Turbine engine (10) with a contra-rotating turbine for an aircraft, the turbine engine comprising a contra-rotating turbine (22) whose first rotor (22a) is configured to rotate in a first direction of rotation and is connected to a first turbine shaft (36), and a second rotor (22b) is configured to rotate in an opposite direction of rotation and is connected to a second turbine shaft (38), the first rotor comprising turbine wheels (36a) inserted between turbine wheels (38a) of the second rotor, the turbine engine further comprising a mechanical reduction gear (42) with an epicyclic gear train which comprises a sun gear (44) driven in rotation by said second shaft, a ring gear (40) driven in rotation by said first shaft, and a planet carrier (46) fixed to a stator casing (28) of the turbine engine situated downstream from the contra-rotating turbine with respect to direction of flow of the gases in the turbine engine, the turbine engine further comprising a bearing (56) for guiding the second shaft with respect to the first shaft, and a bearing for guiding the second shaft with respect to said stator casing, characterised in that said bearings are all situated downstream from the trailing edge of the last turbine wheel of the contra-rotating turbine and upstream from the reduction gear.

A turbine engine having counter-rotating rotors comprising a first rotor, rotating in a first rotational direction, defining a first rotor set of blades axially spaced to define a gap, and a second rotor, rotating in a second rotational direction counter the first rotational direction. The second rotor further including a second set of blades received within the gap of the first rotor. A plurality of fluid passages is formed in the first rotor with an outlet facing the gap.

An interdigitated turbine assembly for a gas turbine engine, the interdigitated turbine assembly including a first turbine rotor assembly interdigitated with a second turbine rotor assembly. A first static frame is positioned forward of the first turbine rotor assembly and the second turbine rotor assembly. The first turbine rotor assembly is operably coupled to an inner rotatable component of a gear assembly. The second turbine rotor assembly is operably coupled to an outer rotatable component of the gear assembly. The static structure is connected to the first static frame. A driveshaft is operably coupled to the outer rotatable component. A first bearing assembly is operably coupled to the driveshaft and the first static frame. A second bearing assembly is operably coupled to the first static frame and first turbine rotor assembly. A third bearing assembly is operably coupled to the first turbine rotor assembly and the second turbine rotor assembly.

The invention relates to an aircraft propulsion system (100) intended for being built into the rear of an aircraft fuselage, the propulsion system comprising at least two gas generators (102a, 102b) supplying a power turbine (104) having two counter-rotating turbine rotors (104a, 104b) for driving two fans (112a, 12b), and separate air inlets (106a, 106b) for supplying each gas generator, characterised in that it comprises an electrical drive device (140) configured to rotate at least one of the turbine rotors, at least one electrical generator (142a, 142b) configured to transform part of the energy of the flow from the gas generators into electrical power and an electric motor (146) supplied by said electrical generator and capable of rotating at least one of the turbine rotors, said electrical generator being installed on one of said gas generators, and in that said turbine rotor is capable of being rotated simultaneously by a flow from said gas generators and by the electrical drive device.

The invention relates to an aircraft propulsion system (100) intended for being built into the rear of an aircraft fuselage, the propulsion system comprising at least two gas generators (102a, 102b) supplying a power turbine (104) having two counter-rotating turbine rotors (104a, 104b) for driving two fans (112a, 12b), and separate air inlets (106a, 106b) for supplying each gas generator, characterised in that it comprises an electrical drive device (140) configured to rotate at least one of the turbine rotors, at least one electrical generator (142a, 142b) configured to transform part of the energy of the flow from the gas generators into electrical power and an electric motor (146) supplied by said electrical generator and capable of rotating at least one of the turbine rotors, said electrical generator being installed on one of said gas generators, and in that said turbine rotor is capable of being rotated simultaneously by a flow from said gas generators and by the electrical drive device.

Disclosed is a turbomachine (10) with counter-rotating turbine for an aircraft, comprising: —a high-pressure body, —a low-pressure counter-rotating turbine (22), —a planetary-type mechanical epicyclic reduction gear (42), —a guide bearings (56-62) for the turbine shafts (36, 38), characterised in that said reduction gear (42) and certain of the bearings (60, 62) are housed in a lubrication chamber (86) supplied with oil and comprising dynamic seals (86a-86d), and in that the turbomachine comprises circuits (C1, C2) for pressurising these seals.

Disclosed is a turbomachine (10) with counter-rotating turbine for an aircraft, comprising: —a high-pressure body, —a low-pressure counter-rotating turbine (22), —a planetary-type mechanical epicyclic reduction gear (42), —a guide bearings (56-62) for the turbine shafts (36, 38), characterised in that said reduction gear (42) and certain of the bearings (60, 62) are housed in a lubrication chamber (86) supplied with oil and comprising dynamic seals (86a-86d), and in that the turbomachine comprises circuits (C1, C2) for pressurising these seals.

Methods, apparatus, systems and articles of manufacture for compact compressors are disclosed including a gas turbine engine defining an axial direction and a radial direction, the gas turbine engine including an axial flow compressor and a radial flow compressor, wherein the axial flow compressor is located axially forward of the radial flow compressor, a blade assembly including a splitter shroud to divide incoming air into axial air flow for the axial flow compressor and radial air flow for the radial flow compressor, the blade assembly rotating relative to the axial flow compressor and counter-rotating relative to the radial flow compressor, and wherein the blade assembly is located axially aft of the radial flow compressor.

Methods, apparatus, systems and articles of manufacture for compact compressors are disclosed including a gas turbine engine defining an axial direction and a radial direction, the gas turbine engine including an axial flow compressor and a radial flow compressor, wherein the axial flow compressor is located axially forward of the radial flow compressor, a blade assembly including a splitter shroud to divide incoming air into axial air flow for the axial flow compressor and radial air flow for the radial flow compressor, the blade assembly rotating relative to the axial flow compressor and counter-rotating relative to the radial flow compressor, and wherein the blade assembly is located axially aft of the radial flow compressor.

An apparatus for a turbine engine comprising an outer casing, an engine core provided within outer casing and having a at least one set of blades, and through which gasses flow in a forward to aft direction, an outer drum located within the outer casing to define an annular cavity. A set of seals extending between the first surface and the second surface to define at least one cooled cavity within the annular cavity.