A component and method for minimizing a gap between a blade platform ledge and a vane platform ledge in a turbine engine. The blade platform supports turbine blades attached to a shaft. The movable blade platform is positioned adjacent to a stationary vane platform having vanes mounted on an outer surface. The blade and vane platforms are separated by the gap between the platform ledges. During manufacture, an abrasive coating is applied to the surface of the blade platform ledges so that the coating contacts the vane platform ledge when the engine is started. The abrasive coating on the turbine blade platform cuts the surface of the vane platform ledge to form a gap sufficient to permit unobstructed motion of the blade platform, yet of minimal size to limit the flow of gas between the space within and the space outside the platforms.

An abrasive coating for a substrate of a component in a gas path exposed to a maximum temperature of 1750 degree Fahrenheit, comprising a plurality of grit particles adapted to be placed on a top surface of the substrate; a matrix material bonded to the top surface; the matrix material partially surrounds the grit particles, wherein the grit particles extend above the matrix material relative to the top surface; a film of oxidant resistant coating applied over the plurality of grit particles and the matrix material and a thermal barrier coating material applied over said film of oxidant resistant coating.

A control system includes one or more processors configured to determine when to extend a life span of an engine by applying an additional restorative coating to the engine based on one or more monitored parameters of the engine. The monitored parameters include a condition of a previously applied restorative coating. The one or more processors are configured to determine the condition of the previously applied restorative coating based on an optical response of the previously applied restorative coating. The one or more processors also are configured to direct application of the additional restorative coating based on the one or more monitored parameters of the engine.

The use of especially an aluminum alloy on a metal substrate in a PVD-AlTiN coating results in good corrosion and erosion protection.

Using the systems and methods discussed herein, CMAS corrosion is inhibited via CMAS interception in an engine environment and/or is prevented or reduced by the formation of a metal oxide protective coating on a hot engine section component. The CMAS interception can occur while the engine is in operation in flight or in a testing or quality control environment. The metal oxide protective coating can be applied over other coatings, including Gd-zirconates (GZO) or yttria-stabilized zirconia (YSZ). The metal oxide protective coating is applied at original equipment manufacturers (OEM) and can also be applied in-situ using a gas injection system during engine use in-flight or during maintenance or quality testing. The metal oxide protective coating contains a rare earth element, aluminum, zirconium, chromium, or combinations thereof, and is from 1 nm to 3 microns in thickness.

The thermal barrier coating includes reactive gadolinia in its microstructures and the embedded gadolinia effectively reacts with CMAS contaminant reducing the damage from CMAS. Moreover, a method to produce a CMAS resistant thermal barrier coating can include a post-treatment to the thermal barrier coating with the reactive gadolinia suspension in sol-gel state.

A vacuum device comprising at least one component having a portion which, during operation of the vacuum device, is in contact with a vacuum and which is coated at least in part by a layer which absorbs gas particles, in particular with a layer of a no-evaporable getter (NEG) material.

A gas turbine engine includes an engine static structure extending circumferentially about an engine centerline axis; a compressor section, a combustor section, and a turbine section within the engine static structure. At least one of the compressor section and the turbine section includes at least one airfoil and at least one seal member adjacent to the at least one airfoil. A tip of the at least one airfoil is metal having a wear resistant coating and the at least one seal member is coated with an abradable coating. The wear resistant coating is formed as a layer in a base metal surface of the airfoil, has a thickness less than or equal to 10 mils (254 micrometers) and includes metal boride compounds.

A gas turbine engine includes: a turbine section including a casing extending circumferentially about a plurality of turbine blades and having at least one seal member coated with an abradable coating. At least one turbine blade has sides and a tip and at least one seal member is located adjacent to the tip of the at least one turbine blade. The tip of the at least one turbine blade has a wear resistant layer and an abrasive coating disposed on the wear resistant layer. The wear resistant layer has a thickness less than or equal to 10 mils (254 micrometers) and includes metal boride compounds.

A component and method for minimizing a gap between a blade platform ledge and a vane platform ledge in a turbine engine. The blade platform supports turbine blades attached to a shaft. The movable blade platform is positioned adjacent to a stationary vane platform having vanes mounted on an outer surface. The blade and vane platforms are separated by the gap between the platform ledges. During manufacture, an abrasive coating is applied to the surface of the blade platform ledges so that the coating contacts the vane platform ledge when the engine is started. The abrasive coating on the turbine blade platform cuts the surface of the vane platform ledge to form a gap sufficient to permit unobstructed motion of the blade platform, yet of minimal size to limit the flow of gas between the space within and the space outside the platforms.