Abstract: A method for repairing run-in coatings is provided. The method includes the steps of filling a damaged site of the run-in coating with a filling material having a material composition that corresponds to a material composition of the run-in coating or is comparable to the material composition of the run-in coating or having material properties that are comparable to material properties of the run-in coating; drying the filling material that has been filled into the damaged site; depositing a donor diffusion layer over an area of the damaged site and onto the dried filling material; and thermally treating the run-in coating at least in the area of the damaged site to locally diffuse at least one metallic element from the donor diffusion layer into the filling material.

Abstract: The invention relates to a method for joining two components (10, 12) made of a metal material, which are connected on two mutually associated joining surfaces (14, 16) by means of a joined connection, wherein at least one of the components (10) is strengthened in at least a partial region of the joining surface (14) thereof prior to joining. The invention further relates to a joined connection of two components (10, 12) made of a metal material.

Abstract: A turbomachine including at least one blade-row group that is arranged in the main flow path and at least two rows of blades that are adjacent to each other in the main flow direction, each row having a plurality of blades, whereby the trailing edges of the blades of the upstream row of blades and the leading edges of the blades of the downstream row of blades are arranged at an axial edge distance that decreases from the center of the main flow path in the direction of at least one main flow limiter.

Abstract: A flow device having a cooling-air injection system for injecting cooling air into cavities (20, 24) between sealing elements (14a, 14b) or sealing bodies (16a, 16b), respectively, on the side of the guide vanes and on the side of the rotor is disclosed.

Abstract: A method of producing holes in a component, in particular of turbomachines, wherein each hole extends from a first, outer surface to a second, inner surface of the component and wherein the method has, for example, the following steps: (i) producing a 3D model of the actual geometry of the component, at least for the region of the holes; (ii) adapting each hole on the basis of the actual geometry of the component; and (iii) generating a production program for each individual hole. In this way, the process quality and with it the quality of the holes increases, because the offset of holes caused by component tolerances is avoided and the drilling funnels are formed according to specification. Furthermore, drilling defects on account of the offset of holes and/or cores can be avoided. Overlapping holes caused by component tolerances are likewise avoided.

Abstract: A blade-disk unit for a turbine engine includes a disk and a plurality of blades which are integrally connected to the disk, a welding seam being disposed in the joining region between the blade and the disk, which welding seam is situated entirely either in the blade or in the disk.

Abstract: A mounting device and method for mounting a component is disclosed. The mounting device includes a component carrier and an adapter. The adapter is insertable into a guide of the component carrier. The method uses the mounting device to mount a component on the mounting device.

Abstract: The present invention relates to a method for producing a component, in particular a blade for a gas turbine, wherein a main body (2) is provided, to which a running-in layer (6) is applied, which can be worn away at least partially during operation to form an accurately fitting surface (11), wherein the running-in layer is applied by kinetic cold-gas compacting, and a component, in particular a blade for a gas turbine, comprising a main body, to which a running-in layer is applied, which can be worn away at least partially during operation to form an accurately fitting surface, wherein the running-in layer is a porous layer made of a Ti alloy.

Abstract: A method for removing material from a component that is connected as an anode is disclosed. In an embodiment, an electrode that is connected as a cathode is guided to the component such that a gap is formed, an electrolyte is introduced into the gap, and a closed system is formed for the electrolyte by the formation of a duct. The electrolyte is continuously guided from an inlet opening to an outlet opening of the duct. Forming the duct, e.g., by guide elements that are mounted on the electrode, ensures that only those surface parts of the component to be machined from which material is to be removed enter in contact with the electrolyte while the other surface parts do not enter in contact with the electrolyte. Since the electrolyte is continuously guided across the surface, used electrolyte is continuously discharged along with residual matter while fresh electrolyte is delivered.

