Abstract
The use of thermal barrier coatings (TBCs) on turbine blades and vanes located in the hot sections of gas turbine engines has allowed higher engine operating temperatures leading to temperatures of the order of 1200°C at the surface of the ceramic coating. At such temperatures, under service conditions, thermal barrier coatings are susceptible to corrosion by molten calcium–magnesium–alumino-silicates (CMAS) resulting from the ingestion of siliceous mineral debris (dust, sand, ash) by the engine.This study consists in a microstructural analysis of CMAS induced degradation of standard 8YPSZ EB-PVD thermal barriers observed on high pressure turbine blades of military engines removed from service. The CMAS/TBC interactions are mainly observed in the hottest zones of the blade pressure side. CMAS infiltration in the TBC porous microstructure (inter-columnar gaps, pores, cracks) down to the thermally grown oxide TGO interfacial layer is observed as well as dissolution of the 8YPSZ into the CMAS melt, TBC transformation from tetragonal 8YPSZ to monoclinic Y-depleted zirconia and formation of a Zr-bearing phase at the interface between CMAS and TBC. CMAS not only turns out to be mainly constituted of CaO, MgO, Al2O3, and SiO2 but also contains a large amount of iron oxide Fe2O3. Comparison with previous published data shows that CMAS composition depends on the flight conditions to a large extent. A part from the loss of column integrity and the modification of the porous morphology resulting from the CMAS chemical attack, large vertical separations between highly sintered columns are observed in the CMAS infiltrated TBC as well as delamination cracks in the upper part of the top coat possibly leading to progressive TBC spallation. These results are discussed in the light of similar studies on CMAS/TBC interaction and of existing sintering and delamination mechanisms.
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