Dans le secteur aéronautique en pleine expansion, les préoccupations environnementales prennent une place de plus en plus importante. Les motoristes recherchent des solutions innovantes pour augmenter les rendements tout en diminuant les coûts. Dans cette perspective, de nouveaux systèmes de barrières thermiques synthétisés par la voie sol-gel à partir de poudres commerciales, de céramiques avec différents facteurs de forme et d'agents porogènes ont été mis en œuvre et évalués. Certains systèmes présentent une durée de vie de plus de 1000 cycles en oxydation cyclique. Malgré tout, cet accroissement des températures de fonctionnement des moteurs, induit une élévation des températures de surfaces des barrières thermiques et peut générer de nouvelles dégradations du système complet : la corrosion à hautes températures par les CMAS. Pour pallier ces inconvénients, il est possible de développer des revêtements anti-CMAS, susceptibles de réagir avec les composés CMAS avant qu'ils n'aient un effet néfaste sur l'intégrité de la barrière thermique. Dans cette étude, nous nous sommes intéressés particulièrement aux revêtements sacrificiels anti-CMAS à base d'yttrine et de systèmes pyrochlore, qui ont été testés sur des barrières thermiques industrielles de type EBPVD. Par ailleurs, les procédés que nous avons développés, basés sur la voie sol-gel, nous permettent, de par leur facilité de mise en œuvre, d'envisager des perspectives prometteuses en termes de réparabilité des barrières thermiques endommagées. En effet, compte tenu du coût élevé de fabrication des pièces, les aubes devraient être réparées plusieurs fois avant d'être mises au rebut. Dans ce travail, un procédé de mise en forme a été évalué dans ce sens. Il s'agit de l'électrophorèse qui est une technique bien adaptée au dépôt sur pièces complexes. L'objectif de ces investigations a donc été double : tout d'abord créer de nouveaux systèmes de barrières thermiques avec des propriétés anti-CMAS par électrophorèse puis réparer les barrières thermiques EBPVD endommagées et leur déposer une couche protectrice anti-CMAS par ce même procédé. Cet aspect " procédé " sera abordé en dernière partie de ces travaux. / In the aeronautics sector, environmental concerns are becoming increasingly important. Engine manufacturers are looking for innovative solutions to increase efficiency while lowering costs. The objective is to optimize thermal conductivity and durability with the cyclic oxidation resistance. In this perspective, new thermal barrier systems synthesized by the sol-gel route from commercial powders, ceramics with various form factors and pore-forming agents have been implemented and evaluated. Some systems are a lifetime higher than 1000 cycles in cyclic oxidation. However, this increase in the operating temperatures of the engines induces an increase in the temperature of the surfaces of the thermal barriers and can generate further degradations of the complete system: the corrosion by CMAS. To overcome these disadvantages, it is possible to develop anti-CMAS coatings capable of reacting with CMAS compounds before they have a detrimental effect on the integrity of the thermal barrier. In this study, we were particularly interested in anti-CMAS protective coatings based on yttria and pyrochlore systems, which were tested on industrial thermal barriers realized by EBPVD. Moreover, the processes we have developed, based on the sol-gel path, allow us, because of their ease of implementation, to envisage promising prospects in terms of repair of damaged thermal barriers. Indeed, given the high cost of manufacturing parts, the blades should be repaired several times before being discarded. In this work, a shaping process has been evaluated in this direction. This is electrophoretic deposition which is a technique allowing to deposit on complex parts. The objective of these investigations was therefore twofold: firstly to create new thermal barrier systems with anti-CMAS properties by electrophoretic deposition and then to repair the damaged EBPVD thermal barriers and to deposit an anti-CMAS protective layer by this same process. This "process" aspect will be discussed at the end of this work.

