Showing papers on "Turbine blade published in 1996"

Experimental Validation of Condensing Flow Theory for a Stationary Cascade of Steam Turbine Blades

Abstract: The paper describes a detailed experimental and theoretical study of non-equilibrium condensing steam flow in a stationary cascade of turbine blades operating transonically. Instrumentation was installed for obtaining colour schlieren photographs of the shock wave structure, the blade surface static pressure distribution, the pitchwise variation of the mean droplet radius downstream of the cascade and the stagnation pressure loss across the cascade. Only one blade profile was tested but a comprehensive set of measurements was acquired covering a wide range of inlet steam conditions and exit Mach numbers. By careful interpretation of the data, it was possible, for the first time, to infer the thermodynamic loss due to irreversible condensation directly from experimental measurements. An elaborate comparison of the experimental data with condensing flow theory was also undertaken using a two-dimensional inviscid time-marching calculation scheme, simulating both steady and unsteady flows. Excellent agreement was obtained throughout and it can be stated with some confidence that the theory and calculation procedures used reproduce accurately all the main features of steady transonic condensing flow in stationary cascades.

Heat Transfer Committee Best Paper of 1995 Award: Distribution of Film-Cooling Effectiveness on a Turbine Endwall Measured Using the Ammonia and Diazo Technique

Abstract: The distribution of adiabatic film-cooling effectiveness on the endwall of a large-scale low-speed linear turbine cascade has been measured using a new technique. This technique is based on an established surface-flow visualization technique, and makes use of the reaction between ammonia gas and a diazo surface coating. A new method of calibration has been developed to relate the result of the reaction to surface concentration of coolant. Using the analogy that exists between heat and mass transfer, the distribution of film-cooling effectiveness can then be determined. The complete representation of the film-cooling effectiveness distribution provided by the technique reveals the interaction between the coolant ejected from the endwall and the secondary flow in the turbine blade passage. Over- and undercooled regions on the endwall are identified, illustrating the need to take these interactions into account in the design process. Modifications to the cooling configuration examined in this paper are proposed as a result of the application of the ammonia and diazo technique.

Unsteady aerodynamics associated with a horizontal-axis wind turbine

Abstract: The wind turbine industry is currently facing many difficulties constructing efficient wind turbine machines caused by the inability to adequately predict structural loading and power output. Available evidence from wind turbines in a field environment suggests that formation of complex unsteady separated flowfields may be responsible for many aspects of wind turbine component failure. To examine this possibility in more detail, the Combined Experiment was developed. A full-scale wind turbine was constructed and operated in a field environment. The environment chosen was subject to wide variations in wind speed and direction and subsequently generated an extensive set of data for a variety of inflow test conditions. A single wind turbine blade was instrumented with pressure transducers and strain gauges with several data sets collected across a wide spectrum of typical and limiting wind turbine operating conditions. Surface pressure data taken at various spanwise locations along the blade demonstrated highly transient and spatially complex aerodynamic behavior for even the most basic operating conditions. Integrated normal force coefficient data showed enhanced lift values significantly beyond that predicted from steady-state two-dimensional wind-tunnel test data. Surface pressure data and integrated moment coefficient data suggested formation of coherent vortices consistent with the dynamic stall process observed in wind-tunnel tests for pitching wings. The unsteady, three-dimensional aerodynamic behavior for this wind turbine was then discussed and summarized.

