Dissertations / Theses: 'Turbomachinery aerodynamics'

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Gostelow, J. P. "Publications in turbomachinery aerodynamics and related fields." Thesis, University of Liverpool, 1987. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.384347.

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He, Li. "Unsteady flows around oscillating turbomachinery blades." Thesis, University of Cambridge, 1990. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.385407.

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Ning, Wei. "Computation of unsteady flow in turbomachinery." Thesis, Durham University, 1998. http://etheses.dur.ac.uk/4819/.

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Unsteady flow analysis has been gradually introduced in turbomachinery design systems to improve machine performance and structural integrity. A project on computation of unsteady flows in turbomachinery has been carried out. A quasi 3-D time-linearized Euler/Navier-Stokes method has been developed for unsteady flows induced by the blade oscillation and unsteady incoming wakes, hi this method, the unsteady flow is decomposed into a steady flow plus a harmonically varying unsteady perturbation. The coefficients of the linear perturbation equation are formed from steady flow solutions. A pseudo-time is introduced to make both the steady flow equation and the linear unsteady perturbation equation time-independent. The 4-stage Runge-Kutta time-marching scheme is implemented for the temporal integration and a cell-vertex scheme is used for the spatial discretization. A 1-D/2-D nonreflecting boundary condition is applied to prevent spurious reflections of outgoing waves when solving the perturbation equations. The viscosity in the unsteady Navier- Stokes perturbation equation is frozen to its steady value. The present time-linearized Euler/Navier-Stokes method has been extensively validated against other well- developed linear methods, nonlinear time-marching methods and experimental data. Based upon the time-linearized method, a novel quasi 3-D nonlinear harmonic Euler/Navier-Stokes method has been developed. In this method, the unsteady flow is divided into a time-averaged flow plus an unsteady perturbation. Time-averaging produces extra nonlinear "unsteady stress" terras in the time-averaged equations and these extra terras are evaluated from unsteady perturbations. Unsteady perturbations are obtained by solving a first order harraonic perturbation equation, while the coefficients of the perturbation equation are forraed from time-averaged solutions. A strong coupling procedure is applied to solve the time-averaged equation and the unsteady perturbation equation simultaneously in a pseudo-time domain. An approximate approach is used to linearize the pressure sensors in artificial smoothing

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Melzer, Andrew Philip. "Aerodynamics of transonic turbine trailing edges." Thesis, University of Cambridge, 2018. https://www.repository.cam.ac.uk/handle/1810/276280.

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5

Tsay, W. C. "The analysis and design methods for turbomachinery flows." Thesis, Cranfield University, 1989. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.233928.

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6

Brock, Jerry S. "A modified Baldwin-Lomax turbulence model for turbomachinery wakes." Thesis, This resource online, 1991. http://scholar.lib.vt.edu/theses/available/etd-09052009-040231/.

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7

Knapke, Clint J. "Aerodynamics of Fan Blade Blending." Wright State University / OhioLINK, 2019. http://rave.ohiolink.edu/etdc/view?acc_num=wright1567517259599736.

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8

Forhad, Md Moinul Islam. "Robustness analysis for turbomachinery stall flutter." Master's thesis, University of Central Florida, 2011. http://digital.library.ucf.edu/cdm/ref/collection/ETD/id/4894.

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As compared with other robustness analysis tools, such as Hsubscript inf], the Mu analysis is less conservative and can handle both structured and unstructured perturbations. Finally, Genetic Algorithm is used as an optimization tool to find ideal parameters that will ensure best performance in terms of damping out flutter. Simulation results show that the procedure described in this thesis can be effective in studying the flutter stability margin and can be used to guide the gas turbine blade design.; Flutter is an aeroelastic instability phenomenon that can result either in serious damage or complete destruction of a gas turbine blade structure due to high cycle fatigue. Although 90% of potential high cycle fatigue occurrences are uncovered during engine development, the remaining 10% stand for one third of the total engine development costs. Field experience has shown that during the last decades as much as 46% of fighter aircrafts were not mission-capable in certain periods due to high cycle fatigue related mishaps. To assure a reliable and safe operation, potential for blade flutter must be eliminated from the turbomachinery stages. However, even the most computationally intensive higher order models of today are not able to predict flutter accurately. Moreover, there are uncertainties in the operational environment, and gas turbine parts degrade over time due to fouling, erosion and corrosion resulting in parametric uncertainties. Therefore, it is essential to design engines that are robust with respect to the possible uncertainties. In this thesis, the robustness of an axial compressor blade design is studied with respect to parametric uncertainties through the Mu analysis. The nominal flutter model is adopted from (9). This model was derived by matching a two dimensional incompressible flow field across the flexible rotor and the rigid stator. The aerodynamic load on the blade is derived via the control volume analysis. For use in the Mu analysis, first the model originally described by a set of partial differential equations is reduced to ordinary differential equations by the Fourier series based collocation method. After that, the nominal model is obtained by linearizing the achieved non-linear ordinary differential equations. The uncertainties coming from the modeling assumptions and imperfectly known parameters and coefficients are all modeled as parametric uncertainties through the Monte Carlo simulation.
ID: 030423207; System requirements: World Wide Web browser and PDF reader.; Mode of access: World Wide Web.; Thesis (M.S.)--University of Central Florida, 2011.; Includes bibliographical references (p. 44-47).
M.S.
Masters
Mechanical, Materials, and Aerospace Engineering
Engineering and Computer Science

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Sharpe, Jacob Andrew. "3D CFD Investigation of Low Pressure Turbine Aerodynamics." Wright State University / OhioLINK, 2017. http://rave.ohiolink.edu/etdc/view?acc_num=wright1495872867696744.

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10

Chernysheva, Olga V. "Flutter in sectored turbine vanes." Doctoral thesis, KTH, Energy Technology, 2004. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-3737.

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In order to eliminate or reduce vibration problems inturbomachines without a high increase in the complexity of thevibratory behavior, the adjacent airfoils around the wheel areoften mechanically connected together with lacing wires, tip orpart-span shrouds in a number of identical sectors. Although anaerodynamic stabilizing effect of tying airfoils together ingroups on the whole cascade is indicated by numerical andexperimental studies, for some operating conditions suchsectored vane cascade can still remain unstable.

The goal of the present work is to investigate thepossibilities of a sectored vane cascade to undergoself-excited vibrations or flutter. The presented method forpredicting the aerodynamic response of a sectored vane cascadeis based on the aerodynamic work influence coefficientrepresentation of freestanding blade cascade. The sectored vaneanalysis assumes that the vibration frequency is the same forall blades in the sectored vane, while the vibration amplitudesand mode shapes can be different for each individual blade inthe sector. Additionally, the vibration frequency as well asthe amplitudes and mode shapes are supposed to be known.

The aerodynamic analysis of freestanding blade cascade isperformed with twodimensional inviscid linearized flow model.As far as feasible the study is supported by non-linear flowmodel analysis as well as by performing comparisons againstavailable experimental data in order to minimize theuncertainties of the numerical modeling on the physicalconclusions of the study.

As has been shown for the freestanding low-pressure turbineblade, the blade mode shape gives an important contributioninto the aerodynamic stability of the cascade. During thepreliminary design, it has been recommended to take intoaccount the mode shape as well rather than only reducedfrequency. In the present work further investigation using foursignificantly different turbine geometries makes these findingsmore general, independent from the low-pressure turbine bladegeometry. The investigation also continues towards a sectoredvane cascade. A parametrical analysis summarizing the effect ofthe reduced frequency and real sector mode shape is carried outfor a low-pressure sectored vane cascade for differentvibration amplitude distributions between the airfoils in thesector as well as different numbers of the airfoils in thesector. Critical (towards flutter) reduced frequency maps areprovided for torsion- and bending-dominated sectored vane modeshapes. Utilizing such maps at the early design stages helps toimprove the aerodynamic stability of low-pressure sectoredvanes.

