A blade for a turbine blade includes a suction-side side wall and a pressure-side side wall that enclose a cavity at least partially in a manner which extends along a profile centre line from a common front edge to a common rear edge and in a span width direction from a root-side end to a tip-side end. A first perforated impingement cooling wall which is provided with openings for the impingement cooling of the front edge and at least one further perforated impingement cooling wall for the impingement cooling of a section of the suction-side and/or pressure-side side wall are provided in the interior along the span width. The impingement cooling openings of the first impingement cooling wall and the at least one second impingement cooling wall are connected in series in terms of flow.

A turbine vane includes a root carrying a blade terminated by a squealer tip, the blade having intrados and extrados walls, a leading edge, a trailing edge, and a tip wall delimiting a bottom of the squealer tip, by which the intrados wall is connected to the extrados wall. The blade also includes: a serpentine median circuit, including a first radial pipe that collects air at the root and is connected by a first bend to a second radial pipe that is connected by a second bend to a third radial pipe; a cavity under the squealer tip running along the extrados wall and extending from a central region of the squealer tip to the trailing edge; and a central radial pipe collecting air at the root and extending between at least two of the three pipes of the median circuit and directly supplying the cavity under the squealer tip.

A turbomachine component includes a platform, a shank, and an airfoil. The platform includes a pressure side slash face and a suction side slash face. The shank extends radially inward from the platform. The airfoil extends radially outward from the platform. The airfoil includes a leading edge and a trailing edge. A cooling circuit is defined within the shank and the airfoil. The cooling circuit further includes a plurality of exit channels disposed along the trailing edge of the airfoil. The cooling circuit further includes at least one bypass conduit that extends from an inlet disposed in the cooling circuit to an outlet positioned on the pressure side slash face. The at least one bypass conduit being positioned radially inward of the plurality of exit channels.

A device for fastening sealing plates between components of a gas turbine engine includes guide vane ring which includes guide vane segments, wherein each guide vane segment includes an outer platform and an inner platform, sealed off from one another at ends by a sealing strip. The device furthermore includes a plurality of sealing plates which seal off the guide vane segments from a component which is adjacent in the upstream or downstream direction. The sealing strips in each case form a sealing section and an extension section, wherein the sealing section serves to seal off two mutually adjoining platforms, the extension section extends axially forward or axially rearward, starting from the sealing section, and projects from the platforms, and the extension section forms a holding element for at least one sealing plate or is connected to a separate holding element.

A chevron ring is disclosed. In various embodiments, the chevron ring includes an attachment ring; and a chevron, the chevron connected to and extending downstream of the attachment ring and having a trailing edge portion, the chevron defining a hollow portion between the attachment ring and the trailing edge portion.

A turbine engine of an aircraft includes: a primary air flow duct; a secondary air flow duct which is located around the primary duct, the secondary duct including a stator including a plurality of blades distributed around a main axis of the turbine engine and inter-blade platforms located between radially internal ends or between radially external ends of two adjacent blades, each platform including a wall partially delimiting the secondary duct; and a fluid circuit which includes a heat exchanger formed by at least one of the platforms. The platform includes a line that has an inlet port of the fluid and a fluid outlet port. The fluid circuit includes a distributor associated with each port of the at least one platform with the rest of the fluid circuit and of which each distributor is axially offset with respect to the platform along the main axis.

A turbine rotor blade includes a blade body including a pressure-side blade wall and a suction-side blade wall. The blade body includes: a serpentine passage composed of a cooling passage separated into a plurality of parts by a partition wall that connects the pressure-side blade wall and the suction-side blade wall and extends along a height direction of the blade body; and a first cooling hole communicating at one end with the cooling passage via a first inlet opening formed in an inner wall surface of the pressure-side blade wall or the suction-side blade wall, and communicating at another end with a first outlet opening formed in an outer wall surface of the pressure-side blade wall or the suction-side blade wall of the blade body. The first cooling hole extends in a leading edge direction from the first inlet opening to the first outlet opening.

A nozzle segment for a gas turbine engine comprises an outer band having a cooling air inlet, an inner band having a first cooling air outlet, and a nozzle airfoil comprising a cooling flow passage arranged to receive the cooling air as a cooling air stream. A first channel and a second channel are arranged within the cooling flow passage. A deflector is arranged to divide the cooling air stream into a first cooling air stream in the first channel and a second cooling air stream in the second channel, respectively. The deflector deflects the first cooling air stream obliquely to a suction sidewall in the first channel, wherein the first channel is configured to transport the first cooling air stream along the first channel in a swirly flow.

A turbine blade includes an airfoil portion, a cooling passage inside the airfoil portion, and a plurality of cooling holes formed in a trailing edge part of the airfoil portion. The cooling holes communicating with the cooling passage and opening in a surface of the trailing edge part. A relation of d_up<d_mid<d_down is satisfied, where d_mid is an index indicating opening densities of the cooling holes in a center region including an intermediate position between a first end and a second end of the airfoil portion in the blade height direction, d_up is an index in a region positioned upstream of a flow of a cooling medium in the cooling passage from the center region in the blade height direction, and d_down is an index in a region positioned downstream of the flow of the cooling medium from the center region in the blade height direction.

An article, such as a turbine blade, includes an airfoil. The airfoil includes a body, the body having an elongated internal cavity extending from a tip of the body. The cavity is defined an internal wall within the body. At least one elongated damping element is disposed in the elongated internal cavity and frictionally engages the internal wall. Thus, the least one elongated damping element is capable of damping vibrations in the article.