Abstract: A lever assembly forming part of a vane actuating mechanism and serving to connect a stator vane of a turbomachine to an actuating ring of an actuator is provided. The lever assembly is adapted such that an actuating lever of the lever assembly is moved in a first mounting direction transversely to a vane stem of the stator vane extending radially relative to a longitudinal machine axis of the turbomachine and is thereby positioned laterally against the vane stem, and that the actuating lever is radially form-fittingly locked to the vane stem by a locking piece in a second mounting direction. An assembly method and a turbomachine is also provided.

Abstract: The present invention relates to a bearing chamber housing for bearing a shaft of a turbomachine, having an additively built-up housing wall and a cover of the oil chamber, wherein, in each case referred to an axis of rotation of the shaft, the housing wall bounds radially outwardly an oil chamber of the bearing chamber housing and the cover of the oil chamber disposed on an axial end of the housing wall axially bounds a region of the oil chamber; the region bounds radially inwardly an inner shell of the bearing chamber housing.

Abstract: A guide vane for a turbomachine is disclosed. The guide vane includes a heat pipe wall that is interpenetrated by a capillary system, which has a continuous material transition from a first surface, which delimits the guide vane toward the outside, to a second surface, which lies opposite to the first surface and delimits an evaporation cavity in the interior of the guide vane. Further disclosed is a method for producing a guide vane having a heat pipe wall, wherein at least the heat pipe wall is formed by additive manufacture.

Abstract: In a method for the generative production of a component (3), in particular of a turbo-engine component, wherein material (4) is bonded layer-by-layer selectively to a layer disposed therebeneath or to a substrate (6), according to the invention a laser (1A; 1B; 2) additionally acts on the material (4) before, during and/or after the bonding.

Abstract: The present invention relates to a housing for a gas turbine compressor, having an operating duct for taking up several rows of rotating blades arranged one behind the other axially, wherein a wall of the operating duct has a first air bleed, wherein the housing has a cooling passage for cooling said operating duct wall by conveying air from the first air bleed to a first outlet passage for supplying at least one first bleed air consumer, wherein the first outlet passage is arranged downstream of the first air bleed, and wherein the operating duct wall has a second air bleed downstream of the first air bleed, and the housing has a discharge passage from said second air bleed to a second outlet passage for supplying at least one bleed air consumer.

Abstract: The invention relates to a method and an associated device, the method including at least the following steps: applying a layer of powder to a component platform in the region of a building and joining area; locally melting and/or sintering the powder layer, wherein, in the region of the building and joining area, at least one high-energy beam is moved in relation to the component platform, selectively impinging the powder layer, at least part of which at least one high-energy beam and the component platform are moved in relation to one another, in the form of a parallel arrangement arranged along a linear feed direction; lowering the component platform by a predetermined layer thickness in a lowering direction; and repeating the above-mentioned steps until the component region is completed.

Abstract: The invention relates to a seal arrangement for a turbomachine, especially for an aircraft engine, for sealing a radial gap between a rotor and a stator, comprising at least one seal carrier for the supporting and/or fastening of at least one seal element, wherein the seal carrier comprises a radial crosspiece extending in a radial direction of extension and an axial crosspiece, formed as a single piece with the latter and extending in an axial direction of extension, wherein the seal element is arranged at a radially inner-lying bearing surface of the axial crosspiece, and a front ring or a front ring segment, viewed in the flow direction, and/or a rear ring or a rear ring segment, each with a radially running crosspiece.

Abstract: A method for detecting continuous injection during the operation of an internal combustion engine with an injection system having a high-pressure accumulator for a fuel, wherein—a high pressure in the injection system is monitored as a function of time, wherein—in order to detect continuous injection it is checked whether the high pressure has dropped by a predetermined continuous injection differential pressure value within a predetermined continuous injection time interval, wherein—it is checked whether a reduction valve which connects the high-pressure accumulator to a fuel reservoir has been triggered, and wherein—continuous injection is detected if—a reduction valve has not been triggered in a predetermined checking time interval before the dropping of the high pressure, and if—the high pressure has dropped by the predetermined continuous injection differential value amount within the predetermined continuous injection time interval.

Abstract: A measuring device for measuring a physical property of a fluid, including a base body which is closed along a peripheral line. The base body surrounds a first recess passing through the base body in the axial direction and the base body includes at least one second recess passing through the base body in the axial direction recess. At least one sensor device, which can be associated with the base body, can be bought into contact with the first recess.

Abstract: A method for starting an internal combustion engine by a compressed air starting system, in which in a first starting sequence the engagement of the starter is brought about by compressed air, a decompression valve for relieving the cylinder working space is acted on in the opening direction, and starting of the internal combustion engine is initiated by pulsed compressed air being applied to the starter. In a second starting sequence the decompression valve is acted on in the closing direction, and constant compressed air is applied to the starter.

