A gas turbine engine is disclosed having a tip clearance control system. The tip clearance control system has a cabin blower system, a casing arranged in use about a rotor of a gas turbine engine and a fluid delivery passage. The cabin blower system having a cabin blower compressor arranged in use to compress fluid used in a cabin of an aircraft and to compress fluid conducted via the fluid delivery passage into heat exchange with the casing.

There is provided a blade tip for a rotary blade. The blade tip is formed of a metal foam and comprises at least one vortex generator. The vortex generator may comprise at least one passageway and/or cavity in the blade tip. In use, a vortex is created between the blade tip and a fan casing adjacent the blade tip.

A dust mitigation system for airfoils includes a plurality of contoured tip turns which curve about at least two axes. This inhibits recirculation areas common within airfoils and further inhibits dust build up within the cooling flow path of the airfoil.

Disclosed is a process for producing a run-in coating (20, 24, 32, 44) on a component of a turbomachine, in particular of a gas turbine. The run-in coating is applied and produced on the component of the turbomachine by a kinetic cold gas compacting process (K3). The invention also encompasses a run-in coating for a static or rotating component of a turbomachine and a static or rotating component of a turbomachine, in particular of a gas turbine, having at least one run-in coating.

An assembly for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, an airfoil including a radial end, a first passageway having an outlet at the radial end, and a second passageway having an inlet at the radial end. The assembly further includes a cover having at least one turning cavity configured to direct fluid expelled from the outlet of the first passageway into the inlet of the second passageway.

A film-cooled gas turbine component for a gas turbine has a surface exposed to a hot gas and a number of film-cooling openings open out, which film-cooling openings combined to form at least one row transverse to a flow direction of the hot gas. Each of the film-cooling openings has a duct section and a diffuser section having an upstream diffuser edge, two diffuser longitudinal edges and a downstream diffuser edge. At least two immediately adjacent film-cooling openings, of the respective row have their duct axes of the respective duct sections laterally inclined relative to the local flow direction of the hot gas and their diffuser sections are formed asymmetrically with respect to a projection of the duct axis, such that immediately adjacent corner regions of the respective film-cooling openings are in alignment without the respective diffuser sections making contact with one another.

A gas turbine engine blade includes a blade portion having a leading edge and a trailing edge. A first surface connects the leading edge to the trailing edge and a second surface connects the leading edge to the trailing edge. A tip section is located at a first end of the blade portion and includes a pocket protruding into the tip section from an outermost end of the tip section. The pocket has a first side wall adjacent the first surface and a second side wall adjacent the second surface. At least one of the first side wall and the second side wall have a curve distinct from a curve of the corresponding adjacent surface.

A tip shroud, comprising a plurality of tip shoes encircling a rotor assembly, in a turbine may deform due to thermal gradients experienced during operation of the turbine. This can make it difficult to remove the tip shroud during disassembly of the turbine. In an embodiment, to facilitate consistent and reliable removal of the tip shroud during each disassembly of the turbine, one or more, including potentially all, of the tip shoes of a tip shroud may be provided with one or more radially protruding puller hooks. Each puller hook enables an axial force to be transferred by one or more tools to an axially inner surface of the puller hook, to thereby produce axial movement of the tip shoe out of the tip shroud.

A turbine bucket includes a leading edge, a trailing edge, a root portion, and a tip portion. The turbine bucket also includes one or more cooling passages extending through a body of the turbine bucket from an inlet to an outlet. The cooling passages are configured to route a cooling flow of fluid through the turbine bucket. The turbine bucket further includes a plenum defined within the tip portion to receive the fluid from the outlet of the cooling passages for expulsion of the cooling flow of fluid into a main flow path via at least one outlet hole proximate the trailing edge of the turbine bucket.

The present disclosure is directed to a rotor blade for a gas turbine engine. The rotor blade includes an airfoil, a tip shroud having a side surface and a radially outer surface, and a transition portion coupling the tip shroud to the airfoil. The airfoil, the transition portion, and the tip shroud collectively define a primary cooling passage therein. The primary cooling passage includes a primary cooling passage outlet defined by the side surface of the tip shroud.