Abstract: A turbomachine blade system, in particular for a compressor or turbine stage of a gas turbine, which includes at least one blade, in particular a moving or guide blade, and at least one moving body for reducing the vibrations of this blade, at least one area of a guide for guiding the body and/or at least one area of a supporting structure for resiliently mounting the body and/or at least one area of the body being or becoming generatively manufactured together with at least one area of the blade, in particular of a vane and/or blade root and/or a shroud situated thereon.

Abstract: A gas turbine component for a gas turbine is provided. The gas turbine component has a shell portion having an abradable portion (2) to allow rotor blade ends (1) of a rotor blade array of the gas turbine to rub into the same. The abradable portion adjoins a neighboring portion (3) of the shell portion axially at both sides and is additively manufactured together therewith.

Abstract: The segmented seal carrier of a guide vane ring of a turbomachine has a number of segments that form two half-rings of the seal carrier when assembled. They are clamped together by their opposite-lying, corresponding lateral contact regions.

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: The invention relates to a guide vane ring with at least one guide vane ring segment for a turbomachine, including a plurality of guide vanes arranged radially around an axis of rotation, an outer shroud arranged radially on the outer side of the guide vanes, and an inner shroud arranged radially on the inner side of the guide vanes, wherein the outer shroud or the inner shroud has at least one expansion joint, wherein the guide vane ring segment is formed in one piece with the guide vanes, the outer shroud, and the inner shroud and the at least one expansion joint is sealed in each case by at least one sealing device arranged in the respective expansion joint.

Abstract: A blade of a turbomachine is provided, including at least one cooling channel in the interior of the blade for cooling the blade with the aid of a fluid flowing through the cooling channel, the cooling channel having at least one inlet and at least one outlet, between which the cooling channel extends along its longitudinal axis, and the cooling channel being radially delimited by at least one wall, at least one displacement body being situated in the cooling channel, so that an annular or tubular gap between the displacement body and the wall of the cooling channel results in the area of the displacement body/bodies, which is available for the through-flow of the fluid, or at least two or multiple subchannels being formed in the area of the displacement body/bodies. The invention also relates to a method for manufacturing a corresponding blade.

Abstract: The present invention relates to sealing element for sealing a radial gap relative to a counter element of a turbomachine, wherein the sealing element has a number of adjacent cells in the peripheral direction and/or in the axial direction, which are joined to one another by common walls, wherein an extension of a least one, in particular front-side, cross section of at least one cell in a first axial section of the sealing element in the peripheral direction and/or in the axial direction for sealing against a radial flange of the counter element is smaller than that in a second axial section of the sealing element that adjoins the first axial section upstream and/or in a second axial section that adjoins the first axial section downstream.

Abstract: A shaft seal system which is disposed axially between a first chamber and a second chamber is disclosed. The shaft seal system includes a shaft, a casing that surrounds the shaft, and a seal which is disposed axially closer to the second chamber where the first chamber has a fluid and the second chamber is to be protected from the fluid. In this arrangement, the seal includes a pressure booster.

Abstract: A turbomachine blade system, in particular for a compressor or turbine stage of a gas turbine, which includes at least one blade, in particular a moving or guide blade, and at least one moving body for reducing the vibrations of this blade, at least one area of a guide for guiding the body and/or at least one area of a supporting structure for resiliently mounting the body and/or at least one area of the body being or becoming generatively manufactured together with at least one area of the blade, in particular of a vane and/or blade root and/or a shroud situated thereon.

Abstract: The segmented seal carrier of a guide vane ring of a turbomachine has a number of segments that form two half-rings of the seal carrier when assembled. They are clamped together by their opposite-lying, corresponding lateral contact regions.

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 gas turbine component for a gas turbine is provided. The gas turbine component has a shell portion having an abradable portion (2) to allow rotor blade ends (1) of a rotor blade array of the gas turbine to rub into the same. The abradable portion adjoins a neighboring portion (3) of the shell portion axially at both sides and is additively manufactured together therewith.

Abstract: The invention relates to a guide vane ring (10) with at least one guide vane ring segment for a turbomachine, including a plurality of guide vanes (12) arranged radially around an axis of rotation, an outer shroud (14) arranged radially on the outer side of the guide vanes (12), and an inner shroud (16) arranged radially on the inner side of the guide vanes (12), wherein the outer shroud (14) or the inner shroud (16) has at least one expansion joint (18), wherein the guide vane ring segment is formed in one piece with the guide vanes (12), the outer shroud (14), and the inner shroud (16) and the at least one expansion joint (18) is sealed in each case by at least one sealing device (24) arranged in the respective expansion joint (18).

Abstract: A shaft seal system which is disposed axially between a first chamber and a second chamber is disclosed. The shaft seal system includes a shaft, a casing that surrounds the shaft, and a seal which is disposed axially closer to the second chamber where the first chamber has a fluid and the second chamber is to be protected from the fluid. In this arrangement, the seal includes a pressure booster.

Abstract: A turbine engine and a rotor for a turbine engine are provided. Several adjacent turbine blades are positioned in the circumferential direction of the rotor, each turbine blade having a blade root, each turbine being fixed in a retainer groove, running in the circumferential direction of the rotor, by means of the blade root. Each turbine blade may be introduced, by the blade roots thereof, into the retaining groove, by means of a filling groove, whereby the width of the filling groove is matched is matched to the width of the blade roots. The width of the blade roots (13, 14) and the width of the, or each filling groove (17) in the circumferential direction is greater than half the width of a desired nominal blade pitch (18), whereby in the region of the, or each filling groove (17), a first number of turbine blades (11) with desired nominal blade pitch (18), is exchanged for a second number of turbine blades (12), with increased blade pitch (19), the first number being greater than the second number.

Abstract: A turbine engine and a rotor for a turbine engine are provided. Several adjacent turbine blades are positioned in the circumferential direction of the rotor, each turbine blade having a blade root, each turbine blade being fixed in a retainer groove, running in the circumferential direction of the rotor, by means of the blade root. Each turbine blade may be introduced, by the blade roots thereof, into the retaining groove, by means of a filling groove, whereby the width of the filling groove is matched is matched to the width of the blade roots. The width of the blade roots (13, 14) and the width of the, or each filling groove (17) in the circumferential direction is greater than half the width of a desired nominal blade pitch (18), whereby in the region of the, or each filling groove (17), a first number of turbine blades (11) with desired nominal blade pitch (18), is exchanged for a second number of turbine blades (12), with increased blade pitch (19), the first number being greater than the second number.

Abstract: A nozzle ring for a gas turbine, more particularly for an aircraft engine, has at least one shroud with a circumferential surface and at least one blade with a surface. The shroud has at least one opening for fastening the blade. The blade has at least one end section with a platform which projects at least partially over its circumference and which has a transition curve and which is inserted into the opening.