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 blade or vane for a turbomachine, particularly for an aircraft engine, with a blade element for interacting with the flow medium, wherein the blade has different diffusion protective coatings in various regions on its surface for protection against corrosion and/or oxidation, wherein the diffusion protective coatings are produced by chromizing and/or aluminizing, wherein the blade element is divided into two regions along the longitudinal axis of the blade element, wherein the first region extends over 80 to 95% of the length of the blade element, and the second region extends over the remainder of the length of the blade element, and wherein in both regions, an AlCr diffusion protective coating is applied, and wherein in one of the regions, the AlCr diffusion protective coating has a higher Cr content.

Abstract: A method for operating an internal combustion engine with a motor having a number of cylinders and an injection system having a common rail with a number of injectors assigned to the cylinders and similar high pressure components, which is designed to hold fuel from the common rail for the injector, wherein the method has the steps: injecting fuel from the common rail into a cylinder by way of an injector, determining a fuel pressure for a high-pressure component, in particular the common rail, the injector and/or the individual reservoir, having at least one high-pressure sensor measuring the fuel pressure. Provision is made for a defect in the high-pressure sensor to be detected in that a check is made as to whether magnitude of the high-pressure control deviation (ep) during a predetermined time interval (tLimit1SD, tLimit2SD, tLimit3SD) exceeds a predetermined limiting value (eLimit1SD, eLimit2SD, eLimit3SD).

Abstract: A method for implementation with the operation of an internal combustion engine, having an ignition plug, which is arranged on a combustion chamber of a cylinder of the internal combustion engine, wherein: in a first step, a cylinder pressure at the ignition time at the combustion chamber is detected, as well as a breakdown voltage at the ignition plug; and in a second step, a current electrode distance of the ignition electrodes, representing a current ignition electrode wear state, is determined based on the detected cylinder pressure, the detected breakdown voltage and a constant of proportionality.

Abstract: A containerized generator set comprises a container, a generator set, a radiator, and a cooling fan. The generator set, radiator, and cooling fan are disposed within an interior volume defined by the container. The interior volume has a length, width, and height. The generator set comprises an engine and a generator. The cooling fan is mechanically coupled to the engine, and the cooling fan has a diameter that is substantially the same as the width of the interior volume. The cooling fan may further have a swept cross-sectional area that is at least 60% of the cross-sectional area of the interior volume.

Abstract: A method for evaluating the quality of a component produced by means of an additive laser sintering and/or laser melting method, in particular a component for an aircraft engine, includes at least the steps of providing a first data set, which comprises spatially resolved color values, which each characterize the temperature of the component at an associated component location during the laser sintering and/or laser melting of the component, providing a second data set, which comprises spatially resolved color values corresponding to the first data set, which color values each characterize the temperature of a reference component at an associated reference component location during the laser sintering and/or laser melting of the reference component.

Abstract: A cooling system including a first coolant line and a second coolant line, at least one first component to be cooled, into which the first coolant line opens, and a first ventilation line. The first ventilation line is fluidically connected to the at least one first component and is configured for ventilating the at least one first component. The first ventilation line opens into the second coolant line.

Abstract: The invention relates to a guide vane segment for an aircraft engine, having at least one guide vane element, having at least one flange formed in a radial extension direction (R) of the guide vane element and at least one positioning member projecting from the flange in the radial extension direction (R), and at least one seal carrier, which is arranged at the flange and is aligned relative to the guide vane element by way of the positioning member, wherein the flange has a supporting surface on which the seal carrier is supported, wherein a partial surface of the positioning member is adjacent to the flange at a transition from the flange to the positioning member, and respective surface normal lines of the supporting surface and of the partial surface enclose between them an angle (?) that is different from a zero angle.

Abstract: A method for producing a fuel composition, including the following steps: providing special gas containing combustible substances; reforming a first part of the special gas by producing synthesis gas; producing dimethyl ether from the synthesis gas by producing a reaction mixture containing a dimethyl ether; separating methanol from the reaction mixture and producing a methanol-reduced dimethyl ether mixture; and bringing together a second part of the special gas with the methanol reduced dimethyl ether mixture in order to obtain the fuel composition.

Abstract: An engine simulation system and method for replacing an original engine in an existing system of a piece of equipment, including an engine controller from the original engine, a new engine having an engine controller, and a simulator module connected between the engine controller of the new engine and the engine controller of the original engine. The simulator module is configured to simulate that the original engine is still in the existing system.

