Abstract: The invention relates to a high-pressure compressor of a gas turbine having at least one blisk, specifically a disk with a plurality of blades that are disposed on the disk, wherein the blisk has at least one titanium material from the group that comprises Ti-6246 and Ti-6242, wherein the materials of the disk and the blades of the blisk differ in their chemical composition and/or their microstructure. The blades are welded to the disk, in particular, by means of a pressure welding method supported by high-frequency current, whereby the current is conducted through the joint surfaces and leads to local limited melting therein.

Abstract: The present invention relates to methods for joining a rotationally asymmetrical part, preferably a turbine blade, which is made of a monocrystalline material, to a part, preferably a rotor core, which is made of polycrystalline material, having the method steps of: —joining the rotationally asymmetrical part to an adapter piece made of polycrystalline material by means of rotational friction welding, and —joining the adapter piece to the part made of polycrystalline material. The present invention also relates to a turbine blisk having a rotor core made of polycrystalline material and turbine blades made of monocrystalline material.

Abstract: A turbine of a gas turbine, in particular of a gas turbine aircraft engine, having a rotor that has at least one moving blade ring and having a stator that has at least one guide blade ring, moving blades of the, or each, moving blade ring and/or guide blades of the, or each, guide blade ring being fashioned as hollow blades having at least one cavity, wherein on at least one side of at least one moving blade and/or guide blade fashioned as a hollow blade, in a blade wall holes are made that connect the, or each, cavity to the surrounding environment of the respective moving blade and/or guide blade, so that the respective cavity can be used as a resonator or sound muffler in order to reduce the sound radiated by the turbine during operation.

Abstract: A method for joining at least two components by inductive high-frequency pressure welding is disclosed. The first component has a first material structure with a first hardness and the second component has a second material structure with a second hardness which is smaller than the first hardness. Both components are inductively heated in the region of the joining surfaces and are subsequently pressed together by a compressive force. The joining surface of the first component with the first hardness, which is greater than the second hardness of the second component, is pre-contoured in a spherical, conical or convex manner prior to joining the two components.

Abstract: A turbine blade assembly is disclosed. The turbine blade assembly includes a turbine blade and an outer shroud section at the radially outer end of the turbine blade. The turbine blade intersects or joins the shroud section. For increased stiffness, the shroud section has a reinforcing rib extending along its outer face and extending substantially in the direction of the principal axis of inertia.

Abstract: A system for injecting a fluid into a wall boundary layer of a flow in a turbomachine is disclosed. The system has a plurality of nozzles which are disposed in a side wall limiting the flow and are oriented diagonally in the direction of flow. The nozzles each have a rectangular, flat nozzle cross-section. A compressor having such a system, as well as a turbomachine having such a compressor, are also disclosed.

Abstract: A damping system for damping vibrations of a rotor blade of a rotor of a turbomachine is disclosed. A damping element is guided on a support such that the damping element is radially outwardly movable during a rotation of the rotor and is contactable with a lower platform area of a rotor blade. A circumferential contact surface is formed by an elevation on the lower platform area, where the circumferential contact surface is a limit stop for a movement of the damping element.

Abstract: The invention relates to a clearance control system and a method for adjusting a running clearance between a rotor having rotor blades of a turbomachine, and a casing that surrounds at least sections thereof. At least one adjusting gear unit, which can be coupled to at least one casing segment allows for movement of at least one segment radially in relation to the rotational axis of the rotor. An adjusting element can be arranged around the rotor and coupled to at least one adjusting gear unit and can be moved in relation to it for actuating the adjusting gear unit, whereby axial movement and/or pivoting of the adjusting element in relation to the rotor actuates adjustment of the running clearance, with each adjusting gear unit converting an at least predominantly axial movement of the adjusting element into at least predominantly radial movement of the assigned segment of the casing.

Abstract: A sealing device for an integrally bladed rotor basic body of a turbomachine to suppress a gas exchange between a hot gas stream and a cooling air stream in the radial direction in contact areas of adjacent moving blades, including a plurality of sealing elements activatable by centrifugal force, which may be situated in the area of channels which extend from a high-pressure side to a low-pressure side and are peripherally delimited by the moving blades, an integrally bladed rotor basic body having such a sealing device, and a turbomachine having such an integrally bladed rotor basic body.

Abstract: The invention relates to a housing system (10) for an axial turbomachine, in particular for an aircraft turbine, comprising a housing (12), at least one guide blade (14), which is fastened to the housing (12) by means of at least one fastening element (18),a scaling element (20) for reducing a leakage flow between the listening element (18) and the housing (12), wherein the sealing element (20) is arranged between a downstream wall (22) of the fastening element (18) with respect to a flow direction (I) of the axial turbomachine and a wall (24) of the housing (12), and a pressing means (36), by means of which an axial force can be applied to the scaling element (20) And the scaling element can be pressed against the wall (24) of the housing (12), at least during the operation of the axial turbomachine. The invention further relates to an axial turbomachine having such a housing system (10).

