Abstract: A turbine center frame for a gas turbine, in particular an aircraft gas turbine, having a hot gas-conveying, segmented annular duct (6), a radial outer casing (4) that is disposed at a distance around the annular duct (6), a hub element (7) that is disposed radially inwardly at a distance from the annular duct (6), a plurality of struts (5) that essentially extend in the radial direction (RR) through the annular duct (6) and that are coupled radially inwardly to the hub element (7) and radially outwardly to the outer casing (4), and having at least one connecting element (30), that is formed obliquely to the struts (5), that extends through the annular duct (6), and is coupled radially inwardly to a seal carrier element (20), the seal carrier element (20) being supported axially forwardly on the hub element (7). At least one connecting element (30) is coupled radially outwardly to one of the struts (5).

Abstract: The invention relates to a module for a gas turbine comprising a rotationally symmetrical outer casing of a first material having a first linear thermal expansion coefficient, a rotationally symmetrical component for guiding hot gas of a second material having a linear thermal expansion coefficient which is lower than the first coefficient, at least one annular structure arranged radially within the component, and at least three struts having a radially inner end secured on the outer casing and a radially outer end secured on the annular structure. The component also comprises at least one first fixing which is free in radial direction and fixed in axial direction and arranged between the component and the outer casing and/or the annular structure, as well as at least three second fixings having an elastic effect in radial, axial and/or circumferential direction and being arranged between the component and the strut and/or the annular structure.

Abstract: The invention relates to a blade or vane, a blade or vane segment, and an assembly for a turbomachine wherein the blade or vane has a blade or vane element with a blade or vane element profile and a radially inner end of the blade or vane element and a radially outer end of the blade or vane element, and wherein the blade or vane has, in addition to the blade or vane element profile, at least one first guide profile in the region of the radially inner end of the blade or vane element (EI) and/or of the radially outer end of the blade or vane element and is spaced apart from the associated blade or vane element end in the radial direction, and extends at least partly in the axial direction and at least partly in the peripheral direction and/or at least partly in the tangential direction.

Abstract: A stator vane (3) for a turbine (50c) of a turbomachine (50), the stator vane having a stator vane airfoil (3c), an inner shroud (3a) and an outer shroud (3b), the inner shroud (3a) and the outer shroud (3b) bounding an annular space (2), in which working gas (51) is conveyed during operation, radially with respect to a longitudinal axis (52) of the turbomachine (50), and the stator vane airfoil (3c) having a stator vane airfoil channel (3d) extending through its interior between a radially inner inlet (6) and a radially outer outlet (7). A characteristic features is that the inlet (6) is disposed in such a manner that a gas (8) flowing through the stator vane airfoil channel (3d) during operation is at least partially formed of the working gas (51) conveyed in the annular space (2), and thus the working gas is redistributed from radially inward to radially outward.

Abstract: Disclosed is a component with a component section and a flange that is arranged at the component section for connection of the component another element. Arranged at the flange is one or a plurality of oblique flange channels, that is, channels for respectively accommodating a section of a connection element that runs obliquely to a surface portion of the component section that lies opposite to the flange. Further disclosed is a flange connection with at least one component of this kind, another component, and at least one connection element. Disclosed, in addition, is a turbomachine with a housing that is assembled from two or more components by means of such a flange connection.

Abstract: An adjustable guide vane ring of a turbomachine having an inner ring, the guide vanes of which each have a radially inner vane disk with disk thickness increasing in the direction of flow and acting as a bearing pin; a guide vane, as well as an inner ring for such a guide vane ring, as well as a turbomachine are disclosed.

Abstract: A rotor is provided, in particular a blisk of a gas turbine, including a rotor disk or a rotor ring and rotor blades disposed on a periphery of the rotor disk or rotor ring. At least one radial slot is provided in a peripheral surface of the rotor disk or rotor ring between two rotor blades, which radial slot opens into a rounded opening extending between axial end faces of the rotor disk or rotor ring. A method for manufacturing the rotor is also provided.

Abstract: A locking element for a row of blades of a turbomachine includes a number of blades that are inserted by their roots in an anchoring groove on the rotor side and are thereby held in the radial direction in form-fitting manner in the anchoring groove. The locking element has a basic body that has a front bearing surface and a back bearing surface for form-fitting with opposite-lying groove surfaces, a front lock projection and a back lock projection for engaging in opposite-lying recesses of the anchoring groove, and a threaded region for interacting with a tightening element. Access to the threaded region for the tightening element is blocked on one side.

Abstract: An axially split inner ring for a fluid flow machine is disclosed. The inner ring is constructed for connecting to guide vanes and includes at least a first ring segment arranged upstream and a second ring segment arranged downstream. The second ring segment has a first sealing segment, where at least one section of the second ring segment is arranged in a radially inward manner relative to the first ring segment. A guide vane ring of a fluid flow machine and an aircraft engine are also disclosed.

