Phases in Plasma Sprayed Thermal Barrier Coatings from Yttria Partially Stabilized Zirconia

The application of thermal barrier coatings (TBC) is increasing in aeroengines. Surface temperatures up to 1450°C require the application of ceramic TBCs because the temperature capability of metallic substrate materials is not high enough. The service life of turbine components could be improved by the use of yttria partially stabilized zirconia top coatings. The most successfull TBCs are made from 7–9 wt‐% yttria partially stabilized zirconia. One of the most discussed reasons of damages of such TBCs is the transformation between monoclinic and tetragonal phase in zirconia in connection with a dramatic change in volume. Thus in this work resulting phases of plasma sprayed zirconia coatings were investigated. It was found that no monoclinic phase could be detected after heat treatments at 1300, 1400 and 1466°C with cooling rates > 2°/min. Only with cooling rates < 2°/min monoclinic phases occured. It can be concluded that the metastable tetragonal high temperature configuration of yttria partially stabilized zirconia is “very stable”. The conditions in aeroengines with cooling rates > 2°C prevent the formation of the monoclinic phase in zirconia.     

Comparative analysis of steady state heat transfer in a TBC and functionally graded air cooled gas turbine blade

Internal cooling passages and thermal barrier coatings (TBCs) are presently used to control metal temperatures in gas turbine blades. Functionally graded materials (FGMs), which are typically mixtures of ceramic and metal, have been proposed for use in turbine blades because they possess smooth property gradients thereby rendering them more durable under thermal loads. In the present work, a functionally graded model of an air-cooled turbine blade with airfoil geometry conforming to the NACA0012 is developed which is then used in a finite element algorithm to obtain a non-linear steady state solution to the heat equation for the blade under convection and radiation boundary conditions. The effects of external gas temperature, coolant temperature, surface emissivity changes and different average ceramic/metal content of the blade on the temperature distributions are examined. Simulations are also carried out to compare cooling effectiveness of functionally graded blades with that of blades having TBC. The results highlight the effect of including radiation in the simulation and also indicate that external gas temperature influences the blade heat transfer more strongly. It is also seen that graded blades with about 70% ceramic content can deliver better cooling effectiveness than conventional blades with TBC.     

Design and modeling of multiple layered TBC system with high reflectance

Ceramic thermal barrier coatings (TBCs) are playing an increasingly critical role in advanced gas turbine engines due to their ability to sustain further increases in operating temperatures. However, these increases in temperature could raise considerable issues associated with increased radiative heat transfer into the TBC systems. This study was conducted to design a ceramic based multiple layered TBC system with high reflectance to radiation. Mathematical modeling was used to calculate the potential temperature reduction on the substrate surface when the multiple layered TBC is applied. The result of the simulation shows that a temperature reduction up to 90 °C is possible when utilizing the designed multiple layered TBC coatings.     

Investigation on the thermo-chemical reaction mechanism between yttria-stabilized zirconia (YSZ) and calcium--magnesium--alumino-silicate (CMAS)

Thermal barrier coatings (TBCs) with Y₂O₃-stabilized ZrO₂ (YSZ) top coat play a very important role in advanced turbine blades by considerably increasing the engine efficiency and improving the performance of highly loaded blades. However, at high temperatures, environment factors result in the failure of TBCs. The influence of calcium--magnesium--alumino-silicate (CMAS) is one of environment factors. Although thermo-physical effect is being paid attention to, the thermo-chemical reaction becomes the hot-spot in the research area of TBCs affected by CMAS. In this paper, traditional two-layered structured TBCs were prepared by electron beam physical vapor deposition (EB-PVD) as the object of study. TBCs coated with CMAS were heated at 1240°C for 3 h. Additionally, 15 wt.% simulated molten CMAS powder and YSZ powder were mixed and heated at 1240°C or 1350°C for 48 h. SEM and EDS were adopted to detect morphology and elements distribution. According to XRD and TEM results, it was revealed that CMAS react with YSZ at high temperature and form ZrSiO₄, Ca_0.2Zr_0.8O_1.8 and Ca_0.15Zr_0.85O_1.85 after reaction, as a result, leading to the failure of TBCs and decreasing the TBC lifetime.     

