A shroud unit has multiple stay members for supporting an electric motor, which rotates a fan unit having multiple blade members. Each of the blade members is inclined with respect to a radial direction of the fan unit in a predetermined circumferential direction. The stay members include a first stay member and multiple second stay members. Each of the second stay members is inclined with respect to the radial direction in a circumferential direction opposite to the predetermined circumferential direction. The first stay members is inclined with respect to the radial direction in the predetermined circumferential direction. The first stay member is located at a position overlapping with one of air-flow areas of a circular opening formed in the shroud unit, when viewed it in a direction along a rotational center axis of the electric motor. A flow amount of the air passing through the air-flow area of the circular opening having the first stay member is smaller than a flow amount of the air passing through another air-flow area of the circular opening having the second stay member.

An assembly includes a variable area inlet and an inlet duct. The variable area inlet includes an inlet structure and a center body structure. The inlet structure extends circumferentially about the center body structure with an outer inlet passage radially between the center body structure and the inlet structure. The center body structure includes an outer body and an inner body. The outer body extends circumferentially about the inner body with an inner inlet passage radially between the inner and outer bodies. The inner body is configured to move along a centerline relative to the outer body between a first position and a second position. The inlet duct is fluidly coupled with the outer inlet passage when the inner body is in the first position. The inlet duct is fluidly coupled with the outer inlet passage and the inner inlet passage when the inner body is in the second position.

A combustion chamber structure for a rocket engine includes a hot gas wall (12) that surrounds a combustion chamber (40) and has a plurality of first coolant channels (50) and a plurality of second coolant channels (52). The plurality of first (50) and second (52) coolant channels extend from a first longitudinal end (16) of the hot gas wall (12) to a second longitudinal end (18) of the hot gas wall (12) opposite to the first longitudinal end (16). The combustion chamber structure (10) further comprises a first manifold (20) forming a first coolant chamber (30) and a second manifold (22) forming a second coolant chamber (32) being fluidly separated from the first coolant chamber (30). The first (20) and second (22) manifolds are provided at the first longitudinal end (16) of the hot gas wall (12) and extend in a circumferential direction of the hot gas wall (12). The first coolant chamber (30) is fluidly connected to each of the plurality of first coolant channels (50) and the second coolant chamber (32) is fluidly connected to each of the plurality of second coolant channels (52).

The invention relates to an aircraft turbine engine (101) which comprises, downstream of the low pressure compressor of its low pressure body (104) and upstream of its combustion chamber (102), an electric motor (206) which is configured to rotate a rotor blade ring so as to generate a flow of air. Moreover, the rotational speed of the electric machine rotor is independent of the rotational speed of the compressor rotor.

A turbocharger includes a first housing configured to house a turbine impeller, a second housing configured to rotatably support a rotating shaft to which the turbine impeller is fixed, and a variable capacity mechanism configured to surround the turbine impeller and to guide a fluid to the turbine impeller. The variable capacity mechanism has a nozzle ring that faces the second housing. A first pin and a second pin extend between the second housing and the nozzle ring and are attached to one of the second housing and the nozzle ring. The other of the second housing and the nozzle ring is provided with a first guide in which an end portion of the first pin is disposed and a second guide in which an end portion of the second pin is disposed.

A turbine engine with a turbine core that includes a compressor section having a compressor coupled to a high speed shaft, a combustion section, a turbine section having a high pressure turbine coupled to the high speed shaft and a low pressure turbine coupled to a low speed shaft, and a nozzle section. The turbine engine also includes an electric machine that can operate in a first starting mode and a second generating mode.

An assembly is provided for a turbine engine. This turbine engine assembly includes a turbine engine shaft, a gearbox and a coupler. The turbine engine shaft is configured to rotate about a rotational axis. The gearbox includes a gear configured to rotate about the rotational axis. The coupler is coupled to the turbine engine shaft by a coupler-shaft connection. The coupler-shaft connection is configured as or otherwise includes a crowned spline connection. The coupler is coupled to the gear by a coupler-gear connection. The coupler-gear connection is configured as or otherwise includes a crowned spline connection.

An assembly is provided for a turbine engine. This turbine engine assembly includes a rotating structure, a stationary structure and an electric machine. The rotating structure is configured to rotate about a rotational axis. The stationary structure circumscribes the rotating structure. The electric machine includes a rotor and a stator. The rotor circumscribes the rotating structure and is coupled to the rotating structure through a spline connection. The stator is connected to the stationary structure.

A variable geometry inlet system of an aircraft engine includes an inlet duct. The inlet duct includes at least first and second sections moveable between extended and retracted positions such that the inlet duct defines a variable axial length of an inlet passage for selective flight conditions. The inclusion of acoustic treatment may assist in controlling noise.

An example of a concentric probe includes an outer shroud having a bore that extends through the outer shroud, an inner shroud located within the outer shroud and having a bore that extends through the inner shroud, the inner shroud joined to the outer shroud via brazing, an annulus defined by a space between the inner shroud and a wall of the bore of the outer shroud, a plenum defined by a space between the inner shroud and the wall of the bore of the outer shroud, the plenum being in fluid communication with the annulus, and a transducer disposed within inner shroud.