Characterization and Laboratory Simulation of Turbine Airfoil Surface Roughness and Associated Heat Transfer

Abstract: The physical characteristics of surface roughness observed on first-stage high-pressure turbine vanes that had been in service for a long period were investigated in this study. Profilometry measurements were utilized to provide details of the surface roughness formed by deposits of foreign materials on different parts of the turbine vane. Typical measures of surface roughness such as centerline average roughness values were shown to be inadequate for characterizing roughness effects. Using a roughness shape parameter originally derived from regular roughness arrays, the turbine airfoil roughness was characterized in terms of equivalent sand-grain roughness in order to develop an appropriate simulation of the surface for laboratory experiments. Two rough surface test plates were designed and fabricated. These test plates were evaluated experimentally to quantify the heat transfer rate for flow conditions similar to that which occurs on the turbine airfoil. Although the roughness levels on the two test plates were different by a factor of two, both surfaces caused similar 50 percent increases in heat transfer rates relative to a smooth surface. The effects of high free-stream turbulence, with turbulence levels from 10 to 17 percent, were also investigated. Combined free-stream turbulence and surface roughness effects were found to be additive, resulting in as much as a 100 percent increase in heat transfer rate.

Performance Improvement Through Indexing of Turbine Airfoils: Part 1—Experimental Investigation

Abstract: This paper describes the results of an experimental/analytical study to determine the performance improvements achievable by circumferentially indexing successive rows of turbine stator airfoils. A series of tests was conducted at the National Aeronautics and Space Administration`s (NASA) Marshall Space Flight Center (MSFC) to experimentally investigate stator wake clocking effects on the performance of the Space Shuttle Main Engine Alternate Fuel Turbopump Turbine Test Article. Extensive time-accurate Computational Fluid Dynamics (CFD) simulations have been completed for the test configurations. The CFD results provide insight into the performance improvement mechanism. Part one of this paper describes details of the test facility, rig geometry, instrumentation, and aerodynamic operating parameters. Results of turbine testing at the aerodynamic design point are presented for six circumferential positions of the first stage stator, along with a description of the initial CFD analyses performed for the test article. It should be noted that first vane positions 1 and 6 produced identical first to second vane indexing. Results obtained from off-design testing of the best and worst stator clocking positions, and testing over a range of Reynolds numbers are also presented. Part two of this paper describes the numerical simulations performed in support of the experimental test programmore » described in part one. Time-accurate Navier-Stokes flow analyses have been completed for the five different turbine stator positions tested. Details of the computational procedure and results are presented. Analysis results include predictions of instantaneous and time-average midspan airfoil and turbine performance, as well as gas conditions throughout the flow field. An initial understanding of the turbine performance improvement mechanism is described.« less

Application of a Genetic Algorithm to Wind Turbine Design

Abstract: This paper presents an optimization procedure for stall-regulated horizontal-axis wind-turbines. A hybrid approach is used that combines the advantages of a genetic algorithm and an inverse design method. This method is used to determine the optimum blade pitch and blade chord and twist distributions that maximize the annual energy production. To illustrate the method, a family of 25 wind turbines was designed to examine the sensitivity of annual energy production to changes in the rotor blade length and peak rotor power. Trends are revealed that should aid in the design of new rotors for existing turbines. In the second application, a series of five wind turbines was designed to determine the benefits of specifically tailoring wind turbine blades for the average wind speed at a particular site. The results have important practical implications related to rotors designed for the Midwest versus those where the average wind speed may be greater.

Gas turbine blade platform cooling concept

Abstract: A turbine blade with a cooling air flow path specifically directed toward cooling the platform portion of the blade root. Cooling air passages are formed in the blade root platform just below its radially outward facing surface on an overhanging portion of the platform opposite the convex surface of the blade airfoil. Each of these passage extends radially outward from an inlet that receives a flow of cooling air, and then extends through the platform. Cavities are formed in a radially inward facing surface of an over hanging portion of the platform opposite the concave surface of the blade airfoil. An impingement plate directs cooling air as jets into these cavities. A passage is connected to the cavities and directs this cooling air through the overhanging portion of the platform opposite the concave surface.