Measurements of the thermal gradient over EB-PVD thermal barrier coatings

Conventional two-layered structure thermal barrier coatings (TBCs), graded thermal barrier coatings (GTBCs) and graded thermal barrier coatings with micropores were prepared onto superalloy DZ22 tube by electron beam physical vapor deposition (EB-PVD). Thermal gradient of the TBCs was evaluated by embedding two thermal couples in the surfaces of the tube and the top coat at different surrounding temperatures with and without cooling gas flowing through the tube. The results showed that higher thermal gradient could be achieved for the GTBCs with micropores compared to the two-layered structure TBCs and GTBCs. However, after the samples were heated at 1050°C, the thermal gradient for the GTBCs with or without micropores decreased with the increase of heating time. On the other hand, the thermal gradient for the TBCs increased with the increase of heating time. Cross-section observations by scanning electron microscopy showed that the change in microstructure was the main reason for the change of the thermal gradient.     

EB-PVD process and thermal properties of hafnia-based thermal barrier coating

Thermal barrier coatings (TBCs) are being developed for the key technology of gas turbine and diesel engine applications. In general, 8 mass% Y₂O₃–ZrO₂ (8YSZ) coating materials are used as the top coating of TBCs. The development of hafnia-based TBC was started in order to realize the high reliability and durability in comparison with 8YSZ, and the 7.5 mass% Y₂O₃–HfO₂ (7.5YSH) was selected for coating material. By the investigation of electron-beam physical vapor deposition (EB-PVD) process using 7.5YSH ceramic ingot, 7.5YSH top coating with about 200 µm thickness could be formed. The microstructure of the 7.5YSH coated at coating temperature of 850 °C showed columnars of laminated thin crystals. On the other hand, the structure of the 7.5YSH coated at coating temperature of 950 °C showed solid columnars. From the result of sintering behavior obtained by heating test of 7.5YSH coating, it was recognized that the thermal durability of 7.5YSH coating was improved up to about 100 °C in comparison with 8YSZ coating. This tendency was confirmed by the experimental result of the thermal expansion characteristics of sintered 7.5YSH and 8YSZ.

©2003 Elsevier Science Ltd. All rights reserved.     

EB‐PVD Thermal Barrier Coatings for Aeroengines and Gas Turbines

Ceramic thermal barrier coatings (TBCs) offer the potential to significantly improve efficiencies of aero engines as well as stationary gas turbines for power generation. On internally cooled turbine parts temperature gradients of the order of 100 to 150 °C can be achieved. Today, state‐of‐the‐art TBCs, typically consisting of an yttria‐stabilised zirconia top coat and a metallic bond coat deposited onto a superalloy substrate, are mainly used to extend lifetime. Further efficiency improvements require TBCs being an integral part of the component which, in turn, requires reliable and predictable TBC performance. Presently, TBCs fabricated by electron beam physical vapor deposition are favoured for high performance applications. The paper highlights critical research and development needs for advanced TBC systems, such as reduced thermal conductivity, increased temperature capability, lifetime prediction modelling, process modelling, bond coat oxidation, and hot corrosion resistance as well as improved erosion behaviour.     

Performance of thermal barrier coatings in industrial gas turbine conditions

Abstract

The effect of aerofoil geometry on the oxidative degradation mechanisms experienced by thermal barrier coatings (TBCs) used on industrial turbine blades has been investigated. Modified aerofoil-shaped samples (CMSX4 coated with high-velocity oxy-fuel sprayed AMDRY 995 and air plasma sprayed TBC) were oxidised at five temperatures in furnaces from 900 to 1000°C. Scanning electron microscopy and energy dispersive X-ray analysis were used to characterise details of the microstructural evolution of the thermally grown oxide and to monitor inter-diffusion between the bond coating and substrate. Additionally, a novel non-destructive examination technique (flash thermography) was used to detect and track the spread of cracks beneath the TBCs. Multiple samples cracking in identical locations suggested an effect of geometry in the failure of coatings. Furthermore, it was observed that coating curvature influenced spinel formation.     

The degradation of thermal barrier coatings by molten deposits: introducing the concept of basicity

Airborne particles such as dust, sand and volcanic ash, otherwise termed ‘CMAS’ (Calcia Magnesia Alumina-Silicate) are ingested into the engine where at the elevated temperatures this deposit can melt and induce degradation of the thermal barrier coating (TBC). The columnar, high porosity microstructure of electron beam physical vapour deposited (EB-PVD) TBCs, typically found on HP turbine blades, makes them particularly susceptible to CMAS molten deposit attack. This paper investigates the effect of modifying the chemical composition of molten CMAS deposits and, therefore, the degradation mechanisms of EB-PVD TBCs. This paper identifies a parameter termed ‘Basicity Index’, which modifies the severity and damage caused by these molten deposits to TBCs. It is believed that with the higher temperatures likely to result in next generation aero-engines and with sand chemistry varying around the world, basicity driven TBC degradation will become more prevalent.     

