A blade includes an airfoil section extending between leading and trailing edges, first and second opposed sides each joining the leading and trailing edges, and an inner end and a free end. The blade also includes an abrasive tip at the free end of the airfoil section. The abrasive tip includes particles diposed in a matrix material. The matrix material is a polymeric material that has a glass transition temperature greater than or equal to about 225 degrees C. (487 degrees F.). A gas turbine engine and a method of fabricating a blade are also disclosed.

A propeller blade comprises a root, a tip distal from the root, a trailing edge extending from the root to the tip, a trailing edge, e.g. foam, insert, a shell forming an outer surface of the propeller blade and a plurality of stitches of yam extending through two parts of the shell adjacent the trailing edge, wherein the yarns do not extend through the trailing edge insert.

A coated substrate is provided that includes an environmental barrier coating on (e.g., directly on) a surface of a substrate (e.g., a ceramic matrix composite). The environmental barrier coating can include a barrier layer having a refractory material phase and a silicon-containing glass phase. The silicon-containing glass phase may be a continuous phase within the barrier layer (e.g., a breathable grain boundary of the barrier layer), or may be a plurality of discontinuous layers dispersed throughout the refractory material phase. The refractory material phase can include a rare earth silicate material having a rare earth component at a first atomic percent, while the silicon-containing glass phase comprises the rare earth component at a second atomic percent that is less than the first atomic percent. Methods are also provided for forming a barrier layer on a substrate.

A composite airfoil having a non-metallic leading edge protective wrap is provided. In one aspect, the airfoil has a composite core having a pressure sidewall and a suction sidewall each extending between a core leading edge and a core trailing edge. A leading edge protective wrap protects the core leading edge and includes a trailing wrap and a leading wrap. The trailing wrap wraps around the core leading edge and is connected to the composite core. The leading wrap wraps around the core leading edge and is connected to the trailing wrap. The trailing and leading wraps both have leading edges that are spaced from one another. A filler is positioned between the leading edges of the trailing and leading wraps. A protective nose is connected to the leading edge of the leading wrap. The components of the leading edge protective wrap are formed of non-metallic materials.

A hybrid vane for a gas turbine engine. The hybrid vane comprises an airfoil having an inner core composed of a fiber-reinforced thermoplastic composite. A longitudinal axis of the hybrid vane extends between a vane root and a vane tip. The hybrid vane further comprises a metallic outer layer at least partially covering the inner core.

A compressor wheel includes a compressor wheel body, and a thermal insulating coating layer disposed so as to cover at least a part of a back surface of the compressor wheel body.

A gas turbine engine includes a circumferential row of blades, with the blades having respective blade tips. A seal is disposed about the blades. The seal has an abradable layer which the tips of the blades, at times, rub against when the blades rotate. The rubbing produces a maximum temperature at the abradable layer. The abradable layer includes a metal matrix and microballoons dispersed in the metal matrix. The microballoons are formed of a glass that has a glass transition temperature that is approximately 50° F. to 300° F. greater than the maximum temperature.

An engine shaft assembly for an engine is provided. The engine shaft assembly includes a shaft and a thermal distribution layer. The thermal distribution layer is provided on the shaft, and is configured to minimize the effect of distortion of the shaft caused by asymmetric cooling on shutdown of the engine.

A heat insulating sheet member capable of enhancing a heat insulating performance of an exhaust gas introduction path by an easy operation, including such a path in a turbocharger. The heat insulating sheet member is a bendable member formed from an inorganic flexible material. The heat insulating sheet member includes a first region corresponding to an inlet of a bottom wall portion, a second region corresponding to at least a terminating end portion of the scroll portion, a third region provided between the first region and the second region and corresponding to a coupling wall portion, and a fourth region corresponding to an outer peripheral wall portion. The first region and the third region, the third region and the second region, and the first region and the fourth region are coupled to each other with the inorganic flexible material in a continuous state.

A method for forming a composite part of a gas turbine engine. The method includes assembling the composite part of a first composite material and a second composite material. The second composite material defines an outer surface of the composite part, and is selected to be curable at a cure temperature generated by heat from operation of the engine. The first composite material is selected to have an operating temperature limit less than the cure temperature. The method includes placing the composite part within the engine so that, in use, the second composite material is cured by exposure to the heat generated from operation of the engine. The second composite material thermally shields the first composite material from the heat generated from operation of the engine. The method includes operating the engine to cure the second composite material.