An aircraft propulsion system apparatus includes a first skin, a second skin, an intermediate layer between the first skin and the second skin, a first cellular core and a second cellular core. The first cellular core is connected to the first skin and the intermediate layer. The first cellular core includes a plurality of first core chambers, where a first of the first core chambers is fluidly coupled with one or more first perforations in the first skin and one or more first perforations in the intermediate layer. The second cellular core is connected to the intermediate layer and the second skin. The second cellular core includes a plurality of second core chambers and a plurality of corrugations, where a first of the second core chambers is fluidly coupled with the first of the first core chambers through the one or more first perforations in the intermediate layer.

An anti-bird strike protection net for aircraft jet engine, comprising fixed circular truncated cone, conical protection net body, connection section, and double helix electrical heating wire, wherein the fixed circular truncated cone is secured on the cowl inlet's outside of the aircraft jet engine via which the conical protection net body is connected to the jet engine with the connection section: hinge, catches, and gas springs; The conical protection net body comprises the base of the conical protection net body, the conical support frame, the net which covers the outside of the conical support frame; the double helix electrical heating wire spirally surrounding the surface of the net is used as deicing device.

This invention is an aftermarket add-on device and can be an original integration part as jet engine manufacturing. It's simple and easy to be secured on the cowl inlet without changing the structure of the aircraft jet engine.

A seal carrier for a turbomachine, in particular a gas turbine, including a carrier base and at least one seal member, the at least one seal member being connected to the carrier base, and the at least one seal member being formed by a plurality of cavities arranged adjacent one another, in particular in a regular fashion, in the circumferential direction and the axial direction, the cavities extending from the carrier base in the radial direction, is provided. At least one stiffening element on the carrier base, the stiffening element extending along the circumferential direction and at least partially covering the at least one seal member at one of its axial end regions is provided.

An airfoil for a turbine engine having first and second radially extending cooling chambers separated by a radially extending rib in which rows of crossover holes pass through the rib and fluidly coupling the first and second chambers.

A damping device for being situated between a housing wall of a housing of a thermal gas turbine and a casing ring is provided. The casing ring includes an area radially internal with regard to a rotation axis of a rotor of the thermal gas turbine and facing rotating moving blades of the gas turbine. The damping device includes at least sectionally a porous damping structure. A method for manufacturing this type of damping device as well as to a thermal gas turbine, in particular an aircraft engine, in which this type of damping device is situated in a housing of the gas turbine between a housing wall and a casing ring are also provided.

A host gas flow path component includes a body including a leading edge, a trailing edge, a first side edge, a second side edge, and a pair of opposed lateral sides. A first lateral side is configured to interface with a cavity having a cooling fluid. The hot gas flow path component includes a supply channel disposed within the body and extending from the cavity to adjacent the leading edge or the trailing edge. The hot gas flow path component includes a channel disposed within the body adjacent the trailing edge or the leading edge. The channel extends across the body in a direction from the first side edge toward the second side edge. The channel is configured to receive the cooling fluid from the cavity to cool the trailing edge or the leading edge via an intermediate channel extending between the supply channel and the channel.

A lubricant nozzle for a planetary gear set speed reducer of a turbomachine, the nozzle having a generally elongate shape and including a body with a longitudinal axis B, the body having a longitudinal inner cavity in fluid communication with a lubricant inlet located at a longitudinal end of the body and with lubricant outlet apertures that are provided in an annular wall of the body and that extend substantially radially relative to axis B, wherein the apertures are formed in at least one boss of the body, which boss projects radially outwards on the wall and extends, about the axis, at an angular extent that does not exceed 180°.

A guide vane within an annular inlet duct of a gas-turbine engine provides for generating swirl within an annular inlet duct so as to provide for reducing the rate of deceleration of the inlet air flow within the annular inlet duct while providing for diffusion of the meridional component of velocity thereof.

A case for a gas turbine engine includes a case wall and a boss that extends from the case wall. The boss includes a perimeter step.

A turbomachine assembly, comprising a distributor with a tubular platform, an inner radial flange extending radially inwardly from the platform, and a plurality of vanes extending radially outwardly from the platform; a casing comprising an outer radial flange; an abutment axially holding the outer radial flange in contact with the inner radial flange; a support ring attached to the casing and radially supporting the abutment; wherein the abutment is formed of a head and a body releasably connected to the head.