热障涂层热冲击试验研究

随着先进发动机的研制和生产,对高温涂层的要求也不断提高。热障涂层（TBCs) 作为一种新型隔热涂层,在发动机涡轮叶片上的应用得到更多的重视。对利用电子束物理气相沉积（EB－PVD）制备的TBCs进行了高温热冲击试验,并对试验前后的TBCs进行了形貌、SEM和XRD分析。     

Multidisciplinary analysis of the operational temperature increase of turbine blades in combustion engines by application of the ceramic thermal barrier coatings (TBC)

The improvement of the temperature resistance of the aircraft engine elements can be obtained by application of a single ceramic thermal barrier coating (TBC) (e.g. Noda [1]) or several composite layers (e.g. Sadowski [2]). Engine elements protected by TBC can work safely in elevated temperature range above 1000 °C. Continuous endeavour to increase thermal resistance of engine the elements requires, apart from laboratory investigations, also numerical study of the different aero-engine parts. The most important are turbine blades, where high temperatures and stress concentrations during thermal shocks or thermal fatigue can be observed during engine exploitation. The high temperatures and stress concentrations can act as the local sources of damage initiation and defects propagation in the form of cracks.The present paper deals with the solution of the transient temperature transfer problem in bare and thermal barrier coated alloy Inconel 713 for the temperature range up to 1000 °C. The computational fluid dynamics (CFD) part of analysis was performed by application of ANSYS Fluent code receiving the temperature field of combustion gas, whereas computational structural mechanics (CMS) part concerning the temperature distribution inside the turbine blade was done by ABAQUS. Finally, the efficiency of the TBC layer (0.5 mm thickness) protecting and cooling channels was discussed in order to explore the operational temperature increase in the aero-engines.     

Thermocyclic Behavior of Differently Stabilized and structured EB-PVD thermal barrier coatings

The electron-beam physical vapor deposition (EB-PVD) process provides distinctive coatings of a unique columnar microstructure for gas turbine components. Main advantage of this structure is superior tolerance against straining, erosion and thermoshock, thus giving it a major edge in lifetime. This paper outlines the interaction between chemical composition and microstructural evolution EB-PVD zirconia-based thermal barrier coatings (TBCs) and their respective lifetimes in cyclic burner rig and furnace tests. Customizing TBC microstructure by adjusting EB-PVD processing parameters is emphasized. A structural zone diagram for PVD is modified by interconnecting the influence of substrate rotation with microstructural evolutions. Finally, some basic aspects of single source and dual source evaporation are compared.     

Optimizing Heat Treatment of Gas Turbine Blades with a Co–C Fixed Point for Improved In-service Thermocouples

Improvement of energy efficiency of jet aircraft is achieved by operating gas turbine engines at higher temperatures. To facilitate this, gas turbine engine manufacturers are continuously developing new alloys for hot-zone turbine blades that will withstand the increased in-service temperatures. A critical part of the manufacture of these blades is heat treatment to ensure that they attain the necessary metallurgical characteristics. Current heat-treatment temperature-control requirements are at the limit of what is achievable with conventional thermocouple calibrations. A project that will allow thermocouple manufacturer CCPI Europe Ltd. to realize uncertainties of ± 1°C, or better, in the calibration of its noble metal thermocouples is described. This will be realized through implementing a Co–C eutectic fixed point in CCPI’s calibration chain. As this melts at 1,324°C, very close to the heat-treatment temperatures required, low uncertainties will be obtained. This should yield an increase in effectiveness of the heat-treatment process performed by Bodycote Heat Treatments Ltd., allowing them to respond effectively to the increasingly stringent demands of engine manufacturers. Outside the current project, there is a strong requirement by industry for lower uncertainties at and above 1,300°C. Successful implementation of the current fixed point in an industrial setting is likely to result in rapid take-up by other companies, probably through the supply of ultra-low uncertainty thermocouples, looking to improve their high-temperature processes.     

