A gas turbine engine is provided having a static structure including a flowpath wall. A fluid circuit is extended through the flowpath wall and includes a first inlet opening in fluid communication with a first cavity to receive a first flow of fluid through the fluid circuit. The static structure includes an ejector positioned at the fluid circuit, in which the ejector includes a second inlet opening in fluid communication with a second cavity to receive a second flow of fluid through the ejector and into the fluid circuit.

A turbine rotor includes a base and a plurality of blades. The base and the blades curve such that radially outward portions of the base and the blades extend in a direction with a greater component in a radial direction than in an axial direction. Radially central portions of the base and the blade extend in a direction with the two components being closer. Radially inner sections of the base and the blades extend in a direction with a greater component in the axial direction than in a radial direction. There is a cooling channel arrangement in the turbine rotor. The cooling channel arrangement includes impingement cooling for a nose and serpentine passages for cooling sections of the platform circumferentially intermediate the blades, and distinct serpentine passages for cooling the plurality of blades. A turbomachine and method are also disclosed.

An airfoil includes an airfoil section that is formed of a fiber-reinforced composite that has fiber plies. The fiber plies include at least one overwrap fiber ply, first and second support fiber plies, and a serpentine fiber ply. The overwrap fiber ply circumscribes an internal cavity and defines first and second sides and leading and trailing ends of the airfoil section. The first and second support fiber plies define respective first and second radial tubes in the internal cavity. The serpentine fiber ply winds from the first side and around the first radial tube to the second side, then from the second side back to the first side, and then from the first side and around the second radial tube back to the second side.

A turbine rotor blade root is additively manufactured and includes a shank having a radially extending chamber defined therein. A blade mount is at a radial inner end of the shank. The blade mount has a hollow interior defined therein with the hollow interior in fluid communication with the radially extending chamber. A lattice support structure is disposed within the hollow interior of the blade mount.

A gas turbine engine includes a blade outer air seal, a ceramic vane, and a cooling passage circuit that extends through a first internal passage in the blade outer air seal and a second internal passage in the ceramic vane.

A blade of a turbo machine, having a blade leaf, with a flow leading edge, a flow trailing edge, and flow conducting surfaces, and a cooling passage integrated in the blade leaf. In the region of the blade leaf cooling passage portions extend substantially in the radial direction. Adjacent cooling passage portions merge into one another via a diversion passage portion having a material web extending between the adjacent cooling passage portions. The respective material web ends in the region of the respective diversion passage portion. The respective material web has a defined axial width between the respective adjacent cooling passage portions and the respective material web in the region of the respective diversion passage portion has a material thickening enlarging the axial width by at least 20%.

In one exemplary embodiment, a gas turbine engine is provided. The gas turbine engine defines a radial direction, an axial direction, and an axis extending along the axial direction of the gas. The gas turbine engine includes: a shaft configured to rotate about the axis; an electric machine comprising a rotor coupled to and rotatable with the shaft and a stator, the rotor defining an end along the axial direction; and a cooling manifold rotatable with the rotor and positioned at the end of the rotor, the cooling manifold configured to receive a flow of cooling fluid and provide the cooling fluid to the stator during operation of the gas turbine engine.

An airfoil includes an airfoil wall that defines a leading end, a trailing end, and first and second sides that join the leading end and the trailing end. A rib connects the first and second sides of the airfoil wall. The rib defines a tube portion that circumscribes a rib passage, and first and second connector arms that solely join the tube portion to, respectively, the first and second sides of the airfoil wall.

Devices for distributing oil from a rolling bearing for an aircraft turbine engine include a rolling bearing including two rings, respectively an inner ring and an outer ring, an oil distribution ring configured to be mounted on a turbine engine shaft, said distribution ring including a first outer cylindrical surface for mounting the inner ring of the bearing, an oil recovery scoop supplying a lubricating circuit of the bearing, and an annular track of a dynamic seal. The distribution ring and the track are formed by a single-piece body, and the lubricating circuit is formed in said body and extends into the distribution ring and the track.

A rotor blade for a turbomachine is provided. The rotor blade includes a platform, an airfoil extending from the platform, and a cooling circuit extending within the platform and the airfoil. The cooling circuit includes a plurality of cooling passages defined by a plurality of ribs. The plurality of ribs includes an offset rib.