Abstract: The present invention relates to a turbomachine assembly with an annular flow cascade, which has a plurality of blades, which are arranged distributed in a peripheral direction, and a detuning device for different detuning of natural frequencies of the blades, wherein the detuning device has a ring that revolves in the peripheral direction, or wherein the detuning device has all detuning elements that are arranged at the blade, which follow one another in a row in the peripheral direction, and are designed in such a way that, during operation, first blades of the flow cascade each contact at most one, in particular, no movable detuning element of the detuning device.

Abstract: The invention relates to a nickel-based alloy having a microstructure with a matrix of ?-phase and precipitates of ??-phase. The ??-phase comprises a percentage by volume of from 50 vol % to 80 vol % in the temperature range of from 1000° C. to 1100° C. The nickel-based alloy comprises 8 to 13 at % aluminum, 3 to 14 at % cobalt, 4 to 12 at % chromium, 0.6 to 8 at % molybdenum, 0 to 6 at % rhenium, 0.5 to 4 at % tantalum, 0.5 to 4 at % titanium, 0.3 to 3.5 at % tungsten, 0 to 4 at % germanium, 0 to 0.6 at % hafnium, 0 to 4 at % ruthenium, balance nickel and unavoidable impurities. The concentrations of molybdenum and tungsten are selected such that the percentage X of molybdenum and tungsten in the ?-phase, X=0.84 CMo+CW, is greater than 5.5 at % at a temperature of from 1000° C. to 1100° C., CMo and CW being the concentrations of molybdenum and tungsten in at %.

Abstract: The invention relates to a run-in coating (10) for an outer air seal of a turbomachine, especially of an aero engine, comprising a substrate (12) of segmented form which carries a honeycomb structure (14) with an run-in surface (15) to face rotor blades of a rotor of the turbomachine, wherein a radial thickness of the honeycomb structure (14) varies in the circumferential direction. The invention furthermore relates to an outer air seal for a turbomachine and also to a turbomachine with an outer air seal.

Abstract: An exhaust-gas aftertreatment system for an internal combustion engine, having a catalyst device, which is designed to catalytically react at least one exhaust-gas component with a reactant, and a reactant-metering device, which is arranged upstream of the catalyst device along a flow path of the exhaust gas through the exhaust-gas aftertreatment system. The reactant-metering device has at least one exhaust-gas flow nozzle.

Abstract: A housing element (12) for an intermediate turbine housing of a gas turbine, in particular an aircraft gas turbine; the housing element (12) being installable or installed in the intermediate turbine housing radially outwardly, in each case between a plurality of circumferentially spaced struts; the housing element (12) having a planar form and, relative to the radial outer side (16) thereof, a plurality of depressions (18c), is provided. On the radial inner side (32, 34) thereof, the housing element (12) has at least one recess (26) that is configured outside of the regions in which depressions (18c) are formed on the radial outer side (16).

Abstract: The present invention relates to a device for the additive manufacturing of components by selectively irradiating a powder bed, the device having a working chamber in which at least one powder bed chamber and at least one radiation source are arranged such that the radiation source can irradiate a powder in the powder bed chamber, and wherein the device comprises at least one induction coil such that the powder bed and/or a component, which is generated by irradiating the powder bed, can be at least partially inductively heated.

Abstract: The present invention relates to a TiAl alloy for use at high temperatures which has aluminum and titanium as main constituents. The TiAl alloy has an aluminum content of greater than or equal to 50 at. % and a matrix of ?-TiAl and at least one phase of Al and Ti incorporated in the ?-TiAl matrix which is different from ?-TiAl, as well as depositions of oxides and/or carbides and/or silicides. In addition, the invention relates to a method for producing the alloy and to the use of the alloy for components of turbo-machines, in particular aircraft engines.

Abstract: The invention relates to a blade or vane, a blade or vane segment, and an assembly for a turbomachine wherein the blade or vane has a blade or vane element with a blade or vane element profile and a radially inner end of the blade or vane element and a radially outer end of the blade or vane element, and wherein the blade or vane has, in addition to the blade or vane element profile, at least one first guide profile in the region of the radially inner end of the blade or vane element (EI) and/or of the radially outer end of the blade or vane element and is spaced apart from the associated blade or vane element end in the radial direction, and extends at least partly in the axial direction and at least partly in the peripheral direction and/or at least partly in the tangential direction.

Abstract: The present invention relates to a method for producing a component of a composite material comprising a metal matrix and incorporated intermetallic phases, which method comprises providing powders of at least one member of the group which comprises pure chemical elements, alloys, chemical compounds and material composites, the powder corresponding overall to the chemical composition which the composite material to be produced is intended to have, each individual powder being different to the chemical composition of the composite material to be produced, compacting the powders, bonding the powders to one another to form a unit and thermoplastically shaping the unit.

Abstract: A rotor stage (200) for a turbomachine, including a plurality of rotor blades (9), a rotor main body (7) having at least two rotor arms (3) and a balancing assembly (1) for balancing the rotor stage (200). The rotor arm (3) configured downstream from the rotor main body (7) features the balancing assembly (1). The balancing assembly (1) is configured between the axial end region of the rotor arm (3) and the rotor disk (27) of the rotor main body (7). Furthermore, a rotor drum (17, 18) for a turbomachine and a rotor (1000) for a turbomachine.

