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 vane includes an airfoil section which includes an outer wall that defines an internal cavity. A baffle is situated in the internal cavity. The baffle includes a leading edge portion and a trailing edge portion which is bonded to the leading edge portion at a joint. The leading edge portion and the trailing edge portion define an internal cavity therewithin. Both the leading edge portion and the trailing edge portion include a plurality of cooling holes which are configured to provide cooling air to the airfoil outer wall. The trailing edge portion is formed of sheet metal and the leading edge portion is formed of non-sheet-metal. A method of making a baffle for a vane arc segment and a method of assembling a ceramic matrix composite airfoil vane are also disclosed.

A gas turbine engine is provided. The gas turbine engine includes a turbomachine having a low speed spool and a high speed spool; a rotor assembly coupled to the low speed spool; an electric machine rotatable with the low speed spool for extracting power from the low speed spool, for adding power to the low speed spool, or both; and an inter-spool clutch positioned between the low speed spool and the high speed spool for selectively coupling the low speed spool to the high speed spool.

A gas turbine engine includes a ceramic wall for bounding an engine core gas path. The ceramic wall has a ceramic wall first side that faces the engine core gas path and a ceramic wall second side that faces away from the engine core gas path. There is a metallic wall adjacent the ceramic wall second side. The metallic wall has a metallic wall first side that faces the ceramic wall and a metallic wall second side that faces away from the ceramic wall. The metallic wall and the ceramic wall are spaced apart such that there is a channel there between. There is a seal on the ceramic wall second side, and the metallic wall has at least one cooling hole adjacent the seal for emitting cooling air to cool the seal.

A turbine vane is provided. The turbine vane may include an inner shroud having an upper surface and a lower surface, a seal unit disposed in the lower surface of the inner shroud and defining a first region and a second region in the lower surface of the inner shroud, a first impingement unit arranged in the first region and comprising a first impingement plate facing the inner shroud defining a first impingement chamber therebetween, wherein the first impingement plate is configured to receive cooling air and form impingement jet directed to the first impingement chamber, a second impingement unit arranged in the second region and comprising a second impingement plate facing the inner shroud defining a second impingement chamber therebetween, and at least one connector flow channel configured to direct cooling air from the first impingement chamber to the second region, wherein the second impingement plate is configured to receive cooling air from the at least one connector flow channel and form impingement jet directed to the second impingement chamber.

An ejector assembly for a cooling system of a gas turbine engine may comprise: a tail cone having a tail cone outlet in fluid communication with a cooling air flow of the cooling system; an ejector body defining a mixing section, a constant area section, and a diffuser section; and a nozzle section in fluid communication with an exhaust air flow of the gas turbine engine, the ejector assembly configured to entrain the cooling air flow via the exhaust air flow.

An impingement insert for an airfoil of a turbomachine component is provided. The insert includes first and second body parts, each having inner and outer surfaces; and first and second contact parts provided on the outer surfaces of the first and the second body parts. The insert includes a flexible mechanical seal part between the body parts. A flow channel for cooling air is defined by the seal part and the inner surfaces of the body parts. One or both of the body parts include impingement holes. The insert has an elastic part connected to the body parts. When the elastic part is subjected to deformation, the elastic part is configured to apply a force, on the first and/or the second body parts, in a direction of increasing a separation between the first contact part and the second contact part.

An annular ablative gas blocking device provides for automatic altitude compensation of a rocket engine exhaust plume. The nozzle is over expanded at low level launch altitudes and near optimally expanded at the highest altitude at the terminal burnout or staging altitude of the rocket engine. The ablative gas blocking device in the nozzle exit mitigates low altitude launch effects of an over expanded nozzle and inhibits external atmospheric air entrance into the nozzle at launch. The gas blocking means ablatively erodes away from plume impingement as the rocket ascends in a pre-programmed manner to achieve optimum area expansion ratio at all altitudes.

An airfoil vane includes an airfoil section which includes an outer wall that defines an internal cavity. A baffle is situated in the internal cavity. The baffle includes a leading edge portion and a trailing edge portion which is bonded to the leading edge portion at a joint. The leading edge portion and the trailing edge portion define an internal cavity therewithin. Both the leading edge portion and the trailing edge portion include a plurality of cooling holes which are configured to provide cooling air to the airfoil outer wall. The trailing edge portion is formed of sheet metal and the leading edge portion is formed of non-sheet-metal. A method of making a baffle for a vane arc segment and a method of assembling a ceramic matrix composite airfoil vane are also disclosed.

A turbine engine airfoil and method of cooling includes an outer wall defining an exterior surface bounding an interior and defining a pressure side and a suction side extending between a leading edge and a trailing edge to define a chord-wise direction and extending between a root and a tip to define a span-wise direction. The airfoil can also include at least one cooling conduit and an impingement zone located within the at least one cooling conduit.