Abstract: A moving blade for a turbine or compressor stage of a gas turbine is provided, including a vane (10) and a radially outer shroud (20), for at least one contour point (PK) of an outer contour of a profile cross section (P) of the vane at a radial outer gas channel boundary, the wall thickness of the shroud at at least one edge point (PR) of the shroud, whose connecting line (V), including the contour point with a normal (N), which is perpendicular to the outer contour of the profile cross section in the contour point, encloses an angle (?), which is no more than 5°, is less than a wall thickness of the shroud at the contour point, the wall thickness of the shroud decreasing strictly monotonously along the connecting line.

Abstract: The invention relates to a flow arrangement for placing in the hot gas duct of a turbomachine, having a first surrounding-flow structure and a second surrounding-flow structure, the surrounding-flow structures each having, in reference to the surrounding flow in the hot gas duct, a leading edge and, downstream thereof, a trailing edge, wherein the second surrounding-flow structure is provided as a deflecting blade with a suction side and a pressure side and has a lesser profile thickness than the first surrounding-flow structure, which is arranged on the suction side of the second surrounding-flow structure, and wherein, although the second surrounding-flow structure has a partial axial overlap with the first surrounding-flow structure referred to a longitudinal axis of the turbomachine, the trailing edge of the second surrounding-flow structure is, at the same time, displaced axially downstream relative to the trailing edge of the first surrounding-flow structure.

Abstract: In general, techniques of this disclosure are directed to a modular generator set system. A generator set system may comprise a primary module, which may include an engine, a generator coupled to the engine, a primary frame configured to support the engine and the generator, and a first guiding element. The generator set system may also include a secondary module, a secondary frame, and a second guiding element that is reciprocal to the first guiding element such that, during installation of the secondary module, the second guiding element is configured to interface with the first guiding element to align the secondary module for purposes of automatically electrically or fluidically coupling the primary module to the secondary module upon assembling the primary frame and the secondary frame.

Abstract: A method for controlling and regulating an internal combustion engine with exhaust gas recirculation, in which an EGR rate is determined by a Kaiman filter from calculated and measured variables of the gas path and from calculated and measured variables of combustion. A method for the model-based control and regulation of an internal combustion engine includes calculating injection system set values for controlling the injection system actuators as a function of a set torque by a combustion model. Gas path set values for controlling the gas path actuators are calculated as a function of an EGR rate by a gas path model. A measure of quality is calculated by an optimizer as a function of the injection system and gas path set values. The measure of quality is minimized by the optimizer by changing the injection system and gas path set values within a prediction horizon.

Abstract: The invention relates to a method for coating a component, which is provided for the hot gas duct of a turbomachine, wherein the coating material is applied onto the uncoated component surface in the form of particles in mixture with a binding agent, and the component with the particle-treated binding agent thereupon then undergoes thermal treatment in such a way that the binding agent is released and the coating material remains on the component.

Abstract: A method for monitoring a fuel supply system in an internal combustion engine includes the steps of: recording a first pressure value in the fuel supply system in a region of the fuel supply system associated with a pre-filter and upstream of a low pressure pump; recording a second pressure value in the fuel supply system downstream of the low pressure pump in a region of the fuel supply system associated with a primary filter; monitoring the first pressure value and the second pressure value for an error state of the fuel supply system; and recognizing the error state only if the error state is plausible on a basis of both the first pressure value and the second pressure value.

Abstract: The invention relates to a blade or vane, particularly of a turbine stage of a gas turbine, in particular of an aircraft gas turbine, having a blade or vane root and a blade or vane element joined to the blade or vane root, wherein the blade or vane element has a pressure side and a suction side, and wherein the blade or vane root has at least one raised region on its radial outer side facing the blade or vane element. It is proposed according to the invention that the blade or vane has a first raised region on the pressure side and a second raised region on the suction side, wherein the highest point of the first raised region is disposed essentially directly adjacent to the pressure side, and the highest point of the second raised region is disposed essentially directly adjacent to the suction side.

Abstract: The present invention relates to an impulse body module group with at least two impulse body modules, which, in particular, are identical in construction, for a gas turbine blade, wherein each of the impulse body modules comprises a housing, which, in particular, is of one piece and which has a first cavity, which is closed, in particular in an airtight manner, by a first cover, which is joined to the housing, in particular in a material-bonded manner, and in which a single, in particular spherical or cylindrical, first impulse body is accommodated with play of movement in at least one first direction of impact, wherein the housing has no additional cavity or a single additional cavity, in which a single, in particular spherical or cylindrical, additional impulse body is accommodated with play of movement in at least one direction of impact.

Abstract: A guide vane arrangement for a turbomachine, wherein the guide vane arrangement has a plurality of adjustable guide vanes that are rotatably mounted around an adjustment axis, at least one inner ring segment that is formed separately from the guide vanes, and a seal carrier that is formed separately from the at least one inner ring segment, wherein the at least one inner ring segment is fastened to one guide vane or to a plurality of guide vanes, and wherein the seal carrier is mounted in a spoke-centered manner with respect to the at least one inner ring segment, so that the seal carrier is movably mounted with respect to the at least one inner ring segment in the direction of at least one adjustment axis of at least one guide vane.

