Thermal Barrier Coating systems (TBCs) serve as a key component in gas turbines in aerospace engines, isolating the metallic substrate from severe heat flux of the environment. The durability of TBCs has been considered to be a critical issue to determine the service lifespan of hot section components. Comprehensive studies of failure mechanisms benefit the gas turbine industry to develop TBCs with better material properties and stable microstructures, thus potentially enhancing their durability. To date, many failure mechanism analyses have been conducted based on the understanding of critical residual stress developed under different thermal tests. For the present study, using the Finite Element (FE) method with temperature-process-dependent model parameters, the maximum residual stress is calculated with evolution of the localized/global interfacial roughness profile based on Electron Beam-Physical Vapour Deposition Thermal Barrier Coating system (EB-PVD TBCs). With studies of cracking routes from past research, qualitative failure mechanism analysis is conducted for EB-PVD TBCs. In addition, the estimated energy release rates are compared to reveal the effect of different thermal profiles on the crack driving forces for Atmospheric Plasma Sprayed Thermal Barrier Coating systems (APS-TBCs). Using previously observed cracking routes from different thermal cycling experiments, a quantitative failure mechanism analysis is conducted for APS-TBCs with modified analytical expressions. In addition, literature works revealed that physics and mechanics-based models were proposed to evaluate environment induced damage. For the last part of my research, erosion of EB-PVD TBCs is estimated using a modified solid particle erosion model. A stochastic approach is applied to study the erosion of EB-PVD topcoat (TC) under real engine service conditions. The durability of TBCs is affected by both oxidation-induced degradation and environment-induced damage. The combination of “internal” crack driving forces (generated from residual stresses developed upon different stages of thermal cycles) and “external” erosion damage (from temperature-process dependent brittle/ductile erosion) lead to complexity of evaluating durability under different service conditions.

The efficiency of natural gas turbines is directly linked to the turbine inlet temperature, or the combustor exit temperature. Further increasing the turbine inlet temperature damages the turbine components and limits their durability. Advances in turbine vane cooling schemes protect the turbine components. This thesis studies the conjugate effects of internal cooling, film cooling and thermal barrier coatings (TBC) on turbine vane metal temperatures. Two-dimensional thermal profiles were experimentally measured downstream of a single row of film cooling holes on both an adiabatic and a matched Biot number model turbine vane. The measurements were taken as a comparison to computational simulations of the same model and flow conditions. To improve computational models of the evolution of a film cooling jet as it propagates downstream, the thermal field above the vane, not just the footprint on the vane surface must be analyzed. This study expands these data to include 2-D thermal fields above the vane at 0, 5 and 10 hole diameters downstream of the film cooling holes. In each case the computational jets remained colder than the experimental jets because they did not disperse into the mainstream as quickly. Finally, in comparing results above adiabatic and matched Biot number models, these thermal field measurements allow for an accurate analysis of whether or not the adiabatic wall temperature was a reasonable estimate of the driving temperature for heat transfer. In some cases the adiabatic wall temperature did give a good indication of the driving temperature for heat transfer while in other cases it did not. Previous tests simulating the effects of TBC on an internally and film cooled model turbine vane showed that the insulating effects of TBC dominate over variations in film cooling geometry and blowing ratio. In this study overall and external effectiveness were measured using a matched Biot number model vane simulating a TBC of thickness 0.6d, where d is the film cooing hole diameter. This new model was a 35% reduction in thermal resistance from previous tests. Overall effectiveness measurements were taken for an internal cooling only configuration, as well as for three rows of showerhead holes with a single row of holes on the pressure side of the vane. This pressure side row of holes was tested both as round holes and as round holes embedded in a realistic trench with a depth of 0.6 hole diameters. Even in the case of this thinner TBC, the insulating effects dominate over film cooling. In addition, using measurements of the convective heat transfer coefficient above the vane surface, and the thermal conductivities of the vane wall and simulated TBC material, a prediction technique of the overall effectiveness with TBC was evaluated. / text

