Abstract: Disclosed is a turbomachine having an annular flow duct and a housing structure surrounding the flow duct and a multiplicity of guide vanes and rotor blades which are arranged in the flow duct. The rotor blades are rotatably accommodated in the housing structure whereas the guide vanes are fixed in the housing structure, a plurality of guide vanes forming an annular guide vane ring. The housing structure has a seal in the region of the radially inner flow duct boundary in order to prevent hot gas escaping from the flow duct, which seal is arranged on guide vane roots of the guide vanes of the guide vane ring via a seal support and forms a seal against a rotatable seal surface, the seal support being formed from an intermetallic material, in particular a TiAl material.

Abstract: A covering device for an integrally bladed rotor base body of a turbomachine for preventing or reducing a cooling air stream that flows from a high-pressure side to a low-pressure side through channels that are formed between adjacent rotor blades, a plurality of cover elements being provided that are individually insertable into the channels and have at least one peripheral sealing surface. Also an integrally bladed rotor base body having a covering device of this kind, a method for producing such an integrally bladed rotor base body, as well as a turbomachine having such an integrally bladed rotor base body.

Abstract: A fish mouth seal carrier for a guide vane arrangement of a gas turbine comprises a first half-shell element and a second half-shell element bonded to it, which together form a box profile with two axial arms and two radial arms. A sealing element is arranged on one of the axial arms of the box profile. At least one of the two half-shell elements has an integrally formed axial flange for forming, with a guide vane platform, a fish mouth seal accommodating an axial flange of an adjoining moving blade radially between the guide vane platform and the axial flange of the seal carrier.

Abstract: A restraint mechanism for rotor blades of a turbo engine on a rotor disk having blade root receptacles for radial guidance of the rotor blades is disclosed. The restraint mechanism has a plurality of axial securing elements, each having at least one web for positioning in a blade root receptacle, such that each web has a retaining section at one end for establishing a form-fitting design to a mating brace of another axial securing element. A method for axially securing rotor blades on a rotor disk and a turbo engine are also disclosed.

Abstract: The invention relates to a gas turbine blade (10), in particular a compressor blade and/or turbine blade for a turbomachine, wherein the gas turbine blade (10) comprises at least one receptacle (18) for the form-fitting arrangement of a connecting element, by means of which a mechanical load can be introduced into gas turbine blade (10). The invention further relates to a shrouding band segment (16) for arrangement on a gas turbine blade (10), a connection system having a gas turbine blade (10) and having a connecting element, a method for connecting a gas turbine blade (10) to a connecting element, and a rotor for a turbomachine having a rotor disk (12) joined to a blade ring or having a rotor ring joined to a blade ring.

Abstract: A turbo compressor in an axial type of construction is disclosed. The turbo compressor has a bladed stator and a bladed rotor, and has a longitudinally split compressor casing and a guide blade ring with adjustable guide blades. The guide blades are pivotably mounted about radial axes radially within their aerofoil on an inner ring belonging to the stator. The inner ring is split, i.e., segmented, at at least two points of its circumference. Furthermore, the inner ring has for each guide blade at least one bearing bush which can be inserted radially into an opening from inside.

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: A method for producing an integrally bladed rotor, in particular a gas turbine rotor, comprises the following steps:—Providing a blade (12) having a lower blade foot that has a joining surface (18);—Retaining the blade (12) in a retaining unit (20); and—Exciting the blade (12) to oscillate about an axis of rotation. A device for carrying out the method comprises a retaining unit, in which blade (12) can be solidly clamped in place, and an oscillating unit (24) that transmits translational oscillations to the retaining unit in a plane substantially parallel to joining surface (18), and a fixation unit (30) for establishing the axis of rotation.

Abstract: A blade for a turbomachine is disclosed. At least one cooling channel is configured in the blade neck, whose inlet is disposed near a platform projection on the high-pressure side and whose outlet is disposed in the region of a platform projection on the low-pressure side. An integrally bladed rotor base body having a plurality of these types of blades as well as turbomachine with such a rotor base body is also disclosed.

Abstract: An inner ring for forming a guide blade ring for a turbomachine is disclosed, composed of at least two one-part ring segments having a plurality of openings, closed on the peripheral side, for accommodating journal bearings on the blade side, the outer diameter of the inner ring being at least 12 times larger than its height. Also disclosed are a guide blade ring having this type of inner ring and a turbomachine having this type of guide blade ring.

Abstract: The invention relates to a gas turbine blade (10), in particular a compressor blade and/or turbine blade for a turbomachine, wherein the gas turbine blade (10) comprises at least one receptacle (18) for the form-fitting arrangement of a connecting element, by means of which a mechanical load can be introduced into gas turbine blade (10). The invention further relates to a shrouding band segment (16) for arrangement on a gas turbine blade (10), a connection system having a gas turbine blade (10) and having a connecting element, a method for connecting a gas turbine blade (10) to a connecting element, and a rotor for a turbomachine having a rotor disk (12) joined to a blade ring or having a rotor ring joined to a blade ring.

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 producing an integrally bladed rotor, in particular a gas turbine rotor, by means of joining, comprises the following steps:—Providing a blade (10) having a joining surface;—Providing a retaining mechanism having two clamping jaws (28);—Retaining the blade (10) at two opposite points of blade (10) by means of clamping jaws (28) in a purely force-closed manner; and—Pressing the blade (10) retained in a force-closed manner onto a basic rotor body, so that blade (10) is aligned independently in a joining position. A device for producing an integrally bladed rotor comprises a retaining mechanism for a blade (10) having two clamping jaws (28), and by pressing these jaws together, blade (10) is retained at two opposite points in a purely force-closed manner; and a mechanism for pressing blade (10) retained in a force-closed manner onto a basic rotor body, in such a way that blade (10) is aligned independently in a joining position.

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: 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: The invention relates to a rotor (13) for a turbo machine, in particular for an aircraft turbine, having 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). In addition, the invention relates to a blade or vane, in particular a rotating blade (12) for a rotor (13) of a turbo machine, as well as a turbo machine.

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: The invention relates to a rotor (10) for a turbomachine, in particular for a gas turbine, having a basic rotor body (12) and a plurality of blades (14), wherein at least one damping element (24) is provided in the circumferential direction between adjacent blades (14) of the rotor (10).

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: 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).