Abstract: A roller bearing, particularly a needle bearing, for arranging on a pivot pin of a variable guide vane of a turbomachine, is disclosed. The roller bearing includes at least one radially elastic needle which has two opposing end regions and at least one roller region arranged between the end regions and constructed in a cross-sectionally thickened manner. The end regions roll off at associated raceways of a bearing housing of the roller bearing and the roller region is arranged in the region of an associated groove of the bearing housing. A variable guide vane with at least one such roller bearing, a housing of a turbomachine with such a guide vane, as well as a turbomachine, is also disclosed.

Abstract: A guide blade for a gas turbine is disclosed. The guide blade includes a blade leaf having a receptacle in which at least one sealing element is arranged, where the sealing element is movable relative to the blade leaf between a sealing setting, in which the sealing element is at least partially moved out of the receptacle, and a storage setting, in which the sealing element is moved back into the receptacle. The guide blade further includes at least one fluid channel by which fluid under pressure can be routed into the receptacle in order to move the sealing element from the storage setting into the sealing setting. An inlet opening of the fluid channel is formed on a pressure-side surface of the blade leaf. A housing as well as a gas turbine having at least one guide blade is also disclosed.

Abstract: A housing for a gas turbine is disclosed. At least one wall element is housed on the housing so as to move, which limits a flow channel of the housing in the radial direction from a rotational axis of a rotor of the gas turbine toward the exterior. The housing includes at least one variably adjustable guide blade which extends through the wall element into the flow channel. The wall element can be moved between a sealing setting, in which the wall element makes contact at least in a partial area of a side of a blade leaf of the guide blade facing toward the wall element, and an open setting, in which the blade leaf and the wall element are spaced some distance apart from one another. A gas turbine as well as a process for operating a gas turbine is also disclosed.

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); 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 first duct segment, 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 (16), which is arranged in the radial direction on the inside of the first duct segment (14), by means of which second duct segment, the duct (16) is at least partially delimited inward in the radial direction; and having at least one guide vane (28).

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 safety device (10) for a bearing arrangement of a rotor of a turbomachine, whereby the safety device (10) includes at least two support structures (12) between which at least one buckling structure (14) is arranged that is configured to collapse when a predetermined buckling load that acts on at least one of the support structures (12) is exceeded, thereby reducing the volume of the safety device (10). A method for the production of a safety device (10) for a bearing arrangement of a rotor of a turbomachine, as well as to a bearing arrangement of a rotor of a turbomachine having such a safety device (10).

Abstract: A guide vane (300) for a turbomachine, having a sealing device (27, 27?) at the radially inner end region of the guide vane (300) for sealing leakage flows (25) between the guide vane (300) and an inner ring (7) joined thereto. The sealing device (27, 27?) is movably configured relative to the guide vane (300). The sealing device (27, 27?) is positionable in at least one open or in a closed configuration for sealing the leakage flows (25). Also, a guide vane (100), as well as a turbomachine.

Abstract: A turbine or compressor stage of a gas turbine, the rotor blade having a radially outer shroud (1) which has a sealing fin array having a first sealing fin (3.1) and a second sealing fin (3.2) which is adjacent to the first sealing fin and connected thereto by a first groove base (10) having a circumferential region (13) of maximum radial height, which is located at a first circumferential position is provided. The sealing fin array has a third sealing fin (3.3) adjacent to the second sealing fin and opposite to the first sealing fin, the third sealing fin being connected to the second sealing fin by a second groove base (20) having a circumferential region (23) of maximum radial height, which is located at a second circumferential position different from the first circumferential position.

Abstract: The invention relates to a housing system (10) for an axial turbomachine, in particular for an aircraft turbine, comprising a housing (12), at least one guide blade (14), which is fastened to the housing (12) by means of at least one fastening element (18),a scaling element (20) for reducing a leakage flow between the listening element (18) and the housing (12), wherein the sealing element (20) is arranged between a downstream wall (22) of the fastening element (18) with respect to a flow direction (I) of the axial turbomachine and a wall (24) of the housing (12), and a pressing means (36), by means of which an axial force can be applied to the scaling element (20) And the scaling element can be pressed against the wall (24) of the housing (12), at least during the operation of the axial turbomachine. The invention further relates to an axial turbomachine having such a housing system (10).

Abstract: A turbomachine having a cooling device for supplying cooling air onto a compressor region from an air distribution chamber is disclosed, the air distribution chamber being arranged between an outlet diffuser of a compressor and a combustion chamber, where the cooling device has cooling air pipes for supplying cooling air from the air distribution chamber into a cavity between the outlet diffuser and a rotor segment of the compressor.

