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 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: 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 present invention relates to an arrangement, comprising a first turbomachine element with a first turbomachine element through-opening, a second turbomachine element, a sliding element with a sliding element through-opening that is arranged, at least in part, in the first turbomachine element through-opening, a sleeve that engages through the sliding element through-opening along a through-axis, and a fastening member that engages through the sleeve for fastening the sleeve to the second turbomachine element. Furthermore, the invention relates to a method for fastening the first turbomachine element to the second turbomachine element, a turbomachine with the above-mentioned arrangement, and use of this arrangement.

Abstract: A liner element of a hot-gas-conveying duct of a turbine intermediate case of a gas turbine, in particular of an aircraft gas turbine, the liner element includes a central portion having at least one first reinforcement portion projecting in a direction away from the duct and extending substantially straight between an axial forward end and an axial rearward end; at least one of the two axial ends being adjoined by a second reinforcement portion projecting in a direction away from the duct and extending inclinedly or curvedly relative to the straight-line extent of the first reinforcement portion; the first reinforcement portion and the second reinforcement portion together forming a reinforcing element; the entire reinforcing element being disposed within the outer surface of the central portion, in particular in such a way that the reinforcing element is spaced apart from the first connecting portion and from the second connecting portion.

Abstract: The invention relates to a guide vane segment for a gas turbine comprising at least one radially outer shroud and one radially inner shroud, which extend along a respective arc and together form an annular segment, wherein, in the radial direction between the outer shroud and the inner shroud, a plurality of guide vanes are arranged adjacent to one another in the peripheral direction and are joined to the inner shroud and to the outer shroud in a material-bonded manner and, in particular, are joined in a one-piece manner, wherein, in an axial lengthwise direction, the outer shroud comprises an axially front end wall element and an axially rear end wall element in such a way that the outer shroud and the two end walls form a tub-like profile in lengthwise section.

Abstract: A turbomachine includes a seal device for sealing a gap between two components, with a plurality of sealing elements tippable in the direction of a relative displacement of the components. The sealing elements are mounted on one of the components and bridge the gap where the device has prestressing elements, by which the tippable sealing elements are prestressed against a respective sealing edge of the components. The prestressing elements are each supported at a front support point and a back support point of the component on the mounting side, these points being distanced from one another in the axial direction and in the radial direction of the turbomachine where the front support point is disposed radially outside and the back support point is disposed radially inside with respect to the mounting.

Abstract: A turbine blade (100) having an airfoil (1). In its radially outer region, the turbine blade (100) has at least one shroud (200) or at least one shroud segment, as well as at least one cutting tooth (300) which is intended to engage at least once into portions of a turbine casing during use of the turbine blade (100). The cutting tooth (300) is connected to the shroud (200) or the shroud segment.

Abstract: A sealing system (100) for a turbomachine is provided which includes a support (10) designed as a ring part, and a ring-like sealing plate (20, 30) that protrudes from the support in the axial direction (X). The support includes a fastening section (11) for fastening the sealing system to a stator part (200), and a retaining area (12), situated radially within the fastening section, for retaining a radial gap seal (40). The sealing plate (20, 30) is welded to the support (10) radially outside the retaining area (12). Also provided are a turbomachine that includes a sealing system (100), a support (10) for a sealing system, and a method for manufacturing a sealing system.

Abstract: A component system of a turbine engine including a first component segment and a second component segment configurable in a ring segment shape, so that at least one abutment surface of the first component segment and an abutment surface of the second component segment abut against each other; together, the first component segment and the second component segment including at least three overlapping elements for sealing a gap between the abutment surfaces. In the case of mutually abutting abutment surfaces, each overlapping element overlapping radially with the respective other component segment. At least two of the overlapping elements are configured on the first component segment, while at least one of the overlapping elements is configured on the second component segment. In the case of mutually abutting abutment surfaces, the overlapping element of the second component segment is axially configured between the overlapping elements of the first component segment.