Abstract: The present invention relates to a guide vane cascade for a turbomachine, which has guide vanes that are mounted adjustably about an axis of rotation, so as to change an inflow angle each time, wherein a first guide vane and a second guide vane arranged on the pressure side thereof, referred to a longitudinal axis of the turbomachine, are arranged with an axial offset, namely, the axis of rotation of the second guide vane is offset axially toward the back, wherein the first guide vane and the second guide vane are provided such that, in an adjusted state, they form together a tandem configuration in a radially outer region but together they delimit a divergent channel in a radially inner region.

Abstract: A rotor (10) for a turbomachine, in particular for an aircraft engine, having a rotor base body (12), on which at least one sealing fin (14), which is disposed on a base (16), is provided for cooperating with an associated sealing element (20) of the turbomachine; relative to an axial direction of the rotor (10), the base (16) having a base portion (16a) disposed upstream of the sealing fin (14) and a base portion (16b) disposed downstream thereof, for supporting masks during the coating of sealing fins; the upstream base portion (16a) and the downstream base portion (16b) having different radial distances (A1, A2) to a radially outer sealing tip (18) of the sealing fin (14). Also, a turbomachine having at least one such rotor (10).

Abstract: The present invention relates to a blade for a gas turbine, in particular of an aircraft engine, which is produced at least in part from ceramic matrix composite with a plurality of superimposed fabric layers, wherein at least one pair of superimposed fabric layers comprises a first fabric layer that has at least one first point of interruption between two mutually facing first edges of this fabric layer, and a second fabric layer that has at least one second point of interruption adjacent to the first point of interruption between two mutually facing second edges of the second fabric layer, this second point being displaced from the first point of interruption.

Abstract: A stator vane segment for a turbomachine is provided, in particular for a gas turbine, in particular for a turbine stage of a gas turbine. The stator vane segment has at least one stator vane and at least one shroud, in particular an outer shroud, having at least one first profile disposed on the shroud and adapted for attachment of the stator vane segment to the turbomachine casing, the profile extending in the circumferential direction at least partially over a circumferential length of the stator vane segment along the shroud of the stator vane segment and having at least one functional surface which extends at least partially in the axial direction and in the circumferential direction, at least one functional surface of at least one profile having at least two different curvatures in the circumferential direction in at least one radial plane perpendicular to an axis of rotation of the turbomachine in at least one temperature range below a defined operating temperature of the turbomachine.

Abstract: A gas turbine includes a housing and an outer sealing ring. In an assembled state of the housing and the outer sealing ring, a first planar structure of the outer sealing ring is axially moveable with respect to a second planar structure of the housing against a direction of flow through the gas turbine.

Abstract: A covering device (10) for covering at least one region (21) of a component (20) during a high-temperature coating process, the covering device (10) being placed essentially form-fittingly on the component (20) during the high-temperature coating process to prevent the at least one region covered by the covering device from being coated. The elements of the covering device (10) placed essentially form-fittingly on the component (20) are fabricated of a high temperature-resistant plastic which is dimensionally stable during the high-temperature coating process.

Abstract: A guide system for repairing turbomachinery includes a housing; a guide tube have a hollow interior for accepting a working tool module; an axial motor for moving the guide tube axially with respect to the housing; a rotational motor for rotating the guide tube with respect to the housing; and an attachment for securing the housing with respect to a turbomachinery casing. A repair kit and method are also provided.

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 belt-tensioning device, in particular for an internal combustion engine, the belt-tensioning device including a bearing housing having a latching recess, a tensioning roller, a pivoting arm and a locking device. The pivoting arm is mounted on the bearing housing, the pivoting arm being under a torsion spring load. The tensioning roller being coupled to the pivoting arm. The locking device having a first latching device and a second latching device. The first latching device and the second latching device being coupled with the pivoting arm. The latching recess interacts with at least one of the first latching device and the second latching device wherein the locking device is configured in such a way that the pivoting arm can be locked by the locking device in different predefined latching positions and the pivoting arm can be pivoted with predefined freedom in an operating position.

Abstract: A device and a method for producing a blade airfoil from a workpiece which comprises at least two gaps and at least one blank arranged between the two gaps, wherein the blank comprises two opposite lateral faces which are bounded by a base, a top and a first and a second edge. The method comprises: (a) arranging the first and second electrodes in the first and second gaps, the surface of the workpiece forming an annular space surface at the gaps, (b) applying a positive voltage to the blank and applying a negative voltage to the first and second electrodes, (c) moving the first and second electrode in the direction of the first and second lateral faces. Step (b) is preceded by passing electrolyte between the two electrodes over the top toward the base.

Abstract: An airfoil for disposition and exposure to flow in a hot gas duct of a turbomachine is provided. The airfoil has a suction-side wall and a pressure-side wall with respect to its exposure to the flow in the gas duct, the side walls converging in an upstream end region toward a leading edge and converging in a downstream end region toward a trailing edge. The airfoil further has a cavity structure therein which is enclosed between an inner wall surface of the suction-side wall and an inner wall surface of the pressure-side wall when viewed in a cross-sectional plane axially parallel to a longitudinal axis of the turbomachine. When viewed in the cross-sectional plane, at least one of the inner wall surfaces, normalized to a mean camber line of the airfoil as a reference, extends with a change in curvature into at least one of the end regions in such a way that a clearance width of the cavity structure is increased at the at least one end region.