Abstract: A gas turbine blade in an associated receiving recess of a rotor base body includes a slot in the blade root and has the shape of a straight cylinder. A blade retaining plate has a first securing region and a second securing region interconnected by a shaft region; the shaft region having the shape of a straight cylinder configured to conform to the slot of the blade root. The shaft region comes to rest on a bottom of the receiving recess of the rotor base body; the first securing region being reshaped to form a first limit stop for a first end wall of the blade root; and the second securing region being reshaped to allow the blade root to be inserted along the insertion direction into the receiving recess. The second securing region of the blade retaining plate is subsequently reshaped to allow the second securing region to form a second limit stop for a second end wall of the blade root facing opposite the first end wall, and the blade root to be secured in position in the receiving recess.

Abstract: A blade, in particular a rotor blade or a stator vane, for a gas turbomachine, in particular a turbojet engine, the blade having an airfoil (1) for deflecting a flow of working fluid and a first platform (3) connected thereto, in particular integrally connected thereto, to radially bound a flow duct for the working fluid, the airfoil having a suction side and a pressure side (1.1) which are connected at a leading edge (1.2) and at a trailing edge (1.3). The trailing edge has a first minimum wall thickness in a first region (A) of a radial longitudinal extent (R) of the airfoil proximal to the first platform (3), and a maximum wall thickness that is smaller than the first minimum wall thickness in a platform-distal region (C) of the radial longitudinal extent.

Abstract: A turbine stage having a gas channel in which at least one rotor assembly is disposed, is disclosed. A cavity communicates with the gas channel and is delimited by a front side of a rotor element and by a rotor element-mounted seal. A blow-out arrangement with at least one gas passage for blowing out a sealing gas flow into the cavity is provided, where the gas passage for blowing out the sealing gas flow is configured with a swirl in the circumferential direction and has an outlet opening that is offset radially outwardly from the rotor element-mounted seal.

Abstract: A turbine blade (100) having an airfoil (1). In its radially outer region, the turbine blade (100) has at least one shroud (200) or at least one shroud segment, as well as at least one cutting tooth (300) which is intended to engage at least once into portions of a turbine casing during use of the turbine blade (100). The cutting tooth (300) is connected to the shroud (200) or the shroud segment.

Abstract: The invention concerns a moving blade system for a moving blade sequence of a continuous-flow machine, in particular a thermal gas turbine with at least two moving blade segments, of which each moving blade segment encompasses at least two moving blades that are disposed, in reference to a rotation axis of the moving blade sequence, at least primarily radially between a radially interior and a radially exterior shroud band and are coupled to the shroud bands, wherein at least the radially exterior shroud bands of the two moving blade segments encompass contact surfaces that correspond to one another, wherein radial equilibrium axes of neighboring moving blades of the at least two moving blade segments are disposed in a tilted manner by respectively axially an angle (?a, ?b) relative to their respective base inclination, wherein the angles (?a, ?b) in regard to the base inclination feature opposite signs.

Abstract: A turbomachine, especially a gas turbine, includes a rotor having rotating blades and a stator having a housing and guide blades. The rotating blades form at least one rotating blade ring, which at one radially outward lying end adjoins an inner ring or casing ring of the housing, thereby defining a gap therebetween. The casing ring is connected to a support ring via curved walls, which together with the casing ring and the support ring bound a cavity and form a bellowslike structure. By changing the pressure prevailing in the cavity of the respective bellowslike structure, the gap between the casing ring and the radially outward lying ends of the respective rotating blade ring can be pneumatically adjusted.

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: The present invention relates to a strut for a turbine center frame of a jet engine, in particular a gas turbine, with a first end for disposition on an internal structure of the turbine center frame and a second end opposite the first end for attaching the strut on a housing or an internal boundary wall of the turbine center frame, and on the second end at least one flange-like element with at least one first aperture is removably arranged with at least one first aperture for accommodating or passing through an attachment means. The invention further relates to a turbine center frame for a jet engine, in particular a gas turbine as well as a method for producing a turbine center frame of a jet engine, in particular a gas turbine.

Abstract: A moving blade for a turbomachine, in particular an aircraft engine, is disclosed, having an inner shroud which has a front elongation for forming an axial overlap with an upstream guide blade, and on which at least one flow guide element for deflecting a leakage flow of a cooling air flow in the peripheral direction is situated. The at least one flow guide element is guided beyond a leading edge of the elongation. A turbomachine having a plurality of these types of moving blades is also disclosed.

Abstract: A sealing system for a turbomachine, in particular, for a gas turbine, having an annular space between a flow-restricting wall and at least one series of rotor blades including a plurality of rotor blades. The sealing system comprises at least one first sealing fin disposed on an end of the rotor blade facing the wall, at least one second abradable lining disposed following the first sealing fin in the flow direction on the end of the rotor blade facing the wall, at least one first abradable lining disposed on the inside of the wall and opposite the first sealing fin, and at least one second sealing fin disposed in the flow direction following the first abradable lining on the inside of the wall and opposite the second abradable lining. The invention furthermore relates to a gas turbine comprising at least one sealing system.