This dissertation presents an approach to use blade tip timing measurements with finite element analysis to predict the fatigue life of a low pressure steam turbine last stage blade under high backpressure and low flow conditions. Material fatigue properties were determined through the extended universal material law for FV566 material, along with different temper scenarios. A finite element model of a blade with damping pins was developed, using the principle of cyclic symmetry for a perfectly tuned model. Pre-stress modal analysis was conducted, incorporating damping via friction and plasticity for initial 20% overspeed test. The finite element model was verified by two experimental tests: the first being a blade impact test and the second a telemetry strain gauge test in a balance pit. Fatigue life analysis was conducted under the assumption that non-synchronous vibration is experienced by the blade and that only one mode is dominant in the vibration. The results from the fatigue analysis corresponded to the location of the cracks experienced on the blades. The results show twelve orders of magnitude lower life at low load, high backpressure conditions, compared to high load high pressure conditions. The research was further extended to check the same vibratory response on the first three modes, up to their tenth nodal diameters. This was done to analyse fatigue life in a case that a different mode was excited. / Dissertation (MEng)--University of Pretoria, 2020. / Eskom Rotek Industries / Mechanical and Aeronautical Engineering / MEng / Unrestricted

This research developed two innovative test methods that were used to experimentally evaluate the performance of a novel blade tip timing (BTT) system from Prime Photonics, LC. The research focused on creating known blade tip offsets and tip vibrations so that the results from a BTT system can be validated. The topic of validation is important to the BTT field as the results between many commercial systems still are not consistent. While the system that was tested is still in development and final validation is not complete, the blade tip offset and vibration frequency validation results show that this BTT system will be a valuable addition to turbomachinery research and development programs once completed. For the first test method custom rotors were created with specified blade tip offsets. For the blade tip offset alternate measurement, the rotors were optically scanned and analyzed in CAD software with a tip location uncertainty of 0.1 mm. The BTT system agreed with the scanned results to within 0.13 mm. Tests were also conducted to ensure that the BTT system identified and indexed the blades properly. The second developed test method used an instrumented piezoelectric blade to create known dynamic deflections. The active vibration rotor was able to create measureable deflection over a range of frequencies centered on the first bending mode of the blade. The results for the 110 Hz, 150 Hz, 180 Hz first bending resonance, 200 Hz, and 1036 Hz second bending resonance cases are presented. A strain gage and piezoelectric sensor were attached to the active blade during the dynamic deflection tests to provide an alternate method for determining blade vibration frequency. The BTT system correctly identified the active blade excitation frequencies as well as a 120 Hz frequency from the drive motor. This thesis also explored applying BTT methods and testing to more realistic blade geometry and vibration. Blade vibrations are usually classified by their frequency relative to the rotation speed. Synchronous vibrations are integer multiples of the rotational speed and are often excited by struts or vanes fixed to the engine case. For this reason, special probe placement algorithms were explored that use sine curve fitting to optimize the probe placement. Knowing how the blade will vibrate at operation before testing is critical as well. In preparation for future research, ANSYS Mechanical was used to predict the first three modes of a PT6A-28 first stage rotor blade at 1,966, 5,539, and 7,144 Hz. These frequencies were validated to within 4% using scanning laser vibrometry. The simulation was repeated at speed to produce a Campbell Diagram to highlight synchronous excitation crossings. / Master of Science

This thesis presents the results of a study performed in the Virginia Tech low speed linear cascade wind tunnel operating at a Reynolds number of 382,000 designed to model an axial compressor rotor. To simulate the flow created by the junction of a set of inlet guide vanes and the compressor casing, vortex generators were glued to a moving end wall. In this investigation, the tip clearance was varied from 0.83% to 12.9% chord. Measurements of the midspan and the tip blade loading were made using static pressure taps. The tip loading shows that the minimum suction surface pressure coefficient increases in magnitude linearly up to a tip clearance of 7.9% chord. Unsteady pressure was measured on the pressure and suction surfaces at the tip of two cascade blades using an array of 23 microphones mounted subsurface. These measurements reveal that the unsteady pressure at the blade tip is a linear function of tip clearance height. The instantaneous pressure shows that the surface pressure at the blade tip has the same character regardless of whether or not the blade is disturbed by the inflow vortices. This suggests that the vortex generators simply stimulate and organize the existing response of the blade. Single sensor hot-wire measurements were made within the tip clearance on the suction side of the blade 1mm from the tip gap exit. These measurements show that the mass flux through the tip clearance is closely related to the pressure difference across the tip gap. / Master of Science