Full Surface Local Heat Transfer Coefficient Measurements in a Model of an Integrally Cast Impingement Cooling Geometry

Abstract: Cast impingement cooling geometries offer the gas turbine designer higher structural integrity and improved convective cooling when compared to traditional impingement cooling systems, which rely on plate inserts. In this paper, it is shown that the surface that forms the jets contributes significantly to the total cooling. Local heat transfer coefficient distributions have been measured in a model of an engine wall cooling geometry using the transient heat transfer technique. The method employs temperature-sensitive liquid crystals to measure the surface temperature of large-scale perspex models during transient experiments. Full distributions of local Nusselt number on both surfaces of the impingement plate, and on the impingement target plate, are presented at engine representative Reynolds numbers. The relative effects of the impingement plate thermal boundary condition and the coolant supply temperature on the target plate heat transfer have been determined by maintaining an isothermal boundary condition at the impingement plate during the transient tests. The results are discussed in terms of the interpreted flow field.

Study of Electrochemical Jet Machining Process

Abstract: Jet Electrochemical Machining (ECJM) employs a jet of electrolyte for anodic dissolution of workpiece material. ECJM is extensively used for drilling small cooling holes in aircraft turbine blades and for producing maskless patterns for microelectronics parts. ECJM process drills small diameter holes and complex shape holes without the use of a profile electrode. One of the most significant problems facing ECJM user industries is the precise control of the process. A theoretical analysis of the process and a corresponding model are required for the development of an appropriate control system. This paper presents a mathematical model for determining the relationship between the machining rate and working conditions (electrolyte jet flow velocity, jet length, electrolyte properties, and voltage) of ECJM. This model describes a distribution of electric field and the effect of change of conductivity of electrolyte (caused by heating) on the process performance. A maximum dissolution rate is determined from the allowable increase of electrolyte temperature. Experimental verification of theoretical results is also presented.

Performance Improvement Through Indexing of Turbine Airfoils: Part 2—Numerical Simulation

Abstract: An experimental/analytical study has been conducted to determine the performance improvements achievable by circumferentially indexing succeeding rows of turbine stator airfoils. A series of tests was conducted to experimentally investigate stator wake clocking effects on the performance of the space shuttle main engine (SSME) alternate turbopump development (ATD) fuel turbine test article (TTA). The results from this study indicate that significant increases in stage efficiency can be attained through application of this airfoil clocking concept. Details of the experiment and its results are documented in part of this paper. In order to gain insight into the mechanisms of the performance improvement, extensive computational fluid dynamics (CFD) simulations were executed. The subject of the present paper is the initial results from the CFD investigation of the configurations and conditions detailed in part 1 of the paper. To characterize the aerodynamic environments in the experimental test series, two-dimensional, time-accurate, multistage, viscous analyses were performed at the TTA midspan. Computational analyses for five different circumferential positions of the first stage stator have been completed. Details of the computational procedure and the results are presented. The analytical results verify the experimentally demonstrated performance improvement and are compared with data whenever possible. Predictions of time-averagedmore » turbine efficiencies as well as gas conditions throughout the flow field are presented. An initial understanding of the turbine performance improvement mechanism based on the results from this investigation is described.« less

Unsteady Wake-Induced Boundary Layer Transition in High Lift LP Turbines

Abstract: The development of the unsteady suction side boundary layer of a highly loaded LP turbine blade has been investigated in a rectilinear cascade experiment. Upstream rotor wakes were simulated with a moving-bar wake generator. A variety of cases with different wake-passing frequencies, different wake strength and different Reynolds-numbers were tested. Boundary layer surveys have been obtained with a single hot-wire probe. Wall shear stress has been investigated with surface-mounted hot-film gauges. Losses have been measured.The suction surface boundary layer development of a modern highly loaded LP turbine blade is shown to be dominated by effects associated with unsteady wake-passing. Whereas without wakes the boundary layer features a large separation bubble at a typical cruise Reynolds-number, the bubble was largely suppressed if subjected to unsteady wake-passing at a typical frequency and wake strength. Transitional patches and becalmed regions, induced by the wake, dominated the boundary layer development. The becalmed regions inhibited transition and separation and are shown to reduce the loss of the wake-affected boundary layer.An optimum wake-passing frequency exists at cruise Reynolds-numbers. For a selected wake-passing frequency and wake-strength, the profile loss is almost independent of Reynolds-number. This demonstrates a potential to design highly loaded LP turbine profiles without suffering large losses at low Reynolds-numbers.Copyright © 1996 by ASME

