Abstract: A module for a turbomachine, including a guide blade arrangement, a seal carrier that is situated radially within an inner platform of the guide blade arrangement, seal carrier walls, namely, a first seal carrier wall and a second seal carrier wall, and a sliding body as well as a connecting element. The seal carrier walls have a multipart design with respect to one another, and the second seal carrier wall is fastened to the first seal carrier wall and to the seal carrier via the connecting element. The sliding body holds the seal carrier walls at a distance from one another in such a way that they mutually axially delimit a clearance in which the guide blade arrangement engages via a radially inwardly extending guide pin. A radially inner section of the second seal carrier wall has frictional contact relative to the seal carrier.

Abstract: A rotor blade (20) for a rotor blade assembly (10) of a turbomachine (1) is provided, having an inner rotor blade platform (40) which extends axially from the rotor blade (20) with respect to a longitudinal turbomachine axis (2) and has two opposite circumferential end faces (41) and a free axial end (42) whose cross section is radially inwardly and radially outwardly bounded by circular arcs of two concentric circles. The inner rotor blade platform (40) has a cross section of connection (45) with the rotor blade (20) which is bounded radially inwardly by an inner connecting line (46) and radially outwardly by an outer connecting line (47). Each of the connecting lines (46, 47) has a central portion (61) having a convex curvature.

Abstract: A rotor blade for a gas turbine, in particular an aircraft gas turbine, including a blade root, a blade neck that adjoins the blade root in the radial direction, an airfoil that adjoins the blade neck in the radial direction, a radially outer partition wall that forms a radially inner delimiting section of an annular space of a gas turbine, an axially front partition wall and an axially rear partition wall that are connected to the radially outer partition wall so that the partition walls surround the blade neck on three sides, the partition walls protruding beyond the blade neck in the circumferential direction. For placement in a blade root receptacle of a rotor disk, the rotor blade is provided with a blade root protective plate that is situated between the blade root and the rotor disk.

Abstract: The invention relates to a method and a set of balance weights for balancing a rotor particularly of a turbomachine, on which a plurality of balance weights, which are distributed over the periphery, can be arranged for balancing an imbalance. A set of balance weights, in particular plate-shaped weights, with a predetermined number of groups of balance weights, is used for balancing the rotor. Each of the balance weights of every group of balance weights has an identical weight that is assigned to this group and a shape assigned to this group, which differs in basic shape and/or thickness of the balance weights from the balance weights of the other groups.

Abstract: A jacket ring assembly for a turbomachine, the jacket ring assembly including a casing part, a jacket ring segment which is adapted to radially outwardly surround a rotor blade ring and to this end is disposed radially inwardly of the casing part, as considered with respect to a longitudinal axis of the turbomachine, and a segmented ring which is circumferentially divided into segments and by which the jacket ring segment is mounted to the casing part, the segmented ring being axially form-fittingly disposed on a form-fitting element of the casing part, for which purpose each of the respective segments of the segmented ring is radially outwardly assembled with the form-fitting element, and the segmented ring forming a supporting seat on which the jacket ring segment is seated and radially inwardly supported with an axially forward end.

Abstract: Disclosed is a static seal arrangement for sealing a gap between a first static component and a second static component of a turbomachine such as an aircraft engine or a static gas turbine, having a brush seal held distant from a hot gas, this brush seal being arranged standing or vertical with its clamping section that can be surrounded by cooling air in a radially outer clamping space relative to the gap, and forming a radial sealing site with the first component and an axial sealing site with the second component; also disclosed is a turbomachine.

Abstract: The invention relates to a turbine blade which comprises at least one cavity that is defined by a wall with one or more surface discontinuities which preferably are selected from elevations, depressions and undercuts and preferably change at least one eigenmode of the blade. The blade may suitably be produced by a generative production method such as selective laser melting (SLM).

Abstract: The present invention relates to a guide vane arrangement for a turbomachine, comprising a guide vane segment and a housing part that are fastened to one another, for which a back guide vane hook that rises radially toward the outside from an outer shroud of the guide vane segment, when referred to a longitudinal axis of the turbomachine, and a housing hook, which is arranged radially inside circumferentially at the housing part, engage in one another in form-fitting manner, wherein, in a first peripheral segment of the guide vane arrangement the guide vane hook has a front wall and a back wall, and therewith forms a groove open radially toward the outside, in which a ring section of the housing hook is arranged and held axially.

Abstract: A segmented ring for installation in a turbomachine, circumferentially divided into segments, considered with respect to a ring axis of the segmented ring, wherein the segmented ring is adapted for installation from radially inside; i.e., the segments can be assembled radially outwardly to form the segmented ring. At least two immediately circumferentially adjacent segments meet in a joint at which a sealing insert is provided. A pocket which is open toward the joint is formed in each of the at least two immediately circumferentially adjacent segments; i.e., two circumferentially mutually facing pockets are provided at the joint. The sealing insert is disposed in the two mutually facing pockets and axially retained therein and extends circumferentially across the joint, and a pocket of the two mutually facing pockets is additionally also open in a radial direction such that this segment can be radially slid onto the sealing insert.

Abstract: The present invention relates to a housing structure for a turbomachine, a turbomachine, and a method for cooling a housing portion of a housing structure of a turbomachine, wherein the housing structure has an outer housing wall, which is formed by at least one housing part, and an inner wall, wherein the inner wall is arranged in the radial direction inside of the housing wall and spaced apart from the housing wall and is designed to bound the main flow channel, at least partially, wherein the housing structure comprises a cooling air channel for cooling a housing portion to be cooled, and has an upstream end and a downstream end, wherein the cooling air channel extends, at least partially, between the housing wall and the inner wall.