Overview on the Development of Nanostructured Thermal Barrier Coatings

Thermal barrier coatings （TBCs） have successfully been used in gas turbine engines for increasing operation temperature and improving engine efficiency. Over the past thirty years, a variety of TBC materials and TBC deposition techniques have been developed. Recently, nanostructured TBCs emerge with the potential of commercial applications in various industries. In this paper, TBC materials and TBC deposition techniques such as air plasma spray （APS）, electron beam physical vapor deposition （EB-PVD）, laser assisted chemical vapor deposition （LACVD） are briefly reviewed. Nanostructured 7-8 wt pct yttria stabilized zirconia （7-8YSZ）TBC by air plasma spraying of powder and new TBC with novel structure deposited by solution precursor plasma spray （SPPS） are compared. Plasma spray conditions, coating forming mechanisms, microstructures,phase compositions, thermal conductivities, and thermal cycling lives of the APS nanostructured TBC and the SPPS nanostructured TBC are discussed. Research opportunities and challenges of nanostructured TBCs deposited by air plasma spray are prospected.     

Effect of protective coatings on mechanical properties of ZhS32-VI heat-resistant alloy

We discuss the effect of diffusion heat-resistant and condensation gradient thermal barrier coatings used for protection of gas turbine disk operating blades on mechanical characteristics (long- and short-term durability and plasticity at temperature of 1000°C) of specimens from ZhS32-VI heat-resistant alloy. We study the effect of isothermal annealing in air at temperature 1100°C during 240 h on mechanical characteristics of uncoated and coated specimens. We note the expediency of application of gradient thermal barrier coating consisting from external heat-insulating ceramic layer and thermostable heat-resistant metal layer, which provide temperature reduction of metal walls of cooled blades made from ZhS32-VI heat-resistant alloy with the purpose of achieving the required residual life.     

Erosion of thermal barrier coatings

Abstract

Thermal barrier coatings have been used within gas turbines for over 30 years to extend the life of hot section components. Thermally sprayed ceramics were the first to be introduced and are widely used to coat combustor cans, ductwork, platforms and more recently turbine aerofoils of large industrial engines. The alternative technology, electron beam physical vapour deposition,(EB-PVD) has a more strain-tolerant columnar microstructure and is the only process that can offer satisfactory levels of spall resistance, erosion resistance and surface finish retention for aero-derivative engines.

Whatever technology is used, the thermal barrier must remain intact throughout the turbine life. Erosion may lead to progressive loss of TBC thickness during operation, raising the metal surface temperatures and thus shortening component life. Ballistic damage can lead to total TBC removal.

This paper reviews the erosion behaviour of both thermally sprayed and EB-PVD TBCs relating the observed behaviour to the coating microstructure. A model for the erosion of EB-PVD ceramics is presented that permits the prediction of erosion rates. The model has been validated using a high velocity erosion gas gun rig, both on test coupons and samples removed from coated components. The implications of erosion on component life are discussed in the light of experimental results and the model predictions.     

Functionally gradient ceramic/metallic coatings for gas turbine components by high-energy beams for high-temperature applications

Failure of turbine blades generally results from high-temperature oxidation, corrosion, erosion, or combinations of these procedures at the tip, and the leading and trailing edges of a turbine blade. To overcome these limitations, functionally gradient ceramic/metallic coatings have been produced by high-energy beams for high-temperature applications in the aerospace and turbine industries to increase the life of turbine components. Thermal spray processes have long been used to apply high-temperature thermal barrier coatings to improve the life of turbine components. However, these processes have not met the increased demand by the aerospace and turbine industries to obtain higher engine temperatures and increased life enhancement as a result of the inhomogeneous microstructure, unmelted particles, voids, and poor bonding with the substrate. High-energy beams, i.e. electron beam-physical vapour deposition (EB-PVD), laser glazing, laser surface alloying, and laser surface cladding, have been explored to enhance the life of turbine components and overcome the limitations of the thermal spray processes. EB-PVD has overcome some of the disadvantages of the thermal spray processes and has increased the life of turbine components by a factor of two as a result of the columnar microstructure in the thermal barrier coating (TBC). Laser glazing has been used to produce metastable phases, amorphous material, and a fine-grained microstructure, resulting in improved surface properties such as fatigue, wear, and corrosion resistance at elevated temperatures without changing the composition of the surface material. Laser surface alloying and laser surface cladding have shown promising results in improving the chemical, physical, and mechanical properties of the substrate's surface. Metal-matrix composite coatings have also been produced by a laser technique which resulted in increased wear and oxidation-resistant properties. The advantages and disadvantages of thermal spray processes, EB-PVD, laser glazing, laser surface alloying, and laser surface cladding will be discussed. Microstructural evolution of thermal barrier coatings, recent advancements in functionally gradient coatings, laser grooving, and multilayered textured coatings will also be discussed.     

