Abstract: A blade/vane cascade segment for a blade/vane cascade of a turbomachine includes a stage and at least two blade/vane elements that define a blade/vane intermediate strip with an axial cascade span on the stage surface by their leading and trailing edges. A stage edge on the inflow side has a contour with a depression. In the axial direction, this depression extends at most by 10% of the cascade span into the blade/vane intermediate strip. Also provided is a corresponding stage, a blade/vane cascade, a blade/vane channel, and a turbomachine.

Abstract: Described is a method for operating a compressor of a turbomachine, in which, when considered in the direction of a main flow, an, in particular, radially averaged degree of reaction has dropped in a front compressor area from a maximum to a minimum, is held constant or virtually constant across a central compressor area up into a rear compressor area, an, in particular, radially averaged degree of reaction being adjusted in the rear compressors area which is closer to the minimum than to the maximum, and a residual swirl of at least 47° is present in the middle section, and a compressor and a turbomachine.

Abstract: A method for operating an internal combustion engine having a number of cylinders and an injection system having an injection system that has a common rail and a number of injectors associated with the cylinders, wherein an individual accumulator is associated with each injector and stores fuel from the common rail for the injector. The method has the following steps: starting the internal combustion engine, operating the internal combustion engine, shutting off the internal combustion engine. The following steps are also provided: a state indicating an engine standstill is detected, in particular after the internal combustion engine has been shut off, a high-pressure limit value is defined and a target high pressure is specified, a leakage is produced in the common rail without injection, the fuel pressure in the common rail is reduced to the defined high-pressure limit value below the target high pressure by way of the leakage.

Abstract: The invention relates to a device for producing, repairing and/or replacing a component, particularly an aircraft component, by means of a powder that can be solidified by energy radiation of an energy radiation source, characterized in that the device comprises an application unit that is designed such that the powder can be applied onto an uneven surface by means of the application unit.

Abstract: The invention relates to a method for manufacturing a housing of a turbomachine, in particular a gas turbine. The method comprises at least the steps: providing a housing blank, manufacturing a housing element, producing an assembly opening corresponding to the housing element in the housing blank, arranging the housing element in the assembly opening, and joining the housing element to the housing blank by means of a welding method. In addition, the invention relates to a turbomachine housing.

Abstract: An inner ring according to the invention for a guide vane cascade of a turbomachine has a radially outer inner ring surface as well as a plurality of bearing mounts for a respective guide vane plate of a guide vane. The bearing mounts each have an opening in the outer inner ring surface as well as a bottom face lying radially opposite the opening. At least two of the bearing mounts are separated from each other by a separating wall and are connected to each other in a region of their bottom face by a through-opening. A guide vane cascade according to the invention for a turbomachine has an inner ring according to the invention and a plurality of guide vanes inserted into the bearing mounts.

Abstract: A method for execution with an exhaust-gas particle filter which is operated with an exhaust-gas aftertreatment system, wherein the exhaust-gas particle filter has a filter wall along which exhaust gas is conducted for filtering purposes; wherein the method includes a regeneration phase with the steps: a) setting a soot load on the exhaust-gas particle filter, wherein the set soot load effects the formation of a soot layer on ash deposited on the filter wall; and b) subsequently mobilising the deposited ash by burning off the formed soot layer during the course of an active regeneration of the exhaust-gas particle filter.

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 compressor section driven by the turbine section, 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: A guide vane element for a gas turbine, the guide vane element having a first guide vane, a second guide vane distanced by one division in the peripheral direction, and at least one band joining these guide vanes, in particular, a radially inner band and/or a radially outer band, wherein at least one band joining these guide vanes has a vane-side surface having a contouring and a first front side in the peripheral direction having a groove, which is particularly straight in the axial direction, for the uptake of a sealing element.

Abstract: A dual-fuel injector for an internal combustion engine, including a first injection device for injecting a first fuel and at least one second injection device, arranged radially outward from the first injection device, for injecting a second fuel, wherein the second fuel is different from the first fuel, and further including a first connection for supplying the first fuel and a second connection for supplying the second fuel. The first connection is fluidically connected to the first injection device in order to supply the gaseous first fuel and the second connection is fluidically connected to the second injection device in order to supply the liquid second fuel.

Abstract: A device for influencing the temperatures in inner ring segments of a gas turbine is provided, which includes in the circumferential direction of the gas turbine, at least one outer shell assembly having a plurality of outer ring segments and at least one inner shell assembly having a plurality of inner ring segments. The inner ring segments are equal to or greater in number than the outer ring segments. An outer ring segment gap is between each two adjacent outer ring segments. An inner ring segment gap is between each two adjacent inner ring segments. An outer ring segment gap, or at least a radial cooling air passage, is located opposite each inner ring segment gap in the radial direction of the gas turbine.