Le système barrière thermique (BT) est utilisé pour protéger les aubes de turbines à gaz aéronautiques. Aux températures de fonctionnement, une des causes de l'endommagement du système barrière thermique est la dégradation de la couche céramique isolante en zircone yttriée (8YPSZ : ZrO2 - 4% mol. Y2O3) par corrosion. Celle-ci est due à des dépôts d'oxydes à base de Ca, Mg, Al, Si, appelés CMAS provenant de diverses particules ingérées par le moteur. A haute température (~1200°C), le CMAS fond et s'infiltre dans la microstructure poreuse de la BT, se rigidifie au refroidissement provoquant, à terme, la délamination de la BT. A haute température, la BT subit une corrosion chimique induisant sa dissolution dans le CMAS liquide. L'ensemble de ces deux phénomènes conduit à la perte d'intégrité de la barrière thermique. Le présent travail s'est focalisé sur la compréhension des mécanismes de dégradation chimique en vue de proposer une solution de protection contre l'infiltration par les CMAS. Après expertise d'aubes de turbines de retour de vol, une reproduction de la corrosion de la barrière thermique par un CMAS modèle de type CAS et une étude thermodynamique et cinétique de la dissolution de différents oxydes des systèmes ZrO2 - Y2O3 et ZrO2 - Nd2O3 ont été menées dans le verre silicaté CAS pour comprendre le processus de dissolution de Zr et Y et définir une nouvelle composition de barrière thermique anti-CMAS. Le comportement en corrosion par le CAS de matériaux céramiques denses de compositions ZrO2 - 12% mol Nd2O3 et Zr2Nd2O7 ainsi qu'un revêtement déposé par EB-PVD ((La, Nd)2Zr2O7) a été testée. Les résultats obtenus font apparaître que : i) le CAS réplique le mécanisme de corrosion en service, soit la dissolution-re-précipitation. ii) l'oxyde ZrO2 se dissout progressivement et forme le zircon (ZrSiO4) dans le verre CAS, dès 30 min iii) les dopants (Nd2O3 et Y2O3) conduisent à la formation très rapide, de la phase apatite X8Ca2(SiO4)6O2 (X = Nd ou Y) après réaction avec le verre silicaté. En plus de la phase apatite, Y2O3 forme la phase Ca3Y2Si6O18, qui est instable entre 1300°C et 1400°C. iv) les composés dopés au néodyme (ZrO2 - 12% mol Nd2O3 et Zr2Nd2O7) se dissolvent et conduisent, quasi-spontanément, à la phase apatite Nd8Ca2(SiO4)6O2 ainsi qu'à la re-précipitation de grains de ZrO2 appauvris en néodyme. v) malgré la présence de Y2O3, les composés ZrO2 - 4% mol Y2O3, ZrO2 - 10% mol Y2O3 ne conduisent qu'à la re-précipitation de la zircone appauvrie en Y2O3. L'absence de phases secondaires notamment, la phase apatite, pourrait expliquer l'infiltration facile du CMAS dans la microstructure de la barrière thermique en zircone yttriée. vi) l'inhibition avérée de l'infiltration du CAS dans la microstructure poreuse de couches céramiques de nouvelles compositions semble être due à la formation rapide d'une couche superficielle fine et dense, constituée de zircone appauvrie en dopant et de phase apatite / Thermal barrier coating (TBC) system is used to protect aeronautical gas turbine blades. At operating temperatures, one of the damaging causes of thermal barrier system is the degradation of the insulating ceramic layer in zirconia (8YPSZ: ZrO2 - 4 mol %. Y2O3) by corrosion. The corrosion is due to calcium - magnesium alumino-silicates (CMAS) deposits from various particles ingested by the engine. At high temperature (~ 1200°C), the molten CMAS infiltrates the porous microstructure of the thermal barrier leads to i) the chemical dissolution of the thermal barrier zirconia and ii) the delamination of the TBC after cracking at low temperature due to the mismatch of CTE of the solid oxides constituting the CMAS and TBC. This study has contributed to understanding the mechanisms of chemical degradation in order to propose a solution to protect against infiltration by CMAS. After expertise of ex-service turbine blades, a reproduction of the thermal barrier corrosion by model CMAS (CAS) and thermodynamic and kinetic study