Abstract: A housing for a turbomachine includes a peripheral wall delimiting an annular space and a fluid conveying system for redirecting a partial flow of a main flow, which has an axial channel, a front annular channel and a rear annular channel in fluid connection with the axial channel and having a front annular space opening and a rear annular space opening. The fluid conveying system has a front valve device and a rear valve device controllable independently of the front valve device, for opening and closing the annular space openings, and an outlet opening for tapping the partial flow from the fluid conveying system. In the closed state of the rear annular space opening, the partial flow is directable through the rear annular channel and, in the closed state of the front annular space opening, the partial flow is directable through the front annular channel.

Abstract: A method for removing the coating from a gas turbine component, namely for the complete or partial removal of a multilayer wear protection coating from the surface of the gas turbine component, the wear protection coating having at least one relatively hard ceramic layer and at least one relatively soft metallic layer, wherein, in order to remove the multilayer wear protection coating, the gas turbine component is alternately positioned in two different chemical baths, a first bath being used exclusively for the removal of each relatively hard ceramic layer, and a second bath being used exclusively for the removal of each relatively soft metallic layer of the wear protection coating.

Abstract: A blading for a turbomachine, particularly for a gas turbine, wherein thickened areas and depressions are formed and disposed on a lateral wall having a plurality of blades such that at least one depression or thickened area is disposed at a blade pressure side and at least one thickened area or depression is disposed at a blade suction side for each blade of the plurality of blades.

Abstract: The invention relates to a method of nondestructive and contactless testing of components (3), wherein ultrasonic waves (6) are irradiated onto the surface of the component (3) at a predefinable, non-perpendicular angle of incidence (9) using an ultrasonic transmission sound transducer (1) arranged spaced apart from the surface of the component (3) and the intensity of the ultrasonic waves (7) reflected from the surface of the component (3) is detected with time resolution and/or frequency resolution by the antenna array elements (2n) of an ultrasonic antenna array (2) configured for detecting ultrasonic waves (7) and the phase shift of the ultrasonic waves guided at the surface of the test body is determined therefrom with respect to the ultrasonic waves (7) directly reflected at the surface of the component (3).

Abstract: A compressor is disclosed, which has an inflow channel for guiding a compression medium into a compressor housing, a pressure chamber formed within the compressor housing, and a liner segment with nozzles that transport the compression medium from the pressure chamber to a rotor. The nozzles are formed in the liner segment as a plurality of nozzles arranged in a group in fan-like manner.

Abstract: A rotor disk for a turbomachine, which is connectable to at least one rotor blade and/or a shaft of the turbomachine, having at least one borehole, which has an elliptical inlet opening having a first passage cross-sectional area and an elliptical outlet opening having a second passage cross-sectional area, so that the second passage cross-sectional area is smaller than the first passage cross-sectional area.

Abstract: The turbomachine has a ring of radial blades which are connected to an annular end wall. Situated on the end wall upstream of each blade is a deflector wedge with two divergent guide surfaces. The deflector wedge generates pressure-side and suction-side vortices of which the vortices caused by the blade cascade are altered such that the transverse flow and backflow areas on the blade are reduced, and therefore the edge zone losses are decreased.

Abstract: A method for replacing a blade (12) of a rotor (2) having integral blades, wherein a new blade (22) is joined to a base (34) arranged on a separation surface (28) on the main rotor body, and to a rotor repaired or mended according to said method.

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 method for producing a rotor or stator blade (1a) of a gas turbine, in particular an aircraft gas turbine (2), comprising forming a airfoil (8a) made of at least one first material and attaching a blade root (4a) made of at least one second material to the airfoil (8a) by means of a thermal spraying method. The invention further relates to a rotor or stator blade (1a) produced according to such a method.

Abstract: A gas turbine including at least one compressor, one combustion chamber, and at least one turbine including at least one rotor and at least one generator coupled to the at least one rotor is provided. The at least one turbine is coupled to the at least one compressor. Once the gas turbine is shut down, the at least one generator can be used as a motor in order to drive the at least one rotor for a predetermined time period following shutdown of the gas turbine and thereby effect a uniform cooling of the rotor. A method of operating a gas turbine is also provided.