A special emphasis in the present work is put on theimportance for the chosen unsteady inviscid flow model to bewell-posed during numerical calculations. The necessity for thecorrect simulation of the far-field boundary conditions indefining the stability margin of the blade rows isdemonstrated. Existing and new-developed boundary conditionsare described. It is shown that the result of numerical flowcalculations is dependent more on the quality of boundaryconditions, and less on the physical extension of thecomputational domain. Keywords: Turbomachinery, Aerodynamics,Unsteady CFD, Design, Flutter, Low-Pressure Turbine, Blade ModeShape, Critical Reduced Frequency, Sectored Vane Mode Shape,Vibration Amplitude Distribution, Far-field 2D Non-ReflectingBoundary Conditions. omain.

Keywords:Turbomachinery, Aerodynamics, Unsteady CFD,Design, Flutter, Low-Pressure Turbine, Blade Mode Shape,Critical Reduced Frequency, Sectored Vane Mode Shape, VibrationAmplitude Distribution, Far-field 2D Non-Reflecting BoundaryConditions.

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Heinlein, Gregory S. "Aerodynamic Behavior of Axial Flow Turbomachinery Operating in Transient Transonic Flow Regimes." The Ohio State University, 2019. http://rave.ohiolink.edu/etdc/view?acc_num=osu1573149943024303.

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12

Kurtulus, Berkin. "Development of a Tool for Inverse Aerodynamic Design and Optimisation of Turbomachinery Aerofoils." Thesis, KTH, Flygdynamik, 2021. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-293353.

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The automation of airfoil design process is an ongoing effort within the field of turbo-machinery design, with significant focus on developing new reliable and consistent methods that can meet the needs of the engineers. A wide variety of approaches has been in use for inverse airfoil design process which benefit from theoretical inverse design, statistical methods, empirical discoveries and many other ways to solve the design problem. This thesis work also develops a tool in Python to be used in airfoil aerodynamic design process that is simple, fast and accurate enough for initial design of turbo-machinery blades with focus on turbine airfoils used for operation in aircraft engines. To convey the decision-making process during development a simplified case is presented. The underlying considerations are discussed. Other available methods in the literature used for similar problems, are also evaluated and compared to demonstrate the advantages and limitations of the methods used within the tool. The inverse design problem is formulated as a multi-objective optimization problem to handle various different objectives that are relevant for aerodynamic design of turbo-machinery airfoils. Test runs are made and the results are discussed to assess how robust the tool is and how the current capabilities can be modified or extended. After the development process, the tool is verified to be a suitable option for real-life design optimization tasks and can be used as a building block for a much more comprehensive tool that may be developed in the future.
Automatisering av processen för design av vingprofiler kräver fortlöpande insatser inom området turbomaskindesign, med stort fokus på att utveckla nya tillförlitliga och konsekventa metoder som kan tillgodose ingenjörernas behov. Ett stort antal olika tillvägagångssätt har provats för omvänd design av vingprofiler såsom teoretisk invers design, statistiska metoder, empiriska upptäckter och många andra sätt att lösa designproblemet. Detta avhandlingsarbete är också ett lyckat försök att utveckla ett verktyg i Python som ska användas i den aerodynamiska designprocessen; det är enkelt, snabbt och noggrant för den initiala designen av turbomaskinblad med fokus på turbinblad som för användning i flygmotorer. För att förmedla beslutsprocessen under utvecklingen presenteras ett förenklat fall. De underliggande övervägandena diskuteras. Andra tillgängliga metoder i litteraturen som används för liknande problem utvärderas och jämförs för att visa fördelarna och begränsningarna med de metoder som används i verktyget. Det omvända designproblemet formuleras som ett multi-objektivt optimeringsproblem för att hantera olika mål som är relevanta för aerodynamisk design av turbomaskiner. Testkörningar görs och resultaten diskuteras för att bedöma hur robust verktyget är och hur de nuvarande funktionerna kan modifieras eller utökas. Efter utvecklingsprocessen verifieras verktyget som ett lämpligt alternativ för verkliga designoptimeringsuppgifter och kan användas som en byggsten för ett mycket mer omfattande verktyg som kan utvecklas i framtiden.

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13

Sheard, A. G. "Aerodynamic and mechanical performance of a high-pressure turbine stage in a transient wind tunnel." Thesis, University of Oxford, 1989. http://ora.ox.ac.uk/objects/uuid:73ecb15e-efde-474d-ae30-3f8f7e1d6f4e.

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Unsteady three-dimensional flow phenomena have major effects on the aerodynamic performance of, and heat transfer to, gas-turbine blading. Investigation of the mechanisms associated with these phenomena requires an experimental facility that is capable of simulating a gas turbine, but at lower levels of temperature and pressure to allow conventional measurement techniques. This thesis reports on the design, development and commissioning of a new experimental facility that models these unsteady three-dimensional flow phenomena. The new facility, which consists of a 62%-size, high-pressure gas-turbine stage mounted in a transient wind tunnel, simulates the turbine design point of a full-stage turbine. The thesis describes the aerodynamic and mechanical design of the new facility, a rigorous stress analysis of the facility’s rotating system and the three-stage commissioning of the facility. The thesis concludes with an assessment of the turbine stage performance.

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Flegel, Ashlie Brynn. "Aerodynamic Measurements of a Variable-Speed Power-Turbine Blade Section in a Transonic Turbine Cascade." Cleveland State University / OhioLINK, 2013. http://rave.ohiolink.edu/etdc/view?acc_num=csu1387437733.

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15

DiPietro, Anthony Louis. "Design and experimental evaluation of a dynamic thermal distortion generator for turbomachinery research." Thesis, This resource online, 1993. http://scholar.lib.vt.edu/theses/available/etd-09292009-020206/.

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16

Zhang, Luying. "Rotating instability on steam turbine blades at part-load conditions." Thesis, University of Oxford, 2013. http://ora.ox.ac.uk/objects/uuid:cf8ecad1-0fd2-49b7-8e28-6d00c62c173e.

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A computational study aimed at improving the understanding of rotating instability in the LP steam turbine last stage working under low flow rate conditions is described in this thesis. A numerical simulation framework has been developed to investigate into the instability flow field. Two LP model turbine stages are studied under various flow rate conditions. By using the 2D simulations as reference and comparing the results to those of the 3D simulations, the basic physical mechanism of rotating instability is analysed. The pressure ratio characteristics across the rotor row tip are found to be crucial to the inception of rotating instability. The captured instability demonstrates a 2D mechanism based on the circumferential variation of unsteady separation flow in the rotor row. The 3D tip clearance flow is found not a necessary cause of the instability onset. Several influential parameters on the instability flow are also investigated by a set of detailed studies on different turbine configurations. The results show that the instability flow pattern and characteristics can be altered by the gap distance between the stator and rotor row, the rotor blading and the stator row stagger angle. Some flow control approaches are proposed based on the observations, which may also serve as design reference. The tip region 3D vortex flow upstream to the rotor row is also captured by the simulations under low flow rate conditions. Its appearance is found to be able to suppress the inception of rotating instability by disrupting the interaction between the rotor separation flow and the incoming flow. Finally, some recommendations for further work are proposed.

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Wilde, Daniel G. "Validation of a CFD Approach for Gas Turbine Internal Cooling Passage Heat Transfer Prediction." DigitalCommons@CalPoly, 2015. https://digitalcommons.calpoly.edu/theses/1384.

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This report describes the development and application of a validated Computational Fluid Dynamics (CFD) modelling approach for internal cooling passages in rotating turbomachinery. A CFD Modelling approach and accompanying assumptions are tuned and validated against academically available experimental results for various serpentine passages. Criteria of the CFD modelling approach selected for investigation into advanced internal cooling flows include accuracy, robustness, industry familiarity, and computational cost. Experimental data from NASA HOST (HOt Section Technology), Texas A&M, and University of Manchester tests are compared to RANS CFD results generated using Fluent v14.5 in order to benchmark a CFD modelling approach. Capability of various turbulence models in the representation of cooling physics is evaluated against experimental data. Model sensitivity to boundary conditions and mesh density is also evaluated. The development of a validated computational model of internal turbine cooling channels with bounded error allows for the identification of particular shortcomings of heat transfer correlations and provides a baseline for future CFD based exploration of internal turbine cooling concepts.