of the solubility of different oxides of both ZrO2-Y2O3 and ZrO2-Nd2O3 systems were performed in the silicate glass (CAS) in order to understand the mechanism of Zr and Y dissolution and to define a new composition of TBC. The corrosion by the CAS of dense ceramic (ZrO2 - 12 mol% Nd2O3 and Zr2Nd2O7) and of a EB-PVD coating (La, Nd)2Zr2O7)was studied. The results obtained show that: i) CAS replicates the corrosion mechanism, i.e. dissolution-re-precipitation reaction ii) ZrO2 oxide dissolves gradually and forms zircon (ZrSiO4) in the glass after 30 min iii) (Nd2O3 and Y2O3) oxides lead very rapidly to the apatite X8Ca2(SiO4)6O2 (X = Nd, Y) phase formation, after reaction with silicate glass. In addition to the apatite phase, Y2O3 forms Ca3Y2Si6O18 phase, which is unstable at 1300°C and 1400°C iv) the compounds doped with Nd2O3 (ZrO2 - 12 mol% Nd2O3 and Zr2Nd2O7) dissolve and form almost spontaneously, the apatite Nd8Ca2(SiO4)6O2 phase and the ZrO2 depleted in Nd2O3 grains v) Although Y2O3 is a constitutent of the compounds ZrO2 - 4 mol% Y2O3, ZrO2 - 10 mol% Y2O3, the chemical corrosion of these compounds leads only to the re-precipitation of zirconia depleted Y2O3. The absence of secondary phases, particularly the apatite phase may explain the easy CMAS infiltration in the microstructure of the 8YPSZ thermal barrier vi) inhibition of CAS infiltration into the porous microstructure of ceramic layers of new compositions seems to be due to the rapid formation of a thin and dense layer, consisting in Nd-depleted zirconia and apatite phase

This document describes work performed related to the determination of how elemental additions affect the interfacial fracture toughness of thermal barrier coatings at the bond coat/thermally grown oxide interface in gas turbines. These turbines are exposed to cyclical thermal loading, therefore a simulation was designed to model this interface in a temperature cycle between 200 K and 1000 K that included oxide growth between 2 μm and 27 μm. The fracture toughness of this interface was then determined to elucidate the function of elemental additions. It was shown that minimal concentrations of atomic species, such as hafnium and yttrium cause notable increases in the toughness of the bond coat/thermally grown oxide interface, while other species, such as sulphur, can dramatically reduce the toughness. Furthermore, it was shown that, contrary to some empirical results, the addition of platinum has a negligible effect on the fracture toughness of this interface.

Because of their good mechanical resistance at high temperature, Ni-based superalloys are used for aero-engine and land-based turbines but undergo "dry" oxidation between 900 and 1500°C. These materials are thus coated with nickel-aluminide coatings (BC). An additional thermal barrier coating (TBC) is generally applied in the hottest sections of the turbines (T>1050°C) to lower the impact of the temperature on the substrate. In the framework of the European research programme "PARTICOAT", this PhD work was focused on the growth mechanisms of a full protective coating system (BC+TBC) in a single step process, using a water-based slurry containing a dispersion of Al micro-particles to satisfy the European environmental directives. The rheological and physico-chemical characterizations showed the slurry stability up to seven days. After depositing the latter by air spraying, a tailored thermal treatment resulted in a nickel-aluminide coating (β-NiAl) similar to the conventional industrial ones but through an intermediate Al liquid phase stage. Simultaneously, the oxidation of the Al micro-particles brought aboutthe formation of a top alumina "foam" (PARTICOAT concept). After a validation step of the mechanisms involved in pure Ni substrate, the extrapolation of the process to several Ni-based superalloys (René N5 (SX), CM-247 (DS), PWA- 1483 (SX) and IN-738LC (EQ)) revealed different coating compositions and microstructures. A particular attention was therefore paid onto the effect of alloying elements (Cr, Ta, Ti) as well as their segregation in the coating. The high temperature behaviour of the coated samples has been studied through isothermal oxidation (1000h in air between 900 and 1100°C) and showed that the oxidation and interdiffusion phenomena ruled the degradation mechanisms. Besides, the electrodeposition of ceria before the application of the PARTICOAT coating allowed to strongly limit interdiffusion phenomena and stabilized the nickel aluminide coating.