Endwall heat transfer measurements in a transonic turbine cascade

Abstract: Turbine blade endwall heat transfer measurements are given for a range of Reynolds and Mach numbers. Data were obtained for Reynolds numbers based on inlet conditions of 0.5 and 1.0 x 106, for isentropic exit Mach numbers of 1.0 and 1.3, and for freestream turbulence intensities of 0.25% and 7.0%. Tests were conducted in a linear cascade at the NASA Lewis Transonic Turbine Blade Cascade Facility. The test article was a turbine rotor with 136' of turning and an axial chord of 12.7 cm. The large scale allowed for very detailed measurements of both flow field and surface phenomena. The intent of the work is to provide benchmark quality data for computational fluid dynamics (CFD) code and model verification. The flow field in the cascade is highly three-dimensional as a result of thick boundary layers at the test section inlet. Endwall heat transfer data were obtained using a steady-state liquid crystal technique.

Impact of Rotor–Stator Interaction on Turbine Blade Film Cooling

Abstract: The goal of this study is to quantify the impact of rotor-stator interaction on surface heat transfer of film cooled turbine blades. In Section 1, a steady-state injection model of the film cooling is incorporated into a two-dimensional, thin shear layer, multiblade row CFD code. This injection model accounts for the penetration and spreading of the coolant jet, as well as the entrainment of the boundary layer fluid by the coolant. The code is validated, in the steady state, by comparing its predictions to data from a blade tested in linear cascade. In Section 2, time-resolved film cooled turbine rotor heat transfer measurements are compared with numerical predictions. Data were taken on a fully film cooled blade in a transonic, high pressure ratio, single-stage turbine in a short duration turbine test facility, which simulates full-engine non-dimensional conditions. Film cooled heat flux on the pressure surface is predicted to be as much as a factor of two higher in the time average of the unsteady calculations compared to the steady-state case. Time-resolved film cooled heat transfer comparison of data to prediction at two spanwise positions is used to validate the numerical code. The unsteady stator-rotor interaction results in the pulsation ofmore » the coolant injection flow out of the film holes with large-scale fluctuations. The combination of pulsating coolant flow and the interaction of the coolant with this unsteady external flow is shown to lower the local pressure side adiabatic film effectiveness by as much as 64% when compared to the steady-state case.« less

A Novel Technique for the Internal Blade Cooling

Abstract: Development of an adequate air cooling system for the thermally highly loaded leading edge and tip of the blade, that is cost effective and also relatively insensitive to manufacturing tolerances and operating environment continues to be one of the major challenges in advanced gas turbine design. Extensive studies on the convective (including impingement) and film cooling techniques produced remarkable progress in achieving a high cooling effectiveness level for turbine airfoils. However, in the case of turbine blades, application of these techniques has severe limitations. Highly effective impingement cooling needs to be combined with film discharge of the spent air to avoid a negative impact of crossflow on internal heat transfer and also provide additional thermal protection of the surface downstream of the discharge holes. Noticeable aerodynamic penalties, stress concentration and significant increase in manufacturing cost limit application of blade film cooling, particularly for moderately high operating temperatures.Search for a highly effective robust design of internal airfoil cooling which can delay the use of film cooling resulted in the creation of a new technique which is described in this paper. This technique is based on generation of a swirling flow structure in the blade internal leading edge passage. Significant heat transfer augmentation can be achieved when the cooling air is delivered into the leading edge plenum tangentially to the inner concave surface. The best results can be obtained when the swirling flow is allowed to move radially, creating a three-dimensional screw-shaped flow in the plenum.The presented results of the flow and heat transfer studies performed for the practical range of Reynolds numbers for the internal flow show that the leading edge screw-shaped cooling technique provides internal heat transfer rate comparable with impingement coupled with film discharge of the spent air, is more effective than impingement with cross flow and is almost five times higher than heat transfer in the smooth channel.Copyright © 1996 by ASME

