Aircraft engines include a turbine engine comprising a compressor section, a burner section, and a turbine section arranged along a shaft, with a core flow path through the turbine engine such that exhaust from the burner section passes through the turbine section, a condensing assembly arranged downstream of the turbine section of the turbine engine along the core flow path, and an exhaust compressor arranged downstream of the condensing assembly along the core flow path. The condensing assembly is configured to reduce a mass flow of the exhaust compressor by condensing water vapor from the core flow and removing liquid water from the core flow.

A cooling structure for a trailing edge of a turbine blade is provided. The cooling structure for the trailing edge of the turbine blade comprising an airfoil shaped blade part including a leading edge, a trailing edge, a pressure surface and a suction surface connecting the leading edge and the trailing edge, and a cavity channel formed in the blade part and through which a cooling fluid flows, the cooling structure including slots and lands arranged alternately on the trailing edge along a span direction of the pressure surface by cutting a portion of the pressure surface, the slots communicating with the cavity channel and defined by adjacent lands where the pressure surface remains, wherein a pin-fin structure is disposed in the cavity channel on an upstream side of the slot, and wherein the cooling fluid is introduced through a micro-channel formed inside the pin-fin structure and is discharged through film cooling holes formed in the pressure surface.

An integral cooling system for a turbine casing and guide vanes in an aeroengine is provided, belonging to the field of research on flow and heat exchange of a turbine casing in an aeroengine. An inner guide ring and multiple of guide vanes are mounted on the turbine casing; the cooling system includes an electromagnetic pump, a heat exchanger, an expansion joint and a cooling pipeline; an annular cavity is provided in the turbine casing, the cooling pipeline is mounted on the inner wall of the annular cavity and periodically and uniformly distributed along the circumferential direction of the turbine casing, and the cooling pipeline is filled with cooling liquid; a mounting cavity is further provided in the turbine casing, and the mounting cavity communicates with the annular cavity; the electromagnetic pump, the expansion joint and the heat exchanger are all mounted in the mounting cavity.

A gas turbine engine includes a compressor section and a turbine section operably coupled to the compressor section. The gas turbine engine further includes a means for selectively releasing a cooling fluid flow produced at a cryogenic temperature and a plumbing system in fluid communication with the means for selectively releasing the cooling fluid flow. The plumbing system is configured to route the cooling fluid flow to one or more of the compressor section and the turbine section.

A turbofan engine including an exhaust casing traversed by a primary flow and surrounded by a secondary flow and an outlet cone carried by this exhaust casing, the exhaust casing and the outlet cone together defining an internal space. The exhaust casing includes a hollow radial arm traversing the primary flow in order to convey part of the secondary flow so as to form a cooling flow for supplying the internal space, and the outlet cone is terminated by an opening for discharging the cooling flow.

A turbine component includes a hot wall, a cold wall spaced apart from the hot wall and a conduit defined between the hot wall and the cold wall. A cooling system is defined in the conduit. The cooling system includes a plurality of cooling pins, each including a first end having a first cross-sectional area and a second end having a second cross-sectional area. Each cooling pin includes a body extending between the first end and the second end, with a pin leading edge defined along the body from the first end to the second end. The pin leading edge is defined by a first diameter and a pin trailing edge is defined by a second diameter. At least one first cooling pin has the first end coupled to the hot wall and the second end coupled to the cold wall with a fillet.

An aspect of the present disclosure is directed to a system for reducing thermal gradient at a heat engine. The heat engine includes a rotor assembly with a rotor disk and a seal assembly is provided. An interfacing structure at least partially surrounds the rotor assembly at the seal assembly. The seal assembly and the interfacing structure together form a first cavity defining a first environmental condition and a second cavity defining a second environmental condition. A fluid supply manifold is connected to the rotor assembly and is extended at least partially along a radial direction from the first cavity to an outlet opening in thermal communication with the rotor disk of the rotor assembly.

An engine includes: a first rotating component; a second rotating component separate from the first rotating component; and an electric machine, the electric machine including a first rotor rotatable with the first rotating component; a second rotor rotatable with the second rotating component; and a stator assembly arranged between the first rotor and the second rotor, the stator assembly including a first set of windings arranged adjacent to the first rotor, a second set of windings arranged adjacent to the second rotor, and a non-ferromagnetic inner housing arranged between the first set of windings and the second set of windings.

A gas turbine engine component with a core includes a first core portion configured to provide a first cooling passage. A second core portion is spaced from the first core portion and configured to provide a second cooling passage. The second core portion includes a longitudinal leg and a resupply leg transverse to and intersecting the longitudinal leg. The resupply leg has a terminal end and is configured to provide a resupply channel. A connector interconnects the terminal end to the first core portion. The connector is configured to provide a resupply hole.

There is provided a gas turbine engine comprising a blower system for supplying pressurised air to an airframe via an airframe port. The blower system comprises a compressor configured to receive air from a bypass duct or a core of the gas turbine engine and to discharge compressed air into a delivery line extending from the compressor to the airframe port. The blower system also comprises a heat exchanger configured to transfer heat from the compressed air to a coolant and a valve arrangement configured to switch between operation of the blower system in a baseline mode and a cooling mode, the valve arrangement being configured to: selectively divert the compressed air within the delivery line to the heat exchanger for operation in the cooling mode; and/or selectively provide the coolant to the heat exchanger for operation in the cooling mode.