Abstract: A blade arrangement for a turbo engine, in particular a gas turbine, with a rotor and several blades fastened thereto, which are configured to be systematically different, wherein at least two adjacent blades have systematically different shrouds (12, 22) and/or inner blade platforms (11, 21).

Abstract: An anti-wear coating, in particular an anti-erosion coating, which is applied to a surface of a component that is stressed under fluid technology and is to be protected, in particular a gas turbine part, is disclosed. The anti-wear coating includes one or more multilayer systems applied in a repeating order to the surface to be coated, and the/each multilayer system includes at least one relatively soft metallic layer and at least one relatively hard ceramic layer. All the layers of the/each multilayer system are based on chromium, and a diffusion barrier layer is applied between the surface to be protected and the multilayer system(s).

Abstract: A method and device for surface peening, particularly ultrasonic shot peening, of a component in the region of a passage opening is disclosed. A vibration device impinging the blasting material is disposed on one side and a counter-element is disposed on an opposite side relative to a component region having the passage opening. The counter-element is disposed at a distance relative to the corresponding side of the component region.

Abstract: A pulsed machining method for the optimized machining of a contour which is characterized in that the instantaneous frequency ratio of ? of the pulse frequency of tool fL and of the rotational frequency of tool fCNC is optimized to achieve a highest possible degree of coverage of the individual machining pulses in the context of a shortest possible machining duration and without a direct concatenation of the machining pulses. It is especially preferred that this be achieved in that frequency ratio ? is an irrational number, thus cannot be expressed by the ratio of two whole numbers. By applying the method according to the present invention, a considerable time savings is achieved since there is no longer a need for time-consuming preliminary trials to ascertain an at least sufficient frequency ratio ?.

Abstract: A method for the milling machining of components is disclosed. The method produces recesses with one or more lateral walls, in particular for the production of integrally bladed rotors for gas turbines, the recesses forming flow channels and the lateral walls forming blade surfaces of an integrally bladed rotor. The component to be machined by milling is clamped in a locating device for the milling machining. To adjust the vibrational properties of the component to be machined by milling, areas of the component are embedded in and/or filled with a machinable plastic.

Abstract: A gas turbine compressor is disclosed. The compressor includes a compressor housing, guide vanes, rotor blades, and valve-controlled blow-in openings for stabilizing the compressor flow by air that is blown in. The air flow is detected by pressure sensors and the valves are controlled as a function thereof.

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 device for fixing a component (4) in position on a component carrier (6) using a radiation-curing adhesive agent (8), in which case the radiation (30, 32) for curing the adhesive agent (8) located between the component carrier (6) and the component (4) is laterally introducible; as well as a method for fixing a component (4) in position in a component carrier (6) using an adhesive bond.

Abstract: Disclosed is a process for producing a run-in coating (20, 24, 32, 44) on a component of a turbomachine, in particular of a gas turbine. The run-in coating is applied and produced on the component of the turbomachine by a kinetic cold gas compacting process (K3). The invention also encompasses a run-in coating for a static or rotating component of a turbomachine and a static or rotating component of a turbomachine, in particular of a gas turbine, having at least one run-in coating.

Abstract: A guide blade of a turbomachine, in particular of a compressor, a blade of the guide blade without an inner shroud having a flow inlet edge, a flow outlet edge, a suction side, and a pressure side. The blade is formed by a plurality of blade sections stacked one on top of the other in the radial direction, and the centers of gravity of the blade sections extend along a stacking axis. The blade sections are stacked one on top of the other in the radial direction in such a way that, in a radially inner section of the blade adjacent to a radially outer section of the blade, the stacking axis has its single inflection point in its radial curvature, namely, between a first, radially inner subsection of the radially inner section in which the stacking axis has a concave curvature toward the pressure side, and a second, radially outer subsection of the radially inner section in which the stacking axis has a concave curvature toward the suction side.

Abstract: An arrangement for optimizing the running clearance for turbomachines of the axial type, such as turbocompressors, gas turbines, and steam turbines, in particular for compressors of stationary gas turbines, by controlling the inner diameter, which is relevant to the running clearance, of at least one stator structure that surrounds a rotor blade ring, including: the stator structure has a closed, circular inner ring, a circular outer ring that is situated concentric to the inner ring at a radial distance therefrom, and a plurality of links that integrally connect the inner ring to the outer ring, the links being circumferentially inclined at a defined angle (?) to the radial direction and distributed around the circumference of the stator structure, and the arrangement includes an adjustment device for rotating the inner ring relative to the outer ring with elastic modification of the running clearance-relevant inner diameter (D) of the inner ring.