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Fawcett, Richard James. "Coherent unsteadiness in film cooling." Thesis, University of Oxford, 2011. http://ora.ox.ac.uk/objects/uuid:57ec3da6-4946-4f66-8421-b01d53d7e0fc.

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Film cooling is vital for the cooling of the blades and vanes in the high temperature environment of a jet engine high pressure turbine stage. Previous research into film cooling has typically concentrated on its time-mean performance. However, results from other studies upon more simplified geometries, suggest that coherent unsteadiness is likely to also be present in film cooling flows. The research presented in this thesis, therefore, aims to characterise what coherent unsteadiness, if any, is present within film cooling flows. Cylindrical and shaped cooling holes, located upon the pressure surface of a turbine blade within a large scale linear cascade, have been investigated. A blowing ratio range of 0.5 to 2.0 has been investigated, with either a plenum or perpendicular crossflow at the cooling hole inlet. Particle Image Velocimetry, high speed photography and Hot Wire Anemometry have been used to investigate the jet downstream of both cooling holes. The impact of crossflow at the hole inlet upon the flowfield inside both cooling holes has been investigated using Hot Wire Anemometry and a further numerical model solved by applying TBLOCK. The results presented in the current thesis, show the existence of two coherent unsteady structures in the jet downstream of both the cylindrical and the shaped holes. These structures are called shear layer vortices and hairpin vortices, and their formation is dependent on the velocity difference across the jet shear layer. Inside the cooling hole coherent hairpin vortices also appear to occur, with their formation dependent on the direction and magnitude of the crossflow at the hole inlet. The coherent unsteadiness presented here is shown for the first time for film cooling flows, and recommendations to build on the current study, in what is potentially an interesting research area, are made at the end of this thesis.

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Tang, Brian M. T. "Unshrouded turbine blade tip heat transfer and film cooling." Thesis, University of Oxford, 2011. http://ora.ox.ac.uk/objects/uuid:f8479e89-9cd1-4aa7-b5c8-8068ad80de54.

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This thesis presents a joint computational and experimental investigation into the heat transfer to unshrouded turbine blade tips suitable for use in high bypass ratio, large civil aviation turbofan engines. Both the heat transfer to the blade tip and the over-tip leakage flow over the blade tip are characterised, as each has a profound influence on overall engine efficiency. The study is divided into two sections; in the first, computational simulations of a very large scale, low speed linear cascade with a flat blade tip were conducted. These simulations were validated against experimental data collected by Palafox (2006). A thorough assessment of turbulence models and minimum meshing requirements was performed. The standard k-ω and standard k-ϵ turbulence models significantly overpredicted the turbulence levels within the tip gap. The other models were very similar in performance; the SST k-ω and realisable k-ϵ models were found to be the most suitable for the flow environment. The second section documents the development and testing of a novel hybrid blade tip design, the squealet tip, which seeks to combine the known benefits of winglet and double squealer tips. The development of the external geometry was performed primarily through engine-representative CFD simulations at a range of tip gaps from 0.45% to 1.34% blade chord. The squealet tip was found to have a similar aerodynamic sensitivity to tip clearance as a baseline double squealer tip, with a tip gap efficiency exchange rate of 2.03, although this was 18% greater than the alternative winglet tip. The squealet tip displayed higher predicted stage efficiency than the winglet tip over the majority of the range of tip clearances investigated, however. The overall heat load was reduced by 14% compared with the winglet tip but increased by 28% over the double squealer tip, primarily due to the change in wetted surface area. The predicted local heat transfer coefficients were similar across all geometries. A realistic internal cooling plenum and an array of blade tip cooling holes were subsequently added to the squealet tip geometry and the cooling configuration refined by the selective sealing of cooling holes. Film cooling performance was largely assessed by the predicted adiabatic wall temperature distributions. A viable cooling scheme which reduced the cooling air requirement by 38% was achieved, compared to the initial case which had all cooling holes open. This was associated with just a 7% increase in blade tip heat flux and no penalty in peak temperature on the blade tip. Film cooling air ejected from holes on the blade suction side was swept away from the blade tip region, making the squealet rim at the crown of the blade particularly challenging to cool. It was demonstrated that this region could be cooled effectively by ballistic cooling from holes located on the blade tip cavity floor, although this was expensive in terms of the mass flow rate of cooling air required. The computational results were reinforced with experimental data collected in a transonic linear cascade. Downstream aerodynamic loss measurements were taken for a linearised version of the squealet tip design without cooling at nominal tip gaps of 0.45%, 0.89% and 1.34% blade chord, which was compared to similar data taken by O’Dowd (2010) for flat and winglet tips. The squealet was seen to have a similar aerodynamic loss to the flat tip and a reduced loss compared with the winglet tip. Full surface heat transfer measurements were taken for the uncooled squealet tip, at tip gaps of 0.89% and 1.34% blade chord, and for two configurations of the cooled squealet tip, at a tip clearance of 0.89% blade chord. The qualitative similarity between the measured heat transfer distributions and the those predicted by the engine-representative CFD simulations was good. A CFD simulation of the uncooled linear cascade environment at the 1.34% blade chord tip clearance was performed using a single blade with translationally periodic boundary conditions. The predicted size of the over-tip leakage vortex was smaller than had been measured, resulting in a large underprediction in the magnitude of the downstream area-averaged aerodynamic loss. The magnitudes of the predicted blade tip Nusselt number distribution were similar to those produced by the engine-representative CFD simulations and lower than that measured experimentally. Differences in the shape of the Nusselt number distribution were observed in the vicinity of regions of separated and reattaching flow, but other salient features were replicated in the computational data. The squealet tip has been shown to be a promising, viable unshrouded blade tip design with an aerodynamic performance similar to the double squealer tip but is more amenable to film cooling. It is significantly lighter than a winglet tip and incurs a reduced thermal load. The squealet tip design can now be developed into a blade tip geometry for use in real engines to provide an alternative to shrouded turbine blades and current unshrouded blade tip designs. A commercial CFD solver, Fluent 6.3, was shown to capture blade tip heat transfer and over-tip leakage flow sufficiently well to be a useful design guide. However, the sensitivity of the flow structure (and hence, heat transfer) in the forward part of the blade tip cavity suggests that physical testing cannot be eliminated from the design process entirely.

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Sanz, Luengo Antonio. "Experimental Investigation of the Influence of Local Flow Features on the Aerodynamic Damping of an Oscillating Blade Row." Licentiate thesis, KTH, Kraft- och värmeteknologi, 2014. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-145179.

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The general trend of efficiency increase, weight and noise reduction has derived in the design of more slender, loaded, and 3D shaped blades. This has a significant impact on the stability of fan, and low pressure turbine blades, which are more prone to aeroelastic phenomena such as flutter. The flutter phenomenon is a self-excited, self-sustained unstable vibration produced by the interaction of flow and structure. These working conditions will induce either blade overload, or High Cycle Fatigue (HCF) produced by Limited Cycle Oscillation (LCO). The main objectives of the present work are on the investigation of the aeroelastic properties of a high-lift low-pressure in the light of the local flow features present in such profiles, in nominal and extreme off-design conditions both in high and low subsonic Mach number, for three dif-ferent rigid body modes. In addition, the validity of the linearity assump-tion of the influence coefficient technique has also been investigated, in order to expand the understanding of the physical limits of this assumption. This work has been designed as experimental investigation in the influence coefficient domain focused on a high-lift low-pressure turbine designed by ITP within the framework of the European FP7 project FU-TURE. These experiments have been carried out in the Aeroelastic test rig (AETR), at KTH Stockholm, which consist of an instrumented annular sector cascade with a single oscillating blade. The results acquired have been supported by numerical results provided by a non-propietary commercial software package (ANSYS CFX). The results suggest that the typical three-dimensional effects associated secondary flow features and tip leakage flows have a significant influence on the aeroelastic performance and the cascade stability. However the major influence appears as a consequence of the separation surface on the pressure side which appears at extreme off-design operating conditions. The contribution to stability of this local feature depend on the oscillation mode showing for the axial and torsion mode a neutral stability contribution, which is directly associated with the geometrical properties of the cascade. However, on the circumferential mode this separation surface has a stabilizing effect much more independent of the blade geometry. The study of the linearity assumption of the influence coefficient domain has revealed, that an apparent linear relation between the integrated unsteady response and the vibrational amplitude, does not necessary imply that the local unsteady response is linear with respect to the oscillation amplitude. The results also suggest that the validity of the linearity as-sumption is more sensitive to high oscillation amplitudes at high Mach conditions.

