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 cascade assembly for a thrust reverser of an aircraft engine. The cascade broadly comprises a number of vanes formed of a pliable material and shiftable between a collapsed position when the thrust reverser is in a stowed configuration and a distended position when the thrust reverser is in a deployed configuration to redirect fan duct flow in a reverse thrust flow opening created by the thrust reverser.

A cascade assembly for a thrust reverser of an aircraft engine. The cascade broadly comprises a number of vanes formed of a pliable material and shiftable between a collapsed position when the thrust reverser is in a stowed configuration and a distended position when the thrust reverser is in a deployed configuration to redirect fan duct flow in a reverse thrust flow opening created by the thrust reverser.

A fire protection structure to be affixed to a substrate adapted to be exposed to a heat source, has: an expandable layer containing an intumescent material expandable when exposed to the heat source, the expandable layer expandable, when exposed to a temperature above a temperature threshold generated by the heat source, from a first thickness to a second thickness greater than the first thickness; and a compliant textile structure retaining the expandable layer, at least a portion of the expandable layer disposed relative to the compliant textile structure such that the portion of the expandable layer is disposed between the substrate and the compliant textile structure when the fire protection structure is attached to the substrate, the compliant textile structure deformable with expansion of the expandable layer from the first thickness to the second thickness.

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.

The present invention relates to a blade for a gas turbine, in particular of an aircraft engine, which is produced at least in part from ceramic matrix composite with a plurality of superimposed fabric layers, wherein at least one pair of superimposed fabric layers comprises a first fabric layer that has at least one first point of interruption between two mutually facing first edges of this fabric layer, and a second fabric layer that has at least one second point of interruption adjacent to the first point of interruption between two mutually facing second edges of the second fabric layer, this second point being displaced from the first point of interruption.

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 heat insulating sheet member capable of enhancing a heat insulating performance of an exhaust gas introduction path by an easy operation, including such a path in a turbocharger. The heat insulating sheet member is a bendable member formed from an inorganic flexible material. The heat insulating sheet member includes a first region corresponding to an inlet of a bottom wall portion, a second region corresponding to at least a terminating end portion of the scroll portion, a third region provided between the first region and the second region and corresponding to a coupling wall portion, and a fourth region corresponding to an outer peripheral wall portion. The first region and the third region, the third region and the second region, and the first region and the fourth region are coupled to each other with the inorganic flexible material in a continuous state.

The present disclosure relates to a variable stator vane and a method of fabricating the variable stator vane of a gas turbine engine. The method includes providing at least one fibre sheet. The method further includes rolling the at least one fibre sheet around a mandrel to form a spindle section of the variable stator vane. An excess of material of the at least one fibre sheet remains after forming the spindle section. The method further includes using the excess of material of the at least one fibre sheet to form the at least one aerofoil section of the variable stator vane.

A rotor assembly is provided for a piece of rotational equipment. This rotor assembly includes a rotor disk, a rotor blade and a seal element. The rotor disk is configured to rotate about a rotational axis. The rotor blade includes an airfoil, a platform and a mount attaching the rotor blade to the rotor disk. The seal element is seated in a groove of the rotor disk. The seal element is configured to sealingly engage the platform and the mount.