A vane arc segment includes an airfoil fairing that has first and second fairing platforms and an airfoil section therebetween. The airfoil section has a pressure side, a suction side, and an internal cavity. The first fairing platform defines suction and pressure side circumferential mate faces, forward and aft faces, a gaspath side, a non-gaspath side, and a flange that projects from the non-gaspath side. The flange extends along one of the suction or pressure side circumferential mate faces. There is a rib that has a rib section in the internal cavity that spans the pressure and suction sides and a rib extension section that extends from the rib section in the internal cavity and along the non-gaspath side of the first fairing platform. The rib extension intersects the flange to form a gusset for reinforcing the flange.

A coated gas turbine engine part includes a substrate and a calcium-magnesium-alumino-silicate CMAS protection layer present on the substrate. The layer includes a first phase of a calcium-magnesium-alumino-silicate CMAS protection material and a second phase including particles of an anti-wetting material dispersed in the first phase.

An airfoil for a fan section of a turbine engine may include a fan blade or an outlet guide vane formed of a first material, and an edge guard disposed about an edge of the fan blade. The edge guard may include a matrix composite that has a toughness that is greater than a toughness of the first material. The airfoil may include a fan blade or an outlet guide vane. The first material of the airfoil may include a metal alloy and/or a matrix composite. A method of manufacturing an airfoil for a fan section of a turbine engine may include manufacturing an edge guard, attaching the edge guard to the airfoil.

A turbine vane assembly adapted for use in a gas turbine engine includes a support and a turbine vane arranged around the support. The support is made of metallic materials. The turbine vane is made of ceramic matrix composite materials to insulate the metallic materials of the support.

Airfoils for gas turbine engines are provided. In one embodiment, an airfoil formed from a ceramic matrix composite material includes opposite pressure and suction sides extending radially along a span and defining an outer surface of the airfoil. The airfoil also includes opposite leading and trailing edges extending radially along the span. The pressure and suction sides extend axially between the leading and trailing edges. The leading edge defines a forward end of the airfoil, and the trailing edge defining an aft end of the airfoil. Further, the airfoil includes a trailing edge portion defined adjacent the trailing edge at the aft end of the airfoil; a plenum defined within the airfoil forward of the trailing edge portion; and a cooling passage defined within the trailing edge portion proximate the suction side. Methods for forming airfoils for gas turbine engines also are provided.

A preceramic resin formulation comprising a polycarbosilane preceramic polymer and an organically modified silicon dioxide preceramic polymer. A ceramic material comprising a reaction product of the polycarbosilane preceramic polymer and organically modified silicon dioxide preceramic polymer is also described. Articles comprising the ceramic material are also described, as are methods of forming the preceramic resin formulation and the ceramic material.

An airfoil includes a ceramic airfoil that defines a leading edge, a trailing edge, a pressure side, a suction side, a first radial end, and a second radial end. The ceramic airfoil section has an internal cavity and a rib that divides the internal cavity into a first radial passage and a second radial passage. The first radial passage is open at both the first radial end and the second radial end, and the second radial passage is open at least at the second radial end. A cooling passage circuit includes a first radial leg through the first radial passage, a second radial leg though the second radial passage, and a turn leg outside of the internal cavity at the second radial end. The turn leg connects the first radial leg and the second radial leg.

A method for coating a ceramic matrix composite part with an environmental barrier, the method including a) applying, to a surface of the part, a coating composition including a first powder of a rare earth silicate and a second powder including boron, the coating composition having a ratio R=[mass of the second powder]/[mass of the first powder] of between 0.1% and 5%, and b) sintering the first and second powders to obtain the environmental barrier on the part.

A multilayer protective coating system includes a turbine engine component substrate formed of a ceramic matrix composite material, an environmental barrier coating layer including a rare earth disilicate material deposited directly on the substrate, and a plurality of pairs of alternating layers of the rare earth disilicate material and a rare earth monosilicate material deposited and sintered directly on the environmental barrier coating layer. Each layer of the plurality of pairs of alternating layers is relative less thick as compared with the environmental barrier coating layer.

An airfoil vane assembly according to an exemplary embodiment of this disclosure, among other possible things includes a vane piece having a first vane platform, a second vane platform, and a hollow airfoil section joining the first vane platform and the second vane platform, a spar piece having a spar platform and a spar extending from the spar platform into the hollow airfoil section, and at least one seal arranged at a sealing surface of the spar platform and sealing between the spar platform and the first vane platform. The airfoil vane assembly also includes a thermal barrier coating disposed on the spar piece. The sealing surface is free from the thermal barrier coating. A gas turbine engine and a method of making a spar piece for an airfoil vane assembly are also disclosed.