Abstract: Described is a material sample for testing material properties under biaxial load as well as a method therefor. The material sample has a circular disk shape and, for rotation about the center line thereof, is provided with an integrally formed hub. The edge of the material sample and the hub are thicker than the ring section of the material sample located in between. In one embodiment, the ring section is designed with an annular concave fillet concentric to the hub and the concave fillet may extend from the edge to the hub and have a continuous curvature over its width. In the method, it is rotated about its center line at least at 100,000 rpm, preferably at 150,000 rpm±10,000 rpm.

Abstract: An aircraft engine, in particular a helicopter engine, having a one or multi-stage compressor system (V), a combustion chamber (BK) connected downstream therefrom, and a one- or multi-stage turbine system (HT, NT) connected downstream therefrom, a compressor heat exchanger system (WV), a turbine heat exchanger system (WT), and a bypass means for optionally guiding the working medium through or past at least one turbine heat exchanger (WT) of the turbine heat exchanger system.

Abstract: A device (10) for suspending gas channel elements, in particular for suspending guide blades or guide blade segments or gas channel segments, on a housing of a gas turbines is provided. The device comprises first plate-shaped elements (11, 12, 13, 14) and second plate-shaped elements (15, 16, 17), whereby the first plate-shaped elements (11, 12, 13, 14) and the second plate-shaped elements (15, 16, 17) are connected together by web-like elements (18) which extend in an essentially perpendicular manner in relation to the first and second elements and form a meandering or crenelated profile.

Abstract: A shroud for the rotating blades of a turbo machine, particularly a gas turbine, in which shroud is arranged along the circumference of a row of several rotating blades disposed on a rotor and has at least one separation gap along its circumference. The separation gap is formed in zigzag shape and has at least three damping gaps that are distanced from one another and extend at an angle relative to an axis of rotation of the rotor, and adjacent to these, has connection gaps each connecting damping gaps or extending the latter in the direction of shroud edges, wherein, when the rotor rotates, the gap width of damping gaps is reduced until the gap walls forming the damping gaps come to rest against each other.

Abstract: A tandem blade design for an axial turbomachine, comprising a front blade and a rear blade disposed with an offset thereto in the circumferential direction and in the axial direction. The rear blade is profiled and positioned with respect to the front blade such that it raises the speed level at the trailing edge of the front blade in a predetermined working range in interaction with the front blade.

Abstract: A damper (2) for damping a blade movement of a turbomachine (1), and to a method for producing the damper (2). The damper (2) has at least one side surface (21, 21?) which can be brought into frictional contact with a friction surface of the turbomachine (1) in order to damp a blade movement. The side surfaces (21, 21?) are asymmetrically convex in shape.

Abstract: The present technology relates generally to a device for the detection of a shaft fracture in a rotor of a turbine, for example a medium pressure turbine such as a gas turbine, in particular an aircraft engine, whereby at least one stator-side sensor element is positioned downstream from the turbine in the area of a stator-side guide vane ring of another turbine in particular, a low pressure turbine; and whereby a radially inner section of the last rotor-side moving blade ring (as seen in the direction of flow) of the turbine cooperates with at least one sensor element directly and/or proximally in the event of a shaft fracture in the rotor of the turbine to generate an electric signal corresponding to the shaft fracture.

Abstract: In a moving blade system for a turbomachine, in particular a gas turbine, having at least one moving blade (1), the moving blade system having at least one cavity (3) in which at least one tuning mass (2) is movably situated, the tuning mass and/or the cavity is/are adapted in such a way that the tuning mass rests against an inner wall (3.1) of the cavity in a predefined first operating state of the turbomachine and at least temporarily moves away from the inner wall in a second predefined operating state of the turbomachine.

Abstract: Solder alloy based on nickel is composed of a mixture with a first soldering material, a second soldering material, and a base material, wherein the base material is a nickel-based material which corresponds to the material to be soldered and is present in a proportion of 45-70% by weight in the mixture, the first soldering material is a nickel-based material including chromium, cobalt, tantalum, aluminum and boron, and is present in a proportion of 15-30% by weight in the mixture, and the second soldering material is a nickel-based material including chromium, cobalt, molybdenum, tungsten, boron and hafnium, and is present in a proportion of 15-25% by weight in the mixture.