A particle separator system for a turbine engine having an engine inlet. The particle separator system includes an inlet particle separator located within the engine inlet and configured to remove particles from an incoming airflow. The particle separator system also includes a barrier filter located within an enclosure of the turbine engine downstream of the inlet particle separator, the barrier filter being configured to intercept particles not scavenged by the inlet particle separator.

A fan shroud and a blower device are provided. The fan shroud includes a shroud wall surface and a plurality of ventilation openings. The shroud wall surface extends around a fan along a radial direction orthogonal to a rotation axis. The plurality of ventilation openings are formed through the shroud wall surface in a thickness direction of the shroud wall surface to allow passing of wind generated by a fan motor. A width of a wall surface side opening on a downstream side of the wind in the ventilation opening is smaller than a width of a lateral wall side opening on an upstream side of the wind in the ventilation opening.

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.

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 cooling system includes: a high pressure bleed line configured to bleed high pressure compressed air from a first bleed position of a compressor and to send the air to a first hot part; a low pressure bleed line configured to bleed low pressure compressed air from a second bleed position of the compressor and to send the air to a second hot part; an orifice provided in the low pressure bleed line; a connecting line configured to connect the high pressure bleed line and the low pressure bleed line; a first valve provided in the connecting line; a bypass line configured to connect the connecting line and the low pressure bleed line; and a second valve provided in the bypass line.

Discharge grate intended to be mounted inside or at the outlet of a conduit of a discharge valve of a turbine engine of an aircraft, the discharge grate comprising an upstream face intended to receive a gas flow, a downstream face parallel to the upstream face and intended to deliver the gas flow received on the upstream face, and orifices passing through the perforated plate from the upstream face to the downstream face and intended to convey the gas flow through the perforated plate. The discharge grate comprises for each orifice of the perforated plate a tubular channel, coaxial with the orifice with which it is associated, and projecting from the downstream face of the perforated plate.

A turbine module (2) for a turbomachine (1). The turbine module (2) includes a main channel (26) to guide a main flow (36) through the turbine module (2), a rotor blade (21) and a stator vane (22), the stator vane (22) including a stator airfoil (22) and a platform (23), with the stator airfoil (22) arranged downstream of the rotor blade (21) in the main channel (26), and a cavity (30) including an inlet (31) for injecting a part (36.2) of the main flow (36) into the cavity (30), an outlet (32) for a reinjection of the part (36.2) of the main flow (36) from the cavity (30) into the main channel (26), wherein the cavity (30) is arranged at an axial position of the stator vane (20) and is radially offset from the stator airfoil (22).

An aircraft engine, has: a low-pressure compressor and a high-pressure compressor located downstream of the low-pressure compressor; a gaspath valve upstream of the high-pressure compressor, the gaspath valve movable between an open configuration and a closed configuration; and a bypass flow path having in flow series a bypass inlet, a bypass valve, and a bypass outlet, the bypass inlet fluidly communicating with the gaspath upstream of at least one stage of the low-pressure compressor, the bypass valve having an open configuration in which the bypass valve allows a bypass flow and a closed configuration in which the bypass valve blocks the bypass flow, the bypass outlet fluidly communicating with the bypass inlet via the bypass valve and with the gaspath at a location in the gaspath fluidly downstream of the gaspath valve, downstream of the low-pressure compressor, and upstream of the high-pressure compressor.

A turbofan engine is provided. The turbofan engine includes a fan comprising a plurality of rotatable fan blades, each fan blade defining a fan tip speed; a turbomachine operably coupled to the fan for driving the fan, the turbomachine comprising a compressor section, a combustion section, and a turbine section in serial flow order and together defining a core air flowpath; and a gear box, wherein the turbomachine is operably coupled to the fan through the gear box, wherein a gear ratio of the gear box is greater than or equal to 1.2 and less than or equal to 3.0; wherein during operation of the turbofan engine at a rated speed the fan tip speed is greater than or equal to 1000 feet per second. In exemplary embodiments, during operation of the turbofan engine at the rated speed the fan pressure ratio is less than or equal to about 1.5.

A gas turbine engine includes a fan rotor driven by a fan drive turbine about an axis through a gear reduction to reduce a speed of the fan rotor relative to a speed of the fan drive turbine. A fan case surrounds the fan rotor, and a core engine with a compressor section, including a low pressure compressor. The fan rotor delivers air into a bypass duct defined between the fan case and the core engine. A rigid connection is between the fan case and the core engine includes three triangular-frame connecting members rigidly connected to the fan case at a fan case connection point, and to the core engine at a core engine connection point. The triangular-frame connecting members each are defined by two rigid legs which extend between the fan case and to the core engine, along directions each have a component extending radially inwardly and a component in opposed circumferential directions to each other. A plurality of non-structural fan exit guide vanes and the non-structural fan exit guide vanes are provided with an acoustic feature to reduce noise. The non-structural fan exit guide vanes are rigidly mounted to at least one of the fan case and the core engine.