QC 20140609

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Saroch, Michael F. (Michael Frank) Carleton University Dissertation Engineering Mechanical and Aerospace. "Contributions to the study of turbomachinery aerodynamics; Part I: Design of a fish-tail diffuser test section, Part II: Computations of the effects of AVDR on transonic turbine cascades." Ottawa, 1996.

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Jöcker, Markus. "Numerical Investigation of the Aerodynamic Vibration Excitation of High-Pressure Turbine Rotors." Doctoral thesis, KTH, Energy Technology, 2002. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-3416.

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The design parameters axial gap and stator count of highpressure turbine stages are evaluated numerically towards theirinfluence on the unsteady aerodynamic excitation of rotorblades. Of particular interest is if and how unsteadyaerodynamic considerations in the design could reduce the riskofhigh cycle fatigue (HCF) failures of the turbine rotor.

A well-documented 2D/Q3D non-linear unsteady code (UNSFLO)is chosen to perform the stage flow analyses. The evaluatedresults are interpreted as aerodynamic excitation mechanisms onstream sheets neglecting 3D effects. Mesh studies andvalidations against measurements and 3D computations provideconfidence in the unsteady results. Three test cases areanalysed. First, a typical aero-engine high pressure turbinestage is studied at subsonic and transonic flow conditions,with four axial gaps (37% - 52% of cax,rotor) and two statorconfigurations (43 and 70 NGV). Operating conditions areaccording to the resonant conditions of the blades used inaccompanied experiments. Second, a subsonic high pressureturbine intended to drive the turbopump of a rocket engine isinvestigated. Four axial gap variations (10% - 29% ofcax,rotor) and three stator geometry variations are analysed toextend and generalise the findings made on the first study.Third, a transonic low pressure turbine rotor, known as theInternational Standard Configuration 11, has been modelled tocompute the unsteady flow due to blade vibration and comparedto available experimental data.

Excitation mechanisms due to shock, potential waves andwakes are described and related to the work found in the openliterature. The strength of shock excitation leads to increasedpressure excitation levels by a factor 2 to 3 compared tosubsonic cases. Potential excitations are of a typical wavetype in all cases, differences in the propagation direction ofthe waves and the wave reflection pattern in the rotor passagelead to modifications in the time and space resolved unsteadypressures on the blade surface. The significant influence ofoperating conditions, axial gap and stator size on the wavepropagation is discussed on chosen cases. The wake influence onthe rotorblade unsteady pressure is small in the presentevaluations, which is explicitly demonstrated on the turbopumpturbine by a parametric study of wake and potentialexcitations. A reduction in stator size (towards R≈1)reduces the potential excitation part so that wake andpotential excitation approach in their magnitude.

Potentials to reduce the risk of HCF excitation in transonicflow are the decrease of stator exit Mach number and themodification of temporal relations between shock and potentialexcitation events. A similar temporal tuning of wake excitationto shock excitation appears not efficient because of the smallwake excitation contribution. The increase of axial gap doesnot necessarily decrease the shock excitation strength neitherdoes the decrease of vane size because the shock excitation mayremain strong even behind a smaller stator. The evaluation ofthe aerodynamic excitation towards a HCF risk reduction shouldonly be done with regard to the excited mode shape, asdemonstrated with parametric studies of the mode shapeinfluence on excitability.

Keywords:Aeroelasticity, Aerodynamics, Stator-RotorInteraction, Excitation Mechanism, Unsteady Flow Computation,Forced Response, High Cycle Fatigue, Turbomachinery,Gas-Turbine, High-Pressure Turbine, Turbopump, CFD, Design

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Newman, Timothy James. "Towards a silent fan : an investigation of low-speed fan aeroacoustics." Thesis, University of Cambridge, 2015. https://www.repository.cam.ac.uk/handle/1810/251318.

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The noise (unwanted sound) from fans of all sizes, operating in close proximity to people, can be a design constraint due to annoyance or, in the worse cases, health damage. Of the total noise, aeroacoustic noise - produced by unsteadiness in the air - often represents a significant source and is intrinsically linked to the aerodynamic features of the flow field. In this work, the aeroacoustics of low-speed fans are investigated using a compact mixed-flow fan as a test case. The low-speed regime is less developed compared to large-scale, high-speed machines and is increasingly relevant to applications such as micro air vehicles, small wind turbines, and other environmental comfort technologies found in buildings or vehicles. The test case fan Reynolds number is of the order of 104 which is a couple of orders lower than those generally found in gas turbines. Its main sources are therefore best identified experimentally in the absence of proven alternative methods. In order to do this, a way of quantifying fan noise is developed in tandem with control of the aerodynamic operating point. Following a study of sources of the significant broadband and tonal noise, a low-order noise prediction scheme is developed and applied to predict tonal noise with reference to Reynolds number effects. The new, duct-based rig and method has several advantages over the existing sound power measurement rig built to the ISO 5136 standard at Dyson. The approach, which makes no assumptions about the relative power of different modes, has resulted in a rig that is much shorter. Unlike the ISO rig, it is capable of accurate narrow-band tone measurements with sources which excite strong non-plane-wave duct modes (as the modal structure of the sound is determined) for the frequencies of interest. Tests have been carried out at different operating points with a range of geometry modifications produced with 3D printing techniques. In terms of tonal sources which particularly impact sound quality, the mixed-flow impeller alone produces tones due to very high sensitivity to inflow distortion of the mean flow (giving unsteady blade loading). This means that the product inlet must be designed very carefully to optimally condition the flow. Periodicity in the impeller outlet flow produces rotor-stator interaction tones even with a number of guide vanes chosen to satisfy the Tyler-Sofrin theory cut-off criteria. This is thought to be due to abrupt radius change after the guide vanes in the rig (while the theory assumes constant radius). In the product, abrupt radius change also occurs. The sensitivity of the broadband level to inflow turbulence was confirmed to be low in the rig, although the in-product inflow appears much less ideal. The main broadband noise source in rig tests is suggested to be impeller self-noise as only small reductions in rotor-stator interaction noise are achieved with far fewer vanes. The low-order modelling scheme to understand the fundamental unsteady loading noise mechanism compares well to experiments for sample rotor-stator interaction tones. The velocity fluctuations which induce this noise, measured experimentally with a 2D hotwire, are shown to increase in intensity as Reynolds number is reduced towards 104. This is due to a higher importance of viscosity which can give boundary layers that are thicker and liable to laminar separation. Surface treatments such as boundary layer trips could be used to prevent such separation and potentially reduce noise. Based on the thesis findings, further tests, simulations and possible design modifications are suggested to understand and reduce the important noise sources.

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Ahmadi, Majid. "Aerodynamic inverse design of transonic turbomachinery cascades." Thesis, National Library of Canada = Bibliothèque nationale du Canada, 1998. http://www.collectionscanada.ca/obj/s4/f2/dsk1/tape11/PQDD_0003/NQ40321.pdf.

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25

Choi, Myeonggeun. "Thermal control of gas turbine casings for improved tip clearance." Thesis, University of Oxford, 2015. https://ora.ox.ac.uk/objects/uuid:14a9ce6a-2af6-4187-afe7-8c6f8e113855.