Aeroacoustic investigations regarding different types of helicopter noise have indicated that the most annoying noise is caused by impulsive blade surface pressure changes in descent or forward flight conditions. Blade Vortex Interaction (BVI) is one of the main phenomena producing significant impulsive noise by the unsteady fluctuation in blade loading due to the rapid change of induced velocity field during interaction with vortices shed from previous blades. The tip vortex core structure and the blade vortex miss distance were identified as having a primary influence on BVI. In this thesis, piezoelectrically modulated and/or vectored blowing at the rotor blade tip is theoretically investigated as an active technique for modifying the structure of the tip vortex core as well as for increasing blade vortex miss distance. The mechanisms of formation and convection of rotor blade tip vortices up to and beyond 360 degrees wake age are described based on the CFD results for the baseline cases of a hovering rotor with rounded and square tips. A methodology combining electromechanical and CFD modeling is developed and applied to the study of a piezoelectrically modulated and vectored blowing two-dimensional wing section. The thesis is focused on the CFD analysis of rotor flow with modulated tangential blowing over a rounded blade tip, and with steady mid-plane blade tip blowing, respectively. Computational results characterizing the far-wake flow indicate that for steady tangential blowing the miss distance can be doubled compared to the baseline case, which may lead to a significant reduction in BVI noise level if this trend shown in hover can be replicated in low speed forward flight. Moreover, near-wake flow analysis show that through modulated blowing a higher dissipation of vorticity can be obtained.

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.

A novel combined piezoelectric-composite actuator configuration is proposed and analytically modeled in this work. The actuator is a low complexity, active compliant mechanism obtained by coupling a modified star cross sectional configuration composite beam with a helicoidal bimorph piezoelectric actuator coiled around it. This novel actuator is a good candidate as a hinge tension-torsion bar actuator for a helicopter rotor blade flap or blade tip and mirror rotational positioning. In the wing tip case, the tip deflection angle is different only according to the aerodynamic moment depending on the hinge position of the actuator along the chord and applied voltage because there is no centrifugal force. For an active blade tip subject to incompressible flow and 2D quasi steady airloads, its twist angle is related not only to aerodynamic moment and applied voltage but also to coupling terms, such as the trapeze effect and the tennis racquet effect. Results show the benefit of hinge position aft of the aerodynamic center, such that the blade tip response is amplified by airloads. Contrary to this effect, results also show that the centrifugal effects and inertial effect cause an amplitude reduction in the response. Summation of these effects determines the overall blade tip response. The results for a certain hinge position of Xh=1.5% chord aft of the quarter chord point proves that the tip deflection target design range[-2,+2] can be achieved for all pitch angle configurations chosen.

No description available.

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.

