A turbine for a gas turbine engine includes, among other things, a shaft rotatable about a longitudinal axis, a turbine rotor including one or more rows of turbine blades and a disk assembly coupled to the shaft. The disk assembly includes one or more disks each having an attachment region extending radially between an inner boundary and an outer boundary, the outer boundary is established by an outer periphery of the respective disk, the attachment region defines an array of slots distributed about the outer periphery, each of the slots extends radially inwardly from the outer boundary to the inner boundary, and each of the slots is dimensioned to receive a root section of a respective one of the turbine blades to mount the turbine blades to the disk assembly.

A rotor blade for a gas turbine engine including an inner ply layer group that at least partially defines a base of a root region; a platform around the inner ply layer group that at least partially defines the base; and a platform over-wrap around the platform, the platform over-wrap at least partially defines the base.

A rotor disk assembly for a gas turbine engine includes a rotor disk with a multiple of blade slots, the rotor disk defined about an axis; a sleeve received within each of the multiple of blade slots; and a ceramic matrix composite rotor blade received within each of the multiple of sleeves.

A rotor assembly is provided for a gas turbine engine. This rotor assembly includes a first rotor disk, a second rotor disk and a plurality of rotor blades. 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. The first rotor blade includes an attachment projecting axially along the rotational axis into a first pocket in the first rotor disk and a second pocket in the second rotor disk. The attachment has a dovetail cross-sectional geometry when viewed in a plane perpendicular to the rotational axis. A portion of the first rotor disk extends circumferentially across and covers the attachment.

A rotor blade for a gas turbine engine including an inner ply layer group includes a flared region and an airfoil region, a neck region between the flared region and the airfoil region; and a platform shell that comprises a platform shell root region, a platform shell platform region, and a platform shell neck region between the platform shell root region and the platform shell platform region, the platform shell root region sheathes the flared region, and the platform shell neck region flares outwardly away from the neck region to intersect the platform shell platform region at a platform shell neck region perimeter around an airfoil opening perimeter.

A turbine blade is secured with a turbine disk by a retention assembly. The turbine blade includes a blade stem and the disk includes at least one mount post for mounting of the turbine blade. The retention assembly is secured with the disk by forming a diffusion joint with the disk.

An airfoil includes an endwall section and an airfoil section that defines, at least in part, an airfoil profile. The airfoil section includes an internal passage and a rib that sub-divides the internal passage. At least one of the rib or the endwall section includes a seal cavity. A seal is disposed in the seal cavity.

Flow path assemblies of gas turbine engines are provided. For example, a flow path assembly comprises an inner wall and a unitary outer wall that includes a combustor portion extending through a combustion section of the gas turbine engine and a turbine portion extending through at least a first turbine stage of a turbine section of the gas turbine engine. The combustor portion and the turbine portion are integrally formed as a single unitary structure. The flow path assembly further comprises a plurality of nozzle airfoils, each nozzle airfoil having an inner end radially opposite an outer end. The inner wall or the unitary outer wall defines a plurality of openings therethrough, and each opening is configured for receipt of one of the plurality of nozzle airfoils. Methods of assembling flow path assemblies also are provided.

A rotor blade for a gas turbine engine including an inner ply layer group includes a flared region and an airfoil region, a neck region between the flared region and the airfoil region; and a platform shell that comprises a platform shell root region, a platform shell platform region, and a platform shell neck region between the platform shell root region and the platform shell platform region, the platform shell root region sheathes the flared region, and the platform shell neck region flares outwardly away from the neck region to intersect the platform shell platform region at a platform shell neck region perimeter around an airfoil opening perimeter.

A rotor blade for a gas turbine engine including an inner ply layer group that at least partially defines a base of a root region; a platform around the inner ply layer group that at least partially defines the base; and a platform over-wrap around the platform, the platform over-wrap at least partially defines the base.