Abstract: The present invention relates to a device as well as a method for the additive manufacture of components by deposition of material layers by layer-by-layer joining of powder particles to one another and/or to an already produced pre-product or substrate, via selective interaction of the powder particles with a high-energy beam, wherein, for smoothing a surface of the component being produced running crosswise to the deposited material layers in between the deposition of two layers of the component, the complete edge region of the last layer that is applied and that runs along a surface of the component being produced is compacted in a direction of action that has a directional component parallel to the build-up direction of the layers, and/or at least one edge region of a surface of the component is also compacted.

Abstract: A blade channel of a turbomachine that is delimited in the circumferential direction of the turbomachine by a pressure side of an airfoil and by an opposite suction side of an adjacent airfoil that, in the radial direction of the turbomachine, is delimited by two opposing side walls, and whose extent in the axial direction of the turbomachine is delimited by leading edges and by trailing edges of airfoils; at least one of the side walls being provided with localized contours, of which at least two are formed as elevations and at least two as depressions; a saddle surface being formed between the contours that, in one rotation, alternately merges into an elevation and a depression; a blade cascade having such blade channels, as well as a turbomachine having such a blade cascade.

Abstract: A gas turbine compressor having at least one airfoil tip (10) and a flow duct wall (20) which is disposed radially opposite thereto and has a circumferential groove (31-33) therein in which is disposed at least one web (40) having a radial cutback (44) is provided. An upstream beginning (41) of the cutback is located axially downstream of an upstream groove edge (21) between this groove edge and an upstream leading edge (11) of the airfoil tip, and a downstream end (42) of the cutback is located in an airfoil-tip-proximal half (34) of a radial height (35) of the circumferential groove.

Abstract: The invention relates to a guide vane system for a turbomachine with at least one guide vane, which can be rotatably mounted around an adjustment axis and is arranged with a radially inner end region in a corresponding recess of an inner ring, wherein the arrangement of the guide vane on the inner ring is secured by a securing element in a form-fitting manner. For an especially advantageous securing of this arrangement, it is provided that the securing element is designed as an oblong element, which is arranged at least in a recess and/or through-opening of the radially inner end region of the guide vane directed in the peripheral direction of the turbomachine and is arranged at least in a recess and/or through-opening of the inner ring directed in the peripheral direction of the turbomachine. In addition, the invention relates to a turbomachine, in particular an aircraft engine, and to a method for assembling a guide vane system.

Abstract: A blade for a turbomachine, in particular a compressor or turbine stage of a gas turbine, having at least one matrix having a first impact chamber (10) in which at least one impulse element (11) is disposed with play, at least one second impact chamber (20) whose volumetric centroid is offset from a volumetric centroid of the first impact chamber (10) along a first matrix axis (A) and in which at least one impulse element (21) is disposed with play, and at least one third impact chamber (30) whose volumetric centroid is offset from the volumetric centroid of the first impact chamber (10) along a second matrix axis (B) transversely to the first matrix axis (A) and in which at least one impulse element (31) is disposed with play, the first matrix axis (A) and an axis of rotation (R) of the turbomachine forming an angle of at least 60° and no more than 120° is provided.

Abstract: The invention relates to a guide means (10) for a gas turbine, in particular for an aircraft engine, having at least one casing element (12); having at least one first duct segment (14), which is arranged in the radial direction on the inside of the casing element (12), by means of which at least one duct (16), through which gas can flow, is at least partially delimited outward in the radial direction; having at least one second duct segment (26), which is arranged in the radial direction on the inside of the first duct segment (14), by means of which the duct (26) is at least partially delimited inward in the radial direction; and having at least one guide vane (28).

Abstract: A method for manufacturing a turbine casing part having a bearing chamber system including a shell (1) and at least one bearing receptacle (2) and further having a protective heat shield (30) that at least partially encompasses the shell (1) radially outwardly, the protective heat shield (30) being additively manufactured integrally with the bearing chamber system by selectively solidifying layer-by-layer a feedstock material (5) is provided.

Abstract: A seal carrier for a turbomachine, in particular a gas turbine, including a carrier base and at least one seal member, the at least one seal member being connected to the carrier base, and the at least one seal member being formed by a plurality of cavities arranged adjacent one another, in particular in a regular fashion, in the circumferential direction and the axial direction, the cavities extending from the carrier base in the radial direction, is provided. At least one stiffening element on the carrier base, the stiffening element extending along the circumferential direction and at least partially covering the at least one seal member at one of its axial end regions is provided.

Abstract: The invention relates to a device and a method for the additive manufacture of components through the layered bonding of powder particles to one another and/or to a semi-finished product or substrate already produced, using selective interaction of the powder particles with a high-energy beam, wherein, during the bonding of the powder particles into a layer made of powder particles with the aid of the high-energy beam, a gas flow, which has a flow direction having a directional component directed at least partially parallel to the layer of powder particles, is provided across the layer of powder particles and/or the bonding region in the layer of powder particles, wherein the directional component of the gas flow directed at least partially parallel to the layer of powder particles during the bonding of the powder particles in a layer is generated in at least two directions, which have oppositely directed directional components.

Abstract: A blade outer air seal (BOAS) for a gas turbine engine includes, among other things, a seal body having a radially inner face and a radially outer face that axially extend between a leading edge portion and a trailing edge portion. The BOAS includes a trough disposed on the radially inner face and an abradable seal received within the trough. The trough is open to expose a leading edge of the abradable seal to a core flow path of the gas turbine engine.