Heat Transfer Measurements to a Gas Turbine Cooling Passage With Inclined Ribs

Abstract: The local heat transfer coefficient distribution over all four walls of a large-scale model of a gas turbine cooling passage have been measured in great detail. A new method of determining the heat transfer coefficient to the rib surface has been developed and the contribution of the rib, at 5 percent blockage, to the overall roughened heat transfer coefficient was found to be considerable. The vortex-dominated flow field was interpreted from the detailed form of the measured local heat transfer contours. Computational Fluid Dynamics calculations support this model of the flow and yield friction factors that agree with measured values. Advances in the heat transfer measuring technique and data analysis procedure that confirm the accuracy of the transient method are described in full.

Method for forming an environmentally resistant blade tip

Abstract: A method for forming a blade tip on a turbine blade of a gas turbine engine includes providing a powder alloy from a family of environmentally-resistant powder alloys and forming the blade tip by melting and fusing the powder alloy to the turbine blade. The blade tip alloys preferably have a chemical composition of, in weight percent, about 14 to about 18 percent chromium, about 6.45 to about 6.95 percent aluminum, about 9.75 to about 11.45 percent cobalt, about 5.95 to about 6.55 percent tantalum, about 1.85 to about 2.35 percent rhenium, about 0.05 to about 1.75 percent hafnium, about 0,006 to about 0.03 percent zirconium, about 0.02 to about 0.11 percent carbon, up to about 1.1 percent silicon, up to about percent 0.01 percent boron, with the balance being nickel and typical impurities.

Three-Dimensional Inverse Method for Turbomachinery Blading Design

Abstract: An iterative procedure for 3D blade design is presented. The three-dimensional blade shape is modified using a physical algorithm, based on the transpiration model. The transpiration flux is computed by means of a modified Euler solver, in which the target pressure distribution is imposed along the blade surfaces. Only a small number of modifications is needed to obtain the final geometry.The method is based on a high resolution three-dimensional Euler solver. An upwind biased evaluation of the advective fluxes allows for a very low numerical entropy generation, and sharp shock capturing.The method is first validated, by redesigning an existing geometry, starting from a different one. It is further used to redesign a transonic compressor blade, to achieve, for the same mass flow and outlet flow angle, a shock free deceleration along the suction side. The last example concerns the design of a low aspect ratio turbine blade, with a positive compound lean to reduce the intensity of the passage vortices. The final blade is designed for an optimized pressure distribution, taking into account the forces resulting from the blade lean angle.Copyright © 1996 by ASME