Approximately 10% of all energy consumed in the United States is derived from high temperature gas turbine engines. As a result, a 1% increase in engine efficiency would yield enough energy to satisfy the demands of approximately 1 million homes and savings of over $800 million in fuel costs per year. Efficiency of gas turbine engines can be improved by increasing the combustor temperature. Modern engines now operate at temperatures that far exceed the material limitations of the metals they are comprised of in the pursuit of increased thermal efficiency. Various techniques to thermally protect the turbine components are used to allow for safe operation of the engines despite the extreme environments: film cooling, internal convective cooling, and thermal barrier coatings. Historically, these thermal protection techniques have been studied separately without account for any conjugate effects. The end goal of this work is to provide a greater understanding of how the conjugate effects might alter the predictions of thermal behavior and consequently improve engine designs to pursue increased efficiency. The primary focus of this study was to complete the first open literature, high resolution experiments of a modeled first stage turbine vane with both active film cooling and a simulated thermal barrier coating (TBC). This was accomplished by scaling the thermal behavior of a real engine component to the model vane using the matched Biot number method. Various film cooling configurations were tested on both the suction and pressure side of the model vane including: round holes, craters, traditional trenches and a novel modified trench. IR thermography and ribbon thermocouples were used to measure the surface temperature of the TBC and the temperature at the interface of the TBC and vane wall, respectively. This work found that the presence of a TBC significantly dampens the effect of altering film cooling conditions when measuring the TBC interface temperature. This work also found that in certain conditions adiabatic effectiveness does not provide an accurate assessment of how a film cooling design may perform in a real engine. An additional focus of this work was to understand how contaminant deposition alters the cooling performance of a vane with a TBC. This work focused on quantifying the detrimental effects of active deposition by seeding the mainstream flow of the test facility with simulated molten coal ash. It was found that in most cases, except for round holes operating at relatively high blowing ratios, the performance of film cooling was negatively altered by the presence of contaminant deposition. However, the cooling performance at the interface of the TBC and vane wall actually improved with deposition due to the additional thermal resistance that was added to the exterior surface of the model vane. / text

This document describes work performed related to the determination of how elemental additions affect the interfacial fracture toughness of thermal barrier coatings at the bond coat/thermally grown oxide interface in gas turbines. These turbines are exposed to cyclical thermal loading, therefore a simulation was designed to model this interface in a temperature cycle between 200 K and 1000 K that included oxide growth between 2 μm and 27 μm. The fracture toughness of this interface was then determined to elucidate the function of elemental additions. It was shown that minimal concentrations of atomic species, such as hafnium and yttrium cause notable increases in the toughness of the bond coat/thermally grown oxide interface, while other species, such as sulphur, can dramatically reduce the toughness. Furthermore, it was shown that, contrary to some empirical results, the addition of platinum has a negligible effect on the fracture toughness of this interface.

