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.

Counter-rotating turbine of a turbomachine extending about an axis of rotation and comprising an inner rotor having 5 at least one inner moving blade supported in rotation by a first shaft, an outer rotor rotating in an opposite direction to the inner rotor, and comprising at least one outer moving blade supported in rotation by a second shaft coaxial with the first shaft, the first 10 and second shafts extending axially from upstream to downstream of the turbine, the first shaft being guided in rotation by a first bearing disposed between the first shaft and an upstream casing of the turbine, and the second shaft being guided in rotation by a second bearing disposed between the second 15 shaft and said upstream casing of the turbine, the first bearing and the second bearing being disposed upstream of a first stage of the inner rotor.

Turbomachinery comprising first and second sets of rotors configured to operate on a working fluid. The machinery further comprises first and second sets of electric machines coupled to the respective first and second rotors, and a coupling arrangement arranged to couple adjacent rotors of the first and second rotor sets to provide for fixed ratio, contra-rotation of the first and second rotor sets.

The invention relates to a turbomachine for an aircraft, comprising: a first rotor comprising a first shaft, a second rotor comprising a second shaft, a mechanical reduction gearing having an epicyclic gear set comprising a sun gear connected to the second shaft, a ring gear connected to the first shaft, and planet gears located between the sun gear and the ring gear and borne by a planet carrier attached to a stator of the turbomachine, rolling element bearings for guiding said first shaft and second shaft in rotation, an annular gutter which extends around the ring gear of the reduction gearing and which is configured to recover oil for lubricating the reduction gearing that is sprayed by centrifugal action out from the ring gear during operation, and an annular bearing support which is attached, with the gutter, to a stator of the turbomachine and which supports at least one of said bearings, characterized in that it also comprises: at least one device for conveying oil recovered by said gutter, which device is borne by said annular support and extends as far as said at least one bearing in order to lubricate the latter.

The main purpose of the invention is an air circulation device (1) for a turbomachine (10), comprising an air conveyance circuit (2, 4b, 9, 4a, 3) adapted to bring hot bleed air (A1) from the turbomachine (10) to a part to be heated (38), comprising a first segment fixed in rotation to a rotating part (31, 24) and comprising at least one hot air (A2) conveyance conduit (3, 9), and a hot air passage device (4a, 4b), comprising an annular compartment fixed in rotation to the first segment, characterise in that the annular compartment comprises a heat exchanger in contact with external air, and in that the hot air passage device (4a, 4b) comprises a hot air bypass system to deviate air entering into the device and to make it circulate along the heat exchanger when the temperature of this intake air is above a predetermined threshold.

A counter-rotating shaft assembly of a gas turbine engine includes an outer shaft rotatable in a first direction about a virtual rotational axis, an inner shaft counter-rotatable about the virtual rotational axis in a second direction that is opposite to the first direction, a differential bearing rotatably connecting the two shafts, a centering spring connecting the inner shaft to the differential bearing, and a squeeze film damper between the differential bearing and the inner shaft.

A power transmission apparatus has a housing, a first shaft rotatably mounted within the housing, a second shaft rotatably mounted within the housing and extending around at least a portion of the first shaft, a third shaft exterior of the first and second shafts and positioned within the housing, a first transmission connected to the second shaft and to the third shaft such that a rotation of the second shaft causes a rotation of the third shaft, a second transmission connected to the first shaft and to the third shaft such that a rotation of the first shaft applies rotational energy to the third shaft, and a power receiver connected to the third shaft so as to convert the rotational energy of the third shaft into energy or motion.

A gas turbine engine having an interdigitated turbine assembly including a first turbine rotor and a second turbine rotor, wherein a total number of stages at the interdigitated turbine assembly is between 3 and 8, and an average stage pressure ratio at the interdigitated turbine assembly is between 1.3 and 1.9. A gear assembly is configured to receive power from the interdigitated turbine assembly, and a fan assembly is configured to receive power from the gear assembly. The interdigitated turbine assembly and the gear assembly are together configured to allow the second turbine rotor to rotate at a second rotational speed greater than a first rotational speed at the first turbine rotor. The fan assembly and the gear assembly are together configured to allow the fan assembly to rotate at a third rotational speed less than the first rotational speed and the second rotational speed. The interdigitated turbine assembly, the gear assembly, and the fan assembly together have a maximum AN.sup.2 at the second turbine rotor between 30 and 90.

A rotor for a contrarotating turbine comprising a drum and a blading mounted inside, the drum comprising a hook delimiting a housing having an outer wall and an inner wall, the blading comprising a blade and an outer platform provided with spoiler placed inside the housing, wherein the rotor comprises a foil comprising an elastic inner wing and an outer wing, the outer wing being arranged radially between the spoiler and the outer wall, the inner wing having a first support with the inner wall and a second support with the spoiler, the inner wing being arranged in the housing so as to exert a force on the spoiler so as to press the spoiler against the outer wall via the outer wing.

An airfoil assembly defining a primary airflow path for a turbine engine comprising a platform, an airfoil extending from the platform and into at least a portion of the primary airflow path, a secondary airflow path comprising air from the primary airflow path, and a deflector provided within the secondary airflow path.