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A thermal tip clearance control system provides a robust and flexible means of manipulating the closure between the casing and the rotating blade tips in a jet engine, reducing undesirable tip leakage flows. This may be achieved using an impingement cooling scheme on the external casing of the engine in conjunction with careful thermal management of internal over-tip seal segment cavity. For a reduction in thrust specific fuel consumption, the mass flow rate of air used for cooling must be minimised, be at as low a pressure as possible and delivered through a light weight structure surrounding the rotating components in the turbine. This thesis first characterises the effectiveness of a range of external impingement cooling arrangements in typical engine casing closure system. The effects of jet-to-jet pitch, number of jets, inline and staggered alignment of jets, arrays of jets on flange, on an engine representative casing geometry are assessed through comparison of the convective heat transfer coefficient distributions in a series of numerical studies. A baseline case is validated experimentally. The validation data allowed the suitability of different turbulence closure models to be assessed using a commercial RANS solver. Importantly for each configuration the thermal contraction of an idealised engine casing is predicted using thermo-mechanical finite element models, at a series of operating conditions representing engine idle to maximum take-off conditions. Cooling is provided by manifolds attached to the outside of the engine. The assembly tolerance of these components leads to variation in the standoff distance between the manifold and the casing. For cooling arrangements with promising performance, the study is extended to characterise the variation in closure with standoff distance. It is shown that where a sparse array of non-interacting jets is used the system can be made tolerant of large build misalignments. The casing geometry itself contributes to the thermal response of the system, and, in an additional study, the effect of casing thickness and circumferential thermal control flanges are investigated. Restriction of the passage of heat into the flanges was seen to be dramatically change their effectiveness and slight necking of the flanges at their root was shown to improve the performance disproportionally. High temperature secondary air flowing past the internal face of the engine casing tends to heat the casing, causing it to grow. Experimental and numerical characterisation of a heat transfer within a typical over-tip segment cavity heat transfer is presented in this thesis for the first time. A simplified modelling strategy is proposed for casing and a means to reduce the casing heat pickup by up to 25 % was identified. The overall validity of the modelling approach used is difficult to validate in the engine environment, however limited data from a test engine temperature survey became available during the course of the research. By modelling this engine tip clearance control system it was shown that good agreement to the temperature distribution in the engine casing could be achieved where full surface external heat transfer coefficient boundary conditions were available.

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Sy, Birame. "Adaptabilité en espace d'un schéma volumes finis d'ordre élevé pour la CFD/CAA des turbomachines." Thesis, Paris, HESAM, 2020. http://www.theses.fr/2020HESAE045.

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A l’ère du numérique, le cycle de développement d’un produit se fait dans sa quasi totalité sur ordinateur. Il n’est plus nécessaire de produire physiquement des versions préliminaires. Leurs caractéristiques peuvent être testées avec une précision dépendante de la maturité des méthodes de simulation. Les acteurs de la recherche en simulation numérique ont donc pour défi de transposer leurs récentes avancées vers l’industrie. En mécanique des fluides, les codes de calcul doivent gagner en adaptabilité afin de prendre en compte la morphologie du problème et du maillage. Le paramétrage des méthodes avancées doit pouvoir être délégué par l’utilisateur non-expert à la machine. Ce travail de recherche porte sur l’adaptabilité en espace d’un schéma volumes finis d’ordre élevé (FV-MLS). L’ordre élevé est un élément indispensable afin de capter les phénomènes fortement instationnaires.Pour augmenter l’ordre de précision, le schéma FV-MLS fait intervenir une reconstruction polynomiale d’ordre élevé par Moindres Carrés Mobiles. MLS affiche un fort potentiel en terme de flexibilité pour traiter des géométries complexes. Elle possède par ailleurs un nombre important de paramètres pouvant être intégrés dans un procédé d’optimisation. Ces travaux ont tout d’abord apporté des réponses concernant la sensibilité de la méthode vis-à-vis des paramètres MLS. À un second niveau, une série d’algorithmes de choix pertinent de ces paramètres a été mise au point, tout en améliorant nettement la robustesse, la précision et l’efficacité de calcul. La charge de l’utilisateur a ainsi été réduite de manière conséquente, lui permettant de se recentrer sur son cœur de métier. Cette méthodologie a été validée jusqu’à l’ordre 6. Pour améliorer la robustesse au schéma numérique vis-à-vis de l’anisotropie du maillage, un nouveau cadre de reconstruction locale d’ordre élevée a été défini.Cette reconstruction locale permet de réduire drastiquement les effets de l’anisotropie. Plusieurs cas de validation et exemples d’applications ont été réalisé afin de démonter l’intérêt des méthodes proposées
In the digital age, almost all of a product’s development cycle is done on a computer. There is no longer a need to physically produce drafts. Their characteristics can be tested with precision that depends on the maturity of the simulation methods. Researchers in digital simulation therefore have the challenge of transferring their recent advances to industry. In fluid mechanics, the computer codes must gain in adaptability in order to take into account the morphology of the problem and the mesh.The configuration of advanced methods should be delegated by the non-expert user to the machine.This research work focused on the adaptability in space of a high order finite volume scheme (FV-MLS). The high order is an essential element in order to capture highly unsteady phenomena. To increase the order of precision, the FV-MLS scheme involves a high order polynomial reconstruction by Least Mobile Squares. MLS has great potential in terms of flexibility for handling complex geometries. It also has a large number of parameters that can be integrated into an optimization process.This work first provided answers concerning the sensitivity of the method regarding the MLS parameters. At a second level, a series of algorithms for the relevant choice of these parameters has been developed, while clearly improving the robustness, the precision and the calculation efficiency. The user’s load has therefore been reduced significantly, allowing him to focus on his core business. This methodology has been validated up to order 6. To improve the robustness of the numerical scheme vis-à-vis the anisotropy of the mesh, a new high-order local reconstruction framework has been defined. This local reconstruction makes it possible to reduce or even annihilate the effects of anisotropy. Several validation cases and examples of applications have been carried out in order to demonstrate the value of the proposed methods

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Wang, Dingxi. "Turbomachinery aerodynamic and aeromechanic design optimization using the adjoint method." Thesis, Durham University, 2008. http://etheses.dur.ac.uk/2057/.

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The thesis documents the investigation of the application of the adjoint method to turbomachinery blading design optimization, with emphasis on blading aerodynamic design optimization in a multi-bladerow environment and concurrent blading aerodynamic and aeromechanic design optimization for a single bladerow. Based on the nonlinear flow equations, a steady adjoint system has been developed using the continuous adjoint approach. The capability of the conventional adjoint system has been augmented by the introduction of an adjoint mixing-plane treatment. This treatment is a counterpart of the flow mixing-plane treatment, enabling the steady adjoint equations to be solved in multi-bladerow computational domains. This allows turbomachinery blades to be optimised to enhance their aerodynamic performance in a multi-bladerow environment with matching between adjacent bladerows dealt with in a timely manner. The Nonlinear Harmonic Phase Solution method, a neat frequency domain method catered specifically for turbomachinery aeromechanics prediction, has been chosen to integrate with the adjoint method to calculate objective function sensitivities efficiently for concurrent aeromechanic and aerodynamic design optimization for single row turbomachinery blades. The Nonlinear Harmonic Phase Solution method, unlike the time-linearized methods, solves the unsteady flow equations at two or three carefully selected phases of a period of unsteadiness. This approach not only can conveniently turn a steady flow solver to one solving the unsteady flow equations efficiently, but also provides a good basis on which the corresponding adjoint system can be formulated and solved in a similar manner by extending a steady adjoint system. In order to resolve the issue of having a good blading performance over a whole operating range at a given operation speed, a multi-operating-point design optimization is implemented by formulating an objective function of a weighted sum of performance at more than one operating point

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Ramer, Becky E. (Becky Ellen). "Aerodynamic response of turbomachinery blade rows to convecting density distortions." Thesis, Massachusetts Institute of Technology, 1997. http://hdl.handle.net/1721.1/49964.

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Wijesinghe, Hettithanthrige Sanith 1974. "Aerodynamic response of turbomachinery blade rows to convecting density wakes." Thesis, Massachusetts Institute of Technology, 1998. http://hdl.handle.net/1721.1/50476.