Oxidation behaviour of an intermetallic Ti–Al–Cr–Zr bond coat on a γ–TiAl based TNB alloy with 7YSZ thermal barrier coating

Abstract

Intermetallic titanium aluminide alloys are attractive light-weight materials for high temperature applications in automotive and aero engines. The development of γ-TiAl alloys over the past decades has led to their successful commercial application as low pressure turbine blades. The operating temperatures of γ-TiAl based alloys are limited by deterioration in strength and creep resistance at elevated temperatures as well as poor oxidation behaviour above 800 °C. Since improvement in oxidation behaviour of γ-TiAl based alloys without impairing their mechanical properties represents a major challenge, intermetallic protective coatings have aroused increasing interest in the last years.

In this work, a 10 μm thick intermetallic Ti–46Al–36Cr–4Zr (in at.-%) coating was applied on a TNB alloy using magnetron sputtering. This layer provided excellent oxidation protection up to 1000 °C. Microstructural changes in this coating during the high temperature exposure were extensively investigated using scanning and transmission electron microscopy. The coating developed a three-phase microstructure consisting of the hexagonal Laves-phase Ti(Cr,Al)₂, the tetragonal Cr₂Al phase and the cubic τ-TiAl₃ phase. After long-term exposure the three-phase microstructure changed to a two-phase microstructure of the hexagonal α₂-Ti₃Al phase and an orthorhombic body-centred phase, whose crystal structure has not yet been definitely identified. On the coating, a thin protective alumina scale formed. Applying this intermetallic layer as bond coat, thermal barrier coatings (TBCs) of yttria partially stabilized zirconia were deposited on γ-TiAl based TNB samples using electron-beam physical vapour deposition. The results of cyclic oxidation testing (1 h at elevated temperature, 10 min. cooling at ambient temperature) revealed a TBC lifetime of more than 1000 h of cyclic exposure to air at 1000 °C. The ceramic topcoat exhibited an excellent adhesion to the thermally grown alumina scale which contained fine ZrO₂ precipitates.     

Numerical study on modification of ceramic coatings by high-intensity pulsed ion beam

ZrO₂ ceramic coatings, which often call thermal barrier coatings (TBCs), fabricated by electron beam physical vapor deposition (EB-PVD), are widely used in high-temperature environment of aircraft and industry gas-turbine engines, because of the excellent strain tolerance imparted by the columnar structure. However, channels separating the columnar grains in EB-PVD TBCs provide paths for oxygen or other aggressive species from ambient atmosphere into the bond coat, resulting in the premature spallation-failure during high-temperature service. In our previous study, high-intensity pulsed ion beam (HIPIB) technique has been proposed to modify the EB-PVD TBCs, where a melted, densified top layer can be produced as a result of extremely thermal effect induced by the HIPIB irradiation. In this paper, HIPIB melting process is investigated numerically using a physical model based on experimental data, taking into account the surface morphology of HIPIB-melted TBCs to explore the mechanism of interaction between HIPIB and the coatings. Deposition process of the beam energy in TBCs was simulated by Monte Carlo method, and the non-linear equations describing the thermal conducting process were solved numerically based on the deposited energy to obtain the evolution of the temperature field of TBCs. The calculated melting depth of irradiated EB-PVD TBCs is consistent with results obtained in the HIPIB irradiation experiments.     

Anisotropic Thermal Diffusivities of Plasma-Sprayed Thermal Barrier Coatings

Thermal barrier coatings (TBCs) are used to shield the blades of gas turbines from heat and wear. There is a pressing need to evaluate the thermal conductivity of TBCs in the thermal design of advanced gas turbines with high energy efficiency. These TBCs consist of a ceramic-based top coat and a bond coat on a superalloy substrate. Usually, the focus is on the thermal conductivity in the thickness direction of the TBC because heat tends to diffuse from the surface of the top coat to the substrate. However, the in-plane thermal conductivity is also important in the thermal design of gas turbines because the temperature distribution within the turbine cannot be ignored. Accordingly, a method is developed in this study for measuring the in-plane thermal diffusivity of the top coat. Yttria-stabilized zirconia top coats are prepared by thermal spraying under different conditions. The in-plane and cross-plane thermal diffusivities of the top coats are measured by the flash method to investigate the anisotropy of thermal conduction in a TBC. It is found that the in-plane thermal diffusivity is higher than the cross-plane one for each top coat and that the top coats have significantly anisotropic thermal diffusivity. The cross-sectional and in-plane microstructures of the top coats are observed, from which their porosities are evaluated. The thermal diffusivity and its anisotropy are discussed in detail in relation to microstructure and porosity.