Abstract: The invention relates to a method and a set of balance weights for balancing a rotor, particularly of a turbomachine, on which a plurality of balance weights, which are distributed over the periphery, can be arranged for balancing an imbalance. A set of balance weights, in particular plate-shaped weights, with a predetermined number of groups of balance weights, is used for balancing the rotor. Each of the balance weights of every group of balance weights has an identical weight that is assigned to this group and a shape assigned to this group, which differs in basic shape and/or thickness of the balance weights from the balance weights of the other groups.

Abstract: A rotor blade for disposition in a hot gas duct of a turbomachine, the rotor blade having an airfoil and an inner platform radially inwardly thereof, the inner platform having a chamber formed therein which is radially outwardly bounded by an inner shroud of the inner platform and is axially bounded between a forward chamber wall and a rearward chamber wall of the inner platform. A forward sealing plate is disposed at the forward chamber wall and extends axially forwardly away from the chamber, and a rearward sealing plate is disposed at the rearward chamber wall and extends axially rearwardly away from the chamber. Furthermore, an axial partition is disposed in the chamber, which axial partition divides the chamber into a radially inner chamber wall region and a radially outer chamber wall region and which, when viewed in an axial cross section, extends into the forward chamber wall at the level of the forward sealing plate and into the rearward chamber wall at the level of the rearward sealing plate.

Abstract: A sealing assembly for sealing a first sealing bank of a first static component and a second sealing bank of a second static component of an axial turbomachine, the static components axially movable relative to each other, the sealing assembly including the two static components and a sealing element segmented in the circumferential direction, a first end of the sealing element being axially coupled in a groove at the first sealing bank facing an opposite second end bearing axially against the second sealing bank, as well as a stator vane assembly including a casing component of an axial turbomachine, which has at least one casing hook, and further including a stator vane component having at least one stator vane disposed on a platform and at least one stator vane hook which projects from the platform and bears against the casing hook, the stator vane hook having formed therein at least one through-hole for supply of cooling air.

Abstract: The invention relates to a guide vane segment for an aircraft engine, having at least one guide vane element, having at least one flange formed in a radial extension direction (R) of the guide vane element and at least one positioning member projecting from the flange in the radial extension direction (R), and at least one seal carrier, which is arranged at the flange and is aligned relative to the guide vane element by way of the positioning member, wherein the flange has a supporting surface on which the seal carrier is supported, wherein a partial surface of the positioning member is adjacent to the flange at a transition from the flange to the positioning member, and respective surface normal lines of the supporting surface and of the partial surface enclose between them an angle (?) that is different from a zero angle.

Abstract: A component joint including a turbomachine component (10) having at least one component flange (11); a mating turbomachine component (20) connected thereto and having at least one mating component flange (21); and at least one seal (30) to provide a seal against an overpressure (p2?p1) on the component side, the at least one seal (30) having a sealing surface (31) that engages into a groove (12) of the turbomachine component (10), the groove (12) being disposed on a mating-component-flange-facing side of the component flange (11) and having a contact surface (13) for sealingly axially supporting the sealing surface (31).

Abstract: A fastening assembly having a component flange (10), in particular of a turbomachine, having at least one bore (11) in which a sleeve (30) is placed that extends through a bolt (20) which has a head (21) that is axially secured between a rim (12) of the bore and a component flange-mounted axial limit stop (13).

Abstract: Disclosed is a static seal arrangement for sealing a gap between a first static component and a second static component of a turbomachine such as an aircraft engine or a static gas turbine, having a brush seal held distant from a hot gas, this brush seal being arranged standing or vertical with its clamping section that can be surrounded by cooling air in a radially outer clamping space relative to the gap, and forming a radial sealing site with the first component and an axial sealing site with the second component; also disclosed is a turbomachine.

Abstract: An impulse body module a turbine stage of a gas turbine includes a holder component of one-piece construction, with a base and surrounding side walls arranged at the base, wherein the side walls and the base define a holding chamber; an insert component, of one-piece construction, which is inserted into the holding chamber of the holder component, wherein the holder component and the insert component accommodated therein define a number of mutually separated cavities, and wherein an impulse body, in particular a sphere, is accommodated in each cavity; and a closure component of one-piece construction, which is joined to the holder component in a material-bonded manner where the holding chamber is closed and the insert component is surrounded by the holder component and the closure component. The holder component has at least one base projection in the region of its base, which extends away from the holding chamber.

Abstract: A rotor blade for disposition in a hot gas duct of a turbomachine, the rotor blade having an airfoil and an inner platform radially inwardly thereof, the inner platform having a chamber formed therein which is radially outwardly bounded by an inner shroud of the inner platform and is axially bounded between a forward chamber wall and a rearward chamber wall of the inner platform. A forward sealing plate is disposed at the forward chamber wall and extends axially forwardly away from the chamber, and a rearward sealing plate is disposed at the rearward chamber wall and extends axially rearwardly away from the chamber. Furthermore, an axial partition is disposed in the chamber, which axial partition divides the chamber into a radially inner chamber wall region and a radially outer chamber wall region and which, when viewed in an axial cross section, extends into the forward chamber wall at the level of the forward sealing plate and into the rearward chamber wall at the level of the rearward sealing plate.

Abstract: A sealing assembly for sealing a first sealing bank of a first static component and a second sealing bank of a second static component of an axial turbomachine, the static components axially movable relative to each other, the sealing assembly including the two static components and a sealing element segmented in the circumferential direction, a first end of the sealing element being axially coupled in a groove at the first sealing bank facing an opposite second end bearing axially against the second sealing bank, as well as a stator vane assembly including a casing component of an axial turbomachine, which has at least one casing hook, and further including a stator vane component having at least one stator vane disposed on a platform and at least one stator vane hook which projects from the platform and bears against the casing hook, the stator vane hook having formed therein at least one through-hole for supply of cooling air.