Thermal barrier coatings (TBC) are used in the hot sections of gas turbine engine in order to insulate the substrate at high temperature. Molten salt infiltration retards the durability of TBCs. The current standard material, i.e. 8YSZ is susceptible to molten salt infiltration. Therefore, alternate TBC materials are desirable. In addition to material composition, the TBC microstructure plays an important role in mitigating molten salt infiltration. Therefore, in this work, three different TBC variations were investigated. The first variation was a columnar microstructured 48YSZ TBC processed by SPS (48YSZ-SPS). The second variation was a columnar microstructured 8YSZ TBC processed by SPS (8YSZ-SPS), and the third variation was a lamellar microstructured 8YSZ TBC deposited by APS (8YSZ-APS). The as-sprayed TBC specimens were characterized by SEM/EDS, porosity analysis and XRD measurements. Later, the TBC specimens were exposed to hot corrosion test and their interaction with the molten salts were investigated using SEM (EDS and XRD). It was shown that an increase in stabilizer content (yttria content) in zirconia (in the case of 48YSZ) leads to an improved hot corrosion resistance due to the adequate amount of yttria content, which restricts the molten salt infiltration by forming needle like YVO4 phase. In terms of microstructure comparison, the infiltration behavior was similar for columnar microstructured 8YSZ and lamellar microstructured 8YSZ-APS as the molten salts infiltrated the coatings completely compared to the 48YSZ TBC. Furthermore, it seems that the molten salt infiltrates the TBC through globular pores, delamination cracks and splat boundaries in the case of APS-TBCs whereas the column gaps favor easier infiltration of molten salts in the case of columnar microstructured SPS processed TBCs.

<p>Thermal barrier coatings (TBC) are used for reduction of component temperatures in gas turbines. The service temperature for turbines can be as high as 1100ºC and the components are exposed to thermal cycling and gases that will cause the component to oxidize and corrode. The coatings are designed to protect the substrate material from this, but eventually it will lead to failure of the TBC. It is important to have knowledge about when this failure is expected, since it is detrimental for the gas turbine.</p><p>The scope of this thesis has been to see if an existing life model for thin TBC also is valid for thick TBC. In order to do so, a thermal cycling fatigue test, a tensile test and finite element calculation have been performed. The thermal cycling fatigue test and finite element calculation were done to find correlations between the damage due to thermal cycling, the number of thermal cycles and the energy release rate. The tensile test was preformed to find the amount accumulated strain until damage.</p><p>The thermal cycling lead to failure of the TBC at the bond coat/top coat interface. The measurment of damage, porosity and thickness of thermally grown oxide were unsatisfying due to problems with the specimen preparation. However, a tendency for the damage development were seen. The finite element calculations gave values for the energy release rate the stress intensity factors in mode~I and mode~II that can be used in the life model. The tensile test showed that the failure mechanism is dependent of the coating thickness and it gave a rough value of the maximum strain acceptable.</p>

From many years YSZ is used as the top coat material for TBC's, as it has good phase stability up to 1200°C, higher fracture toughness, lower thermal conductivity, erosion resistance &amp; higher coefficient of thermal expansion. But, it has a drawbacks at high temperature such as sintering and transformation of phases. For this reason new ceramic materials with pyrochlores crystal structure such as Gd2Zr2O7 are being considered as it has high melting points, phase stability, lower thermal conductivity and CMAS resistance. But it has low fracture toughness when compared to YSZ. In order to take advantage of low thermal conductivity and high thermal stability of gadolinium zirconate and avoiding the drawbacks of low coefficient of thermal expansion and low toughness using YSZ, a double/multi-layer coatings approach is being used. Therefore, multi-layer TBCs are sprayed and compared with single layer coating in this work. These coatings are processed by suspension plasma spraying. For single layer coating YSZ is used, for double layer coating YSZ as the intermediate coating and Gd2Zr2O7 as the top coat is used. Additionally, a triple layer coating system comprising YSZ, Gd2Zr2O7 and dense Gd2Zr2O7 as top coat is also sprayed. The as sprayed coatings are characterized for microstructure analysis using optical microscope and scanning electron microscope (SEM), elemental analysis of TGO using Energy-Dispersive Spectrometer (EDS). XRD analysis was done to identify various phases in the coating. Porosity analysis using Archimedes principle was carried out. Thermal cyclic fatigue (TCF) test of the sprayed coatings was carried out at 1100°C. Failure analysis of the TCF specimens was carried out using SEM/EDS. TCF results showed that the triple layer coatings (dense Gd2Zr2O7/Gd2Zr2O7/YSZ) had higher thermal cyclic fatigue life and lower TGO thickness when compared to single layer (YSZ) and double layer (Gd2Zr2O7/YSZ) TBCs.