Abstract: A turbofan aircraft engine having a primary duct (C), including a combustion chamber (BK), a first turbine (HT) disposed downstream of the combustion chamber, a compressor (HC) disposed upstream of the combustion chamber and coupled (W1) to the first turbine, and a second turbine (L) disposed downstream of the first turbine and coupled (W2) via a speed reduction mechanism (G) to a fan (F) for feeding a secondary duct (B) of the turbofan aircraft engine. A square of a ratio of a maximum blade diameter (DF) of the fan to a maximum blade diameter (DL) of the second turbine is at least 3.5, in particular at least 4.

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 preferably selected from elevations, depressions and undercuts. The blade may suitably be produced by a generative production method.

Abstract: The present invention relates to a blade or vane for a turbomachine, having at least one impulse element housing with a first impact cavity, in which an impulse element is arranged with play of movement, wherein the impulse element housing has at least one second impact cavity, which is in alignment with the first impact cavity in a first matrix direction and in which an impulse element is arranged with play of movement, and has at least one third impact cavity, which is in alignment with the first impact cavity in a second matrix direction crosswise to the first matrix direction and in which an impulse element is arranged with play of movement.

Abstract: A turbine ring of a turbomachine is described, having a one-piece ring structure which has an outer ring which is continuous in the circumferential direction of the turbine ring and has a fastening section for fastening the turbine ring on a housing section of the turbomachine, an inner ring which is segmented in the circumferential direction and has a plurality of inner ring segments which are mutually displaceable in the circumferential direction for guiding the hot gas and has front sealing segments and rear sealing segments, as seen in the axial direction of the turbine ring, these sealing segments being spaced a distance apart from one another in the axial direction and extending between the inner ring and the outer ring, as well as a turbomachine having such a turbine ring.

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 a turbomachine, including an airfoil (1) for flow deflection, a blade root (2) for attachment to a rotor of the turbomachine, an inner platform (3) between the airfoil and the blade root, and at least one pocket (4, 5) defined by two axially spaced walls (4.1, 4.2; 5.1, 5.2) extending from the side of the platform opposite the airfoil toward the blade root. A first (4.1, 5.1) of these walls has an outer side (4.1A, 5.1A) that faces away from the pocket and, in at least one cross section perpendicular to a radial longitudinal axis of the rotor blade, slopes outwardly.

Abstract: The present invention relates to a sealing element (100) of an axial turbomachine for the sealing of regions at or in at least two static components of the turbomachine, wherein the components can be moved axially relative to one another. The sealing element (100) has a first radial region (1) with a first stiffness or rigidity and a second radial region (5) with a second stiffness or rigidity, wherein the first stiffness and the second stiffness are different from one another, and the sealing element (100) is configured as segmented in the peripheral direction of the turbomachine.

Abstract: A liner element of a hot-gas-conveying duct of a turbine intermediate case of a gas turbine, in particular of an aircraft gas turbine, the liner element includes a central portion having at least one first reinforcement portion projecting in a direction away from the duct and extending substantially straight between an axial forward end and an axial rearward end; at least one of the two axial ends being adjoined by a second reinforcement portion projecting in a direction away from the duct and extending inclinedly or curvedly relative to the straight-line extent of the first reinforcement portion; the first reinforcement portion and the second reinforcement portion together forming a reinforcing element; the entire reinforcing element being disposed within the outer surface of the central portion, in particular in such a way that the reinforcing element is spaced apart from the first connecting portion and from the second connecting portion.

Abstract: A liner element (10), in particular a heat shield element or casing element, of a turbomachine, in particular a gas turbine, including a fastening portion (12) having at least one bore (14-1, 14-2) through which a connection unit (42) having, in particular a pin-type or bolt-type form, is introducible along a bore axis (BA); a liner portion (16) that adjoins the fastening portion (12); the fastening portion (12) and the liner portion (16) being formed in one piece, and having an annular or ring segment shape. The bore (14-1, 14-2) is provided in a depression (24-1, 24-2) formed in the fastening portion (12), the depression (24-1, 24-2) being essentially formed concentrically about the bore (14-1, 14-2).