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Les barrières thermiques avec zircone EB-PVD pour les turbines aéronautiques sont soumises à des conditions extrêmes qui conduisent à l'écaillage du dépôt. La prévention de leur endommagement est donc nécessaire pour assurer l'intégrité des pièces. Afin de comprendre et de reproduire leur évolution dans une turbine, les barrières thermiques actuelles sont soumises à des essais longs de cyclage thermique. L'essai LASAT est un essai d'adhérence rapide à mettre en oeuvre qui se place en complément du cyclage thermique. L'impulsion laser appliquée sur la face nue de l'AM1 génère une onde de choc de compression qui se propage jusqu'à la surface libre de la zircone. La réflexion forme une onde de traction qui effectue le trajet inverse et peut rompre les interfaces qu'elle traverse. Le décohésion génère une tache blanche dans la zircone directement visible à l'oeil. Ce phénomène optique est élucidé en relation avec la microstructure de la zircone et la présence d'une fissure à l'interface. Pour connaitre le potentiel de l'essai, une large gamme d'échantillons avec différentes orientations du superalliage, quatre préparations de sous-couche, cinq microstructures de zircone et deux vieillissements thermiques ont été utilisés.Leur caractérisation a permis de les classer et de comparer leurs évolutions et leurs endommagements par cyclage thermique ou par LASAT. Le dimensionnement des fissures interfaciales par des méthodes non destructives a été réalisé par piézospectroscopie en exploitant les cartographies associées au signal defluorescence, par profilométrie et à partir de la tache blanche. Une approche simple et innovante exploitant et optimisant le comportement optique de la zircone est mise en place. Les tailles des fissures relevées ont mis en évidence le rôle des ondes 2D et permis la réalisation de l'essai LASAT-2D. Ici, ce n'est plus l'apparition de la fissure qui est recherchée, mais sa taille qui peut directement informer de l'adhérence à partir d'un seul choc laser. La modélisation numérique a confirmé le rôle de ces ondes 2D et leur potentield'utilisation par des abaques LASAT-2D. Ces courbes permettent de distinguer différentes préparations de barrières thermiques brutes d'élaboration ou vieillies. Un protocole complet est ainsi fourni pour le contrôle, la mesure et le suivi de la tenue mécanique de barrières thermiques sur des éprouvettes usuelles industrielles. Dans des essais complémentaires, le LASAT-2D a été appliqué en "face avant", avec le choc coté zircone, sur des éprouvettes et des pièces industrielles. Les mêmes tendances que pour le LASAT-2D développé dans cette thèse sont observées. Ceci autorise la perspective de l'application de cet essai et de cette méthodologie sur des formes complexes et fermées, telles les aubes de turbine. / EB-PVD thermal barrier coatings used in aircraft turbines are subjected to extreme conditions that lead to their spallation. To ensure the integrity of the parts, it's necessary to prevent coating damages. In order to understand and reproduce their evolution in a turbine, current thermal barrier coatings are subjected to long thermal cycling. The LASAT is a rapid adhesion test that could complement the thermal cycling. The laser pulse applied to the AM1 face generates a compressive shock wave which propagates towards the free surface of the zirconia. The reflection of this wave generates a tensile shock wave which can damage the interface and create a white spot in zirconia top coat. To determine the potential of the test, a wide range of samples with different superalloy orientations, bondcoat preparations, zirconia microstructures and thermal aging were used. Characterizations were carrying out to classify and compare their evolution and their damage by thermal cycling or LASAT. The size of interfacial cracks by non-destructive tests was achieved by piezospectroscopie maps associated with the fluorescent temporal signal, by profilometry and from the white spot. A simple and innovative approach by optimizing the optical behavior of zirconia is developed. These results highlight the role of 2D waves and allow the realization of the testLASAT-2D. Here, it is not the appearance of the crack that is looking for, but its size, which can directly inform about the mechanical adhesion from a single laser shock. Numerical modeling has confirmed the influence of 2D shock waves and their potential using LASAT-2D charts. These curves are used to distinguishdifferent preparations of thermal barrier coatings as produced and after thermal aging. A complete protocol is thus provided for monitoring, measuring and determine the interface strength of thermal barrier coatingson industrial specimens. On further testing, LASAT-2D was applied directly on the coating of specimens and industrial parts. The same results as for the 2D-LASAT developed in this thesis are observed. This allows the prospect of the application of this test and the methodology on complex and closed shapes, suchas turbine blades.