Abstract: Described is a blade for a high-speed turbine stage of an aircraft gas turbine, in particular of an aircraft engine, the blade including a radially inner blade root, a radially outer shroud, and an airfoil extending between the blade root and the shroud. It is provided that the outer shroud have only a single sealing element, which projects radially from the shroud, in particular only a single sealing fin.

Abstract: The invention relates to a wear-resistant shield for a rotating blade root of a rotating blade of a gas turbine, especially an aircraft gas turbine, having a base and two side walls connected to the base, wherein the side walls lie opposite each other and are shaped so as to be substantially complementary to an outer contour of a particular rotating blade root, and wherein the wear-resistant shield for this purpose is set up in such a way that, in an installed state, it is to be taken up between the respective rotating blade root and a rotating blade root mount of a rotor, especially between the respective rotating blade root and an axial securing element arranged in the rotating blade root mount.

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: 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.

Abstract: A guide vane segment 10 for a turbomachine includes a radially inner shroud plate 13 having a shroud plate surface 14 that is adapted to be configured in the turbomachine to face a rotor blade 20 adjacent to the guide vane segment, and thereby essentially extend along an outer conical surface K1 whose cone axis coincides with the axis of rotation A of a rotor shaft 30. In a radially inner region, a rotor blade 20 for a turbomachine has a base plate 23 having a base plate surface 24 that is adapted to be configured in the turbomachine to face a shroud of a guide vane row 10 adjacent to the rotor blade and thereby essentially extend along an outer conical surface K2 whose cone axis coincides with the axis of rotation A of a rotor shaft 30.

Abstract: A cooling system for cooling a turbine blade with a cooling fluid via an internal flow passage formed in the turbine blade extending from an inlet to an outlet edge having a first passage section defining a first flow direction, a second passage section defining a second flow direction, a wall between the first and second passage section and a diverter, between the first and the second passage section. The wall in a region of the diverter forms a pier head which extends into the region of the first passage section and thereby reduces the flow cross section of the flow passage.

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: The invention relates to a guide vane cluster for a turbomachine, in particular for a turbojet engine, having at least two vanes that are joined together by way of at least one common platform for the radial delimiting of a flow channel of the turbomachine, wherein each vane has a vane element with a suction side and a pressure side that are joined together by a leading edge and a trailing edge, wherein the at least two vanes have different trailing edge wall thicknesses at least in a corresponding radial lengthwise extension region.

Abstract: A shroud segment (100) for disposition on a blade of a turbomachine is provided, in particular a gas turbine, the shroud segment having a stiffening structure raised above a shroud segment surface (15), the stiffening structure including at least three interconnected ribs (3, 5, 7), and a first end portion of the ribs (3, 5, 7) being connected to an upstream sealing tip (11) of the shroud segment (100), and a second end portion of these ribs (3, 5, 7) being connected to a downstream sealing tip (13) of the shroud segment (100). The angles (W1, W2, W3) between the direction of the axis of rotation (a) of the blade and the longitudinal orientations (21, 23, 25) of the ribs (3, 5, 7), as viewed in the direction of flow (9) through the turbomachine, are between zero degrees and eighty degrees. The present invention also relates to a blade (200) of a turbomachine.

Abstract: The present invention relates to a rotating blade for a turbomachine, in particular a compressor stage or a turbine stage of a gas turbine, particularly of an aircraft engine, having a blade element for deflecting the flow, with a pressure side and a suction side, these sides being joined by a leading edge and a trailing edge, wherein a stacking axis of the blade element, in the radial direction over a radius r from a root of a blade element at r=0 to a tip of a blade element at r=H, has a course x(r) in a first downstream direction perpendicular to the radial direction and parallel to a principal axis of the turbomachine and has a course y(r) in a second direction perpendicular to the radial direction and to the first direction.

Abstract: A blade (10) for a turbomachine is provided. The blade (10) includes an airfoil (16) having at least one thickening (22) on a pressure side (20) of the airfoil (16). The thickening (22) has a region (B) of constant thickness (D). A turbomachine, a compressor and a turbine having stator vanes and/or rotor blades including at least one such blade (10) is also provided.

Abstract: The invention relates to a wear-resistant shield for a rotating blade root of a rotating blade of a gas turbine, especially an aircraft gas turbine, having a base and two side walls connected to the base, wherein the side walls lie opposite each other and are shaped so as to be substantially complementary to an outer contour of a particular rotating blade root, and wherein the wear-resistant shield for this purpose is set up in such a way that, in an installed state, it is to be taken up between the respective rotating blade root and a rotating blade root mount of a rotor, especially between the respective rotating blade root and an axial securing element arranged in the rotating blade root mount.

Abstract: The present invention relates to a turbomachine blade (100; 200) with a base element, which has a blade part (1) for flow diversion and a blade root (2), a first guide (110; 210), fixed on the base element, in which a first element (111; 211) is movably guided, and a second guide (120; 220), fixed on the base element, in which a second element (121; 221) is movably guided, wherein a dynamic of the first element in the first guide and a dynamic of the second element in the second guide are designed differently, and the first guide (110; 210) is arranged in a half, nearer to the base root, of a radial height (H) of the base element, and the second guide is arranged in a half, more remote from the blade root, of the radial height of the base element.

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.

Abstract: A blade cascade for a gas turbine, which has a first group with at least two, in particular at least three, first blade arrangements of the same type to one another, on which first element arrangements that are identical to one another and have at least one first element are movably mounted; and a second group with at least two, in particular at least three, second blade arrangements, which are of the same type to one another and are of different type from the first blade arrangements, on which second element arrangements, which are identical to one another and have at least one second element are movably mounted, wherein first and second element arrangements are different from each other.

Abstract: A blade cascade for a turbomachine, having a number of blades which include a monocrystalline material, each blade having a crystal orientation value, which is dependent on a crystal of first blades being less than a first limiting value and the crystal orientation values of second blades being at least equal to the first limiting value; and the blade cascade having at least one first sector, which includes at least three successive first blades, and having at least one second sector, which includes at least three successive second blades; and/or at least one first blade has a first body arrangement having at least one first body mounted so it is movable on the first blade and at least one second blade has a second body arrangement, which differs therefrom, having at least one second body mounted so it is movable on the second blade.

Abstract: A securing plate assortment including multiple different securing plates, the securing plates having different geometric codings. A turbine and method is also disclosed.

Abstract: The invention relates to a guide vane cluster (10) for a turbomachine, in particular for a turbojet engine, having at least two vanes (14) that are joined together by way of at least one common platform (12, 16) for the radial delimiting of a flow channel of the turbomachine, wherein each vane (14) has a vane element (17) with a suction side (18) and a pressure side (20) that are joined together by a leading edge (22) and a trailing edge (24), wherein the at least two vanes (14) have different trailing edge wall thicknesses at least in a corresponding radial lengthwise extension region.

Abstract: A guide vane segment 10 for a turbomachine includes a radially inner shroud plate 13 having a shroud plate surface 14 that is adapted to be configured in the turbomachine to face a rotor blade 20 adjacent to the guide vane segment, and thereby essentially extend along an outer conical surface K1 whose cone axis coincides with the axis of rotation A of a rotor shaft 30. In a radially inner region, a rotor blade 20 for a turbomachine has a base plate 23 having a base plate surface 24 that is adapted to be configured in the turbomachine to face a shroud of a guide vane row 10 adjacent to the rotor blade and thereby essentially extend along an outer conical surface K2 whose cone axis coincides with the axis of rotation A of a rotor shaft 30.

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.