Thermal barrier coatings (TBC) are used for reduction of component temperatures in gas turbines. The service temperature for turbines can be as high as 1100ºC and the components are exposed to thermal cycling and gases that will cause the component to oxidize and corrode. The coatings are designed to protect the substrate material from this, but eventually it will lead to failure of the TBC. It is important to have knowledge about when this failure is expected, since it is detrimental for the gas turbine. The scope of this thesis has been to see if an existing life model for thin TBC also is valid for thick TBC. In order to do so, a thermal cycling fatigue test, a tensile test and finite element calculation have been performed. The thermal cycling fatigue test and finite element calculation were done to find correlations between the damage due to thermal cycling, the number of thermal cycles and the energy release rate. The tensile test was preformed to find the amount accumulated strain until damage. The thermal cycling lead to failure of the TBC at the bond coat/top coat interface. The measurment of damage, porosity and thickness of thermally grown oxide were unsatisfying due to problems with the specimen preparation. However, a tendency for the damage development were seen. The finite element calculations gave values for the energy release rate the stress intensity factors in mode~I and mode~II that can be used in the life model. The tensile test showed that the failure mechanism is dependent of the coating thickness and it gave a rough value of the maximum strain acceptable.

Superalloys used in aeroengines are designed to offer superior strength at increasingly higher operating temperatures. In order to optimise the working efficiency and provide additional protection to the components such as turbine blades; a thermal barrier coating (TBC) system is applied. The TBC is a multilayer system consisting of mainly two layers i.e. bond coat (BC) and topcoat (TC). In addition, a third layer grows between the TC and BC during oxidation known as a reaction layer or thermally grown oxide (TGO). The function of the TC (usually, yttria stabilised zirconia (YSZ)) is to provide thermal insulation to aeroengine parts or reduce their surface temperatures; whereas, the BC provides binding between the TC and the substrate, and oxidation resistance to the underlying alloy by forming an adherent and continuous oxide i.e. α-Al2O3. During service, in the absence of mechanical damage to the TBC, most failures are attributed to the BC performance. The most frequently adopted BCs are; β-(Pt, Ni)Al, Pt-γ-Ni/γ’-Ni3Al and MCrAlY. In addition, reactive elements (REs) are incorporated in the BCs due to their ability to enhance oxidation resistance significantly. In the present study βNiAl based coatings/BCs and alloys with and without REs (Zr and Hf) and Pt were prepared. For the coatings CMSX-4 single crystal superalloy was used as a substrate material and pack aluminising/cementation or in-situ chemical vapour deposition (CVD) as a coating process. The isothermal oxidation testing was carried out at 1150oC for 50 and 100 hours in air. The preparation and oxidation performance of a δNi2Al3 coating was carried out, as, this is a starting material for βNiAl matrix based coatings/or BCs. The oxidation of δNi2Al3 coating showed large volumetric changes (thickness variations), multiphase TGO, TGO/coating interface melting and spallation during oxidation. In contrast, the ‘simple βNiAl’ coating (or βNiAl matrix) was found to exhibit comparably enhanced thermal stability than that of the δNi2Al3 coating. Moreover, a