inner surface. A slot 30 is machined at the root of the groove intersecting the slots, and providing a smooth flow path for the cooling air discharging along the blade surface.