Abstract: A method for operating an internal combustion engine having an injection system which has a high-pressure accumulator, wherein a high pressure in the high-pressure accumulator is regulated via a suction throttle on the low-pressure side as a first pressure control member in a first high-pressure control loop, wherein in a normal operation a high-pressure disturbance variable is produced via a pressure control valve on the high-pressure side as a second pressure control member, via which fuel is redirected from the high-pressure accumulator to a fuel reservoir. For this purpose, the high pressure in a safety operation is regulated by the pressure control valve via a second high-pressure control loop, or, in the safety operation, a maximum fuel volume flow is continuously redirected from the high-pressure accumulator to the fuel reservoir via the pressure control valve.

Abstract: The disclosure describes a generator set that includes a generator, an engine mechanically coupled to the generator, an enclosure housing the engine and the generator, a heating system, and a renewable energy-powered energy source. The generator is configured to supply power to an electrical system. The heating system is configured to heat at least one of the engine or the enclosure to at least a startup temperature. The renewable energy-powered energy source is configured to supply energy to the heating system to heat the at least one of the engine or the enclosure.

Abstract: The present invention relates to a method for producing and/or repairing wear-stressed recesses or openings on components (22) of a turbomachine, especially of elements of a flow passage boundary, and also to corresponding components, wherein the method comprises: producing an at least two-layer molded repair part (15), one layer (2) of which is formed by an Ni-solder and a further layer (3) of which is formed from a mixture of an Ni-solder (4) and hard material particles (5) of hard alloys on a base of cobalt or nickel and which at least partially has an outer shape which is complementary to the inner shape of the recess (20) or opening which is to be repaired, inserting the molded repair part (15) into the recess (20) or opening and at least partially heat-treating the component (22) for soldering the molded repair part (15) onto the component.

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a turbine section including a fan drive turbine, a geared architecture driven by the fan drive turbine, and a fan driven by the fan drive turbine via the geared architecture. At least one stage of the turbine section includes an array of rotatable blades and an array of vanes. A ratio of the number of vanes to the number blades is greater than or equal to about 1.55. A mechanical tip rotational Mach number of the blades is configured to be greater than or equal to about 0.5 at an approach speed.

Abstract: The invention relates to a turbine center frame for an aircraft gas turbine, with a plurality of first components and a plurality of second components, which, in the peripheral direction, are arranged alternately next to one another and, in a radial direction, bound a flow duct conveying hot gas, wherein the first components each have two first overlapping segments and the second components each have two second overlapping segments with a respective region of the transitioning from a first component to a second component or vice versa, a first overlapping segment and a second overlapping segment are arranged overlapping, and with at least one centering element, at which the first components and the second components are supported along the peripheral direction and in a radial direction with the first components and the second components essentially centered with respect to a central axis of the turbine center frame.

Abstract: A power system includes a synchronous electrical generator having a rotor driven by a shaft; a permanent magnet signaling generator, coupled to the shaft; and an angle computation unit configured to calculate a rotor angle or load angle based on a voltage from the permanent magnet signaling generator and a voltage from the synchronous electrical generator.

Abstract: A method for operating an internal combustion engine. The method includes detecting a structure borne sound signal in a time-dependent manner for the at least one combustion chamber during operation of the internal combustion engine, and determining, in a predetermined measuring window, at least one evaluation parameter from the detected structure borne sound signal. The method also includes obtaining at least one comparative result by comparing the at least one evaluation parameter with at least one predetermined comparison value, and assigning to the structure borne sound signal, on the basis of the at least one comparative result, one of a knocking event in the at least one combustion chamber and an interference signal.

Abstract: A method for inspecting a rotor blade unit of a gas turbine, the rotor blade unit including a radially outwardly disposed outer shroud connected by a material-to-material bond to a radially inwardly adjoining rotor blade. The method includes: providing a rotor blade unit; processing the outer shroud at its axially forward or axially aft side, the side selected to be processed being the one that has a greater distance from the adjacent axial leading edge or axial trailing edge of the rotor blade, a contact surface being formed by processing the selected side; providing the rotor blade unit having the contact surface in a measuring device, in particular a measuring device for determining the torsion or twist of the rotor blade unit about an axis.

Abstract: A blade/vane cascade segment of a blade/vane cascade for a turbomachine is disclosed, which comprises a stage with a stage surface as well as a first blade/vane element and a second blade/vane element. The stage surface comprises a first elevation reaching as far as the pressure side of the first blade/vane element and a second elevation reaching as far as the suction side of the second blade/vane element. A furthest downstream point of a boundary of the second elevation has an axial position which differs from the axial position of at least one highest point of the first elevation by a maximum of 10% of an axial cascade span. Furthermore, a blade/vane cascade, a blade/vane channel, a stage, a turbomachine and an aircraft engine are disclosed.

Abstract: The invention relates to a wear-resistant shield for a rotating blade root of a rotating blade of a gas turbine, especially an aircraft gas turbine, having a base and two side walls connected to the base, wherein the side walls lie opposite each other and are shaped so as to be substantially complementary to an outer contour of a particular rotating blade root, and wherein the wear-resistant shield for this purpose is set up in such a way that, in an installed state, it is to be taken up between the respective rotating blade root and a rotating blade root mount of a rotor, especially between the respective rotating blade root and an axial securing element arranged in the rotating blade root mount.