An electric motor assembly includes an electric motor, a fan assembly coupled to the electric motor and configured to rotate therewith about an axis. The fan assembly includes a hub including a cylindrical portion having an inlet end and an outlet end, the hub further including an inlet surface coupled to the inlet end. The fan assembly also includes a plurality of blades coupled to an outer periphery of the cylindrical portion, wherein the inlet surface is tapered to direct an inlet airflow toward the plurality of blades.

A gearbox assembly for a gas turbine engine includes a hollow gear. The hollow gear has one or more reservoirs configured to store a lubricant and one or more passages configured to supply the lubricant to one or more gear meshes in the gearbox assembly. The hollow gear provides a dampening system configured to absorb vibrations and loading experienced in the gearbox assembly and configured to provide impact resistance for the gearbox assembly.

A dual motor system includes a first motor providing a lower speed range and a second motor providing a higher speed range, wherein the motors are coaxially arranged and aligned on and drive a common shaft, and a motor control system controlling the speed of the first motor and engaging the second motor as needed. The first motor is a variable speed motor providing a lower two-thirds of a full speed range, and the second motor is an induction motor providing the upper one-third in the form of one or more discrete fixed speeds. The system may include a transformer including a first winding tap which provides a first higher speed, and a second winding tap which provides a second higher speed. The system may also include a flow control system for automatically controlling the speed of the motors for particular applications, such as flow control in a pool.

An airfoil for a turbine engine includes an array of pins positioned in an internal cavity of the airfoil, such that cooling channels are defined in the interspaces between adjacent pins. Each pin extends lengthwise from a first airfoil wall to a second airfoil wall and is connected thereto at a first intersection and at a second intersection respectively. The pin has a first cross-sectional shape at a respective intersection and a second cross-sectional shape at an intermediate plane located between the first and second intersections. The first cross-sectional shape includes a closed shape defined by relatively sharp corners and the second cross-sectional shape includes a closed shape defined by relatively rounded corners. A cross-sectional area of the pin at the intermediate plane is greater than a cross-sectional area of the pin at the respective intersection.

A cooling structure for a trailing edge of a turbine blade is provided. The cooling structure for the trailing edge of the turbine blade comprising an airfoil shaped blade part including a leading edge, a trailing edge, a pressure surface and a suction surface connecting the leading edge and the trailing edge, and a cavity channel formed in the blade part and through which a cooling fluid flows, the cooling structure including slots and lands arranged alternately on the trailing edge along a span direction of the pressure surface by cutting a portion of the pressure surface, the slots communicating with the cavity channel and defined by adjacent lands where the pressure surface remains, wherein a pin-fin structure is disposed in the cavity channel on an upstream side of the slot, and wherein the cooling fluid is introduced through a micro-channel formed inside the pin-fin structure and is discharged through film cooling holes formed in the pressure surface.

An air-sealing device intended to be inserted between an aircraft dual-flow turbine engine casing element and a nacelle element, the sealing device including an attachment tab at the end of which is located a sealing portion having an outer surface intended to be contacted by the casing element and the nacelle element, and an inner surface defining a cavity. The inner surface defines at least one protuberance extending inside the cavity.

Blades including integrated damping structures are disclosed herein. An airfoil to be disposed within a flow path of a gas turbine engine, the gas turbine engine defining an axial axis, a radial axis and a circumferential axis comprising an airfoil body including a first face, a second face and a recessed portion formed in the second face, and an airfoil cap having a first surface, the airfoil cap disposed within the recessed portion, the first surface substantially flush with the second face.

A cooling structure for a trailing edge of a turbine blade is provided. The cooling structure for the trailing edge of the turbine blade comprising an airfoil shaped blade part including a leading edge, a trailing edge, a pressure surface and a suction surface connecting the leading edge and the trailing edge, and a cavity channel formed in the blade part and through which a cooling fluid flows, the cooling structure including slots and lands arranged alternately on the trailing edge along a span direction of the pressure surface by cutting a portion of the pressure surface, the slots communicating with the cavity channel and defined by adjacent lands where the pressure surface remains, wherein a pin-fin structure is disposed in the cavity channel on an upstream side of the slot, and wherein the cooling fluid is introduced through a micro-channel formed inside the pin-fin structure and is discharged through film cooling holes formed in the pressure surface.

A turbine comprises a casing, an outer metal shroud, an inner metal shroud and an annular distributor having a plurality of CMC ring sectors, each sector comprising a mast, an inner platform, an outer platform and at least one blade having a hollow profile that defines an inner housing, the inner and outer platforms each having an opening communicating with said inner housing, and the mast passing through said openings and the inner housing and being secured to said casing and connected to said annular sector. Each blade comprises at least one first radial shoulder projecting axially towards the inside of the blade, and each mast comprises at least one second shoulder projecting axially towards the outside of the mast configured to radially cooperate with a first shoulder and radially press the blade against the mast.

The present application provides a stator blade, a compressor structure and a compressor. The stator blade comprises a blade body, wherein a cavity is formed inside the blade body, and a gas supply hole is formed on the blade body. The present application forms a jet on the suction surface of the stator blade by supplemented gas, thereby blowing off the low-speed low-energy region formed by the suction surface, reducing the gas flow mixing loss caused by the supplemented gas, thereby improving the aerodynamic efficiency of the centrifugal compressor.