Abstract: Disclosed is a process for coating a component of a TiAl alloy in order to improve the high-temperature resistance of the component. The process comprises depositing a Pt- and Cr-free protective layer alloy comprising Ti, Al, Nb, Mo and B and optionally one or more of W, Si, C, Zr, Y, Hf, Er and Gd on the component by physical vapor deposition at a temperature of less than or equal to 600° C. The protective layer alloy has a higher Al content than the TiAl alloy of the component. A coated component made by this process is also disclosed.

Abstract: Disclosed is a blade/vane cascade segment of a blade/vane cascade of a turbomachine, which comprises at least two blade/vane elements defining an axial cascade width, and a stage with a stage surface, as well as a stage edge on the inflow side. The stage surface has an elevation extending up to the pressure side of a first blade/vane element and a depression extending up to the suction side of the other blade/vane. At least one highest point of the elevation and at least one lowest point of the depression each lie at least 30% and at most 60% of the axial cascade width downstream of the inflow edges of the blade/vane elements. In this case, the elevation and the depression each come close to the stage edge on the inflow side.

Abstract: A method for operating a system having an internal combustion engine and an exhaust aftertreatment system having an SCR catalyst, wherein the internal combustion engine is controlled on the basis of at least one process parameter that influences a nitrogen oxide raw emission, wherein aging detection is performed for the SCR catalyst, wherein in a first operating mode of the internal combustion engine, if aging of the SCR catalyst is detected, the at least one process parameter is changed in the direction of a reduced nitrogen oxide raw emission, wherein the internal combustion engine is controlled on the basis of the changed at least one process parameter.

Abstract: Described is a TiAl alloy which, besides titanium, comprises 42 to 48 at. % aluminum, 3 to 5 at. % niobium, 0.05 to 1 at. % molybdenum, 0.2 to 2.2 at. % silicon, 0.2 to 0.4 at. % carbon, 0.05 to 0.2 at. % boron, and optionally tungsten, zirconium and hafnium, as well as unavoidable impurities, and at room temperature has a microstructure which comprises globular colonies of lamellae of ?2-Ti3Al and ?-TiAl, as well as silicide precipitates, and essentially no ? phase. A method for producing a component made of this alloy is also described.

Abstract: A method for operating an internal combustion engine, device, and an internal combustion engine including a motor which has a crankshaft. A charge air flow is supplied to the motor that is compressed by means of a compressor via a second rotational movement, and a power turbine for producing a first rotational movement is acted on by an exhaust gas flow discharged from the motor. The following steps are provided: in a first operating mode, operating the internal combustion engine in four-stroke operation, and in a second operating mode, operating the internal combustion engine in two-stroke operation. The crankshaft can be driven by the power turbine via the first rotational movement, and the compressor can be driven by the crankshaft via the second rotational movement, wherein the second rotational movement for the compressor can be set differently from the first rotational movement of the power turbine.

Abstract: The present invention relates to a one-piece intermediate sheet having a first layer that has a first outer surface with a first outer surface region, and a second layer that is integrally formed with the first layer and has a second outer surface with a second outer surface region, wherein the first and second outer surfaces form opposite-lying, in particular parallel, outer surfaces of the intermediate sheet, and the hardness of the first outer surface region is smaller than the hardness of the second outer surface region, and/or the modulus of elasticity of the first layer is smaller than the modulus of elasticity of the second layer.

Abstract: A clamping device for joining two adjacent-lying flange pieces of components is provided herein. The clamping device forms a frame that is equipped for the purpose of being penetrated by the two adjacent-lying flange pieces. The clamping device further includes a tensioning device that is equipped for the purpose of placing under tension or clamping the two adjacent-lying flange pieces in the frame. In addition, the disclosure relates to a flange connection of two components having such a clamping device.

Abstract: Disclosed is a method for producing a blade comprising a blade airfoil and a blade root for a turbomachine. The method comprises providing a first workpiece based on a first material and a second workpiece based on a second material which is different from the first material and has a higher temperature resistance than the first material; and connecting the first workpiece and the second workpiece by friction welding to form a composite component having a first region of the first material, and a second region of the second material. Optionally upon material-subtracting further processing, the first region forms the blade root, and the second region forms the blade airfoil.

Abstract: The seal arrangement for a mid turbine frame of a gas turbine, especially an aircraft gas turbine, includes a tubular fluid duct, especially an oil duct, which extends along an axial direction, a bearing sleeve surrounding the fluid duct radially outside and in the peripheral direction, and at least one seal ring arranged between fluid duct and bearing sleeve in the radial direction, this ring being arranged around the fluid duct. The seal ring is taken up at least partly in a groove provided in the bearing sleeve.