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).

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: A turbomachine in the form of a stationary gas turbine or an aircraft engine, respectively a housing structure therefor; the housing structure including an outer housing wall (1) and an inner wall (2) defining the flow channel; and a hollow space (4) being formed between the inner wall and the outer housing wall. The hollow space is separable into at least two regions (5, 6); a movable wire element (slide ring seal) (10, 10?), which is adapted to rest against the contact faces (8, 9), being configured in the hollow space for purposes of the separation.

Abstract: The present invention relates to a balancing body (1) for fastening to a ring (4, 5) of a continuous blade arrangement of a compressor or turbine stage of a gas turbine, wherein the balancing body has a first stop (11) for the form-fitting attachment of the balancing body in one peripheral direction (R) to a first axial shoulder (6; 7) of the ring.

Abstract: The present invention relates to a gas turbine casing arrangement having a gas turbine casing element (10), a guide-vane ring with an outer ring (20), and a coated ring (30), which lies radially opposite a rotor grid (40) adjacent to the guide-vane ring, whereby an intermediate ring (50) by a downstream front face (51) engages behind an upstream stop (61; 61?) fixed in place on the casing element, and by an upstream front face (52) of a radial flange (53) of the intermediate ring engages behind a downstream stop (22) fixed in place on the outer ring, in order to secure the guide-vane ring axially at the gas turbine casing element in the direction of through-flow.

Abstract: The invention relates to a sealing arrangement 15, a guide vane arrangement, and a turbomachine 11 with such a sealing arrangement 15, wherein the sealing arrangement 15 is designed for a guide vane ring 60 of a turbomachine 11, wherein the sealing arrangement 15 comprises a thin-walled annular structure 80 that is substantially closed on all sides, and wherein the annular structure 80 delimits an annular interior space 105, wherein a hollow cell structure 109, which is designed so as to mechanically support the annular structure 80, is provided in the annular interior space 105.

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 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: 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: The invention relates to a guide vane segment (10) for a gas turbine comprising at least one radially outer shroud (14) and one radially inner shroud (12), which extend along a respective arc and together form an annular segment, wherein, in the radial direction (RR) between the outer shroud (14) and the inner shroud (12), a plurality of guide vanes (16) are arranged adjacent to one another in the peripheral direction (UR) and are joined to the inner shroud (12) and to the outer shroud (14) in a material-bonded manner and, in particular, are joined in a one-piece manner, wherein, in an axial lengthwise direction (AR), the outer shroud (14) comprises an axially front end wall element (22) and an axially rear end wall element (24) in such a way that the outer shroud (14) and the two end walls (22, 24) form a tub-like profile in lengthwise section.

Abstract: A stator vane segment (100) of a fluid flow machine, having an upstream casing-side attachment (1) and/or a downstream casing-side attachment (3), and a platform (5). The attachments are provided on the platform (5). The platform (5) has a first sealing section (17) and extends in the direction of flow beyond a connection region (15) that connects the rearward attachment (3) to the platform (5). The first sealing section (17) that is located in a section (18) downstream of the connection region (15) has a first slot (19) which is continuous in the circumferential direction. The present invention also relates to a turbine having at least one stator vane segment (100).

Abstract: A turbomachine having a flow channel (6) and a housing (1, 2) which radially surrounds the flow channel, at least one sealing or lining element (11) being situated between the flow channel and the housing, at least one clamping ring (14) being provided, which is situated circumferentially around the flow channel and which is situated in such a way that it abuts at least one contact surface (22) of the at least one sealing or lining element (11).

Abstract: A sealing device (1) is disclosed for sealing a radially inner gas channel (2) between a guide vane ring (4) and a rotor (6) of a turbomachine, wherein the sealing device (1) has a sealing ring (8) for forming a sealed space (10) with a rear segment with an inner wall structure (30) oriented in the opposite direction, which are joined to each other via an annular arch (32), wherein the radial flange (18) transitions into the outer wall structure (28) and the cylinder (26) forms the sealing ring (8), wherein the inner wall structure (30) transitions, via an annular web (36), into at least one inner body segment (38, 50), wherein the sealing device (1) has a uniform, preferably relatively reduced wall thickness over its individual segments integrally formed with one another, so that the sealing device (1) is resilient within certain limits.

Abstract: The present invention relates to an arrangement 1, comprising a first turbomachine element 10 with a first turbomachine element through-opening 12, a second turbomachine element 20, a sliding element 30 with a sliding element through-opening 32 that is arranged, at least in part, in the first turbomachine element through-opening 12, a sleeve 40 that engages through the sliding element through-opening 32 along a through-axis D, and a fastening member 50 that engages through the sleeve 40 for fastening the sleeve 40 to the second turbomachine element 20. Furthermore, the invention relates to a method for fastening the first turbomachine element 10 to the second turbomachine element 20, a turbomachine with the above-mentioned arrangement 1, and use of this arrangement 1.

Abstract: A casing structure for a fluid flow machine, in particular for a gas turbine or an aircraft engine, including an outer casing wall and an inner casing wall, which annularly surround a flow channel of the fluid flow machine and are spaced apart in a radial direction with respect to the flow channel. At least one cavity is formed between the inner and outer casing walls. The cavity is axially divided into at least two regions which are separated from each other by an axial seal in such a way that different pressure conditions are created according to the axial position of these regions, which different pressure conditions correspond to the pressure conditions in the flow channel. A corresponding fluid flow machine such as, for example, an aircraft engine.

Abstract: A turbomachine having a flow channel and a housing radially surrounding the flow channel, a plurality of stationary and moving blades being situated in the flow channel, stationary blades being situated adjacent to the moving blades in the axial direction, and the stationary blades having at least one stationary blade hook which engages with at least one housing hook for the purpose of connecting the stationary blades to the housing. A sealing and lining element is situated in the radial direction between the moving blades adjacent to the stationary blades and the housing, and has a sealing structure interacting with blade tips of the moving blades. A heat protection element is provided in the area between the sealing and lining element and the housing, and a sealing element being situated on the heat protection element for the purpose of sealing contact with the housing hook.

Abstract: A housing structure for a turbomachine having a stationary blade ring annularly surrounding a flow channel, and having a seal carrier ring which also annularly surrounds the flow channel, the stationary blade ring and the seal carrier ring being connected to each other, and either the stationary blade ring or the seal carrier ring having a one-sidedly open receptacle in the radial direction in relation to the flow channel and a radially projecting engagement area which corresponds to the connected seal carrier ring or stationary blade ring and which is held in a form-fitting manner in the axial direction and in the radial direction in the direction of the closed side of the receptacle but is freely movable with respect to the connected receptacle and the seal carrier ring or stationary blade ring having this receptacle in the radially opposite direction in the direction of the receptacle opening.

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.