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Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1998.
Includes bibliographical references (p. 129-131).
Density wakes have been recently identified as a possible new source for high cycle fatigue failure in the compressor blades of modern turbomachinery. In order to characterize the density wake induced force and moment fluctuations in compressor blades a two-dimensional computational study has been conducted in viscous compressible flows with Mach numbers ranging from M[infinity], = 0.15 to M[infinity] = 0.87 and flow Reynolds number Re(c, U[infinity]) ~~700,000. Parametric tests were conducted at each flow Mach number to establish trends for the change in the maximum fluctuation of the blade force and moment coefficients with the changes in the density wake width 0.1 < w/c < 1.0 and the density ratio 0.25 < P2/P1 < 2.0. Results indicate the magnitude of the blade force and moment fluctuations to scale with (1) the non-dimensional density wake width w/c, (2) a non-dimensional density parameter p* and (3) flow Mach number M[infinity]. The viscous flow simulations have also indicated (1) periodic vortex shedding at the blade trailing edge and (2) separation bubbles on the blade suction surface which generate additional force and moment fluctuations with amplitudes ±(10 - 100%) about the time averaged mean values. These flow features represent possible additional sources for high cycle fatigue failure. Simple functional relationships have also been derived at each flow Mach number to quantify the force and moment fluctuations described above. In addition a simple cascade flow model has been developed in conjunction with the computational study to help determine the trends in the force and moment fluctuations with varying density wake properties and compressor geometries.
by Hettithanthrige Sanith Wijesinghe.
S.M.

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Coppinger, Miles. "Aerodynamic performance of an industrial centrifugal compressor variable inlet guide vane system." Thesis, Loughborough University, 1999. https://dspace.lboro.ac.uk/2134/7263.

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Industrial centrifugal air compressors can require a combination of a large range of mass flow, high efficiency, constant pressure ratio, and constant rotational speed, specifically when used for sewage effluent aeration treatment. In order to achieve this performance it is common to use variable inlet guide vanes (VIGV's). The performance characteristics of an existing VIGV design have been determined using both an experimental test facility and state of art numerical techniques. The results obtained from these techniques are far more comprehensive than earlier fullscale performance testing. Validation of the performance of the existing design using these techniques has led to the development of a new vane design and potential improvements to the inlet ducting geometry. The aerodynamic interaction between the VIGV system and the centrifugal compressor impeller has also been investigated using a 3-D computational model of the complete variable geometry compressor stage. The results of these investigations have been validated by data available from full scale experimental testing. Strong correlation was obtained between numerical and experimental techniques, and a predicted improvement in polytropic efficiency up to 3% at low flow rates using the re-designed variable inlet guide vanes has been achieved. The overall outcome of this research is a usable VIGV design technique for real industrial compressor environments, and confirmation that an acceptable design can be achieved that represents a rewarding improvement in performance.

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Willcox, Karen E. (Karen Elizabeth). "Reduced-order aerodynamic models for aeroelastic control of turbomachines." Thesis, Massachusetts Institute of Technology, 2000. http://hdl.handle.net/1721.1/9265.

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Thesis (Ph.D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2000.
This electronic version was submitted by the student author. The certified thesis is available in the Institute Archives and Special Collections.
Includes bibliographical references (p. 138-143).
Aeroelasticity is a critical consideration in the design of gas turbine engines, both for stability and forced response. Current aeroelastic models cannot provide high-fidelity aerodynamics in a form suitable for design or control applications. In this thesis low-order, high-fidelity aerodynamic models are developed using systematic model order reduction from computational fluid dynamic (CFD) methods. Reduction techniques are presented which use the proper orthogonal decomposition, and also a new approach for turbomachinery which is based on computing Arnoldi vectors. This method matches the input/output characteristic of the CFD model and includes the proper orthogonal decomposition as a special case. Here, reduction is applied to the linearized two-dimensional Euler equations, although the methodology applies to any linearized CFD model. Both methods make efficient use of linearity to compute the reduced-order basis on a single blade passage. The reduced-order models themselves are developed in the time domain for the full blade row and cast in state-space form. This makes the model appropriate for control applications and also facilitates coupling to other engine components. Moreover, because the full blade row is considered, the models can be applied to problems which lack cyclic symmetry. Although most aeroelastic analyses assume each blade to be identical, in practice variations in blade shape and structural properties exist due to manufacturing limitations and engine wear. These blade to blade variations, known as mistuning, have been shown to have a significant effect on compressor aeroelastic properties. A reduced-order aerodynamic model is developed for a twenty-blade transonic rotor operating in unsteady plunging motion, and coupled to a simple typical section structural model. Stability and forced response of the rotor to an inlet ow disturbance are computed and compared to results obtained using a constant coefficient model similar to those currently used in practice. Mistuning of this rotor and its effect on aeroelastic response is also considered. The simple models are found to inaccurately predict important aeroelastic results, while the relevant dynamics can be accurately captured by the reduced-order models with less than two hundred aerodynamic states. Models are also developed for a low-speed compressor stage in a stator/rotor configuration. The stator is shown to have a significant destabilizing effect on the aeroelastic system, and the results suggest that analysis of the rotor as an isolated blade row may provide inaccurate predictions.
by Karen Elizabeth Willcox.
Ph.D.

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32

Fruth, Florian. "Reduction of Aerodynamic Forcing inTransonic Turbomachinery : Numerical Studies on Forcing Reduction Techniques." Doctoral thesis, KTH, Kraft- och värmeteknologi, 2013. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-127967.

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Due to more and more aggressive designs in turbomachinery, assuring the structural integrity of its components has become challenging. Also influenced by this trend is blade design, where lighter and slimmer blades, in combination with higher loading, lead to an increased risk of failure, e.g. in the form of blade vibration. Methods have been proposed to reduce vibration amplitudes for subsonic engines, but cannot directly be applied to transonic regimes due to the additional physical phenomena involved. Therefore the present work investigates numerically the influence of two methods for reducing blade vibration amplitudes in transonic turbomachines, namely varying the blade count ratio and clocking. As it is known that clocking affects the efficiency, the concurrent effects on vibration amplitudes and efficiency are analyzed and discussed in detail. For the computational investigations, the proprietary Fortran-based non-linear, viscous 3D-CFD solver VolSol is applied on two transonic compressor cases and one transonic turbine case. In order to reduce calculation time and to generate the different blade count ratios a scaling technique is applied. The first and main part of this work focuses on the influence of the reduction techniques on aerodynamic forcing. Both the change in blade count ratio and clocking position are found to have significant potential for reducing aerodynamic force amplitudes. Manipulation of the phasing of excitation sources is found herein to be a major contributor to the amplitude variation. The lowest stimulus results are achieved for de-phased excitation sources and results in multiple blade force peaks per blade passing. In the case of blade count ratio variation it was found that blockage for high blade count ratios and the change in potential field size have significant impacts on the blade forcing. For the clocking investigation, three additional operating points and blade count ratios are analyzed and prove to have an impact on the force reduction achievable by clocking. The second part of the work evaluates the influence of clocking on the efficiency of a transonic compressor. It is found that the efficiency can be increased, but the magnitude of the change and the optimal wake impingement location depend on the operating point. Moreover it is shown that optimal efficiency and aerodynamic forcing settings are not directly related. In order to approximate the range of changes of both parameters, an ellipse approximation is suggested.

QC 20130911

TURBOPOWER

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Watson, Brian Christopher. "An investigation into the influence of mistuning on the forced response of bladed disk assemblies." Diss., Georgia Institute of Technology, 1993. http://hdl.handle.net/1853/12463.

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Payer, Florent. "Prédiction et analyse du phénomène de réponse forcée : application à un cas de compresseur haute pression." Phd thesis, Ecole Centrale de Lyon, 2013. http://tel.archives-ouvertes.fr/tel-01063776.

