Turbomachine components and compressors are provided. The turbomachine component includes a platform and a mounting portion that extends from the platform. The mounting portion includes a dovetail received by a slot defined in the turbomachine. The slot includes a floor and a ceiling. The dovetail includes an inner surface and an outer surface. A hole is defined in the dovetail from an inlet at the inner surface to an end wall. The hole has a cylindrical portion and a tapered portion. A mechanical spring is disposed within the hole and in contact with the floor and the end wall such that the outer surface of the dovetail is forced into contact with the ceiling of the slot.

Systems and methods include providing a stator vane bushing for a variable stator vane assembly having a movable stator vane and a stator vane housing disposed annularly about the movable stator vane. The bushing includes a flange, a barrel extending from the flange, and a central aperture extending through the flange and the barrel and is disposed in the housing and annularly about the stator vane. The bushing is formed from a ceramic material having a coefficient of thermal expansion (CTE) lower than or equal to a CTE of one or more of a stator vane and a housing of the variable stator vane assembly.

A turbine rotor wheel for an aircraft turbomachine includes a rotor disk, an annular shroud extending around the disk, and blades arranged between the disk and the shroud. The the root of each of the blades has two tabs configured for attachment to the disk. The tabs are arranged upstream and downstream, respectively, of a wall of the disk, relative to the axis. The tab arranged upstream is engaged in a first recess of the disk and configured to cooperate by abutment with a peripheral edge of the first recess. The tab arranged downstream is engaged in a second recess of the disk and is configured to cooperate by abutment with a peripheral edge of the second recess.

An airfoil assembly defines an axial direction, a radial direction, and a circumferential direction, and includes an airfoil and an outer band disposed on an outer end of the airfoil in the radial direction. The outer band includes an outer attachment structure configured to secure the outer band to an outer support structure on an outer side of the outer band.

A pin member is proposed for a turbo-machine having a shroud arranged to rotate within a turbine housing. The pin member is configured to limit this rotation. It is a one-piece element comprising a cylindrical body and a limit surface for opposing motion of the shroud.

A vane for a gas turbine engine, the vane including a platform with an airfoil extending radially from the upper surface of the platform. The platform includes a joint portion which includes a circumferentially extending flange and a recessed surface both formed on either the upper or lower surface of the platform. The flange and the recessed surface extend from opposing circumferential edges of the joint portion and each include a substantially radially-extending through hole.

A vane assembly for a turbomachine includes a variable vane (10-15), the vane having at least one first, in particular plane or curved, engagement surface (11; 11′) for clamping, in particular without play, an actuator (20) of the vane assembly for adjustment of the vane, this first engagement surface not being inclined toward a longitudinal axis (L) of the vane or being inclined toward it by no more than 15°, and/or the vane assembly including a clamp (30; 31) for clamping the actuator against the first engagement surface by at least partially elastically compressing the clamp transversely to the longitudinal axis of the vane and/or by advancing the clamp in a clamp direction (S) that forms an angle of at least 45° with the longitudinal axis of the vane.

The invention concerns a flow straightener unit (1) for a fan module of a turbomachine, the straightener unit (1) comprising a plurality of blades (2) distributed about an axis of rotation, each blade (2) is made of a composite material and comprises an aerofoil (21) and a root (22) intended to be assembled on a hub (4) of the turbomachine. The unit (1) comprises a centring and attachment plate (3) of the blade (2) on the turbomachine intended to be attached to the hub (4) at a determined azimuthal position and to the root (22) of the blade (2), the plate (3) is designed to be screwed to the hub (4) by screws (51a) that are longitudinal with respect to the axis of rotation of the unit (1) and screwed to the root (22) of the blade (2) by screws (52) that are radial with respect to the axis of rotation of the unit (1).

A turbine guide apparatus, having multiple guide blades. Each guide blade has a first shroud and a second shroud formed on radial ends of a blade leaf having a first carrier for the guide blades. Each guide blade is fastened to the first carrier via a first shroud projection having a second carrier for the guide blades and is fastened on the second carrier via a second shroud projection. The first shroud projection is inserted into a groove of the first carrier in the radial direction and fastened in this groove via a bolt extending in the axial direction through the projection of the first shroud with radial mobility in this groove. The second shroud projection of the guide blade is fastened in the circumferential and radial direction via a pin extending in the axial direction into the projection of the second shroud and the second carrier.

A vane assembly includes a vane that includes an outer platform, an inner platform, and an airfoil. The outer platform defines an outer boundary of a gas path. The inner platform is spaced apart radially from the outer platform relative to an axis and defines an inner boundary of the gas path. The airfoil extends radially between and interconnects the outer platform and the inner platform.