Abstract: A rotor blade for a gas turbine, in particular an aircraft gas turbine, having a blade root element and a stream deflection portion adjoining the blade root element (12) in the longitudinal direction of the blade (RR); respective centroids (24) of blade cross-sectional areas of the stream deflection portion residing on a common stacking axis (26). It is provided that, starting from a first centroid (24) of a first blade cross-sectional area adjoining the blade root element (12), the stacking axis (26) extend within a cone (28) whose apex resides within the first centroid (24), and whose cone height (KH) extends orthogonally to the plane of the blade cross-sectional area; the angle (?) of the cone (28) being greater than 0° and smaller than or equal to 4°; preferably greater than or equal to 0.5° and smaller than or equal to 2°.

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: 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 balancing weight for a rotor blade of a turbine stage of a gas turbine, in particular of an aircraft gas turbine; including a first bent fastening portion that is couplable to an axial leading edge of a shroud of the rotor blade, a second fastening portion that is couplable to an axial trailing edge of the shroud, and a middle portion that joins the first fastening portion and the second fastening portion; the second fastening portion assumes a first position relative to the first fastening portion prior to a mounting of the balancing weight on the rotor blade, and a second position relative to the first fastening portion subsequently to the mounting of the balancing weight on the rotor blade; in the second relative position, the middle portion or/and the second fastening portion being deformed, in particular plastically deformed.

Abstract: A rotor blade for a gas turbine, in particular an aircraft gas turbine, having a blade root element and a stream deflection portion adjoining the blade root element (12) in the longitudinal direction of the blade (RR); respective centroids (24) of blade cross-sectional areas of the stream deflection portion residing on a common stacking axis (26). It is provided that, starting from a first centroid (24) of a first blade cross-sectional area adjoining the blade root element (12), the stacking axis (26) extend within a cone (28) whose apex resides within the first centroid (24), and whose cone height (KH) extends orthogonally to the plane of the blade cross-sectional area; the angle (?) of the cone (28) being greater than 0° and smaller than or equal to 4°; preferably greater than or equal to 0.5° and smaller than or equal to 2°.

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 rotor for a turbomachine, in particular for a jet engine, having a blade ring which includes multiple differently designed rotor blades (10a, 10b) having blade platforms (14a, 14b) engaged flush with one another, the blade ring including at least two groups of differently designed rotor blades (10a, 10b), each group of rotor blades (10a, 10b) being assigned blade platforms (14a, 14b), each of which is engageable flush with a matching blade platform (14a, 14b) of at least one other group of rotor blades (10a, 10b) and not with a blade platform (14a, 14b) of the same group of rotor blades (10a, 10b). A method for manufacturing a blade ring of a rotor for a turbomachine is also provided.

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 (10) for deflecting the flow, with a pressure side and a suction side (11), these sides being joined by a leading edge and a trailing edge (12, 13), wherein a stacking axis (S) of the blade element (10), in the radial direction (R) 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 (X) perpendicular to the radial direction (R) and parallel to a principal axis of the turbomachine and has a course y(r) in a second direction (Y) perpendicular to the radial direction (R) and to the first direction (X).

Abstract: A system for specifying an installation position of adjacent rotor blades of a blade row of a turbomachine is disclosed. The system has a plurality of axial securing elements, which have at least two sections with different profile areas for the arrangement in each case between a groove base of a rotor shaft and a blade root, which are joined together by a connecting web at a groove distance from one another, and which have counter-contours on the blade root side for forming in each case a positive-fit pair with the profile areas. Also disclosed are a securing element and a rotor blade for this type of system, a turbomachine having a rotor which has this type of system, and a method for manufacturing this type of rotor.

Abstract: A blade arrangement for a turbo engine, in particular a gas turbine, with a rotor and several blades fastened thereto, which are configured to be systematically different, wherein at least two adjacent blades have systematically different shrouds (12, 22) and/or inner blade platforms (11, 21).