Abstract: The present invention relates to a module for a turbomachine. The guide vane arrangement has a guide vane, an inner platform, and a first guide pin, the guide vane is arranged radially outside on the inner platform and the first guide pin is arranged radially inside on the inner platform, and wherein the seal carrier is arranged radially inside the inner platform and forms an uptake, which is open radially outward, the first guide pin of the guide vane arrangement being arranged in the uptake. In the uptake, a slide body is arranged, on which the first guide pin of the guide vane arrangement, with respect to a rotation around the longitudinal axis of the module, finds a contact and is guided in a radially movable manner in this contact. The slide body is formed in one piece with a wall of the seal carrier, which axially bounds the uptake.

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 present invention relates to a component of a turbomachine made of a composite material comprising a ceramic matrix and a corrosion protection layer arranged on the composite material. The corrosion protection layer comprises: 50-80 wt % silicon, 12-30 wt % nitrogen, 2-8 wt % oxygen, 2-8 wt % aluminum, and 0-3 wt % of at least one alkaline earth metal. The invention furthermore relates to a method for producing a corresponding component of a turbomachine.

Abstract: The present invention relates to an adjusting device for adjusting guide vanes of a guide vane cascade of a turbomachine, in particular of a turbine stage or a compressor stage of a gas turbine, wherein the adjusting device comprises a coupling ring for coupling the guide vanes and at least two actuators for rotating the coupling ring, said actuators being connected to the coupling ring, wherein at least one of these actuators has a disconnection coupling.

Abstract: The invention is directed to a seal arrangement, in particular a brush seal, for a turbine, in particular an aircraft gas turbine, with a sealing element and with a sealing surface that interacts with the sealing element, wherein the sealing element and the sealing surface are arranged concentrically with respect to each other, in particular concentrically around a machine axis of the turbine is hereby characterized in that the sealing element and the sealing surface are designed in such a way that, in a first operating state, a gap is formed between them, and in that, in a second operating state, a sealing contact is produced between the sealing element and the sealing surface.

Abstract: A containment for a continuous flow machine, in particular for a gas turbine, is disclosed. The containment has at least one annular housing with at least one axial section of lesser radial wall thickness and at least one axial section adjacent thereto of greater wall thickness, where the housing is designed and constructed so that the section of greater wall thickness is at a distance in the axial direction of the continuous flow machine from a trailing edge of a guide vane lying downstream in the direction of flow of the continuous flow machine, and extends downstream in the direction of flow of the continuous flow machine over a leading edge of the next downstream guide vane.

Abstract: Disclosed is a method of repairing damage of an aircraft engine component made of or comprising weldable thermoplastic material, which method comprises welding of the weld filler material into an initially prepared damaged area. Also disclosed are the use of plastic welding technique for repairing damaged aircraft engine components made of weldable thermoplastic materials.

Abstract: The segmented seal carrier of a guide vane ring of a turbomachine has a number of segments that form two half-rings of the seal carrier when assembled. They are clamped together by their opposite-lying, corresponding lateral contact regions.

Abstract: The present invention relates to a blade for a gas turbine, in particular of an aircraft engine, which is produced at least in part from ceramic matrix composite with a plurality of superimposed fabric layers, wherein at least one pair of superimposed fabric layers comprises a first fabric layer that has at least one first point of interruption between two mutually facing first edges of this fabric layer, and a second fabric layer that has at least one second point of interruption adjacent to the first point of interruption between two mutually facing second edges of the second fabric layer, this second point being displaced from the first point of interruption.

Abstract: Disclosed is a method for producing a blade for a turbomachine, in particular for an aero engine. The method comprises providing at least one blade airfoil with a first platform region and at least one blade root with a second platform region and joining the blade airfoil and the blade root at the respective platform regions by a friction welding method at a common joint region of the platform regions, the blade airfoil and the blade root being made of materials which are different from each other. Also disclosed is a blade which is and/or can be obtained by such a method.

Abstract: A turbomachine blade assembly including a turbomachine blade (1), in particular for a gas turbine, and at least one tuning element container including a housing (10) attached to the turbomachine blade and an insert (20) disposed in a recess (11) of this housing. A wall (20; 21) of the insert spaces apart two first cavities (31), which each accommodate at least one tuning element (40) provided for impacting contact with the housing (10) and the insert (20).

Abstract: The present invention relates to a turbomachine blade or vane arrangement having a first turbomachine blade or vane, a second turbomachine blade or vane adjacent to it, and at least one tuning element guide housing with at least one cavity, in which at least one tuning element is arranged with play of movement for impact contact with the tuning element guide housing, with the tuning element guide housing being arranged at least in part in a recess, in particular in a frame, of the first turbomachine blade or vane, where the second turbomachine blade or vane has at least one first rib for securing the tuning element guide housing arranged in the recess.

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 is directed to a method for producing a rotating disk for a turbomachine, forming blade root uptakes on a plurality of circumferential positions, for which, in each case, an axially extending profile groove that is bounded circumferentially by inner wall surfaces of the rotating disk is introduced on a respective circumferential position, in which production method, initially first profile grooves are introduced in a disk on first circumferential positions, and subsequently, a second profile groove is introduced in the disk on a second circumferential position located circumferentially between the first profile grooves.

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: The present invention relates to a method for producing a subassembly having a form-fitting connection with a precipitation-hardened form-fitting region as well as a corresponding subassembly, wherein at least two components of a subassembly are provided that are connected together in form-fitting manner, wherein each of the components has a form-fitting region that can come in contact with at least one other form-fitting region of the other component to be connected, in order to produce a form-fitting connection by limiting at least one degree of freedom of movement of the connected components relative to one another, wherein at least one of the components has at least one deformation form-fitting region for providing the form-fitting connection that is reshaped for producing the form-fitting connection after arranging the components to be connected relative to one another, in order to produce the form-fitting connection.

Abstract: The present invention relates to a run-in coating for a turbomachine, in particular for an aero engine, to be attached to a seal between a rotor and a stator, wherein the run-in coating comprises a metallic scaffold having interstices which are filled with an inorganic-nonmetallic filler material, and wherein the metallic scaffold is formed from an arrangement of metal fibers (9). The invention also relates to a corresponding turbomachine.

Abstract: The invention relates to a device (10) for generative production of at least one component area of a component (12), in particular of a component (12) of a flow machine. Therein, the device (10) can include at least one lowerable component platform (16) with at least one construction and joining zone (20) for receiving at least one powder layer of a component material, at least one heating device (24, 28) movable relative to the component platform (16) and at least one radiation source for generating at least one high-energy beam (22), by means of which the powder layer can be locally melted and/or sintered to a component layer in the area of the construction and joining zone (20), as well as at least one suction and/or gas supply device (36) disposed on the heating device (24, 28). However, the device (10) can also include at least one coater (14).

Abstract: The invention relates to a device for additively producing at least one component region of a component, in particular a component of a turbomachine. The device can comprise at least one coating apparatus for applying at least one powder layer of a component material to at least one build-up and joining zone of at least one component platform, which can be lowered, wherein the coating apparatus can be moved in relation to the component platform. In addition, the device comprises at least one suction device for suctioning of process waste gases and/or particles. The coating apparatus thereby comprises at least one suction channel or at least one suction pipe, which has at least one suction opening oriented in the direction of the component platform.

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.