Aero Engine Test Experience With CMSX-4® Alloy Single-Crystal Turbine Blades

Abstract: A team approach involving a turbine engine company (Rolls-Royce), its single-crystal casting facilities, and a superalloy developer and ingot manufacturer (Cannon-Muskegon), utilizing the concepts of simultaneous engineering, has been used to develop CMSX-4 alloy successfully for turbine blade applications. CMSX-4 alloy is a second-generation nickel-base single-crystal superalloy containing 3 percent (wt) rhenium (Re) and 70 percent volume fraction of the coherent γ' precipitate strengthening phase. Its finely balanced composition offers an attractive range of properties for turbine airfoil applications. In particular the alloy's combination of high strength in relation to creep-rupture, mechanical and thermal fatigue, good phase stability following extensive high temperature, stressed exposure and oxidation, hot corrosion and coating performance, are attractive for turbine engine applications where engine performance and turbine airfoil durability are of prime importance. The paper details the single-crystal casting process and heat treatment manufacturing development for turbine blades in CMSX-4 alloy. Competitive single-crystal casting yields are being achieved in production and extensive vacuum heat treatment experience confirms CMSX-4 alloy to have a practical production solution heat treat/homogenization window. The creep-rupture data-base on CMSX-4 alloy now includes 325 data points from 17 heats including 3630 kg (8000 lb) production size heats. An appreciable portion of this data was machined-from-blade (MFB) properties, which indicate turbine blade component capabilities based on single-crystal casting process, component configuration, and heat treatment. The use of hot isostatic pressing (HIP) has been shown to eliminate single-crystal casting micropores, which along with the essential absence of γ/γ' eutectic phase, carbides, stable oxide, nitride and sulfide inclusions, results in remarkably high mechanical fatigue properties, with smooth and particularly notched specimens. The Re addition has been shown not only to benefit creep and mechanical fatigue strength (with and without HIP), but also bare oxidation, hot corrosion (sulfidation), and coating performance. The high level of balanced properties determined by extensive laboratory evaluation has been confirmed during engine testing of the Rolls-Royce Pegasus turbofan.

Advances in Steam Path Technology

Abstract: For many years, GE has been conducting research to understand better the loss mechanisms that degrade the aerodynamic performance of steam turbine stages, and to develop new computational fluid dynamics (CFD) computer programs to predict these losses accurately. This paper describes a number of new steam path design features that have been introduced in the GE steam turbine product line to improve turbine performance and reliability. These features include diaphragms with contoured sidewalls, advanced vortex blading with compound tangential lean, new continuously coupled last-stage buckets with improved aerodynamic efficiency and reliability, improved downward and axial flow exhaust hoods, and better steam leakage control devices. The benefits of these new features for both new units and retrofits of existing units are discussed. In addition, the paper discusses the new generation of three-dimensional viscous CFD analysis codes being used to develop new design concepts, including codes developed by GE as well as those obtained externally. Also described are the extensive laboratory test programs being conducted to validate the CFD codes and verify the predicted efficiency gains for new design features. Last, the paper describes new and unique state-of-the-art steam path design automation and optimization tools that dramatically reduce the design cycle time for new advanced aerodynamic designs.

Measurements of Heat Transfer Coefficients and Friction Factors in Passages Rib-Roughened on All Walls

Abstract: A liquid crystal technique was used to measure heat transfer coefficients in twelve test sections with square and trapezoidal cross-sectional areas representing blade midchord cooling cavities in a modern gas turbine. Full-length ribs were configured on suction side as well as pressure side walls while half-length ribs were mounted on partition walls between adjacent cooling cavities. Ribs were in staggered arrangements with a nominal blockage ratio of 22 percent and an angle of attack to the mainstream flow, α, of 90 deg. Heat transfer measurements were performed on the roughened walls with full-length as well as half-length ribs. Nusselt numbers, friction factors, and thermal performances of all geometries are compared. The most important conclusion of this study is that the roughening of the partition walls enhances the heat transfer coefficients on those walls but, more importantly, enhances heat transfer coefficients on the primary walls considerably.

Turbulated cooling passages for turbine blades

Abstract: A turbine blade including an airfoil section having contoured surface geometries between the consecutive ribs or turbulators on the leading edge passage walls is described. The contoured surface geometries increase the overall heat transfer surface area between adjacent ribs as compared to the heat transfer surface area of a smooth inter-rib wall. The contoured surface geometries do not, however, change the overall serpentine passage geometry or the large scale flow characteristics of the turbulated cooling passageways. The contoured surface geometries may have many different geometries (e.g., triangular, conical, semi-cylindrical, cylindrical columns or indentations in the side wall such as semi-circular dimples).

