A propulsion system for an aircraft includes a core engine generating an exhaust gas flow that is expanded through a turbine section, a hydrogen fuel system suppling hydrogen fuel to the combustor through a fuel flow path, a condenser extracting water from the exhaust gas flow, an evaporator receiving a portion of the water extracted by the condenser and generating a steam flow. The steam flow is injected into a core flow path upstream of the turbine section and a cooled cooling air system uses water extracted from the exhaust gas flow for cooling a cooling airflow communicated to the turbine section.

A cryogenic cooling system for an aircraft includes a first air cycle machine, a second air cycle machine, and a means for collecting liquid air. The first air cycle machine is operable to output a cooling air stream based on a first air source. The second air cycle machine is operable to output a chilled air stream at a cryogenic temperature based on a second air source cooled by the cooling air stream of the first air cycle machine. An output of the second air cycle machine is provided to the means for collecting liquid air.

A system for an aircraft includes an engine bleed source of a gas turbine engine. The system also includes a means for chilling an engine bleed air flow from the engine bleed source to produce a chilled working fluid. The system further includes a means for providing the chilled working fluid for an aircraft use.

A propulsion system includes an electric fan propulsion motor with a plurality of propulsion motor windings. The propulsion system also includes a means for controlling a flow rate of a working fluid through a cryogenic working fluid flow control assembly to the propulsion motor windings.

A gas turbine engine includes a plurality of rotating components housed within a main compressor section and a turbine section. A first tap is connected to the main compressor section and configured to deliver air at a first pressure. A heat exchanger is connected downstream of the first tap. A cooling air valve is configured to selectively block flow of cooling air across the heat exchanger. A cooling compressor is connected downstream of the heat exchanger. A shut off valve stops flow between the heat exchanger and the cooling compressor. A second tap is configured to deliver air at a second pressure which is higher than the first pressure. A mixing chamber is connected downstream of the cooling compressor and the second tap. The mixing chamber is configured to deliver air to at least one of the plurality of rotating components. A system stops flow between the cooling compressor and the plurality of rotating components. A controller is configured to modulate flow between the heat exchanger and the plurality of rotating components under certain power conditions of the gas turbine engine. The controller is programmed to control the cooling air valve, the shut off valve and the system such that flow is stopped between the heat exchanger and the cooling compressor only after the cooling compressor has been stopped.

A propulsion system for an aircraft includes a core engine that includes a core flow path where a core flow is compressed in a compressor section, communicated to a combustor section, mixed with a hydrogen-based fuel, and ignited to generate a gas flow that is expanded through a turbine section. A fuel system is configured to supply a hydrogen based fuel to the combustor through a fuel flow path. A condenser is arranged along the core flow path and configured to extract water from the gas flow. An intercooling system receives a portion of water from the condenser for cooling a portion of the core flow at a first location within the compressor section. Heated water from the intercooling system is exhausted to a second location within the core flow path downstream of the first location.

A steam injected turbine engine includes a core engine section that generates a first gas flow, a burner where an inlet flow is mixed with fuel and ignited to generate thermal energy, and an evaporator where thermal energy from at least the first gas flow is used to transform water into a first steam flow. The first steam flow is injected into a core flow through the core engine.

A multi-stage turboexpander and a method for generating mechanical power therewith are disclosed. The multi-stage turboexpander includes a casing and a shaft, arranged for rotation in the casing. The shaft is supported by a first bearing at a first shaft end portion and a second bearing at a second shaft end portion. A first radial impeller and a second radial impeller are arranged between the first bearing and the second bearing on the shaft for co-rotation therewith around a rotation axis.

A heat exchanger includes an additively manufactured manifold. The manifold includes an inlet feed manifold and an outlet feed manifold, and a plurality of hypotubes fluidly coupled to the manifold. The hypotubes are round in cross-section, wherein each of the hypotubes has a diameter that has a first value between 0.03 inches and 0.3 inches, and wherein each of the hypotubes has a wall thickness that has a second value between 0.001 inches and 0.0.015 inches.

The occurrence of corrosion on the inner surfaces of pipes is to be minimized in a cooling air system of a gas turbine. Specifically, this gas turbine power generation equipment has: a gas turbine including a turbine connected to a generator, a combustor that supplies combustion gas to the turbine, and a compressor that supplies compressed air to the combustor; a cooling air system that supplies compressed air bled from the compressor to the turbine, the cooling air system being connected at a first end side to an intermediate stage or an outlet of the compressor and connected at a second end side to the turbine; and a drying air system that supplies drying air into the cooling air system when the gas turbine is stopped, the drying air system being connected to the cooling air system.