A gas turbine engine having a fluid transfer connection is provided. The gas turbine engine includes a first component configured to channel a flow of air from a portion of the gas turbine engine; a second component configured to receive the flow of air from the first component, wherein the first component and the second component are movable relative to one another; a sleeve portion disposed between the first component and the second component; and a spoolie device disposed within the sleeve portion, the spoolie device having elliptical shaped opposing ends, wherein the spoolie device bridges the first component and the second component to channel the flow of air from the first component to the second component.

An assembly for a turbine engine includes a combustor wall. The combustor wall includes a shell, a heat shield and an annular land. The heat shield is attached to the shell. The land extends vertically between the shell and the heat shield. The land extends laterally between a land outer surface and an inner surface, which at least partially defines a quench aperture in the combustor wall. A lateral distance between the land outer surface and the inner surface varies around the quench aperture.

Provided are a vacuum pump which can be reduced in size without reducing exhaust performance, and a connection port which is used in the vacuum pump. The connection port which connects the vacuum pump and an auxiliary pump includes an oval first opening connectable to the vacuum pump and a circular second opening connectable to the auxiliary pump, and a diameter dimension of the first opening is set to be smaller than a diameter dimension of the second opening in an axial direction of the vacuum pump.

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.

Baffle inserts for airfoils of gas turbine engines are described. The baffle inserts include a baffle insert body having a first side portion and a second side portion, wherein each side portion has a respective end, a first set of vortex generation elements is arranged at the end of the first side portion, and a second set of vortex generation elements is arranged at the end of the second side portion. The first set of vortex generation elements and the second set of vortex generation elements are arranged at an aft end of the baffle insert body.

An impeller associated with a compressor of a gas turbine engine includes a plurality of impeller blades. Each impeller blade of the plurality of impeller blades has a leading edge and a trailing edge opposite the leading edge in a streamwise direction. Each impeller blade of the plurality of impeller blades extends in a spanwise direction from a hub at 0% span to a tip at 100% span, and each impeller blade of the plurality of impeller blades has a plurality of mean camber lines that each extend from the leading edge to the trailing edge at a respective spanwise location. The leading edge includes a partially swept leading edge surface defined between 70% span to 100% span that extends in the streamwise direction between 3% to 15% of a mean camber line at 100% span.

A turbine blade includes an airfoil body having an outer tip and a platform; and a part-span shroud positioned between the outer tip and the platform of the airfoil body. The part-span shroud has a first opening extending through the airfoil body and having a first inner surface. The airfoil body includes a second opening extending radially from the first opening and having a second inner surface. A first elongated vibration-damping element is disposed in the first opening, and a second elongated vibration-damping element disposed radially in the second opening. The second elongated vibration-damping element includes a free radially outer end and a radially inner end coupled to the first elongated vibration-damping element. The first elongated vibration-damping element frictionally damps vibration, and the second elongated vibration-damping element damps vibration using impact within the second opening.

Impingement insert for an airfoil of a blade/vane of a gas turbine is provided. The impingement insert includes a double-walled section having an outer and an inner walls, that define—an inner channel at an inner surface of the inner wall, an outer channel at an outer surface of the outer wall and a middle channel between the outer and the inner walls. Impingement cooling holes are provided in the outer wall that use the cooling air of the middle channel to eject impingement jets into the outer channel. The impingement insert includes at least one extraction duct that extends between the outer and the inner walls across the middle channel, and has an inlet at the outer channel, and an outlet at the inner channel, for flowing the cooling air, after impingement, from the outer channel into the inner channel.

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 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 vanes. The first rotor disk is configured to rotate about a rotational axis. The first rotor disk is configured from or otherwise includes disk material. 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 vanes are arranged circumferentially around the rotational axis and axially between the first rotor disk and the second rotor disk. The vanes include a first vane, which first vane is configured from or otherwise includes vane material that is different than the disk material.