A seal for sealing a gap in a combustion apparatus of a gas turbine between a first side and an opposite second side. The seal includes a steel strap with a flat shape and a first metal web with a flat shape attached to the steel strap on the first side and a second metal web attached to the steel strap on the second side. Thereby the seal has a seal thickness with at least the steel strap plus the first metal web plus the thickness of the second metal web, wherein the seal thickness is at most 0.2-times of width of the seal. The new seal has an Omega shape of the second metal web, which is therefore attached on both sides with a flat shaped portion to the steel strap.

A method for making a ceramic matrix composite component includes densifying a fibrous preform of the component with a ceramic matrix to form an intermediate component; infiltrating a hole in the intermediate component with an infiltrate material comprising a solid and a metallic alloy whose reaction forms a carbide, silicide, boride or combination thereof, heating the infiltrate material to a temperature in excess of a melting point of the metallic alloy; and sequentially cooling regions of the hole starting from an interior end of the hole to the outer surface of the intermediate component to form a solidified through-thickness reinforcement element. The hole extends in a through-thickness direction and is open to an exterior surface of the intermediate component.

A combustor for a gas turbine engine, the gas turbine engine defining a longitudinal centerline extending in a longitudinal direction, a radial direction extending orthogonally outward from the longitudinal centerline, and a circumferential direction extending concentrically around the longitudinal centerline, the combustor including: a liner at least partially defining a combustion chamber, the liner further defining a gap; and an intermediate member disposed at least partially within the gap of the liner and in fluid communication with the combustion chamber, wherein the intermediate member is retained at a relatively fixed position with respect to the liner by a support extending between the intermediate member and a case of the gas turbine engine.

In a gas turbine engine, coolant (e.g., cooling air) is prone to leak out of the interface between the combustor case, the nozzle of the turbine, and the exhaust diffuser. Embodiments of an interface are disclosed that provide non-fretting sealing using an interference fit between radially facing surfaces of a combustor flange and diffuser flange. In addition, one or more contact sealing lands may be used between the combustor flange and diffuser flange and one or more seals may be provided between various components of the interface to provide additional sealing.

A hydrostatic seal configured to be disposed between relatively rotatable components. The hydrostatic seal includes a seal carrier. The hydrostatic seal also includes a beam extending axially from a forward end to an aft end, the beam cantilevered to the seal carrier at one of the forward end and the aft end, the beam free at the other end. The hydrostatic seal further includes a bellows seal operatively coupled to the seal carrier and in contact with the beam.

A seal includes a ceramic matrix composite ply having woven ceramic-based fibers in a ceramic-based matrix. The ceramic matrix composite ply has at least one bend formed about a bend axis and defines at least one rounded portion. A sealed assembly and a method of making a seal are also disclosed.

A turbine engine having an engine core, a first seal structure, a plurality of circumferentially arranged vanes, a plurality of struts, and a second seal structure. The engine core can define a rotor and a stator having a casing. The first seal structure can include a finger seal projecting from the rotor. The plurality of circumferentially arranged vanes can include an uppers band mounted to the casing, and a lower band located radially above the finger seal. At least some of the struts of the plurality of struts can extend through the plurality of circumferentially arranged vanes. The second seal structure can form a seat with the lower band.

An inner shroud block component for a gas turbine. The inner shroud block has a radially inward facing surface with an abradable material applied thereto. The abradable material includes a zone of ridges that extend radially inwardly from the radially inward facing surface to minimize the clearance between the inner shroud block and the blade tip of a turbine blade. The abradable material may be ceramic and may be abraded by the blade tip if contact occurs between the blade tip and the inner shroud block. The zone of ridges extend along the radially inward facing surface in parallel to a direction of rotation of the turbine blade.

A rotor assembly is provided for a piece of rotational equipment. This rotor assembly includes a rotor disk, a rotor blade and a seal element. The rotor disk is configured to rotate about a rotational axis. The rotor blade includes an airfoil, a platform and a mount attaching the rotor blade to the rotor disk. The seal element is seated in a groove of the rotor disk. The seal element is configured to sealingly engage the platform and the mount.

A probe adapter includes an adapter body including a probe aperture and a slot. The probe adapter further includes a driver slidably mounted within the slot and slidable between a first position and a second position. The driver includes a first end and a second end opposite the first end. The first end includes a ramped recess extending in a direction from the first end toward the second end. The probe adapter further includes a threaded fastener configured to contact the second end of the driver so as to retain the driver in the first position.