Abstract: The invention relates to an axially divided inner ring as well as a variable vane cascade for an aircraft engine. The inner ring and the vane cascade improve the efficiency of the aircraft engine in that differently formed vanes and/or vane mounts are formed respectively in the inner ring or in the vane cascade.

Abstract: The invention relates to a method machining a particularly planar component by means of electrochemical machining, wherein the component has internal stresses resulting particularly from preceding manufacturing steps. In a first step a) of the method, the component to be machined is provided. Subsequently, in step b), at least two tools are provided in the form of electrodes and, in step c), an electrolyte is provided between the component and the at least two electrodes. In step d), a positive voltage is applied to the component and a negative voltage is applied to the at least two electrodes. Thus, in step e), by moving the at least two electrodes along their respective movement paths with respect to the component, electrochemical machining can take place; in the process, the gap between each electrode and the component is flushed with the electrolyte at least intermittently.

Abstract: The invention relates to an air conduction system for an aircraft engine with a primary flow channel and a branch channel, in which, in particular, the branch channel can be a secondary flow channel and/or a bleed air channel, wherein the branch channel branches off the primary flow channel at a connecting region, comprising a closing device arranged at or in the connecting region for separating the branch channel from the primary flow channel, wherein the closing device has at least one closing ring for closing the connecting region.

Abstract: A connecting structure (10) for load transfer, in particular in a gas turbine (1), including a strut (20) and at least one wall element (30) is provided. The strut (20) at one end is integrally joined to the wall element (30), and the strut (20) and the wall element (30) are enclosed by a fillet (40), at least in areas, and integrally joined to same. An elastic deformation of the involved elements of the structure during the load transfer and/or load absorption is improved in that a root section (50) that is formed by a ridge (56) and that extends from the strut (20) to the wall element (30) is situated on the fillet (40).

Abstract: The invention relates to a rotor shaft for an electric motor, in particular for an aircraft engine, wherein the rotor shaft has a forged attachment flange, wherein the attachment flange is provided for attachment to another shaft in order to transmit a force and/or a torque, comprises a first axial end facing the other shaft, and comprises a base plate on a second axial end opposite the first axial end. It is provided that the rotor shaft has a coolant distribution body that is additively manufactured at least in some regions, wherein the additively manufactured coolant distribution body is arranged on the base plate of the forged attachment flange in a radially centered manner, and the rotor shaft has a rotor device, which radially surrounds the additively manufactured coolant distribution body and is joined to the base plate of the forged attachment flange at least in a force-fitting manner.

Abstract: The present invention relates to a method for imaging a surface in an internal space of an axial turbomachine, wherein the surface is captured in a sequence using an imaging device, i.e.—only a portion is captured with the imaging device at a respective time, thus generating only a partial image of the surface; the imaging device and the surface are moved relative to each other such that different portions of the surface are captured over the course of time, thus generating different partial images, and wherein the partial images are stitched together to produce an image of the surface.

Abstract: A blisk has at least one blade, with a blade airfoil and a blade root; a platform, the blade being integrally attached to the platform; and a fillet arranged on the blade root and between the blade airfoil and the platform. The fillet merges into the blade at a blade connection. The fillet merges into the platform at a platform connection. The fillet extends with a longitudinal extent around the blade root and a transverse extent from the platform to the blade airfoil. The fillet has a variable radius along the transverse extent. The variable radius has, at least in a first section of the fillet, a minimum radius which, at least in the first section, is spaced apart from the platform by at least 15% of the transverse extent along the transverse extent of the fillet.

Abstract: A blade has an internal blade profile, and a jet deflection portion adjoining the internal blade profile. Each blade profile of the jet deflection portion has a center of gravity. The centers of gravity are joined by a thread line, which is a space line, starting from the internal blade profile. A total deviation of the space line in relation to the reference line is described by an nth order polynomial based on radial distance from the blade profile to a reference point. The thread line is selected and designed such that a compressive stress is produced in the region of the leading edge of the blade during operation.

Abstract: An adjustment lever adjusts a stator vane of a turbomachine. The adjustment lever has: a first connection site, of a plurality of connection sites, the first connection site being configured to join to an adjustment ring; a second connection site, of the connection sites, the second connection stie being configured to join to the stator vane; and a joining member arranged between the first connection site and the second connection site. The joining member is shaped having at least two struts which adjoin at least one of the connection sites.

Abstract: The invention relates to a blisk segment for a gas turbine, comprising at least one first blade having an airfoil, a leading edge, a trailing edge, a blade root, a suction side and a pressure side, a pedestal, and a first fillet having a device for influencing crack growth. The robustness of the gas turbine is improved in accordance with the invention in that, at the leading edge of the first blade, at least on the first fillet, a first surface structure is arranged and interacts with the crack-influencing device for influencing crack growth and for flow influencing.

Abstract: The invention relates to a rotor for integration in a turbomachine, with a rotor base body, wherein, on the rotor base body in a rotor outer space of the rotor, at least one blade element or airfoil is mounted, wherein the rotor base body extends in an axial direction of the rotor and comprises a connecting section, which extends in an axial direction of the rotor for connection to at least one further rotor and separates the rotor outer space from a rotor inner space of the rotor, wherein the connecting section comprises, in the axial direction, at a front side, preferably in the axial direction at a downstream-lying front side, a balancing flange, wherein the balancing flange is configured for compensating an imbalance of the rotor and extends in the radial direction in the rotor outer space, at least in sections.

