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: 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: 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 invention relates to a layer building process for the additive manufacture of at least one wall region of a component including applying at least one powder layer of a material to at least one building-up and joining zone of at least one movable building platform, carrying out a first solidifying step, in which the material is irradiated selectively with at least one energy beam, wherein irradiation parameters of the at least one energy beam are set so a molten bath is produced and a defect-affected wall region of the wall is produced, without applying a further powder layer, carrying out a second solidifying step, in which the defect-affected wall region produced in the first solidifying step is irradiated selectively with the at least one energy beam, lowering the building platform layer by layer by a predefined layer thickness, and repeating the steps above one or more times.

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.

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: The invention relates to a method for machining a workpiece by an electrochemical machining process in which material is removed from the workpiece in an electrolyte liquid, where the electrolyte liquid is then filtered in a membrane filter system which has a membrane that undergoes a relative movement in the membrane filter system during the filtering process, and the filtered electrolyte liquid is reused for the electrochemical machining process.

Abstract: The invention relates to a rotor arrangement for a low-pressure turbine of a turbomachine with end-side linkage to a shaft of the turbomachine, comprising a plurality of rotating stages, which are arranged in a flow direction and are connected to one another, wherein the last rotating stage in the flow direction can be linked to the shaft by a rotor disk.

Abstract: The invention relates to a housing structure for a turbomachine, with a housing shell and at least one containment ring arranged in therein, wherein the housing structure surrounds a flow channel in a ring-shaped manner, in which at least one rotating stage with a guide vane arrangement and a rotating blade arrangement is arranged, as well as a turbomachine having such a housing structure.

Abstract: The invention relates to a turbomachine for a flight propulsion drive with a compressor, a combustion chamber, a turbine through which a gas flow flows in the flow direction of turbomachine, and a heat exchanger downstream of the turbine, wherein the heat exchanger is equipped to generate steam from water using energy from the gas flow, which can be supplied with fuel to the gas flow for combustion in the combustion chamber.

Abstract: The invention relates to a turbomachine for a flight propulsion drive with a compressor, a combustion chamber, a turbine through which a gas flow flows in a flow direction of the turbomachine, and a heat exchanger downstream of the turbine, wherein the heat exchanger is equipped to generate steam from water using energy from the gas flow, which can be supplied with the fuel to the gas flow for combustion in the combustion chamber.

Abstract: A rotor blade system includes a blade root, a blade neck adjoining the blade root, an airfoil adjoining the blade neck, a radially outer partition wall, an axially forward partition wall and an axially rearward partition wall, connected to the radially outer partition wall such that the partition walls surround the blade neck on three sides, the partition walls projecting beyond the blade neck in the circumferential direction, and further including a blade root protection plate designed to be arranged on the blade root. It is provided that the blade root protection plate have at least one sealing portion that extends in the axial direction from the forward partition wall to the rearward partition wall and whose radially outer side is disposed opposite the radially outer partition wall when the blade root protection plate is arranged on the blade root.

Abstract: The invention is directed to a guide vane device for a guide vane adjustment of a turbomachine, comprising a guide vane, which has a shaft extending along an adjustment axis of the guide vane, as well as an adjusting lever, which forms a connecting element for connection to an outer end of the shaft, wherein, at its outer end, the shaft has a three-dimensional front surface with at least one beveled region, the connecting element surrounds, at least in sections, the outer end of the shaft in a sleeve-like manner and is joined to the outer end in a form-fitting manner, and the guide vane device forms an interference fit between the connecting element and the outer end of the shaft.

Abstract: The present invention is directed to a method for the material-removing machining of a component for a turbomachine that includes the steps of i) recording a distance image by contact-free measurement of at least one surface of the component that is to be machined; ii) comparing the distance image to a computer model of the component and determining the sites of the component that are to be machined based on deviations between the distance image and the computer model; iii) generating a tool path for the sites of the component that are to be machined; and iv) material-removing machining of the component on the basis of the tool path by spark erosion, laser drilling, or conventional drilling.

Abstract: The invention relates to a method for producing at least in regions an oxidation protection coating on a component of a thermal gas turbine. In accordance with the invention, the method comprises the steps of coating the component at least in regions with a lacquer, which comprises at least one UV-curable binder and metal particles, curing the lacquer by exposure to UV light, and thermal treatment of the component at least in the region of the cured lacquer for production of the oxidation protection coating.

Abstract: A description is given of a fuel cell comprising an electrode-membrane unit comprising a cathode and an anode, a cathodal gas diffusion element, an anodal gas diffusion element, the electrode-membrane unit being accommodated between the gas diffusion elements; a cathodal bipolar plate, and an anodal bipolar plate. Provision is made here for the cathodal gas diffusion element or/and the anodal gas diffusion element to have at least one structural element facing the respective bipolar plate and integrally bonded to the relevant gas diffusion element.

Abstract: The invention relates to an adjustment system for a compressor in a gas turbine, in particular in an aircraft gas turbine, for the adjustment of adjustable guide vanes, including an actuator device with a first actuator for the adjustment of the guide vanes and an adjusting device with a first adjusting valve for the hydraulic operation of the first actuator, wherein the adjusting device can adjust and lock in position the first actuator by way of the hydraulic operation, is characterized in that the actuator device is assigned a locking device for additional fixation and can be fixed in a position by the locking device.

Abstract: The present invention relates to an exhaust-gas treatment device for an aircraft engine, comprising an exhaust-gas channel, through which an exhaust gas of the aircraft engine flows, and a first cooling unit for cooling with ambient air, characterized by a second cooling unit, which is downstream of the first cooling unit with respect to an exhaust-gas flow in the exhaust-gas channel.