A rotating assembly for a turbomachine, including a disc having an external periphery with alternating recesses and teeth, blades extending radially from the disc and the feet of which are axially engaged and radially retained in the recesses of the disc, platforms extending on the circumference from the blades and being arranged end to end on the circumference, with respect to each other, axial sealing device upstream and/or downstream of an annular zone extending radially from the platforms up to the disc. The sealing device radially includes an internal annular portion and an external annular portion structurally distinct from each other, and the external annular portion of which is elastically constrained radially inwards by the internal annular portion.

The present disclosure provides for a turbine wheel having a blade platform disposed to compliantly secure ceramic turbine blades to a rotatable disk. The platform includes opposing ends to engage a portion of an airfoil of each turbine blade and radial extensions to engage a portion of a root of each turbine blade.

A rotor disk assembly for a gas turbine engine includes a rotor disk that defines an axis, the rotor disk comprising a slot within a rim of the disk and a spacer within the slot, the spacer having a clearance feature sized to permit rotation of a rotor blade from an insertion position to an installed position.

A vane arc segment includes an airfoil fairing that has a platform and an airfoil section that extends therefrom. The platform defines periphery edge faces, a gaspath face, and a non-gaspath face. The platform has a flange that projects radially from the non-gaspath face. The flange defines a forward flange face, an aft flange face, and a radial flange face. The flange has an end portion adjacent one of the periphery edges. The end portion includes a step that defines a radial step face that is radially intermediate the radial flange face and the non-gaspath face.

The present invention generally relates to rotary turbomachinery methods and integrated processes requiring high-energy efficiency. In one embodiment, the present invention relates to rim-rotor configurations enabling long-term survival under conditions of either high temperature or oxidation resistance or saturated fluid abrasion.

A rotor assembly is provided for a gas turbine engine. This rotor assembly includes a first rotor disk, a second rotor disk, a plurality of rotor blades and a plurality of disk mounts. The first rotor disk is configured to rotate about a rotational axis. The second rotor disk is configured to rotate about the rotational axis. The rotor blades are arranged circumferentially around the rotational axis. Each of the rotor blades is mounted to the first rotor disk and to the second rotor disk. The rotor blades include a first rotor blade. Each of the disk mounts connects the first rotor disk and the second rotor disk together. The disk mounts include a first disk mount that further supports the first rotor blade.

An assembly adapted for use in a gas turbine engine includes a carrier and a blade track segment. The carrier extends at least partway about an axis. The blade track segment is supported by the carrier radially relative to the axis to define a portion of a gas path of the assembly.

A rotor assembly is provided for a gas turbine engine. This rotor assembly includes a first rotor disk, a second rotor disk, a plurality of rotor blades and a plurality of disk mounts. The first rotor disk is configured to rotate about a rotational axis. The second rotor disk is configured to rotate about the rotational axis. The rotor blades are arranged circumferentially around the rotational axis. Each of the rotor blades is axially between and mounted to the first rotor disk and the second rotor disk. The disk mounts connect the first rotor disk and the second rotor disk together. The disk mounts include a first disk mount. The first disk mount is integral with the first rotor disk. The first disk mount projects axially through the second rotor disk.

A ceramic matrix composite gas turbine blade comprising a root portion coupled to a disk, said root portion having a neck; a platform region is disposed along an upper portion of the neck; an airfoil is located opposite the neck relative to the platform and extends outwardly from the platform; and a limiting section fuse formed in the blade proximate the neck.

A highly efficient gas turbine engine is provided. The fan of the gas turbine engine is driven from a turbine via a gearbox, such that the fan has a lower rotational speed than the driving turbine, thereby providing efficiency gains. The efficient fan system is mated to a core that has low cooling flow requirements and/or high temperature capability, and which may have particularly low mass for a given power.