A method of forming a protective sheath for an aerofoil component includes: providing a first sheath portion and a second sheath portion, the first sheath portion and the second sheath portion each comprising an inner surface, an outer surface and an end surface between the inner and outer surfaces and having a sacrificial flange at its distal end; positioning the first sheath portion and second sheath portion so that the inner surface of the first sheath portion abuts against the inner surface of the second sheath portion with the end surfaces of the first and second sheath portions aligned to form a mating edge; and joining the first sheath portion to the second sheath portion by welding along the mating edge, wherein the sacrificial flanges are completely consumed and a curved outer profile is formed.

A seal system is provided. The seal system may comprise a first duct having an annular geometry, a second duct overlapping the first duct in a radial direction, and a seal disposed between the first duct and the second duct. The seal may comprise a groove defined by the first duct and a piston configured to slideably engage the groove.

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.

In accordance with an embodiment of the disclosure, a system includes a combustor liner disposed about a combustion chamber of a combustor of a gas turbine system. The combustor liner includes an inner wall portion exposed to the combustion chamber, an outer wall portion disposed about the inner wall portion, and multiple channels between the inner and outer wall portions of the combustor liner. Each channel of the multiple channels is configured to direct a coolant along the combustor liner to convectively cool the combustor liner, and each channel of the multiple channels includes a cross-sectional area that progressively changes along a length of each channel of the plurality of channels.

A gas turbine engine includes an engine static structure. A rotating structure is configured to rotate relative to the engine static structure. The rotating structure has a target area with first and second directions. The first direction is greater than the second direction. A lubrication system includes a nozzle having a non-circular exit aimed at the target area. The exit provides a width and a height. The width is greater than the height. The width is oriented in the first direction.

A gas turbine engine component comprises an airfoil with a suction side and pressure side extending from a leading edge to a trailing edge. There are a plurality of cooling holes adjacent the leading edge, with the cooling holes having a non-circular shape, with a longer dimension and a smaller dimension. The airfoil defines a radial direction from a radially outer end to a radially inner end, and radially outer of the cooling holes spaced toward the radially outer end, which have the longer dimension extending closer to parallel to the radial direction. Radially inner cooling holes closer to the radially inner end having the longer dimension extend to be closer to perpendicular relative to the radial direction compared to the radially outer cooling holes.

An assembly comprises a first cooling cavity disposed within one or more of a turbine assembly or a combustion chamber of an engine. The first cooling cavity directs cooling air within the one or more of the turbine assembly or the combustion chamber. The assembly comprises a second cooling cavity also disposed within the one or more of the turbine assembly or the combustion chamber. The second cooling cavity receives at least some of the cooling air from the first cooling cavity. A forward facing step nozzle forms a channel that fluidly couples the first cooling cavity with the second cooling cavity. The step nozzle includes steps having elongated first sides and narrow second sides. The elongated first sides of the steps protrude into the channel such that a cross-sectional area of the channel of the step nozzle at the steps is smaller than a cross-sectional area of the channel of the step nozzle outside of the steps.

A steam turbine may include a casing defining an interior cavity, wherein the casing comprises a first casing half and a second casing half, which are connected to one another to form the casing, wherein an interior surface of the first casing half has a first portion with a first curvature and a second portion with a second curvature, wherein the first curvature and the second curvature are different, and wherein an interior surface of the second casing half has a first portion with a first curvature and a second portion with a second curvature, wherein the first curvature and the second curvature are different. An interior surface of the casing may have a substantially oval-shaped cross-section. Internal pressure acting on the oval interior surface allows better sealing at the split line.

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.

A rotor disk for a gas turbine engine is disclosed. In various embodiments, the rotor disk includes a rim portion disposed about a central axis; a blade post disposed proximate the rim portion, the blade post having a first branch and a second branch; and a first scallop disposed within the rim portion, between and radially inward of the first branch and the second branch.