RENl? N4: A FIRST GENERATION SINGLE CRYSTAL TURBINE AIRFOIL ALLOY WITH IMPROVED OXIDATION RESISTANCE, LOW ANGLE BOUNDARY STRENGTH AND SUPERIOR LONG TIME RUPTURE STRENGTH

Abstract: Introduction GE Aircraft Engine’s first generation single crystal (SX) turbine airfoil alloy, Rene N,(l) was extensively tested and utilized in turbine blades of development engines. Factory engine testing then revealed that tips of the Rene N blades suffered from excessive oxidation. It was obvious that a more oxidation resistant single crystal alloy would be required. Two SX development programs were conducted; C. Wukusick and W. King to develop a more oxidation resistant SX alloy, and E. Ross to increase the low angle boundary (LAB) strength of the RenC N, which could, hopefully, be applied to the new more oxidation resistant SX alloy. This paper will describe the development of this improved first generation SX alloy, Rent N4,(2) which evolved from these two programs.

An Experimental Study of Turbine Vane Heat Transfer With Water–Air Cooling

Abstract: This paper presents data showing the improvement in cooling effectiveness of turbine vanes through the application of water-air cooling technology in an industrial/utility engine application. The technique utilizes a finely dispersed water-in-air mixture that impinges on the internal surfaces of turbine airfoils to produce very high cooling rates. An airfoil was designed to contain a standard impingement tube, which distributes the water-air mixture over the inner surface of the airfoil. The water flash vaporizes off the airfoil inner wall. The resulting mixture of air-steam-water droplets is then routed through a pin fin array in the trailing edge region of the airfoil where additional water is vaporized. The mixture then exits the airfoil into the gas path through trailing edge slots. Experimental measurements were made in a three-vane, linear, two-dimensional cascade. The principal independent parameters--Mach number, Reynolds number, wall-to-gas temperature ratio, and coolant-to-gas mass flow ratio--were maintained over ranges consistent with typical engine conditions. Five impingement tubes were utilized to study geometry scaling, impingement tube-to-airfoil wall gap spacing, impingement tube hole diameter, and impingement tube hole patterns. The test matrix was structured to provide an assessment of the independent influence of parameters of interest, namely, exit Mach number, exit Reynolds number,more » gas-to-coolant temperature ratio, water- and air-coolant-to-gas mass flow ratios, and impingement tube geometry. Heat transfer effectiveness data obtained in this program demonstrated that overall cooling levels typical for air-cooled vanes could be achieved with the water-air cooling technique with reductions of cooling air flow of significantly more than 50%.« less

Coupled Shaft-Torsion and Blade-Bending Vibrations of a Rotating Shaft–Disk–Blade Unit

Abstract: A new approach to analyzing the dynamic coupling between shaft torsion and blade bending of a rotating shaft-disk-blade unit is introduced. The approach allows the shaft to vibrate freely around its rotation axis instead of assuming a periodic perturbation of the shaft speed that may accommodate the shaft flexibility only to a limited extent. A weighted residual method is applied, and the receptances at the connections of blades and shaft-disk are formulated. Numerical examples are given for cases with between two and six symmetrically arranged blades. The results show not only coupling between the shaft, disk, and blades, but also coupling between individual blades where the shaft acts as a rigid support and experiences no torsional vibration. The blade-coupling modes occurred only in repeated frequencies. Finally, the effect of shaft speed on the modal frequencies was investigated. Plots illustrating the occurrence of critical speeds and flutter instabilities are presented.