Abstract: The invention relates to a method for determining a balancing removal process for a balancing device for balancing a rotationally symmetrical component, particularly a rotor component, particularly of a turbomachine, a combination of machining lengths and depths being calculated, taking into account a pre-defined maximum machining length and minimum machining depth, in such a way that, with reliable combinations for compensating the same unbalance, the machining length of the calculated combination is longer than the machining length of at least one other permissible combination and, at the same time, the machining depth of the calculated combination is shallower than the machining depth of said other combination.

Abstract: Described is a turbine center frame for a gas turbine, in particular an aircraft gas turbine, the turbine center frame having a radially inner wall and a radially outer wall bounding an annular space through which hot gas flows and each having a contour facing the annular space, the contours describing an inner annular space curve along the inner wall and an outer annular space curve along the outer wall. At least one vane element extends in the radial direction through the annular space and has an axial leading edge and an axial trailing edge, the vane element having an outer axial width.

Abstract: An airfoil array for a turbomachine, in particular a turbine or compressor stage of a gas turbine.

Abstract: Disclosed is a lever connection for a guide vane adjustment for connecting a guide vane of a turbomachine to an adjusting ring of an actuator, which is arranged in such a way that an adjusting lever of the lever connection is pushed radially from a first mounting direction, in relation to a machine longitudinal axis of the turbomachine, onto a vane shaft of the guide vane, and is locked with the vane shaft in a form-fitting and, in particular, rotationally rigid manner by a movement in a second mounting direction; a mounting method; and a turbomachine.

Abstract: The invention relates to a method and device for generatively producing components, said device comprising a radiation device for selectively radiating a powder bed, and an induction device for inductively heating the component produced by radiating the powder bed, Said induction device comprising at least one voltage source which can simultaneously produce alternating voltages with at least two different frequencies.

Abstract: A distributor apparatus of a common rail system for an internal combustion engine. The distributor apparatus is designed with at least one distributor device having a high-pressure line for fuel and a plurality of feed lines branching off from the high-pressure line within the distributor device. Each feed line leads to an individual accumulator and an injector. The feed line is associated with a restriction device and a discharge bore, which is designed to cooperate with a pressure-measuring device and which is arranged downstream of the restriction device, such that the restriction device is arranged in the feed line and the discharge bore, arranged in the distributor device, is connected to the feed line downstream of the restriction device.

Abstract: A method for operating an exhaust after-treatment system including an SCR-catalyst, a metering device for dosing a reducing agent being controlled on the basis of a determining variable that influences a nitrogen-oxide concentration downstream of the SCR-catalyst. A breakthrough identification is carried out for the SCR-catalyst, wherein if a breakthrough is identified, the determining variable is altered to a higher nitrogen-oxide concentration downstream of the SCR-catalyst and the metering device is controlled in order to dose the reducing agent on the basis of the altered determining variable.

Abstract: The invention relates to a sealing assembly for a fluid kinetic machine, in particular for an aircraft engine, for sealing a radial gap between a rotor and a stator, including at least one sealing support for retaining and/or fixing at least one sealing element, wherein the sealing support includes a first and a second radial web extending in a radial extension direction as well as an axial web extending in an axial extension direction, firmly bonded to the radial webs, and the radial webs form a receptacle for receiving an element of the stator. Therein, the radial webs are formed as sheet elements formed elongated in radial direction, wherein a radially interior end of the radial webs is respectively firmly bonded, in particular welded, to a radially exterior surface of the axial web.

Abstract: An axially divided inner ring is provided for fastening to, in particular adjustable, guide vanes (50) of a turbomachine, in particular a compressor or turbine stage of a gas turbine, which includes a first partial ring (10) and a second partial ring (20), which is supported in the axial direction directly or indirectly on two axially facing support surfaces (11, 12) of the first partial ring (10) and is fixed on the first partial ring (10) in the radial direction with the aid of multiple alignment pins (40) distributed in the circumferential direction.

Abstract: The invention relates to a blade for a turbomachine, comprising a blade element with a suction side and a pressure side, which extend between a leading edge and a trailing edge of the blade element, as well as a blade root for connection of the blade at a main rotor body, wherein the blade comprises a crack-affecting device, which, in the radial direction, has an altered cross-sectional geometry in comparison to an aerodynamically optimized blade profile. The invention further relates to a rotor for a turbomachine having at least one such blade, a turbomachine having at least one such blade, and/or with such a rotor as well as a method for producing a blade.