An aircraft including a fuselage, a wing structure, at least one turbomachine running on hydrogen and generating thrust at a propulsion unit distant from the fuselage, at least one fuel tank positioned in the fuselage and configured to store hydrogen in the cryogenic state, at least one hydrogen supply device connecting the fuel tank and the turbomachine and including at least one pump positioned in the fuselage in the vicinity of the fuel tank, at least one hydrogen heating system positioned in the fuselage in the vicinity of the pump. This solution makes it possible to reduce a length of the complex double-walled pipes configured for carrying the hydrogen in the cryogenic state between the fuel tank and the hydrogen heating system.

Embodiments of system and methods for supplying fuel, enabling communications, and conveying electric power associated with operation of a hydraulic fracturing unit of a plurality of hydraulic fracturing units are disclosed and may include a fuel line connection assembly configured to be connected to the first hydraulic fracturing unit and to supply fuel from a fuel source to a gas turbine engine connected to the hydraulic fracturing unit. A system also may include a communications cable assembly configured to be connected to the hydraulic fracturing unit and to enable data communications between the hydraulic fracturing unit and a data center or another hydraulic fracturing unit. A system further may include a power cable assembly configured to be connected to the hydraulic fracturing unit and to convey electric power between the hydraulic fracturing unit and a remote electrical power source or the plurality of hydraulic fracturing units.

A method of operating a fuel delivery system of an aircraft engine of an aircraft includes operating the aircraft engine in a standby mode by maintaining combustion in a combustor of the aircraft engine by supplying fuel to the combustor via a first set of fuel nozzles of a first fuel manifold while providing a trickle flow of fuel via a second set of fuel nozzles of a second fuel manifold into the combustor during engine operation, the trickle flow being defined as a fuel flow rate selected to prevent flame-out of the combustion while providing one of: substantially no motive power to the aircraft, and no motive power to the aircraft, via the combustion of the trickle flow of fuel. An aircraft gas turbine engine is also described.

A method of managing thermal energy in a propulsion system includes diverting a flow of bleed air from a compressor section of the propulsion system. An amount of the flow of bleed air diverted from the compressor section is at least 5% of an inlet flow at an inlet of a high pressure compressor of the compressor section. The flow of bleed air is provided to a thermal management system. The flow of bleed air is passed through an expansion turbine of the thermal management system. The flow of bleed air is provided to a thermal load.

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.

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.

A fuel injection system may comprise a mixer and a fuel injector disposed within the mixer. The mixer may comprise an outer housing with an exit port and a bluff body extending across the exit port of the outer housing. A flared surface of the mixer may match a contour of the bluff body.

A fuel injection system may comprise a mixer and a fuel injector disposed within the mixer. The mixer may comprise an outer housing with an exit port and a bluff body extending across the exit port of the outer housing. A flared surface of the mixer may match a contour of the bluff body.

A tuned mass damper for reducing vibration on a component includes a shaft connector member configured to be coupled to the component and a cable termination member. The tuned mass damper also includes at least one cable coupled to the shaft connector member and to the cable termination member such that vibration of the component is transferred to the at least one cable via the shaft connector member and increased or decreased by the at least one cable.

The present application discloses a method of determining one or more fuel characteristics of an aviation fuel used for powering a gas turbine engine of an aircraft. The method comprises: determining one or more performance parameters of the gas turbine engine during a first time period of operation of the gas turbine engine; and determining one or more fuel characteristics of the fuel based on the one or more performance parameters. A method of operating an aircraft, a fuel characteristic determination system, and an aircraft are also disclosed.