Abstract: An axially divided inner ring for a turbomachine, for fastening to guide vanes of the turbomachine. The inner ring comprises at least one first, solid ring segment disposed upstream, and a second, solid ring segment disposed downstream, wherein the first ring segment is joined to the second ring segment in a detachable manner by means of at least one fastening element. The first ring segment and/or the second ring segment is joined to at least one sealing segment. The inner ring comprises a securing element for securing the fastening element, wherein the securing element is joined to the first ring segment and/or to the second ring segment. In addition, the present invention relates to a guide vane ring of a turbomachine having guide vanes, which have an axially divided inner ring according to the invention.

Abstract: Disclosed is a component with a component section and a flange that is arranged at the component section for connection of the component another element. Arranged at the flange is one or a plurality of oblique flange channels, that is, channels for respectively accommodating a section of a connection element that runs obliquely to a surface portion of the component section that lies opposite to the flange. Further disclosed is a flange connection with at least one component of this kind, another component, and at least one connection element. Disclosed, in addition, is a turbomachine with a housing that is assembled from two or more components by means of such a flange connection.

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: The present invention relates to a sealing system for an axial turbomachine for a gas turbine, including a rotor with a radially outward arranged shroud and a housing that surrounds the rotor, wherein a gap is arranged between the shroud and the housing, and wherein the gap is bounded by a seal, which is joined to the housing, and at least one sealing tip arranged at the shroud opposite to the seal, for reducing the flow losses through the gap. At the downstream end region of the seal, another static sealing tip, which is joined to the housing, is arranged for influencing the flow through the gap and/or for influencing the flow downstream of the static sealing tip. The present invention further relates to an axial turbomachine with at least one low-pressure turbine stage, wherein the low-pressure turbine stage comprises a sealing system according to the invention.

Abstract: A seal system for a moving blade system of a gas turbine includes a seal carrier including a ring section extending in the axial direction and the circumferential direction, on which a radially inner seal element is situated. A first flange section and a second flange section extend outwardly from the ring section in the circumferential direction and the radial direction and are situated at a distance from each other in the axial direction. An inner shroud section of at least one guide blade or of at least one guide blade segment includes a sealing section extending inwardly in the radial direction and in the circumferential direction, the sealing section being accommodated between the first flange section and the second flange section of the seal carrier and being connected thereto. A disk-like first seal means is situated upstream from the first flange section with respect to the main flow direction of a working medium and partially abuts the first flange section radially inwardly.

Abstract: A turbine center frame for a gas turbine, in particular an aircraft gas turbine, having a hot gas-conveying, segmented annular duct (6), a radial outer casing (4) that is disposed at a distance around the annular duct (6), a hub element (7) that is disposed radially inwardly at a distance from the annular duct (6), a plurality of struts (5) that essentially extend in the radial direction (RR) through the annular duct (6) and that are coupled radially inwardly to the hub element (7) and radially outwardly to the outer casing (4), and having at least one connecting element (30), that is formed obliquely to the struts (5), that extends through the annular duct (6), and is coupled radially inwardly to a seal carrier element (20), the seal carrier element (20) being supported axially forwardly on the hub element (7). At least one connecting element (30) is coupled radially outwardly to one of the struts (5).

Abstract: The invention relates to a mid-frame (10) for a gas turbine, having at least one outer casing element (24), having at least one hub element (26) arranged on the inside of the outer casing element (24) in the radial direction, having a least one strut (42) by means of which the outer casing element (24) is connected to the hub element (26), and having at least one fairing element (46) that delimits at least partially a duct (44) at least in the radial direction, through which a gas can flow, and is constructed separately from the casing element (46), said fairing element having a passage opening (48), through which the strut (42) passes, for at least partial cladding of the strut on the outer peripheral side, wherein the fairing element (46) is coupled exclusively to the hub element (26), at least in the radial direction.

Abstract: The invention relates to a mixing device and a turbofan engine having such a mixing device 30 for mixing a first gas flow 40 with a second gas flow 50 in a turbofan engine 20, having an actuating device 95 and walls 60, which bound a channel 65 for the first gas flow 40 and a channel 70 lying radially outside for the second gas flow 50, the actuating device 95 comprising a coupling element 110 that is coupled to the walls (60), the actuating device 95 being designed to pivot the walls 60 between a first position and a second position disposed radially outside relative to the first position, the actuating device 95 comprising an adjusting ring 105 that can be rotated between a first rotating position and a second rotating position in the peripheral direction and that is joined to the coupling element 110.

Abstract: The invention relates to a turbofan aircraft engine that comprises a primary duct including a combustion chamber; a first turbine disposed downstream of the combustion chamber; a compressor disposed upstream of the combustion chamber and coupled to the first turbine; and a second turbine disposed downstream of the first turbine and coupled to a fan for feeding a secondary duct of the turbofan aircraft engine. The bypass ratio of the inlet area of the secondary duct to the inlet area of the primary duct is at least 7 and the second turbine comprises at least two stages. For the first stage the mean radius r of a stator vane expressed in inch divided by the number of stator vanes is at least 0.18.