Introduction

Engineers have recognized since the early nineteenth century that the performance and efficiency of a heat engine improves with increasing operating temperature (ref. 1). However, maximum operating temperature are generally limited by the temperature capabilities of the materials used in the engine. Because of this, "exotic" superalloys and advanced cooling schemes have been developed for components in gas turbine engines. However, those technologies have now been pushed to near their limits. Higher temperatures may be attainable in future designs through the use of metal matrix composite materials but many years will be required before they could be available commercially. Thermal barrier coatings are the one class of materials which have the potential to allow increased temperatures in both the near term and the future. This approach relies on the application of a thin layer of a thermally insulating ceramic to the surface of metal components. Although this concept has been recognized since the early days of gas turbine engines (ref.2), materials development proved to be very challenging. A breakthrough was reported in the mid 70's when a thermal barrier coating system was developed at NASA that was tested successfully on the turbine blades in a J-75 research gas turbine engine. This successful test, detailsof which may be found in references 3 and 4, proved that the concept was feasible. However, when this early coating system was exposed to higher temperatures in a more advanced JT9D research engine, localized spalling was observed (ref. 5).About ten more years of research and development resulted in improved coatings which were sufficiently durable to use in the turbine section of jet engines. Much of the research during this period was conducted either by or in behalf of NASA (e.g. refs. 6-12).

The use of thermal barrier coatings in the turbine section of commercial engines is currently limited to the vane platforms. This is an important but relatively conservative location. The coatings which are flying consist of a thin -- about 0.25mm -- layer of ambient pressure plasma sprayed zirconia-7% yttria over an approximately 0.12mm thick layer of an MCrAlY bond coat. Figure 1 shows one of these coatings schematically. The metallic bond coat for this high temperature application is typically low pressure plasma sprayed for low porosity. It may also be possible to produce dense metallic coatings by shrouded ambient pressure plasma spraying (ref.13). The structure of the ceramic layer that is produced by ambient pressure plasma spraying is porous and highly microcracked. This high-defect structure is believed to be necessary for high strain tolerance and low thermal conductivity. Substrates for this aeronautical application are air-cooled nickel- or cobalt-based superalloys. Figure 2 illustrates schematically a thermal barrier coating insulating an air-cooled component from the hot gases in a heat engine. Additional details are given in reference 14. The.introduction of thermal barrier coatings into commercial gas turbine engines on the vane platforms was an important milestone which is as significant as the initial research engine test of the previous decade. However, further advances in this technology will ,be required to improve the reliability of thermal barrier coatings so that they may be used on entire turbine blades and vanes and exposed to higher gas temperatures. Future coatings may be based on new approaches to plasma spraying or alternative processes such as physical vapor deposition (ref. 12).

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