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L'enjeu de cette thèse est d'améliorer la compréhension et la prédiction du phénomène de réponse forcée des aubages de turbomachines en situation de résonance. L'étude a été menée au moyen de simulations numériques U-RANS 3D et en s'appuyant sur le compresseur d'essai ERECA, dédié au phénomène de réponse forcée. Pour prédire les amplitudes de vibration des aubages excités aérodynamiquement, la méthode de prédiction la plus répandue consiste à effectuer séparément un calcul d'excitation et un calcul d'amortissement aérodynamique ; on parle alors de calcul découplé. C'est cette méthode qui a été mise en œuvre dans un premier temps. Les calculs d'excitation et d'amortissement aérodynamiques ont été comparés individuellement aux résultats d'essais. Pour cela une méthode de traitement du signal fréquence/amplitude a été développée dans le but d'extraire l'amortissement et l'excitation des résultats d'essais. Les analyses des simulations ont permis de mieux comprendre les mécanismes d'excitation et d'amortissement aérodynamique. On a ainsi pu montrer que le phénomène d'interaction rotor/stator s'apparente par son caractère discontinu à une percussion périodique. Quant au phénomène d'amortissement, il se caractérise par le bilan des contributions de chaque zone d'échange d'énergie sur la paroi de l'aubage. En outre, les amplitudes vibratoires calculées à partir de cette méthode sont très proches des valeurs d'essais. Toutefois, cette procédure de calcul requiert la mise en œuvre de 2 calculs instationnaires différents et ne permet pas à l'heure actuelle d'être utilisée dans un cycle de conception. Dans le but de simplifier et d'améliorer la qualité de prédiction des analyses de réponse forcée, la méthode du couplage dynamique a été mise en œuvre et évaluée. Avec cette méthode, l'aubage répond librement aux sollicitations engendrées par le fluide. Une fois le régime transitoire évacué, l'aubage oscille en régime permanent. Cette méthode permet donc de prédire une amplitude vibratoire à partir d'un seul calcul instationnaire. En revanche, le calcul s'avère bien plus onéreux que la méthode découplée de par l'existence du régime transitoire. Dans le but de rendre accessible cette méthode à un niveau industriel, deux méthodes d'accélération du calcul ont été mises en place. Les résultats obtenus sont très encourageants et devraient permettre de réduire drastiquement les temps de restitution des analyses de réponse forcée. A la connaissance de l'auteur, cette thèse constitue une étude inédite de comparaison entre méthode découplée et couplage dynamique, qui par ailleurs s'appuie sur des résultats d'essais dédiés exclusivement au phénomène de réponse forcée.

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Brouckaert, Jean-François M. "Development of fast response aerodynamic probes for time-resolved measurements in turbomachines." Doctoral thesis, Universite Libre de Bruxelles, 2002. http://hdl.handle.net/2013/ULB-DIPOT:oai:dipot.ulb.ac.be:2013/211406.

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36

Morris, Mary Beth. "Flow randomness and tip losses in transonic rotors." Thesis, This resource online, 1992. http://scholar.lib.vt.edu/theses/available/etd-07212009-040241/.

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37

Monaco, Lucio. "PARAMETRIC STUDY OF THE EFFECT OF BLADE SHAPE ON THE PERFORMANCE OF TURBOMACHINERY CASCADES : PART III A: AERODYNAMIC DAMPING BEHAVIOUR – COMPRESSOR PROFILES." Thesis, KTH, Kraft- och värmeteknologi, 2010. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-131210.

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Ghate, Devendra. "Inexpensive uncertainty analysis for CFD applications." Thesis, University of Oxford, 2014. http://ora.ox.ac.uk/objects/uuid:6be44a1d-6e2f-4bf9-b1e5-1468f92e21e3.

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The work presented in this thesis aims to provide various tools to be used during design process to make maximum use of the increasing availability of accurate engine blade measurement data for high fidelity fluid mechanic simulations at a reasonable computational expense. A new method for uncertainty propagation for geometric error has been proposed for fluid mechanics codes using adjoint error correction. Inexpensive Monte Carlo (IMC) method targets small uncertainties and provides complete probability distribution for the objective function at a significantly reduced computational cost. A brief literature survey of the existing methods is followed by the formulation of IMC. An example algebraic model is used to demonstrate the IMC method. The IMC method is extended to fluid mechanic applications using Principal Component Analysis (PCA) for reduced order modelling. Implementation details for the IMC method are discussed using an example airfoil code. Finally, the IMC method has been implemented and validated for an industrial fluid mechanic code HYDRA. A consistent methodology has been developed for the automatic generation of the linear and adjoint codes by selective use of automatic differentiation (AD) technique. The method has the advantage of keeping the linear and the adjoint codes in-sync with the changes in the underlying nonlinear fluid mechanic solver. The use of various consistency checks have been demonstrated to ease the development and maintenance process of the linear and the adjoint codes. The use of AD has been extended for the calculation of the complete Hessian using forward-on-forward approach. The complete mathematical formulation for Hessian calculation using the linear and the adjoint solutions has been outlined for fluid mechanic solvers. An efficient implementation for the Hessian calculation is demonstrated using the airfoil code. A new application of the Independent Component Analysis (ICA) is proposed for manufacturing uncertainty source identification. The mathematical formulation is outlined followed by an example application of ICA for artificially generated uncertainty for the NACA0012 airfoil.

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Nuckolls, William E. "Fan noise reduction from a supersonic inlet." Thesis, This resource online, 1992. http://scholar.lib.vt.edu/theses/available/etd-08222009-040447/.

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Powers, Laura M. "Computer-aided design of axial-flow fans." Thesis, Virginia Polytechnic Institute and State University, 1986. http://hdl.handle.net/10919/91059.

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This thesis examines the application of computer-aided design techniques to the field of turbomachinery. Specifically, the process of designing low- to medium-speed, axial-flow fans and blowers is discussed, and a Fortran program called FANJD is introduced. The first purpose of F AN3D is to perform the aerodynamic and mechanical calculations needed to establish the basic geometry of an axial-flow fan blade. Next, geometric modeling techniques are used to model the curves and surfaces of the blade, thereby completing the geometric description of the blade. Finally, FAN3D uses the CADCD component of the CADAM (CADAM, Inc.) Geometry Interface to automatically enter the three-dimensional blade model in the CADAM database.
M.S.

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Lewis, Daniel Joseph. "Tip clearance and angle of attack effects upon the unsteady response of a vibrating flat plate in crossflow /." This resource online, 1993. http://scholar.lib.vt.edu/theses/available/etd-06112009-063924/.

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Ugolotti, Matteo. "Implementation and Evaluation of Machine Learning Assisted Adjoint Sensitivities Applied to Turbomachinery Design Optimization." University of Cincinnati / OhioLINK, 2020. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1593267985073912.

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Butler, Bradley D. "AXIAL COMPRESSOR FLOW BEHAVIOR NEAR THE AERODYNAMIC STABILITY LIMIT." UKnowledge, 2014. http://uknowledge.uky.edu/me_etds/34.

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In this investigation, casing mounted high frequency response pressure transducers are used to characterize the flow behavior near the aerodynamic stability limit of a low speed single stage axial flow compressor. Time variant pressure measurements are acquired at discrete operating points up to the stall inception point and during the transition to rotating stall, for a length of time no shorter than 900 rotor revolutions. The experimental data is analyzed using multiple techniques in the time and frequency domains. Experimental results have shown an increase in the breakdown of flow periodicity as the flow coefficient is reduced. Below a flow coefficient of 0.40 a two node rotating disturbance develops with a propagation velocity of approximately 23% rotor speed in the direction of rotation. During rotating stall, a single stall cell is present with a propagation velocity of approximately 35% rotor speed. The stall inception events present are indicative of a modal stall inception.

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Fletcher, Nathan James. "Design and Implementation of Periodic Unsteadiness Generator for Turbine Secondary Flow Studies." Wright State University / OhioLINK, 2019. http://rave.ohiolink.edu/etdc/view?acc_num=wright1560810428267352.

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Kumar, Sandeep. "Non-AXisymmetric Aerodynamic Design-Optimization System with Application for Distortion Tolerant Hybrid Propulsion." University of Cincinnati / OhioLINK, 2020. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1613749886763596.

