According to an embodiment, a turbine stage sealing mechanism reduces leak flow, which bypasses a working fluid flow path, between a rotating part and a stationary part in each turbine stage of a gas turbine. A radially inner side end portion of the stationary part is formed to be on a circle centered at a center axis of the rotating part under a predetermined operating condition of the gas turbine.

Disclosed is an aircraft turbine engine (10), comprising a gas generator (12) and a fan (14) arranged upstream from the gas generator (12) and configured to generate a gas inlet stream (F), part of which flows into a duct of the gas generator to form a primary stream (36), the turbine engine (10) comprising an electrical machine that is mounted coaxially downstream from the fan (14) and that comprises a rotor (62a) surrounded by a stator (62b) carried by an annular shroud (64), this shroud (64) being surrounded by a casing (40) of the gas generator that defines, with this shroud (64), a section of the flow duct for the primary stream (36), stationary vanes (42, 68) for straightening this primary stream (36) extending into this path.

A rotary assembly for a turbomachine including a rotor including at least two consecutive rotor stages provided with a plurality of blades and an annular rotor shroud connecting the two consecutive rotor stages, a stator including: at least one stator stage provided between the two consecutive rotor stages including a plurality of vanes, a turbomachine stator vane root, an annular clamping part and an annular support of abradable material, the root extending radially and being clamped axially between the annular support of abradable material and the annular clamping part, A space separates a radially internal end of the root and the annular support of abradable material. A turbojet engine including a rotary assembly as previously.

An abrasive coating for a substrate including a metallic based bond coat layer; a top layer; and an intermediate layer between the metallic based bond coat layer and the top layer. A method of applying an abrasive coating including applying a metallic based bond coat layer onto a substrate; grading an intermediate layer into the metallic based bond coat layer to form a graded transition between the metallic based bond coat layer and the intermediate layer; and grading a top layer into the intermediate layer to form a graded transition between the intermediate layer and the top layer.

An aircraft turbomachine module comprising a ventilation device configured to have a cooling air flow circulate in the turbomachine module, the ventilation device comprising an air outlet, blocking means fixed to the air outlet and mobile between a blocking position of the air outlet and an opening position of the air outlet, and locking means configured to maintain the blocking means in one of the blocking position and the opening position when the temperature within the module is less than a predetermined threshold value, the blocking means being configured to adopt the other of the blocking position and the opening position when the temperature within the module is greater than said predetermined threshold value.

A secondary air system for an aircraft engine comprises an air flow path communicating between a source of pressurized cooling air and an air consuming component. A filter is disposed in the air flow path upstream from the air consuming component. The filter has at least one of: openings of a size selected for capturing suspended particles; and a filter surface material for binding with chemical contaminants.

The gas turbine engine includes a casing assembly located proximate a turbine section of the gas turbine engine, and a tangential on-board injector (TOBI) having a body defining a plurality of air passages extending in a radial direction, the plurality of air passages circumferentially distributed and directing cooling air toward a turbine rotor of the turbine section of the gas turbine engine. An interference fit is provided between a face of the body and a face of a member of the casing assembly, the interference fit defining a fastener-free engagement between the bearing housing and the TOBI to prevent relative movement between the member of the casing and the TOBI.

A gas turbine engine defining a longitudinal direction and a radial direction is provided. The gas turbine engine includes a turbine section comprising a stationary outer portion, a first rotating component, and a second rotating component, wherein the first rotating component includes an aft airfoil defining a hollow passageway and coupled to a radially extended first rotor portion, wherein the hollow passageway is in fluid communication with the stationary outer portion, wherein the second rotating component is coupled to a radially extended second rotor portion; and a seal assembly between a portion of the first rotating component and the second rotating component, the seal assembly defining a balance piston cavity therebetween, wherein the balance piston cavity is in fluid communication with the hollow passageway, wherein a flow of air is routed through the stationary outer portion and the hollow passageway to the balance piston cavity, and wherein a pressure of the flow of air within the balance piston cavity is controlled to provide an axial thrust that counteracts a thrust load on the second rotor portion.

A turbine includes a rotor surrounded by a stator, a first movable stage including a series of rotating vanes, a second movable stage including a second series of rotating vanes, a distributor including a series of stationary vanes, the vanes including platforms jointly delimiting a separation between a main space in which a hot flow circulates through the vanes, and a secondary space surrounding a hub of the rotor. The rotor includes an aperture through which air is blown towards the secondary space, this air being discharged towards the main space. Straightening fins are carried by the distributor in the secondary space for straightening the air flow in order to change its gyration.

To increase the inertia of a sealing lip of a blade for an aircraft turbine engine, and thus improve the service life of such a sealing lip, the sealing lip is conformed so as to have a trough in the outer surface thereof and a corresponding boss in the inner surface thereof, the trough and the boss being defined based on a connection cross section of the sealing lip to a blade body, and being formed at a distance from a free axial end of the sealing lip.