Abstract: The present invention relates to a turbine module for a turbomachine, comprising a first flow bypass structure and a second flow bypass structure, said flow bypass structures being arranged in a hot gas duct, which is bounded by the turbine module and is designed to convey a hot gas and, namely, being arranged in succession in relation to a longitudinal axis of the turbine module in a direction of rotation, wherein, in relation to the bypass flow in the hot gas duct, the flow bypass structures each have a leading edge, and, downstream thereto, a trailing edge, and the second flow bypass structure is provided as a deflecting blade, wherein the second flow bypass structure has a smaller profile thickness than the first flow bypass structure, and wherein the hot gas duct is enclosed by a radial width.

Abstract: The present invention relates to a method for smoothing a surface of a component, in which a component is placed in a liquid-solids mixture; a relative movement is produced between the liquid-solids mixture and the component; thus there is a flow of the liquid-solids mixture along the surface; wherein there is provided in the liquid-solids mixture a guide surface, along which the liquid-solids mixture flows, wherein a directional component toward the surface is imposed on the flow.

Abstract: The present invention relates to a deflector for a gas turbine, wherein the deflector is integrally made by a nut portion having an internal thread and a deflector portion extending radially outwards from the nut portion. The present invention also relates to a deflector assembly comprising such a deflector, and to a turbine center frame comprising such a deflector assembly.

Abstract: A sealing assembly for a gas turbine is provided, especially an aircraft gas turbine, including: a first component having a first flange portion and a second component having a second flange portion; in the area of the first flange portion and the second flange portion, the first component and the second component being detachably intercoupled by a bolt connection; between the first component, especially a rounded portion, and the second component, especially the second flange portion, a variable interspace being provided that is at least partially sealed by at least one sealing element. The sealing element has a first sealing portion that rests against the first component, especially against the rounded portion, and a second sealing portion that rests against the second component, especially against the second flange portion; the first sealing portion and the second sealing portion being interconnected and extending obliquely relative to each other.

Abstract: A method for producing a forged component from a TiAl alloy is provided, in particular a turbine blade (10), in which method a blank of a TiAl alloy is provided and deformed by forging into a forged, semi-finished part (9). A usable volume is defined within the forged, semi-finished part, the usable volume corresponding to the forged component to be produced. The shape of the blank is selected such that within the usable volume of the forged, semi-finished part, the degree of deformation resulting from forging deviates by no more than ±1 from a defined value.

Abstract: A component joint including a turbomachine component (10) having at least one component flange (11); a mating turbomachine component (20) connected thereto and having at least one mating component flange (21); and at least one seal (30) to provide a seal against an overpressure (p2?p1) on the component side, the at least one seal (30) having a sealing surface (31) that engages into a groove (12) of the turbomachine component (10), the groove (12) being disposed on a mating-component-flange-facing side of the component flange (11) and having a contact surface (13) for sealingly axially supporting the sealing surface (31).

Abstract: The invention relates to a guide vane segment for aircraft gas turbine, which extend along a respective circular arc and together form an annular section, wherein, in the radial direction between the outer shroud and the inner shroud, a plurality of guide vanes are arranged adjacently in the peripheral direction, the outer shroud and the two sealing walls form a trough-like profile in longitudinal section, wherein, at the axially front or/and rear sealing wall element(s), at least one relief gap with a main section is provided, which extends substantially radially inward, starting from a radial outer edge of the respective front or/and rear sealing wall element(s) along the sealing wall element. According to the invention, it is proposed that the relief gap has at least one additional section that adjoins the main section radially inward, wherein the additional section is formed by at least one curved subsection.

Abstract: The present invention relates to a guide vane cascade for a turbomachine, which has guide vanes that are mounted adjustably about an axis of rotation, so as to change an inflow angle each time, wherein a first guide vane and a second guide vane arranged on the pressure side thereof, referred to a longitudinal axis of the turbomachine, are arranged with an axial offset, namely, the axis of rotation of the second guide vane is offset axially toward the back, wherein the first guide vane and the second guide vane are provided such that, in an adjusted state, they form together a tandem configuration in a radially outer region but together they delimit a divergent channel in a radially inner region.

Abstract: A method for operating an internal combustion engine including a step of concurrently introducing at least two combustible fuel jets into a combustion chamber of an internal combustion engine. A first combustible fuel jet of the at least two combustible fuel jets is ignited at an ignition time point. In a first operating mode of the internal combustion engine a second combustible fuel jet which is different from the first combustible fuel jet of the at least two combustible fuel jets is ignited after the ignition time point.

Abstract: The present invention relates to a blade arrangement for a turbomachine, in particular a gas turbine, with a first blade, which has a first blade body and a first platform, and a second blade, which is adjacent in the peripheral direction, and has a second blade body and a second platform, wherein a first wall of the first blade and a second wall of the second blade bound a blade cavity, in which a damper with a wall-side contact surface is arranged, wherein this contact surface has at least one first surface portion, which is convexly curved in a first direction, which, in at least one contact position, contacts the first wall in the first surface portion and is parallel to at least one portion of an edge of the first platform, said edge facing the second platform.

