The present invention relates to a protective layer against CMAS, to a CMAS-resistant article comprising the protective layer according to the invention, and to a process for preparing a corresponding article.

An airfoil vane includes an outer ceramic wall that defines an airfoil section and a cavity that extends through the airfoil section. A jumper tube is disposed in the cavity for transferring cooling air. The jumper tube includes a wall, a through-passage circumscribed by the wall, and a thermal barrier coating disposed on the wall.

An article may include a substrate, such as a silicon-containing ceramic matrix composite, an environmental barrier coating (EBC) layer on the substrate, and a CMAS-resistant EBC layer on the EBC layer. The EBC layer may include at least one rare-earth disilicate (REDS). The CMAS-resistant EBC layer may include at least one rare-earth monosilicate (REMS) configured to react with CMAS to form crystalline reaction products. The CMAS-resistant EBC layer may include a plurality of vertical cracks extending from a surface of the CMAS-resistant EBC layer at least partially into the CMAS-resistant EBC layer. Additionally, or alternatively, the EBC layer may include a plurality of vertical cracks extending from a surface of the EBC layer into at least a portion of the EBC layer.

A part includes a substrate having, adjacent to a surface of the substrate, at least a portion that is made of a material that contains silicon, and an environmental barrier formed on the surface of the substrate, the environmental barrier including a first layer including at least one first rare earth silicate and presenting grains with a mean size less than or equal to 1 μm; and a second layer covering the first layer, the second layer including at least one rare earth silicate and presenting grains with a mean size greater than 1 μm.

A method of analyzing layer thickness of a multilayer component is provided. The method includes: creating an opening having a predefined geometry partially into the multilayer component at a selected location on a surface of the multilayer component. The multilayer component includes a plurality of material layers including a substrate and a bond coat. The opening exposes each of the plurality of material layers including the substrate. Contrast of the exposed plurality of material layers can be increased. An image is created of the exposed layers in the opening using a digital microscope, and thickness of a bond coat, thickness of a depletion layer\ and/or thickness of an oxide layer is calculated from the image and based on the predefined geometry of the opening. Repairing the opening, allows the multilayer component to be used for an intended purpose after testing, e.g., re-installed and reused in a gas turbine.

An article includes a substrate, a ceramic barrier coating, and a layer of networked ceramic nanofibers. The ceramic barrier coating is disposed on the substrate and has a porous columnar microstructure. The layer of networked ceramic nanofibers is disposed on the ceramic barrier layer and seals the pores of the porous columnar microstructure.

An environmental barrier coating, comprising a substrate containing silicon; an environmental barrier layer applied to the substrate; the environmental barrier layer comprising an oxide matrix; an oxidant getter phase interspersed throughout the oxide matrix; and a self-healing phase interspersed throughout the oxide matrix.

An internal core profile for a turbine nozzle airfoil of a gas turbine is provided. The turbine nozzle may include an airfoil core having an uncoated nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in Table 1, wherein the X, Y, and Z coordinates are distances in inches measured in a Cartesian coordinate system, the corresponding X and Y coordinates, when connected by a smooth continuous arc, define one of a plurality of airfoil core profile sections at each Z distance, and the plurality of airfoil core profile sections, when joined together by smooth continuous arcs, define an airfoil core shape.

A method comprises inspecting a ceramic matrix composite component assembled in a gas turbine engine to determine an extent of damage to the ceramic matrix composite component, determining a repair technique to repair the damage to the ceramic matrix composite component based on the extent of damage to the ceramic matrix composite component, and repairing the ceramic matrix composite component using the repair technique.

A method for manufacturing a blade in composite material with added metal leading edge for gas turbine aeroengine, the method including producing a batch of plurality of blade bodies in composite material; creating a digital model of a blade body from a blade in the batch of plurality of blade bodies; creating a digital model of a theoretical final blade including a leading edge; generating a digital model of a leading edge from the digital model of a blade body and final blade model; manufacturing at least one leading edge via additive manufacturing from the generated leading edge digital model; bonding each manufactured leading edge onto a blade body from the batch of plurality of blade bodies.