ABSTRACT

In order to increase the gas turbine inlet temperature, thermal barrier coatings (TBCs) applied on gas turbine components have been investigated and developed for over forty years. Thermal barrier coatings generally exhibit nonuniform microstructures that include voids, unmelted particles and oxide inclusions and may also demonstrate poor bonding adhesion between the TBC and the bond coat which lowers the overall thermal shock resistance of thermal barrier coatings. The performance of thermal barrier coatings is dependent on its: (1) chemistry, (2) microstructure, (3) stress magnitude and interactions, and (4) environmental interactions. In a hot corrosion environment, the absorption of corrosion products into the porous structure tends to cause premature fracture of the thermal barrier coatings. The paper discusses the historical and current understanding of thermal barrier coatings and possible directions to improve performance in marine gas turbine engines.

INTRODUCTION

Economic and environmental concerns have forced development of gas turbines with lower fuel consumption. One way of decreasing fuel consumption is to increase turbine inlet temperature. This, to date, has required advanced high temperature superalloys that require massive coolant flow which reduces overall engine efficiency. There have been several efforts to replace superalloys with monolithic ceramics, but the brittle behavior of many ceramics has prevented their use in most turbine applications. The idea of introducing ceramic-based thermal barrier coatings on superalloys was to combine the good mechanical behavior of alloys with the low thermal conductivity, insulating capacity of selected ceramics.

Insulating alloys in gas turbine components with thin layers of low thermal conductivity ceramics is almost as old as gas turbine cooling technology itself. Uncoated metals and superalloys have an upper operating temperature range of about 1100°C. Refractories can operate at temperatures up to approximately 3000°C. Coating superalloys with selected low-conductivity ceramics extends the use of metallic components up to 1400-1500°C.

The principle function of thermal barrier coatings is to insulate the underlying substrate alloy from prevailing high temperature conditions. A second benefit of TBCs is the lowering of thermally induced strains, thereby improving thermal fatigue resistance. The magnitude of the these benefits is determined by several factors including coating thickness, thermal conductivity, surface finish of the ceramic topcoat, temperature and heat transfer coefficient in the gaseous environment, and corrosion and other degradation processes.

Theoretical design evaluations of ceramic coated airfoils were reported in the Russian technical literature as early as 1949 and model systems were evaluated by the U.S. National Aeronautics and Space Administration (NASA) in the early fifties [1]. The early difficulty of obtaining reproducible long-term adherence of ceramic coatings to practical airfoil components impeded widespread use. Practical use of ceramic thermal barrier coatings on less demanding, less complex components began in the late sixties.

Efforts to implement ceramic coating technology for turbine airfoil protection continued in several laboratories at low levels of funding and associated efforts through the early seventies [1]. In 1976, NASA demonstrated significant durability of thermal barrier coatings on first stage blades of a test turbine which sparked notable increases in funding and effort to achieve long-term durability of hot section airfoils. Several concentrated programs from 1976 to 1984 achieved significant improvements in the understanding of the e

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