A process for coating a gas turbine engine disk comprises placing the disk having an outer surface into a chamber, the chamber configured to perform atomic layer deposition; injecting a first reactant into the chamber; forming a first monolayer gas thin film on the outer surface; removing the first reactant from the chamber; injecting a second reactant into the chamber; reacting second reactant with the first monolayer gas thin film; removing the second reactant from the chamber; and forming a protective barrier coating on the outer surface.

A process for coating a gas turbine engine disk comprises placing the disk having an outer surface into a chamber, the chamber configured to perform atomic layer deposition; injecting a first reactant into the chamber; forming a first monolayer gas thin film on the outer surface; removing the first reactant from the chamber; injecting a second reactant into the chamber; reacting second reactant with the first monolayer gas thin film; removing the second reactant from the chamber; and forming a protective barrier coating on the outer surface.

The present invention provides a thermal barrier coated component, monitoring or evaluation of the soundness of which is able to be adequately carried out on the basis of the thermal boundary conditions that are detected by a sensor. A thermal barrier coated component according to the present invention comprises: a base material; a first bond coat layer that is a metal bonding layer formed on the base material; a sensor unit that comprises a sensor and a conductive wire, which are formed on the first bond coat layer; a second bond coat layer that is formed on the first bond coat layer so as to cover at least the sensor unit, while having a surface roughness higher than that of the first bond coat layer; and a top coat layer that is formed on the second bond coat layer.

A turbine system including a sealing component is presented. The sealing component includes a ceramic material. The ceramic material includes grains having an average grain size of less than 10 microns. A turbine shroud assembly including the sealing component is also presented.

High temperature stable thermal barrier coatings useful for substrates that form component parts of engines such as a component from a gas turbine engine exposed to high temperatures are provided. The thermal barrier coatings include a multiphase composite and/or a multilayer coating comprised of two or more phases with at least one phase providing a low thermal conductivity and at least one phase providing mechanical and erosion durability. Such low thermal conductivity phase can include a rare earth zirconate and such mechanical durability phase can include a rare earth a rare earth aluminate. The different phases are thermochemically compatible even at high temperatures above about 1200° C.

Methods for repairing the conventional physical barrier coating barrier on a component having a damaged portion. Applying a coating on the outer surface of the damaged portion of the component. The coating containing a reactive oxide. Initiating a reaction between the coating and the molten sulfates within the outer surface of the component. The reaction catalytically decomposes molten sulfates at the outer surface of the damaged portion of the component.

A thermal barrier coated component, such as a turbine blade formed from a superalloy substrate, includes a thermal barrier coating applied onto the substrate. A metallic bond coat layer is on the substrate and includes rare-earth luminescent dopants. A ceramic top coat layer is on the bond coat layer. A temperature sensing thermally grown oxide (TGO) layer is formed at the interface of the bond coat layer and ceramic top coat layer. The temperature sensing TGO layer includes grown rare-earth luminescent ions migrated from the metallic bond coat layer in an amount sufficient to enable luminescence sensing of the TGO layer for real-time phosphor thermometry temperature measurements at the TGO layer.

The disclosure relates to a method for operating a gas turbine at a high temperature and to a gas turbine assembly. In the method, a gas turbine having a structural material and a thermal barrier layer disposed on the structural material is cooled down in a decelerated manner after operation at an operating temperature above 1000° C., so that damage to the structural material and/or the thermal barrier layer is minimized. In this way, the gas turbine can be operated permanently at temperatures above 1500° C.

A coated substrate comprises: a metallic substrate; a bondcoat atop the substrate; and a ceramic barrier coat atop the bondcoat. The bondcoat has a combined content of one or more of molybdenum, chromium, and vanadium of at least 50 percent by weight.

An article has a metallic substrate having a plurality of recesses. A first coating is at least at the recesses and has: a splatted layer; and a columnar layer atop the splatted layer. A second coating is away from the recesses and has: a columnar layer atop the substrate without an intervening splatted layer.