Abstract: The present invention relates to a module for a turbomachine, with a guide vane arrangement, a seal carrier, which is arranged radially inside an inner platform of the guide vane arrangement, seal carrier walls, namely, a first seal carrier wall and a second seal carrier wall, and a sliding member as well as a connecting element, wherein the first seal carrier wall and the seal carrier are formed in a one-piece manner with each other, and wherein the seal carrier walls are multipiece in construction relative to each other, and the second seal carrier wall is fastened to the first seal carrier wall and the seal carrier by way of the connecting element, wherein the sliding member holds the seal carrier walls spaced apart from each other in such a manner that they bound axially an intervening space with each other.

Abstract: The present invention relates to a rotor blade (20) for arrangement in a gas duct (2) of a turbomachine (1), having a rotor blade airfoil (23), which, viewed in a tangential section, has a blade airfoil profile (24) with a leading edge radius RVK and a rotor blade airfoil thickness d, wherein the blade airfoil profile (24) is thickened, at least in sections, specifically the blade airfoil thickness d is specified, in relation to the front edge radius RVK, such that (2d/Rvk2)?d?5.5.

Abstract: A rotor for a turbomachine has at least one blade and has at least one rotor main part, which has at least one recess, in which a blade root of the least one blade is interlockingly received, wherein the blade root comprises at least one depression, in which at least one protrusion of the at least one rotor main part, which protrusion delimits the at least one recess in regions is received, wherein the at least one depression is delimited by a first delimiting face on the blade root side and the at least one protrusion is delimited by a second delimiting face on the rotor main part side. At least the first delimiting face has at least one elevation which narrows a gap at least in regions, which extends between the first delimiting face and the second delimiting face.

Abstract: A rotor blade (20) for a rotor blade assembly (10) of a turbomachine (1) is provided, having an inner rotor blade platform (40) which extends axially from the rotor blade (20) with respect to a longitudinal turbomachine axis (2) and has two opposite circumferential end faces (41) and a free axial end (42) whose cross section is radially inwardly and radially outwardly bounded by circular arcs of two concentric circles. The inner rotor blade platform (40) has a cross section of connection (45) with the rotor blade (20) which is bounded radially inwardly by an inner connecting line (46) and radially outwardly by an outer connecting line (47). Each of the connecting lines (46, 47) has a central portion (61) having a convex curvature.

Abstract: A stator vane for a turbomachine, including an airfoil having a leading edge and a trailing edge, which are interconnected by a suction side and a pressure side, the airfoil extending substantially in a radial direction between an inner platform and an outer platform, wherein, when viewed in an axial direction of the turbomachine toward the trailing edge, the trailing edge extends so as to be inclined at a first angle to the pressure side radially on an outside in relation to the radial direction at a circumferential position at which the trailing edge meets the outer platform, and wherein the first angle between the trailing edge and a tangent to the outer platform at a transition from the outer platform to the trailing edge is between 72° and 84°.

Abstract: The guide vane assembly as well as to a method for mounting a guide vane assembly of a turbomachine, including a number of adjustable guide vanes, provides a guide vane platform of which has a guide vane journal that is mounted in a receiving opening of a housing of the turbomachine, wherein a guide vane head of the guide vane has a bearing pin, which is mounted on an inner ring arranged on a rotor of the turbomachine. In the method, a guide vane journal, which is arranged on a guide vane platform of a guide vane, is brought into a receiving opening of a housing, which is arranged radially with respect to a rotor axis of the turbomachine, and an inner ring is provided on a rotor of the turbomachine.

Abstract: The invention relates to a compressor for an engine, wherein the compressor has compressor stages arranged in succession in a flow direction of the compressor and each compressor stage has a rotating blade cascade and a guide vane cascade arranged downstream of the rotating blade cascade and the rotating blade cascade and the guide vane cascade each have an aspect ratio.

Abstract: A turbine module (2) for a turbomachine (1). The turbine module (2) includes a main channel (26) to guide a main flow (36) through the turbine module (2), a rotor blade (21) and a stator vane (22), the stator vane (22) including a stator airfoil (22) and a platform (23), with the stator airfoil (22) arranged downstream of the rotor blade (21) in the main channel (26), and a cavity (30) including an inlet (31) for injecting a part (36.2) of the main flow (36) into the cavity (30), an outlet (32) for a reinjection of the part (36.2) of the main flow (36) from the cavity (30) into the main channel (26), wherein the cavity (30) is arranged at an axial position of the stator vane (20) and is radially offset from the stator airfoil (22).

Abstract: An aircraft system includes, among other things, an aircraft and a gas turbine engine coupled to the aircraft. The gas turbine engine includes a propulsor section including a propulsor, a compressor section, a turbine section including a first turbine and a second turbine, and a gear reduction between the propulsor and the second turbine. The second turbine includes a number of turbine blades in each of a plurality of rows of the second turbine. The second turbine blades operating at least some of the time at a rotational speed. The number of blades and the rotational speed being such that the following formula holds true for a majority of the blade rows of the second turbine: 5500 Hz?(number of blades×speed)/60 sec?10000 Hz. The gas turbine engine is rated to produce 15,000 pounds of thrust or more.