detailed study of the simple βNiAl coating was also carried out in order to understand the oxidation performance. The coating before oxidation in the as-deposited condition was found to contain residual compressive stresses of 140 – 200 MPa. In contrast, after oxidation analysis exhibited substantial interdiffusion between the coating and the substrate resulting in a large reduction of the Al content and influx of substrate elements into the coating. This in turn caused coating transformation from βNiAl to the γ’-Ni3Al phase and formation of a multiphase TGO (TiO2, NiAl2O4, and ϴ-Al2O3 intrusion in α-Al2O3). Moreover, the degree of the TGO spallation and residual stresses increased with the oxidation time. In order to enhance the oxidation performance of the βNiAl coatings, the substrate pre-treatment was carried out i.e. CMSX-4 superalloy was electrolytically etched to remove the γ-Ni phase and fabricate βNiAl coatings on the remaining γ’-Ni3Al. This coating is termed as E-βNiAl. In comparison to simple βNiAl, the E-βNiAl coating showed improved spallation resistance. However, E-βNiAl revealed increased surface area due to etching of the substrate and triggered fast TGO growth rates when tested in an un-polished condition. Furthermore, simple βNiAl coatings were doped with Zr and Hf separately using a two-step aluminising method. The appropriate addition of either Zr or Hf was found to reduce the substrate elements (W, Ta, Cr and Ti etc.) in the coating before and after oxidation. After oxidation, examination of the presence of Zr or Hf in the coating was found to confirm the commonly reported beneficial effects. The TGOs grown on these coatings were almost pure α-Al2O3 which subsequently reduced growth and stresses. In addition to Zr/& Hf doped coatings, a study on Hf and Zr doped βNiAl bulk alloys was also carried out in order to understand the dopant effects on the oxidation resistance of βNiAl alloys in the absence of interdiffusion (as in case of coatings). In general, the commonly reported oxidation benefits were confirmed by the addition of these elements such as reduced TGO growth, oxide pegging, a columnar morphology of the TGO and segregation of REs at alumina grain boundaries etc. In addition, two more beneficial effects are suggested to be the ‘TGO crack filling up (or crack-healing)’ and formation of the ‘dense-TGO’. Within this study, the investigation of commercially available Pt-βNiAl BC was also carried out in air and vacuum atmospheres. The results demonstrated that the initial chemistry and elemental distribution (particularly Al/& Pt) was found to affect the TGO growth and phases significantly. In addition to its well established beneficial effects, the main effect of a Pt addition is suggested to be the stabilisation of the βNiAl structure even at a lower Al content.

Etwa 27000 Flugzeuge durchqueren täglich den Luftraum über Europa. Dieser weiter steigende Flugverkehr erfordert neue Richtlinien für die Luftfahrzeuge. Im Besonderen stehen CO2- und NOX-Emission, Kerosinverbrauch und Lärmbelastung unter Optimierungsbedarf. Diese Anforderungen wurden bis 2050 vom Advisory Council for Aerospace Research in Europe (kurz: ACARE) festgelegt und werden wissenschaftlich unterstützt [3, 4]. Um diese Ziele zu erreichen, gibt es verschiedene Forschungsprogramme, Clean Sky ist ein EU-Technologieprogramm davon. In diesem Projekt werden sechs Demonstrator-Programme entwickelt, von denen MTU Aero Engines eines gestaltet. Im Rahmen dieses Projektes wurde eine Weiterentwicklung des Getriebefan (Geared Turbofan-GTF) erreicht, bei dem Fan und Niederdruckturbine durch