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.

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.

An air cooled integrally bladed rotor with single crystals turbine rotor blades having cooling air passages formed into an equiax rotor disk, where a mold having expendable Molybdenum tooling and reusable molybdenum tooling is used to form the IBR. An annular blade ring with openings is used to secure the single crystal rotor blades within the mold, and a number of tube tools are inserted into a bottom end of each blade that forms a cooling air supply passage within the rotor disk. Two molybdenum circular shaped hubs are used to secure a bottom end of the tube tools with the mold. The mold is filled with metal powder and high pressure is used to solidify the powder to form the IBR. Expendable tooling is removed using sublimation when exposed to oxygen. Reusable tooling is reused to form additional IBRs.

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.

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.

The present disclosure relates to a novel iron-based austenitic alloy for a turbocharger housing and to methods of its preparation.

The invention relates to a blade for use in a turbine of an aircraft engine. The blade is made of (a) a Mo-based alloy strengthened by intermetallic silicides or (b) a Ni-based single crystal superalloy. An aircraft engine and in particular, a turbofan aircraft engine including a corresponding turbine blade is also disclosed.

When a compressor vane or blade for an engine is used in an environment containing abundant foreign substances, deposits originated from the foreign substances are likely to deposit on surfaces of the vane or blade. The compressor vane or blade according to the present disclosure has a base body of the compressor vane or blade; and a coating covering the base body, which consists essentially of one or more selected from the group of molybdenum disulfide and tungsten disulfide.

The bearing arrangement includes a one-piece gas bearing sleeve (32) configured to rotatably support the drive shaft (4) and made in molybdenum or in a molybdenum alloy, the one-piece gas bearing sleeve (32) including a radial bearing surface (33) configured to Surround the drive shaft (4).

A split ring seal has: a first circumferential end and a second circumferential end; an inner diameter surface and an outer diameter surface; a first axial end face and a second axial end face. A circumferentially distributed first plurality of open channels are along the first axial end face. A circumferentially distributed second plurality of open channels are along the second axial end face.