The Effect of Vortex Shedding on the Unsteady Pressure Distribution Around the Trailing Edge of a Turbine Blade

Abstract: The wakes behind turbine blade trailing edge are characterized by large scale periodic vortex patterns known as the von Karman vortex street. The failure of steady-state Navier-Stokes calculations in modeling wake flows appears to be mainly due to ignoring this type of flow instabilities. In an effort to contribute to a better understanding of the time varying wake flow characteristics behind turbine blades, VKI has performed large scale turbine cascade tests to obtain very detailed information about the steady and unsteady pressure distribution around the trailing edge of a nozzle guide vane. Tests are run at an outlet Mach number of M2,is,=0.4 and a Reynolds number of Rec = 2·106. The key to the high spatial resolution of the pressure distribution around the trailing edge is a rotatable trailing edge with an embedded miniature pressure transducer underneath the surface and a pressure slot opening of about 1.5° of the trailing edge circle. Signal processing allowed or differentiation between random and periodic pressure fluctuations. Ultra-short schlieren pictures help in understanding the physics behind the pressure distribution.Copyright © 1996 by ASME

Sealing apparatus for airfoils of gas turbine engines

Abstract: An improved airfoil tip sealing apparatus is disclosed wherein brush seals are attached to airfoil tips with the distal ends of the brush seal fibers sealingly contacting opposing wall surfaces. Embodiments for variable vanes, stators and both cooled and uncooled turbine blade applications are disclosed.

Criteria for analysis of abradable coatings

Abstract: Performance of turbine blades and abradable seals during rubbing is usually evaluated by rig tests. When rubbing occurs at high incursion rates, performance of the blade/seal during rubbing can also be analyzed using the elastic modulus and/or the yield stress criteria. The analysis of rubbing performance based on the criteria was composed of three steps. Firstly, a thermal model was developed to calculated the temperature changes during blade and seal interaction. Secondly, high temperature elastic moduli and/or yield stresses of blade and seal materials were examined. Finally, rubbing performance of blade and seal was evaluated by considering the temperature distribution and the high temperature mechanical properties. The analysis was carried out for rubbing processes between titanium blades and the NiCrAl-silicate seal, as well as titanium blades and the AlSi seal. The results of this analysis were in agreement with those obtained by the high speed and high incursion rate rubbing tests.

Internal radiation effects in zirconia thermal barrier coatings

Abstract: Using thermal barrier coatings on combustor liners, turbine vanes, and rotating blades is important for reducing metal temperatures in current and advanced aircraft engines. Zirconia is a common coating material, and it is partially transparent to thermal radiation. Radiation becomes more significant as temperatures are raised for higher efficiency in advanced engines. Calculations are often made with radiation effects neglected inside the coating. The effect of radiation is illustrated here, where an analytical procedure is provided by using the two-flux method for the radiative contribution. A detailed study was made of ceramic thermal barrier coatings for diesel engines, and a two-flux analysis was developed for radiation in semitransparent multilayer composites. These efforts provide the basis for the present analysis where illustrative solutions are obtained for typical conditions in an aircraft engine. The formulation and solution of the exact spectral radiative transfer equations including large scattering, as is characteristic of zirconia, are rather complicated. The two-flux method is used here to provide a simplified method.

Unsteady numerical simulations of radial temperature profile redistribution in a single-stage turbine

Abstract: Experimental data taken from gas turbine combustors indicate that the flow exiting the combustor can contain both circumferential and radial temperature gradients. A significant amount of research recently has been devoted to studying turbine flows with inlet temperature gradients, but no total pressure gradients. Less attention has been given to flows containing both temperature and total pressure gradients at the inlet. The significance of the total pressure gradients is that the secondary flows and the temperature redistribution process in the vane blade row can be significantly altered. Experimental data previously obtained in a single-stage turbine with inlet total temperature and total pressure gradients indicated a redistribution of the warmer fluid to the pressure surface of the airfoils, and a severe underturning of the flow at the exit of the stage. In a concurrent numerical simulation, a steady, inviscid, three-dimensional flow analysis was able to capture the redistribution process, but not the exit flow angle distribution. In the current research program, a series of unsteady two- and three-dimensional Navier-Stokes simulations have been performed to study the redistribution of the radial temperature profile in the turbine stage. The three-dimensional analysis predicts both the temperature redistribution and the flow underturning observed in the more » experiments. « less