ein Getriebe voneinander entkoppelt sind. Durch die optimierte Drehzahl beider Komponenten (vergrößerter Fan - langsamer, Niederdruckturbine (LPT) - schneller) wird die Turbinenleistung gesteigert und gleichzeitig die Geräuschemission minimiert. Entwickelt wurde der GTF von Pratt & Whitney in Kooperation mit MTU Aero Engines. Herkömmliche Varianten sehen vor, dass die Niederdruckturbine u.a. den Fan antreibt und zwar nur so schnell, dass der äußere Radius des Fans die zulässige Geschwindigkeit nicht überschreitet. Die herkömmlich verwendeten Nickelbasislegierungen in der Niederdruckturbine haben mit 8 g/cm3 eine zu hohe Dichte um einige Anforderungen im ACARE wirtschaftlich erfüllen zu können. Bereits 1967 hat die US Airforce das große Potential zur Gewichtsreduzierung durch Titanaluminid-Legierungen (TiAl-Legierungen) mit einer Dichte von rund 4 g/cm3 im Hochtemperaturbereich der Flugzeugtriebwerke erkannt. Zwischen 1980 und 1990 entwickelte das General Electric-Forschungscenter die gamma-TiAl-Legierung Ti-48Al-2Cr-2Nb, welche als erste kommerzielle Titanaluminidlegierung in der Niederdruckturbine von Flugzeugtriebwerken eingesetzt wurde. Eine weitere Legierung dieser Werkstoffgruppe kam erst ca. 15 Jahre später zum Einsatz, die TNM-Legierung. Wie man an diesem Beispiel sehen kann, dauert die Integration neuer Werkstoffe in der Luftfahrt aufgrund der notwendigen Vorversuche und Sicherheitsaspekte teilweise 20 Jahre. Seit September 2014 kommt im Triebwerk PW1100G GTF von Pratt & Whitney die geschmiedete Version der TNM-Legierung zum Einsatz. MTU Aero Engines AG München baut hierfür die Niederdruckturbine. Durch die hervorragenden Hochtemperatureigenschaften der gamma-TiAl-Legierungen wie z.B. thermische Stabilität der Mikrostruktur, Resistenz gegen Titanfeuer und hohe spezifische Fes-tigkeit, konnten sich die Titanaluminide in Konkurrenz zu den Nickelbasislegierungen sehr gut platzieren. Deswegen werden die beiden gamma-TiAl-Legierungen (Ti-48Al-2Cr-2Nb, TNMTM) bereits in den letzten Stufen der Niederdruckturbine eingesetzt. Ein Nachteil der gamma-Titanaluminide ist die begrenzte Oxidationsbeständigkeit über 750 °C, wodurch das Einsatzfeld als Hochtemperaturwerkstoff stark begrenzt wird. Um das Anwen-dungspotential der gamma-Titanaluminide weiter zu steigern und auch bei Temperaturen über 750 °C einzusetzen, ist eine Steigerung der Oxidationsbeständigkeit notwendig. Die Oxidationsbeständigkeit kann durch das Aufbringen von Oxidationsschutzschichten wie z.B. Al2O3 erreicht werden. Welche neben der Korrosionsbeständigkeit auch die thermisch-mechanischen Anforderungen des Substrat-Schicht-Verbundes sicherstellen müssen. Zur Erhöhung der Temperaturbelastbarkeit von gamma-TiAl-Schaufeln können zur thermischen Isolation keramische Wärmedämmschichten (WDS) aufgebracht werden. Aufgrund der WDS können höhere Prozesstemperaturen realisiert und die Lebensdauer des Grundwerkstoffs verlängert werden. Die Lebensdauer der Wärmedämmschichten und das Betriebsverhalten werden unter anderem durch eine gute Haftung auf dem Untergrund, eine niedrige Wärmeleitfähigkeit und einen thermisch stabilen Phasenaufbau bestimmt. Die Kombination aus Oxidationsschutz und Wärmedämmung wird bereits für Nickelbasislegierungen in der Brennkammer und Hochdruckturbine der Flugzeugtriebwerke eingesetzt. Um gamma-Titanaluminide in weitere Stufen der Niederdruckturbine oder Hochdruckturbine einzubringen, müssen diese Temperaturen von mindestens 900 °C aushalten und erfordern ebenso Beschichtungen zum Oxidations- und Wärmeschutz. Diese Schutzschichten finden für gamma-Titanaluminide bisher jedoch noch keine Anwendung.

