Control of low pressure recoup air in a gas turbine engine disposed in a gas turbine enclosure with low pressure recoup air piping coupled to a gas turbine combustion exhaust and gas turbine engine enclosure is disclosed. A first valve of the piping controls a flow of the recoup air to the gas turbine combustion exhaust. A second valve of the piping diverts the recoup air to the enclosure for eventual flow to the air intake. A controller controls the flow of the recoup air from the piping to the exhaust and/or the enclosure as a function of ambient and air intake temperature measurements, and a predetermined temperature requirement having an ambient temperature constraint and an air intake temperature differential constraint.

The invention relates to a turbomachine comprising a device for ventilating and pressurising a turbine rotor of a turbomachine of axis (A), the device comprising at least one collection pipe suitable for collecting a fraction of the air circulating in a high-pressure compressor of the turbomachine and conveying it to a first inner chamber (39) inside the turbomachine that communicates with a second inner chamber (40) delimited by the turbine (24) rotor (23), the first and second chambers (39, 40) being at least partially separated by a stationary shroud (56) of axis (A), characterised in that it comprises at least one injector (58) passing through the stationary shroud (56) and having a cross-section that varies in response to a pressure difference between said first and second chambers (39, 40).

A manufacturing method is provided. During this method, a preform component is provided for a turbine engine. The preform component includes a substrate. A meter section of a cooling aperture is formed in the substrate. An internal coating is applied onto a surface of the meter section. An external coating is applied over the substrate. A diffuser section of the cooling aperture is formed in the external coating and the substrate to provide the cooling aperture.

A manufacturing method is provided during which a preform component for a turbine engine is provided. The preform component includes a substrate. An outer coating is applied over the substrate. A characteristic of the outer coating is determined. Instructions for forming a cooling aperture are revised based on the characteristic of the outer coating to provide revised instructions. The cooling aperture is formed in the outer coating and the substrate based on the revised instructions.

The cooling system may comprise: an electric machine; a first conduit including a cable housing and an inlet; a plurality of conductive cables extending from the electric machine, the plurality of conductive cables disposed at least partially in the cable housing; and an electric fan disposed in the first conduit, the cooling system configured to passively flow air through the first conduit to cool the plurality of conductive cables during operation of the gas turbine engine, and the electric fan configured to actively cool the plurality of conductive cables after an engine shutdown of the gas turbine engine.

A propulsor section includes a propulsor having a plurality of blades rotatable about an engine longitudinal axis. A compressor section includes a low pressure compressor and a high pressure compressor. A turbine section includes a low pressure turbine that drives the propulsor through an epicyclic gear arrangement, and includes a second turbine that drives the high pressure compressor. A power ratio is provided by the combination of a first power input of the low pressure compressor and a second power input of the high pressure compressor. The power ratio is defined by the second power input divided by the first power input. The power ratio is equal to, or greater than, 1.0 and less than, or equal to, 1.4.

A fiber-reinforced component for use in a gas turbine engine includes a first braided fiber sleeve forming a cooling channel and a plurality of fiber plies enclosing the first braided fiber sleeve, with the plurality of fiber plies forming first and second walls separated by the first braided fiber sleeve. The fiber-reinforced component further includes a matrix material between fibers of the braided fiber sleeve and the plurality of fiber plies.

A gas turbine engine includes a compressor section that provides first and second compressor stages that are configured to respectively provide first and second cooling fluids. The first compressor stage has a higher pressure than the second compressor stage. The gas turbine engine further includes a component that has platform with an internal cooling passage fed by first and second inlets that respectively receive fluid from the first and second cooling sources. The second inlet is downstream from the first inlet.

Disclosed herein are a ring segment having an air pouch and a first cooling hole formed therein, and a turbine including the same. The air pouch and the first cooling hole are formed in a shield wall, thereby achieving an improvement in cooling performance as well as simplification of production process.

An assembly is provided for an aircraft propulsion system. This assembly includes a first drain tube, a second drain tube, a container and a gas tube. The container fluidly couples the first drain tube to the second drain tube. The container is configured to receive fluid from the first drain tube. The gas tube is fluidly coupled with the container. The gas tube is configured to direct gas into the container for propelling the fluid received within the container into the second drain tube.