An apparatus and method for passing fluid flow through at least a portion of a blade of turbomachinery, such as a gas turbine or the like. The fluid flow is directed through a plurality of flow channels which are interwoven with each other. Each flow channel is non-intersecting with any other flow channel and thus does not contact fluid flowing within any other flow channel. The method and apparatus can be used to reduce heat transfer and thus reduce thermal stresses, particularly in the blade.

A manifold for a gas turbine engine includes oppositely facing first and second walls that extend from an annular wall. An undulating wall oppositely faces the annular wall and couples the first wall to the second wall. The walls collectively form a plurality of first chambers in fluid communication with a plurality of second chambers so that each of the second chambers intersperse adjacent first chambers. Each first chamber has a first volume that is greater that a second volume of each second chamber.

A seal of a turbomachine reduces a leakage flow between a first and second component of the turbomachine. The first component has a first surface and the second component has a second surface, wherein the first component is stiff with regard to a first force exerted perpendicularly thereto and the second component is stiff with regard to a second force exerted perpendicularly thereto. The first surface is opposite the second surface, together defining boundaries of a fluid passage for the leakage flow. The first surface has a first surface riffle. A turbomachine has a seal described above, wherein the turbomachine is a gas turbine engine. A method of manufacturing a first component of a turbomachine with a reduced leakage flow between the first component and a second component of the turbomachine includes fabrication of a first surface riffle, in particular by grinding and/or by electrical discharge machining.

A turbomachine stage includes guide vanes and an airfoil platform forming a guide vane cascade, and rotor blades and an airfoil platform forming a rotor blade cascade. Airfoil platforms have cascade regions extending between circumferentially adjacent airfoils, and gap regions which radially and/or axially bound an axial gap extending axially between the guide vane cascade and the rotor blade cascade. A contour of at least one of these gap regions varies in the radial and/or axial direction around the circumference. A maximum extent of this contour in the radial direction toward the spoke-like pattern is circumferentially spaced from an airfoil edge of this cascade by no more than 50% of the cascade pitch, a maximum variation in the radial direction being no more than 50% of the cascade pitch and/or a maximum extent in the axial direction away from the spoke-like pattern is circumferentially spaced from an airfoil edge of this cascade by no more than 50% of the cascade pitch, a maximum variation in the axial direction being no more than 50% of the cascade pitch.

A case for a gas turbine engine incudes a radial flange with a partial scallop along an inner diameter of the radial flange. A case assembly for a gas turbine engine incudes a first case with a first radial flange with a partial scallop along an inner diameter of the first radial flange, the partial scallop adjacent to a first aperture through the first radial flange and a second case with a second radial flange with a second aperture through the second radial flange, the second radial flange mountable to the first radial flange at an interface such that the second aperture is axially aligned with the first aperture and a seal lip that extends from the second case interfaces with said first case at a longitudinal interface.

A turbomachine having a wall that extends circumferentially to a central axis of the turbomachine at least in certain areas, and having at least one insert device that can be connected in a releasable manner to a wall in the area of a recess which is arranged inside the wall. The insert device has a connection device by means of which the insert device can be brought into operative connection with a connection device of the wall. The connection device of the insert device has a projection that extends in the radial direction of the turbomachine at least in certain areas in the mounted state of the insert device, and that in the mounted state of the insert device meshes in a groove of the connection device of the wall. The connection device of the wall has a projection that extends in the radial direction of the turbomachine at least in certain areas in the mounted state of the insert device, and that, in the mounted state of the insert device, meshes in a groove of the connection device of the insert device.

An assembly for a turbocharger can include a pinched nozzle cartridge that includes a base component, an insert component and vanes disposed at least in part between the base component and the insert component, where the insert component includes a shroud surface and an upper pinch profile, where the base component includes an annular surface that defines a turbine wheel opening and a lower profile and where the upper pinch profile and the lower profile form a pinched nozzle. Various other examples of devices, assemblies, systems, methods, etc., are also disclosed.

An airfoil is provided. The airfoil comprises a crossover and an impingement cavity in fluid communication with the crossover and having an internal surface. At least a portion of the internal surface comprises an undulating internal surface. A plurality of trip strips may be disposed on the at least a portion of the internal surface to define the undulating internal surface. A gas turbine engine and an internally-cooled engine part are also provided.

A gas turbine engine includes a planetary gear system including a sun gear, intermediate gears, and a ring gear. A lubricant recovery system for the planetary gear system includes fluid passages that extend through the planetary gear system. A gutter is located radially outward from the planetary gear system for collecting lubricant. At least a portion of the gutter is rigidly attached to the ring gear.

The gas turbine engine (S1) includes: turbine blades (7b); and a cooling air supply unit (11) to supply cooling air to the turbine blades (7b). A flow path surface (31) is formed so as to be positioned in an upstream side of the turbine blades (7b) and so as to be connected to a base surface (32) in which the turbine blades (7b) are provided. The flow path surface (31) includes: depression portions (31a) depressed relative to the base surface (32), each depression portion (31a) including at least an area overlapping with a front end (7b1) of each turbine blade (7b), when viewed from a direction of the turbine axis (L); and protrusion portions (31b) protruding relative to the base surface (32), each protrusion portion (31b) being at least part of each area positioned between front ends (7b1) of the turbine blades (7b), when viewed from the above direction.