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O'Dowd, Devin Owen. "Aero-thermal performance of transonic high-pressure turbine blade tips." Thesis, University of Oxford, 2010. http://ora.ox.ac.uk/objects/uuid:e7b8e7d0-4973-4757-b4df-415723e7562f.

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Ceylanoglu, Arda. "An Accelerated Aerodynamic Optimization Approach For A Small Turbojet Engine Centrifugal Compressor." Master's thesis, METU, 2009. http://etd.lib.metu.edu.tr/upload/12611371/index.pdf.

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Centrifugal compressors are widely used in propulsion technology. As an important part of turbo-engines, centrifugal compressors increase the pressure of the air and let the pressurized air flow into the combustion chamber. The developed pressure and the flow characteristics mainly affect the thrust generated by the engine. The design of centrifugal compressors is a challenging and time consuming process including several tests, computational fluid dynamics (CFD) analyses and optimization studies. In this study, a methodology on the geometry optimization and CFD analyses of the centrifugal compressor of an existing small turbojet engine are introduced as increased pressure ratio being the objective. The purpose is to optimize the impeller geometry of a centrifugal compressor such that the pressure ratio at the maximum speed of the engine is maximized. The methodology introduced provides a guidance on the geometry optimization of centrifugal impellers supported with CFD analysis outputs. The original geometry of the centrifugal compressor is obtained by means of optical scanning. Then, the parametric model of the 3-D geometry is created by using a CAD software. A design of experiments (DOE) procedure is applied through geometrical parameters in order to decrease the computation effort and guide through the optimization process. All the designs gathered through DOE study are modelled in the CAD software and meshed for CFD analyses. CFD analyses are carried out to investigate the resulting pressure ratio and flow characteristics. The results of the CFD studies are used within the Artificial Neural Network methodology to create a fit between geometric parameters (inputs) and the pressure ratio (output). Then, the resulting fit is used in the optimization study and a centrifugal compressor with higher pressure ratio is obtained by following a single objective optimization process supported by design of experiments methodology.

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48

Stein, Alexander. "Computational analysis of stall and separation control in centrifugal compressors." Diss., Georgia Institute of Technology, 2000. http://hdl.handle.net/1853/11884.

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49

Green, Brian Richard. "Time-Averaged and Time-Accurate Aerodynamic Effects of Rotor Purge Flow for a Modern, Rotating, High-Pressure Turbine Stage and Low-Pressure Turbine Vane." The Ohio State University, 2011. http://rave.ohiolink.edu/etdc/view?acc_num=osu1322535026.

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50

Gougeon, Pierre. "Interactions aérodynamiques entre une turbine haute pression et le premier distributeur basse pression." Thesis, Ecully, Ecole centrale de Lyon, 2014. http://www.theses.fr/2014ECDL0026/document.

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L’amélioration des performances des turboréacteurs actuels est un enjeu crucial dans un contexte de contraintes économiques et environnementales fortes. Au sein du turboréacteur, le canal inter-turbines, localisé à l’interface entre la turbine Haute Pression (HP) et le premier distributeur Basse Pression (BP), est le siège d’écoulements très complexes. Ainsi, les structures aérodynamiques issues de la turbine HP (sillages, tourbillons et ondes de choc) interagissent fortement entre elles et impactent l’écoulement du distributeur BP, engendrant ainsi des pertes de rendement de l’ensemble de la configuration. Ce travail de thèse s’attache à étudier les phénomènes d’interactions aérodynamiques entre une turbine HP et le premier distributeur BP et à analyser les mécanismes à l’origine des pertes aérodynamiques dans le distributeur BP. Une campagne expérimentale antérieure, réalisée sur un banc d’essai comprenant une turbine HP couplée à un distributeur BP, avait permis de recueillir des mesures de l’écoulement dans des plans situés dans le canal inter-turbines et à l’aval du distributeur BP. En lien avec ces résultats expérimentaux, les simulations numériques menées dans cette étude avec le logiciel elsA s’attachent à restituer précisément la nature tridimensionnelle, instationnaire et turbulente de l’écoulement au sein de cette même configuration. Ces travaux se développent alors en trois étapes principales. Dans un premier temps, une étude stationnaire avec traitement plan de mélange permet de comprendre et quantifier les aspects généraux de l’écoulement. Une évaluation de l’effet de la modélisation turbulente RANS (Reynolds-Averaged Navier-Stokes) et du schéma numérique spatial sur les structures aérodynamiques présentes dans la configuration est réalisée. Dans un deuxième temps, une modélisation turbulente avancée de type ZDES (Zonal Detached-Eddy Simulation) est employée pour la résolution de l’écoulement dans le distributeur BP. Les structures aérodynamiques instationnaires issues de la roue HP amont sont modélisées par une condition limite à l’entrée du domaine de calcul. L’approche ZDES est comparée à une approche Unsteady RANS (URANS) sur la même configuration. La formation et la dissipation des sillages et des tourbillons est significativement différente entre les deux modélisations, ce qui impacte de manière importante la génération des pertes aérodynamiques. Enfin, des simulations URANS de plusieurs configurations permettent de mieux comprendre les effets d’interaction entre les différentes rangées d’aubes. Ainsi, les approches instationnaires chorochroniques prenant en compte un seul rotor et un seul stator évaluent des effets instationnaires importants dans le canal inter-turbines. Ces approches conduisent à la mise en oeuvre d’un calcul sur une configuration multipassages-chorochronique prenant en compte les deux stators et le rotor afin de modéliser complètement les interactions déterministes existantes. Afin de quantifier celles-ci avec précision, une décomposition modale du champ instationnaire est mise en place. Les niveaux d’interactions liées aux différentes roues sont alors quantifiés et l’impact sur les pertes aérodynamiques est évalué
Improving the performance of current aeronautical turbines is an important issue in a context of severe economical and environmental constraints. In a turbofan, the inter-turbine channel which is located between the High-Pressure (HP) turbine and the first Low Pressure (LP) vane is characterized by a complex flow. Therefore aerodynamic structures coming from the HP turbine (wakes, vortices and showkwaves) strongly interact between each other and affect the LP vane flow field. This generates efficiency losses of the overall configuration. This PhD thesis aims at studying the aerodynamic phenomena between a HP turbine and the first LP vane and at analyzing the mechanisms creating aerodynamic losses. A previous experimental campaign, which was carried out on a facility including a HP turbine coupled to a LP vane, enabled to gather flow field measurements in planes located in the inter-turbine channel and downstream of the LP vane. In comparison with these experimental data, the numerical simulations done with elsA software intend to reproduce accurately the 3D, unsteady and turbulent nature of the flow within this configuration. The work can be divided into three mains steps. As a first step, steady simulations with a sliding mesh treatment enable to understand the general aspects of the flow. An assessment of the effects of RANS (Reynolds-Averaged Navier-Stokes) turbulent predictions and of spatial numerical schemes on the aerodynamic structures present in the configuration is carried out. As a second step, the advanced turbulence approach ZDES (Zonal Detached-Eddy Simulation) is considered for the LP vane flow prediction. The unsteady aerodynamic structures coming from the upstream HP rotor are set as an inlet boundary condition of the computational domain. The ZDES approach is compared to a URANS (Unsteady RANS) approach on the same computational domain. The generation and dissipation of the wakes and vortices are significantly different on the two simulations, and thus impact the creation of aerodynamic losses. Finally, URANS simulations enable to better understand the interaction effects between the different blade rows. First, the unsteady phase-lagged approaches that take into account a single rotor and stator assess the important unsteady effects in the inter-turbine channel. They finally lead to the implementation of a multipassages phase-lagged computation that takes into account the two stators and the rotor in order to model all the existing determinist interactions. In order to quantify them accurately, a modal decomposition of the unsteady flow field is set up. The interaction levels linked to the different blade rows are therefore quantified and the impact of the aerodynamic losses is evaluated

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Dissertations / Theses on the topic 'Turbomachinery aerodynamics'

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