High-energy X-ray phase analysis of CMAS-infiltrated 7YSZ thermal barrier coatings: Effect of time and temperature
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High-energy X-ray phase analysis of CMAS-infiltrated 7YSZ thermal barrier coatings: Effect of time and temperature Zachary Stein1, Peter Kenesei2, Jun-Sang Park2, Jonathan Almer2, Ravisankar Naraparaju3, Uwe Schulz3, Seetha Raghavan1,a) 1
Department of Mechanical & Aerospace Engineering, University of Central Florida, Orlando, Florida 32816, USA Advanced Photon Source, Argonne National Laboratory, Lemont, Illinois 60439, USA 3 Institute of Materials Research, German Aerospace Centre (DLR), Cologne 51170, Germany a) Address all correspondence to this author. e-mail: [email protected] 2
Received: 6 March 2020; accepted: 23 July 2020
Calcium–magnesium–alumino-silicate (CMAS) particulates enter the aero-engine in a sandy environment, melt and infiltrate into 7 wt% yttria-stabilized zirconia (7YSZ) thermal barrier coatings (TBCs), reducing their lifetime. This leads to chemical degradation in 7YSZ accompanied by tetragonal to monoclinic phase transformation upon cooling. In this work, electron-beam physical vapor deposition coatings were infiltrated with a synthetic CMAS. Synchrotron X-ray diffraction measurements show that CMAS infiltration at 1250 °C has about 43% higher monoclinic phase volume fraction (PVF) at the coating surface compared to 1225 °C and remains consistently higher throughout the coating depth. Additionally, the increase in annealing time from 1 to 10 h results in a 31% higher monoclinic phase at the surface. Scanning electron microscopy revealed the presence of globular monoclinic phases corresponding spatially with the above findings. These results resolve the impact of time and temperature on CMAS infiltration kinetics which is important for mitigation.
Introduction Thermal barrier coatings (TBCs) are applied to superalloy metallic turbine blades within gas turbine engines to lower the temperatures by approximately 200 °C, thereby improving the durability and the lifetime of the blades during operation [1,2]. TBC systems typically comprise of a 150–200-μm thick high-temperature ceramic coating of 7 wt% yttria-stabilized zirconia (7YSZ) consisting of a nontransformable tetragonal prime YSZ phase (tYSZ), a metallic bond coat layer (MCrAlY) that adheres the ceramic coating to the turbine blade, and a thermally grown oxide layer (TGO) of Al2O3 [3]. Electron-beam physical vapor deposition (EB-PVD) coatings are generally applied on rotating parts due to their higher in-plane strain tolerance and lower thermal conductivity. These properties in EB-PVD coatings are a direct result of the columnar microstructure that forms during deposition [4,5]. Calcium–magnesium–alumino-silicate (CMAS) compositions, which are dominant in particulates such as sand or
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volcanic ash, may enter the engine from the surrounding environment during aircraft operation. The increased operating temperatures in gas turbine engines, aided by TBCs, also allows for CMAS to become molten in the combustor section. The melting points of CMAS compositions vary geographically and are heavily base
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