Disclosed is a method for improving the high-temperature stability of a component, in particular a blade of a turbomachine, formed at least partially from a molybdenum alloy that, besides molybdenum, silicon, boron and titanium, comprises iron and/or yttrium. The method comprises depositing a diffusion barrier layer formed from technically pure molybdenum or tungsten or being an alloy based on molybdenum and/or tungsten at least on an outer surface, which comprises the molybdenum alloy, of the component, and depositing silicon on the diffusion barrier layer to form molybdenum silicides and/or tungsten silicides.

A gas turbine engine includes a case disposed around a central axis, a vane structure disposed radially inwards of the case, a seal carriage mounted to the case, and a seal arc segment mounted to the seal carriage. The vane structure exerts an axial load on the seal carriage. The seal carriage defines a load path that radially transmits the axial load to the case.

A seal assembly includes a seal arc segment that defines radially inner and outer sides and first and second axial arc segment sides, a carriage that carries the seal arc segment, and a cover that defines first and second axial cover sides. The second axial cover side is adjacent the first axial arc segment side.

A blade outer air seal support includes, at least one arc body having a first portion and a second portion, a blade outer air seal mounting region defined at least partially between the first portion and the second portion, and an interface feature interfacing the first portion and the second portion. The interface feature is configured such that axially aligned forces are communicated between the first and second portions through the interface feature, bypassing the blade outer air seal mounting region.

An article includes a substrate formed of a molybdenum-based alloy. The molybdenum-based alloy has a composition that includes molybdenum, silicon, and boron. A barrier layer is disposed on the substrate. The barrier layer is formed of at least one noble metal.

An article includes a substrate formed of a molybdenum-based alloy. A barrier layer is disposed on the substrate. The barrier layer is formed of at least one noble metal.

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 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.

An airfoil includes a multi-part body and one or more thermally conductive pins. The multi-part body has an interior region and is formed from multiple pieces joined with each other at an interface. The pieces have multiple cavities and at least one of the pieces defines airfoil cooling channels disposed within the interior region of the body. The one or more thermally conductive pins are within the interior region of the body and extend across the interface. Each of the thermally conductive pins has a first segment disposed within a corresponding cavity of a first piece of the multiple pieces and a second segment disposed within a corresponding cavity of a second piece of the multiple pieces.

A method for forming a secondary component from an original component having an original shape includes separating the original component into a parent component and a replaced portion, and forming a replacement coupon using an additive manufacturing system. The replacement coupon is shaped to substantially match the original shape of the replaced portion. The method further includes coupling the replacement coupon to the parent component to form the secondary component. The method also includes at least one of (i) removing the replacement coupon from a build plate of the additive manufacturing system prior to application of any heat treatment to the as-built replacement coupon, wherein the replacement coupon maintains a near-net original shape of the replaced portion after removal, and (ii) entering the secondary component into normal duty with no hot isostatic press treatment of the replacement coupon having been performed.