Dissertations / Theses on the topic 'Turbomachinery; CFD'

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.

In aeromechanical design one of the major rules is to operate under High Cyclic Fatigue (HCF) margins and away from flutter. The level of dynamic excitations and risk of HCF can be estimated by performing forced response analyses from blade row interaction forces or Low Engine Order (LEO) excitation mechanisms. On the other hand, flutter stability prediction can be assessed by calculation of aerodynamic damping forces due to blade motion. In order to include these analyses as regular practices in an industrial aeromechanical design process, interaction between the fields of fluid and structural dynamics must be established in a rather simple yet accurate manner. Effects such as aerodynamic and structural mistuning should also be taken into account where parametric and probabilistic studies take an important role.

The present work presents the development and validation of a numerical tool for aeromechanical design. The tool aims to integrate in a standard and simple manner regular aeromechanical analysis such as forced response analysis and aerodynamic damping analysis of bladed disks.

Mistuning influence on forced response and aerodynamic damping is assessed by implementing existing model order reduction techniques in order to decrease the computational effort and assess results in an industrially applicable time frame.  The synthesis program solves the interaction of structure and fluid from existing Finite Element Modeling (FEM) and Computational Fluid Dynamics (CFD) solvers inputs by including a mapping program which establishes the fluid and structure mesh compatibility. Blade row interaction harmonic forces and/or blade motion aerodynamic damping forces are inputs from unsteady fluid dynamic solvers whereas the geometry, mass and stiffness matrices of a blade alone or bladed disk sector are inputs from finite element solvers. Structural and aerodynamic damping is also considered.

Structural mistuning is assessed by importing different sectors and any combinations of the full disk model can be achieved by using Reduced Order Model (ROM) techniques. Aerodynamic mistuning data can also be imported and its effects on the forced response and stability assessed. The tool is developed in such a way to allow iterative analysis in a simple manner, being possible to realize aerodynamically and structurally coupled analyses of industrial bladed disks. A new method for performing aerodynamic coupled forced response and stability analyses considering the interaction of different mode families has also been implemented. The method is based on the determination of the aerodynamic matrices by means of least square approximations and is here referred as the Multimode Least Square (MLS) method.

The present work includes the program description and its applicability is assessed on a high pressure ratio transonic compressor blade and on a simple blisk.

QC 20140609

The continuous trend in gas turbine design towards lighter, more powerful and more reliable engines on one side and use of alternative fuels on the other side renders flutter problems as one of the paramount challenges in engine design. Flutter denotes a self-excited and self-sustained aeroelastic instability phenomenon that can lead to material fatigue and eventually damage of structure in a short period of time unless properly damped. The design for flutter safety involves the prediction of unsteady aerodynamics as well as structural dynamics that is mostly based on in-house developed numerical tools. While high confidence has been gained on the structural side unanticipated flutter occurrences during engine design, testing and operation evidence a need for enhanced validation of aerodynamic models despite the degree of sophistication attained. The continuous development of these models can only be based on the deepened understanding of underlying physical mechanisms from test data.

As a matter of fact most flutter test cases treat the turbomachine flow in two-dimensional manner indicating that the problem is solved as plane representation at a certain radius rather than representing the complex annular geometry of a real engine. Such considerations do consequently not capture effects that are due to variations in the third dimension, i.e. in radial direction. In this light the present thesis has been formulated to study three-dimensional effects during flutter in the annular environment of a low-pressure turbine blade row and to describe the importance on prediction of flutter stability. The work has been conceived as compound experimental and computational work employing a new annular sector cascade test facility. The aeroelastic response phenomenon is studied in the influence coefficient domain having one blade oscillating in various three-dimensional rigid-body modes and measuring the unsteady response on several blades and at various radial positions. On the computational side a state-of-the-art industrial numerical prediction tool has been used that allowed for two-dimensional and three-dimensional linearized unsteady Euler analyses.

The results suggest that considerable three-dimensional effects are present, which are harming prediction accuracy for flutter stability when employing a two-dimensional plane model. These effects are mainly apparent as radial gradient in unsteady response magnitude from tip to hub indicating that the sections closer to the hub experience higher aeroelastic response than their equivalent plane representatives. Other effects are due to turbomachinery-typical three-dimensional flow features such as hub endwall and tip leakage vortices, which considerably affect aeroelastic prediction accuracy. Both effects are of the same order of magnitude as effects of design parameters such as reduced frequency, flow velocity level and incidence. Although the overall behavior is captured fairly well when using two-dimensional simulations notable improvement has been demonstrated when modeling fully three-dimensional and including tip clearance.

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