A variable-pitch bladed disc including a plurality of blades, each at a variable pitch in relation to a rotation axis of the blade and each having a root, the plurality of blades including at least one first blade and at least one second blade, a plurality of rotor connecting shafts, each having a root and a tip, each root being mounted at the tip of a corresponding rotor connecting shaft by way of a pivot so as to allow each blade to rotate about the blade rotation axis, the first blade having a first blade inclination, such that the first blade is inclined in a fixed manner with respect to the blade rotation axis of the first blade, and the second blade having a second blade inclination different from the first blade inclination.

The invention relates to an aircraft turboprop engine having two unducted propellers, coaxial about a rotation axis (LL), respectively upstream (12) and downstream (14′), each propeller comprising an annular row of blades, the blades of the downstream propeller (14′) each having a truncated head and each comprising a leading edge (20′) and a trailing edge (22′), the radially external ends (A2, B1) of which are distant from each other and are connected by a terminal edge (21′), characterised in that the terminal edge (21′) of each blade of the downstream propeller (14′) has, at at least two points, tangents in a meridian plane having different slopes, the upstream end (A) of the terminal edge defining a maximum radial dimension (R2) of the blade and its downstream end (B1) defining a minimum radial dimension (R3) of the terminal edge.

The invention relates to an aircraft turboprop engine having two unducted propellers, coaxial about a rotation axis (LL), respectively upstream (12) and downstream (14′), each propeller comprising an annular row of blades, the blades of the downstream propeller (14′) each having a truncated head and each comprising a leading edge (20′) and a trailing edge (22′), the radially external ends (A2, B1) of which are distant from each other and are connected by a terminal edge (21′), characterised in that the terminal edge (21′) of each blade of the downstream propeller (14′) has, at at least two points, tangents in a meridian plane having different slopes, the upstream end (A) of the terminal edge defining a maximum radial dimension (R2) of the blade and its downstream end (B1) defining a minimum radial dimension (R3) of the terminal edge.

A variable pitch bladed disc including a plurality of blades, each being of variable pitch about a blade axis of rotation and having a root, the plurality of blades including at least one first blade and at least one second blade, a plurality of rotor connecting shafts, each shaft having a root and a tip, the root of each blade being mounted on the tip of a corresponding rotor connecting shaft via a pivot so as to allow each blade to be rotated about the blade axis of rotation, the first blade having a first rotation axis inclination such that the rotation axis thereof is inclined in a fixed manner with respect to a radial axis passing through the root of the corresponding shaft, and the second blade has a second rotation axis inclination different from the first rotation axis inclination.

A variable pitch bladed disc including a plurality of blades, each being of variable pitch about a blade axis of rotation and having a root, the plurality of blades including at least one first blade and at least one second blade, a plurality of rotor connecting shafts, each shaft having a root and a tip, the root of each blade being mounted on the tip of a corresponding rotor connecting shaft via a pivot so as to allow each blade to be rotated about the blade axis of rotation, the first blade having a first rotation axis inclination such that the rotation axis thereof is inclined in a fixed manner with respect to a radial axis passing through the root of the corresponding shaft, and the second blade has a second rotation axis inclination different from the first rotation axis inclination.

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 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 turbine engine assembly is provided. The turbine engine assembly includes a core gas turbine engine including a first rotatable drive shaft, a first low-pressure turbine section in serial flow communication with the gas turbine engine, a gear assembly coupled to the first low-pressure turbine section through a second rotatable drive shaft, and a second low-pressure turbine section in serial flow communication with the core gas turbine engine. The first low-pressure turbine section is configured to rotate in a first rotational direction, and the second low-pressure turbine section is configured to rotate in a second rotational direction opposite the first rotational direction. The first and second low-pressure turbine sections are spaced axially apart from each other. The turbine engine assembly also includes a fan assembly coupled to the first low-pressure turbine section through the gear assembly, and coupled to the second low-pressure turbine section through a third rotatable drive shaft.

A turbine engine assembly is provided. The turbine engine assembly includes a core gas turbine engine including a first rotatable drive shaft, a first low-pressure turbine section in serial flow communication with the gas turbine engine, a gear assembly coupled to the first low-pressure turbine section through a second rotatable drive shaft, and a second low-pressure turbine section in serial flow communication with the core gas turbine engine. The first low-pressure turbine section is configured to rotate in a first rotational direction, and the second low-pressure turbine section is configured to rotate in a second rotational direction opposite the first rotational direction. The first and second low-pressure turbine sections are spaced axially apart from each other. The turbine engine assembly also includes a fan assembly coupled to the first low-pressure turbine section through the gear assembly, and coupled to the second low-pressure turbine section through a third rotatable drive shaft.