A gas turbine component, in particular a stator vane cluster (10), having at least one radial flange (12, 13) with a radial slot (20) for spoke-type centering of a gas turbine element, in particular an inner ring (100), the radial slot (20) having a first slot flank (21), a slot base (23) adjoining the same, and a second slot flank (22) adjoining the same, the slot base (23) being asymmetric with respect to a mid-plane (M) which extends radially and perpendicularly to the circumferential direction and centrally between the two slot flanks (21, 22).

A sealing apparatus for a turbomachine, including a stator vane component, which includes an inner shroud element and a flow-directing element connected to the inner shroud element. The sealing apparatus includes a sealing component, which has a seal-carrier ring element coupled to the stator vane component. The seal-carrier ring element includes at least one ring body element and at least one projection, which is connected in one piece to the at least one ring body element, protrudes from the at least one ring body element in the radial direction of the sealing apparatus, and is inserted into at least one opening, which extends through the inner shroud element. Other aspects relate to a seal-carrier ring element for a sealing apparatus, and to a turbomachine which includes at least one sealing apparatus and/or at least one seal-carrier ring element.

A turbine shroud assembly adapted for use with a gas turbine engine includes a carrier, a seal segment, and a mount assembly. The carrier is configured to be coupled to a turbine case. The seal segment is shaped to define a gas path boundary of the shroud assembly. The mounting assembly is configured to couple the seal segment to the carrier.

A vane system includes a plurality of vane assemblies and a continuous support ring. The vane assemblies are arranged circumferentially about an axis and each include a hollow airfoil fairing and a spar. The spar has a spar flange and a spar leg extending radially inwardly from the spar flange and through the hollow airfoil fairing. The continuous support ring has radially inner and outer sides and defines a circumferential row of through-holes between the radially inner and outer sides. The spar legs extend through the through-holes and the spar flanges are affixed at the radially outer side of the continuous support ring.

A vane segment includes a blade ring, a plurality of vanes, and a caulking member. The blade ring includes a blade ring groove and a storage space forming portion. The blade ring groove includes a convex portion projecting from a second groove side surface of the blade ring groove to the first side. The storage space forming portion, in combination with an outer shroud, forms a storage space into which the caulking member enters. The caulking member is accommodated within the storage space, contacts with the blade ring groove and the outer shroud, and is exposed from an inner opening of the storage space.

A turbomachinery stator apparatus includes: a compressor casing including a casing wall defining an arcuate flowpath surface and an opposed backside surface, the flowpath surface defining at least two spaced-apart rotor lands, a stator vane row of stator vanes disposed inside the compressor casing, wherein the casing wall includes a heat shield positioned outboard of the rotor lands immediately upstream or downstream of the stator vane row, and wherein a) the casing wall includes the heat shield and b) the stator vanes form a single monolithic whole.

A flow path component assembly includes a flow path component that has a platform that extends from a first circumferential side to a second circumferential side. The platform has a first radius of curvature between the first circumferential side and the second circumferential side. A coating is arranged on a portion of the platform near the first circumferential side. The coating has a second radius of curvature. The second radius of curvature is larger than the first radius of curvature.

A gas turbine engine includes a fan rotor driven by a fan drive turbine about an axis through a gear reduction. An inner core engine has an inner core engine housing surrounding a compressor section, including a low pressure compressor. A rigid connection between a fan case and the inner core engine includes A-frames rigidly connected at a connection point to the fan case. Fan exit guide vanes rigidly connect to the fan case, and to the inner core engine. A fan intermediate case is positioned forward of a first rotor stage in the low pressure compressor. A rigid structure is connected to the inner core engine and to the fan exit guide vanes. The rigid structure defines a structure moment stiffness. The fan intermediate case defines an intermediate case moment stiffness. A ratio of the structure moment stiffness to the intermediate case moment stiffness is between 5 and 15.

The invention relates to an assembly for an aircraft turbomachine, comprising a central element (1) for the ejection of gas, and a connecting flange (9) interposed between, upstream, a gas outlet (22a) made of metal for a turbomachine, and, at the downstream end, the central element (1). The connecting flange comprises an annular portion (9a) and flexible tabs (11) having axially: a first end (111a) where the tab is connected to the said annular portion, and a second free end (111b), projecting radially inwardly with respect to the first end and to which said tab is attached with the central element (1).

In order to alleviate a mismatch problem of thermal deformation, in all directions, of a connecting and installing structure between a CMC turbine outer annular component and a metal intermediate casing, a connector and an anti thermal mismatch connecting device are provided. The rod part of the connector comprises a subtractive hollow section and a cylindrical section. The subtractive hollow section is composed of a central shaft, a plurality of supporting rib plates extending outwards from a peripheral surface of the central shaft and inclined radially relative to the central shaft, and a plurality of outer annular plates arranged around the central shaft, with a circumferential gap between adjacent outer annular plates. The supporting rib plate is connected with the central shaft and the outer annular plate, and the central shaft is connected with the cylindrical section. The anti thermal mismatch connecting device the connector.