The tips of unshrouded, high-pressure turbine blades are prone to significantly high heat loads. The gap between the tip and over-tip casing is the root cause of undesirable over-tip leakage flow that is directly responsible for high thermal material degradation and is a major source of aerodynamic loss within a turbine. Both must be minimised for the safe working and improved performance of future gas-turbines. A joint experimental and numerical study is presented to understand and characterise the heat transfer and aerodynamics of unshrouded blade tips. The investigation is undertaken with the use of a squealer or cavity tip design, known for offering the best overall compromise between the tip aerodynamics, heat transfer and mechanical stress. Since there is a lack of understanding of these tips at engine-realistic conditions, the present study comprises of a detailed analysis using a high-speed linear cascade and computational simulations. The aero-thermal performance is studied to provide a better insight into the behaviour of squealer tips, the effects of casing movement and tip cooling. The linear cascade environment has proved beneficial for its offering of spatially-resolved data maps and its ability to validate computational results. Due to the unknown tip gap height within an entire engine cycle, the effects of gap height are assessed. The squealer's aero-thermal performance has been shown to be linked with the gap height, and qualitative different trends in heat transfer are established between low-speed and high-speed tip flow regimes. To the author's knowledge, the present work is the first of its kind, providing comprehensive aero-thermal experimental research and a dataset for a squealer tip at engine-representative transonic conditions. It is also unique in terms of conducting direct and systematic validations of a major industrial computational fluid dynamics method for aero-thermal performance prediction of squealer tips at enginerepresentative transonic conditions. Finally, after recognising the highest heat loads are found on the squealer rims, a novel shaped squealer tip has been investigated to help improve the thermal performance of the squealer with a goal to improve its durability. It has been discovered that a seven percent reduction in tip temperature can be achieved through incorporating a shaped squealer and maximising the internal cooling performance.

L’aérodynamique de l’hélicoptère est fortement impactée par les tourbillons générés aux extrémités de pales. La complexité des phénomènes en jeux et l’insuffisance de données expérimentales locales font du design d’extrémité un véritable défi. Cette étude propose une nouvelle approche dédiée à l’étude des extrémités en vol stationnaire. Une méthode numérique rapide et précise est mise au point afin d’étudier une extrémité de pale en rotation comme une extrémité d’aile fixe. Chaque étape de la construction de la méthode est validée par des comparaisons détaillées avec des données expérimentales publiées. Le code CFD elsA est dans un premier temps utilisé pour mettre en place une méthode de calcul basée sur la résolution des équations Reynolds-Averaged Navier-Stokes en stationnaire. La convergence de la solution et l’indépendance au maillage et aux paramètres numériques sont étudiées en détail en deux, puis en trois dimensions. La précision importante de la solution numérique permet d’analyser finement la physique de l’enroulement tourbillonnaire en extrémité. Des géométries tronquée et arrondie sont étudiées en détail, et révèlent la présence de systèmes tourbillonnaires complexes. Puis la nouvelle méthode d’adaptation pale en rotation / aile fixe est présentée. Une méthode de calcul hybride est mise au point entre le code de mécanique du vol HOST et le code elsA. En repère fixe, l’aérodynamique globale sur la pale et locale en extrémité est calculée fidèlement pour toutes les configurations étudiées. Comparée aux méthodes d’adaptation précédemment publiées, cette nouvelle stratégie offre une amélioration considérable concernant la simulation de l’aérodynamique de pale. / Helicopter aerodynamics is strongly influenced by the vortices generated from the rotor-blade tips. The design of efficient tip shapes is a challenging task because of the complexity of the aerodynamic phenomena involved and the lack of local blade-tip flow measurements. This work provides a contribution to the design of helicopter tips in hover. An efficient, relatively simple and quick numerical method is set up to study rotating blade tips in fixed-wing configurations. The accuracy of the method is shown at each step of the construction by comprehensive comparisons with reliable experimental data from the literature. First, an efficient steady Reynolds-Averaged Navier-Stokes method is constructed using ONERA's elsA code. Comprehensive studies of convergence, grid dependence and sensitivity to the numerical method are performed in two and three dimensions. The very good agreement of the solution with measurements and the accuracy of the numerical method allow a physical analysis with unprecedented detail of the vortex generation and roll-up near square and rounded wing tips. The new methodology of framework adaptation is then presented. An uncoupled hybrid strategy is set up using AIRBUS HELICOPTERS' Comprehensive Analysis code HOST and the Computational Fluid Dynamics solver elsA. Global and local performance calculations are validated for all investigated test cases. Comparison with previously published adaptation methods indicates considerable improvement in the prediction of the blade aerodynamics.