Abstract: A fastening assembly having a component flange (10), in particular of a turbomachine, having at least one bore (11) in which a sleeve (30) is placed that extends through a bolt (20) which has a head (21) that is axially secured between a rim (12) of the bore and a component flange-mounted axial limit stop (13).

Abstract: The present invention relates to a method for producing a component of a turbomachine from a metal alloy as well as a correspondingly produced component, wherein the method includes defining at least one first component region that will have a first property profile, and at least one second component region that will have a second property profile which is different from the first property profile; providing at least one powder of the metal alloy or several different powders of constituents of the metal alloy; additive manufacture of the component from the at least one powder, wherein the powder is melted for cohesive joining of the powder particles to each other and to a substrate or to an already produced part of the component.

Abstract: A variable vane cascade for a turbomachine, in particular for a compressor stage or turbine stage of a gas turbine, having at least one first vane, in particular guide vane that has a first distance from a circumferentially adjacent vane, at least one second vane, in particular guide vane that has at least one second distance from at least one circumferentially adjacent vane that is smaller than the first distance, and an actuating device, in particular for jointly and/or reversibly adjusting the first and second vane from a first position where at least one airfoil cross section of the first vane and an airfoil cross section of the second vane each have a first stagger angle, into a second position where these airfoil cross sections have second stagger angles, the second stagger angle of the first vane differing from the second stagger angle of the second vane, in particular being larger than the second stagger angle of the second vane.

Abstract: The invention is directed to a guide vane ring for a turbomachine, in which a vane bearing is produced in the inner ring via platform plates of the vanes and reliability against disintegration is produced via journals that extend radially inward from the platform plates, as well as a turbomachine.

Abstract: A rotor member is described for a gas turbine that is adapted for rotating about a central axis, the rotor member being a blisk having a rotor blade row that extends around the central axis or a rotor disk having a mounting portion for installing rotor blades of a rotor blade row that extends around the central axis, and being axially offset from the rotor blade row and/or the mounting portion and, extending coaxially, having at least one annular and axially asymmetrical sealing fin that has a radially outer tip portion having a front flank facing the rotor blade row and/or the mounting portion, and an opposite flank facing away from the rotor blade row and/or the mounting portion; the front flank being less steep than the opposite flank; a turbine and a compressor having such a rotor member, and a method for manufacturing such a rotor member having at least one sealing fin coating.

Abstract: Disclosed is an airfoil array segment (110, 120, 130, 140, 150) of an airfoil array for a turbomachine, the airfoil array segment including a platform (10) having a platform surface, as well as at least two airfoils (20, 30). The platform surface has a depression (111, 121, 131, 141, 151) which extends up to the first airfoil and contacts the pressure side (21) of the first airfoil downstream of 80% of the axial chord (g) downstream of the leading edges (23, 33) and which contacts the pressure side (21) of the first airfoil downstream up to no more than 80% of the axial chord (g) downstream of the leading edges (23, 33). At least one lowest point (112, 122) of the depression is located at least 90% of the axial chord (g) downstream of the leading edges (23, 33). Also disclosed are an airfoil array, an airfoil passage, a platform, a turbomachine, and an aircraft engine.

Abstract: A rotor blade airfoil (25) for a turbomachine (1) that is adapted for rotation about a longitudinal axis (2) of the turbomachine (1) is provided. The rotor blade airfoil (25) is built from an airfoil material reinforced with a fibrous material (30). At least a portion of the fibers of the fibrous material (30), in particular at least 20%, preferably at least 30% of the fibers, are oriented in a first fiber direction (31), and the first fiber direction (31) is tilted with respect to the stacking axis of the rotor blade airfoil (25).

Abstract: Disclosed is a static seal arrangement for sealing a gap between a first static component and a second static component of a turbomachine such as an aircraft engine or a static gas turbine, having a brush seal held distant from a hot gas, this brush seal being arranged standing or vertical with its clamping section that can be surrounded by cooling air in a radially outer clamping space relative to the gap, and forming a radial sealing site with the first component and an axial sealing site with the second component; also disclosed is a turbomachine.

Abstract: The invention relates to a method for additive production of a component layer of a component including the steps of: generating at least one layer from a powdery component material in the region of a structuring and joining zone; subdividing model data of the layer into virtual sub-regions by a control device; selecting at least one of the virtual sub-regions by the control device; localized heating of at least one heating region in a real sub-region of the layer corresponding with the selected virtual sub-region by a heating device; verifying whether a temperature of the layer has, in a predetermined inspection region, a predetermined minimum temperature; and localized solidifying of the layer in a predetermined solidifying region by selective irradiation by at least one energy beam of an energy source, if the layer has the predetermined minimum temperature in the inspection region.

Abstract: An impulse body module a turbine stage of a gas turbine includes a holder component of one-piece construction, with a base and surrounding side walls arranged at the base, wherein the side walls and the base define a holding chamber; an insert component, of one-piece construction, which is inserted into the holding chamber of the holder component, wherein the holder component and the insert component accommodated therein define a number of mutually separated cavities, and wherein an impulse body, in particular a sphere, is accommodated in each cavity; and a closure component of one-piece construction, which is joined to the holder component in a material-bonded manner where the holding chamber is closed and the insert component is surrounded by the holder component and the closure component. The holder component has at least one base projection in the region of its base, which extends away from the holding chamber.