A bladed disc system for a turbine engine having a disk portion and a plurality of blade portions which are associated with a stator section and an intercavity sealing portion, disc portion shaped such that blade portions are able to fit within firtree slot in disc portion, blade portion having aerofoil section and root section, aerofoil section having portion shaped such that they extend proximate to intercavity sealing portion, disc portion extending from portion that connects with drum to outer edge at which blade portions are connected with disc portion having width transition region in which thickness of disc increases from point at which disc connects to drum to outer edge at which it holds blade portions, and wherein width transition region has curved width transition region with radius r, and an overhanging portion which extends into the intercavity spacing between the width transition region and the intercavity sealing portion.

A bladed disc system for a turbine engine, the bladed disk portion comprising a disk portion and a plurality of blade portions, the disc portion being shaped such that the blade portions are able to fit within fir tree slot in the disc portion, the blade portion comprising an aerofoil section and a root section, with the root section comprising a fir tree profile and a skirt portion and wherein the skirt portions of adjacent blades form an opening that has a maximum separation of between 1-50% of the maximum skirt opening width.

A turbopump includes: a main shaft rotatably supported; a pump section including an impeller attached to one end of the main shaft; and a turbine section including: a disk attached to the other end of the main shaft, rotor blades provided on an outer periphery of the disk, and nozzles provided inclined to an entrance plane of a blade cascade constituted of the rotor blades, the nozzles having axisymmetric cross sections and arranged in at least two rows along a circumferential direction of the main shaft in a plane orthogonal to the main shaft.

Impingement insert for an airfoil of a blade/vane of a gas turbine is provided. The impingement insert includes a double-walled section having an outer and an inner walls, that define—an inner channel at an inner surface of the inner wall, an outer channel at an outer surface of the outer wall and a middle channel between the outer and the inner walls. Impingement cooling holes are provided in the outer wall that use the cooling air of the middle channel to eject impingement jets into the outer channel. The impingement insert includes at least one extraction duct that extends between the outer and the inner walls across the middle channel, and has an inlet at the outer channel, and an outlet at the inner channel, for flowing the cooling air, after impingement, from the outer channel into the inner channel.

A gas turbine engine includes a rotor rotatable about a central axis. The gas turbine engine includes a turbine stage including a stationary portion and a rotating portion made up of a number of rotating blades and a plurality of stationary vanes arranged to define the stationary portion. Each stationary vane includes an inner rail having an inlet face, a suction side face, a pressure side face, and a platform. A vane portion extends along a radial line from the platform and defines one of a first stagger angle and a second stagger angle with respect to the central axis. The platform has an elliptical cross-section in a plane that includes the central axis.

An air circulator which is attached to a sheet member and which is for generating an air flow from one side of the sheet member to another side of the sheet member includes a fan main body and a ring member for attaching the fan main body to the sheet member. The fan main body includes a hollow cylinder unit, a flange formed on the cylinder unit, a motor fixing unit, a motor fixated to the motor fixing unit, a wing attached to a rotating shaft of the motor and first engaging units formed on outer surfaces of first parts of the cylinder unit, the first parts forming a pair and facing each other. The ring member includes second engaging units which are formed on inner surfaces of second parts that form a pair, the second parts facing each other, and which engage with the first engaging units.

Various embodiments of an airfoil and machines with airfoils are disclosed. The airfoils include a thicker leading airfoil portion and a thinner trailing airfoil portion. In one embodiment, the leading airfoil portion is formed by bending a body of the airfoil back toward itself. In another embodiment, the leading airfoil portion has a solid geometry and includes two elliptic surfaces. To prevent detachment of airflow, the leading airfoil portion includes at least two arc portions or surfaces that act to direct the airflow down to the trailing airfoil portion in a manner that stabilizes vortexes that may form in the region of changing thickness.

A variable stiffness damper system including an inner spring positioned between a first wall and a second wall, in which the inner spring includes a first member and a second member each coupled together at a distal end by an inner bumper. The first member and the second member are each contoured toward one another. The first member, the second member, and the inner bumper form a cavity therebetween. An outer spring is positioned between the inner spring and the first wall or the second wall. The outer spring includes a spring arm contoured toward the inner spring. The outer spring includes an outer bumper positioned between the inner bumper and the first wall or the second wall. The inner bumper and the outer bumper are selectively couplable to one another based on a load applied to the damper system.

A manufacturing method is provided. During this method, a preform component is provided for a turbine engine. The preform component includes a substrate comprising electrically conductive material having an outer coating comprising non-electrically conductive material applied over a surface of the substrate. A preform aperture is formed in the preform component using an electrical discharge machining electrode. The preform aperture includes a meter section of a cooling aperture in the substrate. The preform aperture also includes a pilot hole in the outer coating. A diffuser section of the cooling aperture is formed in at least the outer coating using a second machining process.

An impeller associated with a compressor of a gas turbine engine includes a plurality of impeller blades. Each impeller blade of the plurality of impeller blades has a leading edge and a trailing edge opposite the leading edge in a streamwise direction. Each impeller blade of the plurality of impeller blades extends in a spanwise direction from a hub at 0% span to a tip at 100% span, and each impeller blade of the plurality of impeller blades has a plurality of mean camber lines that each extend from the leading edge to the trailing edge at a respective spanwise location. The leading edge includes a partially swept leading edge surface defined between 70% span to 100% span that extends in the streamwise direction between 3% to 15% of a mean camber line at 100% span.