Abstract: A turbomachine blade having a main body (10) that includes a plane, first attachment surface (100) having a first rim contour (110), a cover (20; 21, 22) that has a plane, second attachment surface (200) having a second rim contour (210) that is welded to the first attachment surface, and a tuning body configuration having at least one tuning body (30; 31-33) for contacting an inner wall (220) of the cover by impact therewith, a gap (s) being formed between the first and the second rim contour.

Abstract: A brush seal for a gas turbine, in particular an aircraft engine, is disclosed. The brush seal includes a support ring which has a support plate and a support structure, where the support structure is arranged downstream with respect to the support plate. The brush seal also includes bristles which are arranged upstream with respect to the support ring, where ends of the bristles protrude radially inward beyond the support plate. The support structure yields when a scraping force acting radially outward occurs on a first lateral surface of the support structure, where the first lateral surface is directed radially inward, and the support structure does not yield when an axial operating force occurs on a second lateral surface of the support structure.

Abstract: The invention relates to a rotor (10) for a turbo machine, in particular for an aircraft turbine, with rotating blades (12) that are joined to a basic rotor body (16), whereby at least one channel (22) extending between the high-pressure side (HD) and the low-pressure side (ND) of rotor (10) radially underneath a blade platform (18) of at least one rotating blade (12) is provided, whereby a slope of a principal axis of extension (H) of channel (22) relative to an axis of rotation (D) of rotor (10) has the same sign as a slope of a principal axis of extension (R) of a radially inner boundary of the flow channel of rotor 10. In addition, the invention relates to a turbo machine having a rotor (10) as well as to a method for manufacturing a rotor (10) for a turbo machine.

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: 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 blade cascade for a turbomachine, having a number of blades (11, . . . 14; 21, . . . 25; 31, . . . 37) which include a monocrystalline material, each blade having a crystal orientation value (|?|), which is dependent on a crystal of first blades (11, . . . 14) being less than a first limiting value and the crystal orientation values of second blades (21, . . . 25; 31, . . .

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: The present invention relates to a seal configuration (100) for a turbo machine having a rotor (5), wherein the seal configuration (100) has a sealing element (300) to create a seal with respect to the rotor (5) in an installed position and a supporting device (200). The supporting device (200) has, in at least one section of the supporting device, a material (9) that melts when it comes into contact with the rotating rotor (5) in the installed position. The present invention further relates to a turbo machine having a seal configuration (100).

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 brush seal system for sealing a gap between components of a thermal gas turbine that may be moved relative to one another includes a gas seal housing that receives at least one brush head of a brush seal and a support element by means of which a brush packet projecting from the brush head of the brush seal may be supported against flexing. The support element and the brush seal housing are embodied as separate components. The invention furthermore relates to a method for producing such a brush seal system and to a thermal gas turbine having such a brush seal system.

Abstract: The present invention relates to a brush seal for a turbomachine, having a support (1) that has a recess (2), in which a spring element (3) is clamped axially, which fastens a brush element (4), the spring element being axially locked at an undercut (2.1) in the recess.

Abstract: A sealing arrangement (100) in an axial turbomachine between, on the one hand, a rotor (1) and, on the other hand, a stator (3) and/or a housing, including a support device (7, 7?) and a sealing element, whereby the support device (7, 7?) extends essentially in the radial direction (r). In the axial direction (u) of the turbomachine, the support device (7, 7?) has an axial contour that effectuates a stiffening effect for the support device (7, 7?).

Abstract: A method for mounting an integral inner ring (10) of a turbocompressor stator to a guide vane ring (12) comprises the following steps: -Providing the guide vane ring (12) with an inner radius S; -Providing inner ring (10) with an outer radius R, which corresponds to the inner radius S; -Pre-tensioning inner ring (10) to an outer radius r that is smaller than the outer radius R of the untensioned inner ring (10); and -Relaxing inner ring (10) to the outer radius R.

Abstract: A blade (1) for a turbomachine, including a turbine blade (1.1) which has a channel (1.5), an impact chamber (2) having a constricted cross section being situated in the channel for the purpose of accommodating a single impulse body (3) is provided.

Abstract: A blade (1) for a turbomachine, including an impact chamber (2), a single impulse body (3) being situated in the impact chamber, a clearance (S1+S2) between the impulse body and the impact chamber being at least 10 ?m and/or at most 1.5 mm in at least one direction.

Abstract: The invention relates to a rotor for a turbo machine, having rotating blades (12), which are joined to a basic rotor body (14), whereby a damping element (24) for damping blade vibrations is provided between blade platforms (10) of at least two adjacent rotating blades (12), damping element (24) being arched radially upward along its axial extent relative to an axis of rotation of the rotor. In addition, the invention relates to a method for manufacturing, repairing, and/or overhauling a rotor for a turbo machine, in which rotating blades (12) are joined to a basic rotor body (14), whereby a damping element (24) for damping blade vibrations is disposed between at least two adjacent rotating blades (12). In this case, damping element (24) is arched radially upward along its axial extent relative to an axis of rotation of the rotor.

Abstract: A securement system for securing an attachment portion of at least one rotor blade or rotor blade segment in a mounting site of a rotor base. The securement system has at least one holder element and a second holder element, each of which has a holding portion and a connecting portion. The connecting portions can be introduced into the mounting site so as to establish a clip connection and can be locked together in the mounting site. The holding portions are configured so as to fasten the attachment portion of the at least one rotor blade or of the rotor blade segment in the mounting site in the locked state of the connecting portions. Further disclosed is a method for securing an attachment portion of at least one rotor blade or rotor blade segment in a mounting site of a rotor base and a rotor having a securement system.

Abstract: The invention relates to a rotor (10) for a turbo machine, in particular for an aircraft turbine, having rotating blades (12), which are joined with a basic rotor body (16) and, radially underneath their blade platforms (18), comprise channels (22) for introducing cooling air (KL) into rotating blades (12), whereby a gas diffusor (24) is provided on a high-pressure (HD) side of rotor (10), this diffusor being designed to at least reduce a hot-gas (HL) flow into channels (22) during the operation of the associated turbo machine and to permit an entry of cooling air (KL) into channels (22), whereby gas diffusor (24) is joined to rotor (10) and comprises a blocking element (28) running annularly in the circumferential direction of rotor (10). A related method for manufacturing, repairing and/or overhauling a rotor (10) for a turbo machine is also provided.

Abstract: A component is disclosed. The component includes a base body with a first end section and a second end section. The first end section has a first locking protrusion with a first guide surface and a first guide section with a first support surface. The second end section has a second guide section with a second guide surface and a second locking protrusion with a second support surface. The first guide surface is contactable with the second guide surface with respect to a guide plane and the first support surface is contactable with the second support surface with respect to the guide plane such that the first locking protrusion is engageable with the second locking protrusion in a locked position.

Abstract: An integrally bladed rotor disk (20) for a turbine, including rotor blades (40), which are joined in a substance-to-substance bond to a disk element (30), and a sealing device (60) for preventing or reducing the extent to which cooling air is able to flow from a high-pressure side (12) of the rotor disk (20) through openings (24) on the rotor disk (20) to a low-pressure side (14) of the rotor disk (20).