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The operating efficiency and reliability of modern jet engines have undergone significant improvement largely owing to the advances of the materials science over the past 60 years. The use of both superalloys and TBCs in engine components such as turbine blades has made it possible for jet engines to operate at higher temperatures, allowing an optimal balance of fuel economy and thrust power. Despite the vast improvement in high temperature capability of superalloys, the utilization of TBCs has brought the concern of coating adhesion during their usage. TBCs are prone to spallation failure due to interfacial rumpling, which is driven primarily by thermal coefficient mismatch of the multi-layered structure. Although interfacial degradation of TBCs has been widely studied by detailed numerical and analytical models, the predicted results (i.e. stress state and rumpling amplitude) often deviate from that obtained by experiments. This is largely due to the lack of consideration of the influence of bond coat and substrate chemistry on the interfacial evolution of TBC systems. It is only in recent year that more and more study has been focused on studying the role of chemistry on the interfacial degradation of TBCs. The purpose of this PhD project is to clarify how the bond coat and substrate chemical compositions dictate the mechanisms of interfacial degradation, leading to the final spallation. A cross-sectional indentation technique was utilized to quantitatively characterize the adhesion of oxide-bond coat interface among 5 systematically prepared TBC systems. The adhesion of isothermally exposed oxide-bond coat interface was then correlated with different microstructure parameters, in an attempt to identify the key parameters controlling the TBC spallation lifetime. EBSD and EPMA analyses were conducted on the bond coat near the oxide-bond coat interface, in order to understand the relationship between the key parameters and specific alloying elements. The results clearly demonstrated that the phase transformation of bond coat near the oxide-bond coat interface plays the dominant role in the degradation of interfacial adhesion. Particularly, the co-existence of gamma prime and martensitic phases, each having very different thermomechanical response under thermal exposure, can generate a misfit stress in the TGO layer, and ultimately causes early TBC spallation. In addition, the phase transformation behavior has been closely associated with the inherent chemistry of the bond coat and substrate.

The upward trend in internal combustion engine efficiency is likely driven by the depletion of fossil fuels. Since no replacement in sight can deliver energy comparable to the conventional oil, there is a need to use it more rationally and effectively. Thermal barrier coatings have been seen for a long time as a solutionto increase the thermal efficiency of gas turbine engines but suffer from the lackof strong applicability in internal combustion engines. This is due to the different restrictions when comparing the environment on the gas turbines and in internal combustion engines. To overcome this problem and, at the same time, expand the application field of thermal barrier coatings, more efforts need to be devoted.In this work, different top coat materials using various deposition techniques were evaluated and categorized in three different thermal barrier coating (TBC) architectures. The first was the lamellar yttria-stabilized zirconia (YSZ) top coat deposited by atmospheric plasma spray (APS), used as a reference sample. The second architecture was a columnar suspension plasma spray (SPS) TBC with YSZ and gadolinium zirconate (GZO) top coat. The SPS process can produce avariety of microstructures, and they were, for the first time, tested in an internal combustion engine. The third architecture was an SPS top coat, with an additional layer on the top, called a sealing layer of either metallic or ceramic material, both never investigated in a diesel engine application earlier. For the thermophysical properties investigation, a combination of laser flashanalysis (LFA) and modeling with object-oriented finite element (OOF) was employed to understand the properties in all the applications. The performance of the coatings was evaluated in two different ways, by thermal cyclic tests, basedon the TBCs behavior under cyclic thermal loads and by single-cylinder engine experiment. The characterization of the coatings was done by scanning electron microscope (SEM) before and after the thermal cyclic tests.The performance properties were correlated with coatings microstructure and thermophysical properties. It was shown that a columnar TBC produced by SPS had a superior engine efficiency in the single cylinder engine experiment.