A blade outer airseal has a body comprising: an inner diameter (ID) surface; an outer diameter (OD) surface; a leading end; and a trailing end. The airseal body has a metallic substrate and a coating system atop the substrate along at least a portion of the inner diameter surface. At least over a first area of the inner diameter surface, the coating system comprises an abradable layer comprising a metallic matrix and a solid lubricant; and the metallic matrix comprises, by weight, ≧35% nickel, 12.0-20.0% cobalt, 5.0-15.0% aluminum, and 5.0-15.0% chromium.

A blade outer airseal has a body comprising: an inner diameter (ID) surface; an outer diameter (OD) surface; a leading end; and a trailing end. The airseal body has a metallic substrate and a coating system atop the substrate along at least a portion of the inner diameter surface. At least over a first area of the inner diameter surface, the coating system comprises an abradable layer system comprising a plurality of layers including a relatively erosion-resistant first layer atop a relatively abradable second layer.

The present invention relates to a rotating blade (5), in particular for a compressor or turbine stage of a gas turbine, having a radially outer rib arrangement with at least one rib (2), onto which a coating (3) is disposed, whereby, in a meridian section, the coating (3) has an outer contour (3.1), which extends axially outwardly in the radial direction.

A method of forming a sprayed coating includes preparing a spray target member having a surface on which openings of first ends of holes are formed, preparing a plurality of masking pins each of which comprises metal, and inserting each of the masking pins into a corresponding one of the holes so that each of the masking pins partially protrudes from the surface. The method also includes applying an adhesive agent for fixing the masking pins to the respective holes, to at least one of the holes or the masking pins, forming a ceramic layer by spraying on the surface of the spraying target member, the ceramic layer comprising a ceramic material, while the masking pins are fixed to the respective holes via the adhesive agent, and removing the masking pins from the holes after the spraying step.

A method includes feeding at least one ceramic feedstock into a heating zone of a thermal spray apparatus to form a heated ceramic feedstock. The heated ceramic feedstock is entrained in a plasma gas to form a heated gas stream directed toward a target surface of a CMC substrate. A sacrificial composition is fed with a sacrificial composition feed apparatus into the heated gas stream downstream of the heating zone at a selected injection angle α of about −30° to about +30° with respect to a plane of the target surface of the substrate. The heated ceramic feedstock is deposited from the heated gas stream onto the target surface to form a coating thereon. The thermal spray apparatus and the sacrificial composition feed system are configured to independently control a chemistry and a porosity of the coating.

Disclosed is a thermal barrier coating system for components of a turbomachine, especially for high temperature-stressed or hot gas-stressed components of a turbomachine, comprising a ceramic coating of fully or partially stabilized zirconium oxide, and an oxide cover coating which comprises aluminum and at least one element from the group lanthanum, magnesium, silicon, calcium and sodium. The aluminum oxide exists at least partially as free α-Al.sub.2O.sub.3. Also disclosed is a method for producing a corresponding thermal barrier coating system.

A vane device for a gas turbine is provided. The vane device includes a first airfoil having a first suction side and a first pressure side, a second airfoil having a second suction side and a second pressure side, an inner shroud and an outer shroud. The first airfoil and the second airfoil are arranged between the inner shroud and the outer shroud, wherein the first airfoil and the second airfoil are at least partially coated with a MCrAlY coating. At least the first suction side has a first coated surface section which is coated with a thermal barrier coating and which represents at least a part of the total surface of the first suction side. At least the inner shroud or the outer shroud has a further coated surface section which is coated with a further thermal barrier coating. A corresponding method of manufacturing is also provided.

A turbine part, such as a turbine blade or a distributor fin, for example, including a substrate made of superalloy based on monocrystalline nickel, including rhenium and/or ruthenium, and having a γ′-NisAI phase that is predominant by volume and a γ-Ni phase, the part also including a sublayer made of metal superalloy based on nickel covering the substrate, wherein the sublayer has a γ′-NisAI phase that is predominant by volume and wherein the sublayer has an average atomic fraction of aluminium of between 0.15 and 0.25, of chromium of between 0.03 and 0.08, of platinum of between 0.01 and 0.05, of hafnium of less than 0.01 and of silicon of less than 0.01. A process for manufacturing a turbine part including a step of vacuum deposition of a sublayer made of a superalloy based on nickel having predominantly by volume a γ′-NisAI phase, on a substrate made of superalloy based on nickel including rhenium and/or ruthenium.

The invention relates to a de-icing device designed to supply de-icing air to a turbomachine separation nozzle extending along a longitudinal axis, the turbomachine comprising:—the separation nozzle which is designed to be positioned downstream from a turbomachine fan and comprises an internal casing and an external casing which form a separation between a primary flow vein for a primary stream and a secondary flow vein for a secondary stream, said streams issuing from the fan, the internal casing and the external casing defining an inter-vein space;—turbomachine guide vanes designed to be secured by screws to the internal casing such that the screws extend into the inter-vein space, the de-icing device being intended to be positioned in the inter-vein space and comprising—an air inlet;—an air outlet;—a plurality of channels extending from the air inlet toward the air outlet; the channels being arranged in relation to one another such that they are designed to extend from the air inlet toward the air outlets, passing between the screws for securing the guide vanes.

A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a compressor section, a combustor section, a turbine section, and at least one rotatable shaft. The engine further includes a seal assembly including a seal plate mounted for rotation with the rotatable shaft and a face seal in contact with the seal plate at a contact area. The seal assembly includes an abradable coating adjacent the contact area, and the abradable coating includes magnetic particles.