Abstract: A coupling element 2, 52 is disclosed for arrangement between shrouds 14,16 of adjacent rotor blades 4, 6, the coupling element having force-transmission surfaces 28, 30 for bounding damping gaps 12a, 12b between shrouds 14, 16, and a rotor having at least one coupling element 2, 52 of this type.

Abstract: A method for spraying a coating and a cold gas spray nozzle is disclosed. The method includes spraying a coating by the cold gas spray nozzle. A rinsing gas is fed to the cold gas spray nozzle during an interruption of the spraying or at an end of the spraying. Deposits in the cold gas spray nozzle are cooled and detached by the rinsing gas.

Abstract: A rotor (13) for a turbo machine, in particular for an aircraft turbine, has rotating blades (12), which are joined to a basic rotor body (14), in which at least one rotating blade (12) has at least one inner cooling channel, which extends at least along a predominant region of a blade element (18) of rotating blade (12) and has an inlet opening (23) for introducing cooling air into blade element (18), in which the inlet opening (23) is formed in a blade neck (16) lying between a blade foot (32) and a blade platform (10) of rotating blade (12). Inlet opening (23) is disposed in a cavity (22) formed radially underneath blade platform (10).

Abstract: A seal arrangement for a gas turbine is disclosed. The seal arrangement is used for sealing a gap between radially internally located ends of guide vanes of a guide vane ring and a rotor, in which case the rotor has at least two seal projections positioned at an axial distance relative to each other in a circumferential direction of the rotor. The seal projections effecting a seal of the gap in combination with intake linings associated with the radially internally located ends of the guide vanes. The seal projections are inclined or tilted in the axial direction toward a side of higher pressure, where, in a space limited by the minimum of two seal projections and the corresponding intake linings, at least one recirculation structure is provided. The recirculation structure, or each recirculation structure, is oriented toward the side of the higher pressure.

Abstract: The present invention relates to a blade of a turbomachine, in particular an adjustable guide blade or vane of a gas turbine, having at least one thickened area (18) on a pressure side (D) of the blade profile (P), wherein the thickened area (18) is disposed in a radially outer-lying, housing-side region of the blade (10), wherein the thickened area (18) is designed at a distance from a front edge (12) and a rear edge (14) of the blade (10).

Abstract: A method for maintenance, in particular repair, of gas turbines, in particular aeroengines. Gas turbines are disassembled, with modules and/or assemblies and/or individual parts of the gas turbines then being inspected and/or repaired, and then assembled from inspected and/or repaired and/or new modules and/or assemblies and/or individual parts. The maintenance is subdivided into at least two repair steps, wherein modules and/or assemblies and/or individual parts to be repaired are moved through repair stations in order to move the modules and/or assemblies and/or individual parts of the or each gas turbine to repair stations adapted for this purpose in order to carry out the repair steps. The repair stations may be arranged in repair lines, with gas turbine modules, assemblies and/or individual parts being directed to different repair lines after inspection. The method permits gas turbine maintenance to be performed in a conveyer belt manner.

Abstract: A turbo engine, particularly a gas turbine aircraft engine, has compressor components, turbine components, and at least one combustion chamber. At least one support rib is in flow channel between two turbine components, connected one behind the other. Each support rib diverts a flow through the flow channel. A preferably cylindrical guide element runs within each support rib. Each support rib has a suction side with a greater thickness toward a radially inner flow channel wall as well as toward a radially outer flow channel wall, when viewed in the radial direction. Each support rib has a pressure side with a greater thickness toward a radially inner flow channel wall as well as toward a radially outer flow channel wall, when viewed in the radial direction. The front edge and the rear edge of each support rib are inclined in the meridian direction.

Abstract: The present invention creates a layer system for the rotor/stator seal of a turbomachine, in particular a compressor, which is disposed between components of the turbomachine and can be run in with a movement of the components relative to one another, in such a way that at least one of the components is run into the layer system, having: a first adhesive layer disposed on at least one of the components; a protective layer disposed on the first adhesive layer; a second adhesive layer disposed on the protective layer; and a running-in layer, which is formed softer than the protective layer and which is disposed on the second adhesive layer. The present invention further provides a method for producing a layer system for the rotor/stator seal of a turbomachine, as well as a turbomachine.

Abstract: A blade channel having a not-axially-symmetric end wall contour in a turbomachine is disclosed, the end wall contour having at least one individual contour in the form of an elevation, on the pressure side, and at least three individual contours in the form of two recesses and one elevation, on the suction side, the elevation being situated between the recesses in the flow direction; a turbomachine having a plurality of blade channels of this type is also disclosed.

Abstract: A method for connecting metallic components, in particular components of a gas turbine, including: corresponding connecting surfaces of the components being connected by means of inductive HF pressure welding, and that during or after a sufficiently great heating of the connecting surfaces, the first component is moved by a definite path toward the second component, and is pressed against it and held there.