A device configured to generate an air flow, the device including: a compressor; an air flow duct arranged to convey a flow of air generated by the compressor; a gas-filled cavity disposed beside the air flow duct; a wall separating the air flow duct and the cavity, the wall including at least one aperture; and an acoustic resistive screen covering and held in tension over the aperture of the wall. The screen is in fluid contact with air in the air flow duct and gas in the cavity and is configured to resist air flow between the duct and the cavity. The resistive screen and the cavity together define a noise-damping resonator.

A sealing device is coupled to a first structure and is configured to seal a gap between the first structure and a second structure. The sealing device includes a translatable cartridge including a first wall and a second wall parallel to each other and a third wall perpendicular to and coupled to the first and second walls. A force applying device is configured to apply force on the cartridge in a direction toward the second structure such that the cartridge is exerted against the second structure. A ceramic insulation layer disposed at the third wall such that the ceramic insulation layer is between the third wall and the second structure, and an abrasion layer disposed on the ceramic insulation layer.

A non-metallic engine case inlet compression seal for a gas turbine engine includes a non-metallic longitudinal leg section that extends from the non-metallic arcuate interface section and a non-metallic mount flange section that extends from the longitudinal leg section.

A method of fabricating a fiber preform filled with refractory ceramic particles, includes placing a fiber texture including refractory ceramic fibers in a mold cavity; injecting a slip including a powder of refractory ceramic particles present in a liquid medium, the slip being injected into the pores of the fiber texture present in the mold cavity, injection being performed through at least a first face or a first edge of the fiber texture; and draining the liquid medium of the slip that has penetrated into the fiber texture through the porous material part, the draining being performed at least through a second face or a second edge of the fiber texture different from the first face or the first edge, the porous material part also serving to retain the refractory particle powder in the pores of the fiber texture to obtain a fiber preform filled with refractory particles.

A thrust reverser including doors and at least one flexible deflector obstructing a lateral opening of the thrust reverser when the doors are open. Such a deflector allows better controlling the airflows in thrust reverser configuration and maximizing the counter-thrust force.

A high temperature thermal protection systems for rockets, and associated methods, is disclosed. A representative system includes a launch vehicle having a first end and a second end generally opposite the first end. The launch vehicle is elongated along a vehicle axis extending between the first and second ends and carries a propulsion system having at least one nozzle positioned at the second end of the launch vehicle. A thermal protection apparatus positioned around the nozzle is used to provide cooling and/or insulation to the nozzle during the flight of the launch vehicle. The thermal protection apparatus can include multiple fabric layers and an insulation layer stacked and stitched together. The fabric layers can include metal alloy fibers. In representative systems, the thermal protection apparatus can further include provisions for water that saturates the insulation layer to provide further insulating and/or cooling effects.

A door for a thrust reverser of a propulsion assembly, the door being provided with a flexible baffle. The baffle is configured to change state by moving from a folded position to an unfolded position, and vice versa, depending on the configuration of the reverser. When the reverser is in reverse jet configuration, the door is open and the baffle is in an unfolded position in order to direct a flow of gas upstream of the propulsion assembly, thus generating a counter-thrust. When the reverser is in direct jet configuration, the door is closed and the baffle is in a folded position so as to cancel or minimise interference caused by the baffle in the flow of gas passing through the propulsion assembly and thus generating a thrust.

A method of fabricating a fiber-reinforced airfoil ring includes providing a mandrel that has a mandrel suction side and mandrel pressure side, and forming an endless braid around the mandrel. The endless braid conforms to the mandrel suction side and the mandrel pressure side. The endless braid is then consolidated with a matrix material. The mandrel is then removed, leaving the consolidated endless braid as a fiber-reinforced airfoil ring that has a suction side wall that extends between inner and outer platforms, a pressure side wall that extends between the inner and outer platforms, and suction and pressure side mate faces along, respectively, edges of the suction side wall and the pressure side wall, and at least one of the suction or pressure side mate faces includes protrusions along a trailing edge.

A sealing device is coupled to a first structure and is configured to seal a gap between the first structure and a second structure. The sealing device includes a translatable cartridge including a first wall and a second wall parallel to each other and a third wall perpendicular to and coupled to the first and second walls. A force applying device is configured to apply force on the cartridge in a direction toward the second structure such that the cartridge is exerted against the second structure. A ceramic insulation layer disposed at the third wall such that the ceramic insulation layer is between the third wall and the second structure, and an abrasion layer disposed on the ceramic insulation layer.

The present disclosure is directed to a gas turbine engine defining a radial direction, a longitudinal direction, and a circumferential direction, an upstream end and a downstream end along the longitudinal direction, and an axial centerline extended along the longitudinal direction. The gas turbine engine includes a fan assembly including a plurality of fan blades rotatably coupled to a fan rotor in which the fan blades define a maximum fan diameter and a fan pressure ratio. The gas turbine engine further includes a low pressure (LP) turbine defining a core flowpath therethrough generally along the longitudinal direction. The core flowpath defines a maximum outer flowpath diameter relative to the axial centerline. The gas turbine engine defines a fan to turbine diameter ratio of the maximum fan diameter to the maximum outer flowpath diameter. The fan to turbine diameter ratio over the fan pressure ratio is approximately 0.90 or greater.