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: 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 invention relates to an energy supply device with at least one fuel cell and to a method for operating at least one energy supply device with at least one fuel cell, which has at least one anode that can be supplied with a fuel and at least one cathode that can be supplied with ambient air for generating electrical energy. The proposed energy supply device has a converter device.

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: A heat engine, in particular an aircraft engine, having a first compressor for supplying a combustion chamber of the heat engine with air and a first turbine arranged downstream of the combustion chamber for driving the first compressor, wherein the heat engine also has at least one steam supply line for supplying steam from a steam source into the combustion chamber. The heat engine also has a steam supply device, which has a second compressor and is designed to compress the working gas further by the second compressor as a function of a mass flow conducted through the steam supply line, before the working gas flows into the combustion chamber.

Abstract: The invention relates to a device and a method for cooling a heat exchanger of a fuel cell of a flight propulsion drive outside of a flight phase of an aircraft. The flight propulsion drive has an air line with at least one compressor for feeding compressed air for the operation of the fuel cell. A flow amplifier is arranged in the region of the heat exchanger, and is configured to guide air onto a cooling surface of the heat exchanger, whereby thermal energy is diverted from the cooling surface of the heat exchanger.

Abstract: A rotor assembly may be used in a gas turbine. The rotor assembly includes rotor segments arranged in succession in an axial direction and interconnected in the axial direction by a tie-rod, and a rotor segment disposed forwardly in the axial direction having a first contact surface and a rotor segment disposed rearwardly in the axial direction having a second contact surface. The first and second contact surfaces are at least partially in contact with each other, are substantially annular in shape, and extend in a radial direction and in a circumferential direction. The first contact surface and/or the second contact surface extend at least partially obliquely relative to the radial direction. An angle is formed between the first contact surface and the second contact surface when viewed in a sectional plane defined by the axial direction and the radial direction.

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.

Abstract: Described is a blade for a high-speed turbine stage of an aircraft gas turbine, in particular of an aircraft engine, the blade including a radially inner blade root, and an airfoil extending radially outwardly from the blade root. It is provided that the blade be shroudless and that the airfoil have a radially outer end portion that is positionable opposite a rub surface when the blade is in an installed state, and that the airfoil have a radially inner chord length that is at least 1.1 times, preferably at least 1.2 times, in particular at least 1.3 times a radially outer chord length, the inner chord length being measured at the airfoil directly above the blade root, and the outer chord length being measured at the airfoil in the region of or below the end portion.

Abstract: The invention relates to a guide vane assembly for a turbomachine, comprising a guide vane, which has a guide vane airfoil; and a guide vane holder. The guide vane is mounted in the guide vane holder such that the guide vane can be moved about an axis of rotation. For this purpose, the guide vane has at least one axle element, which is inserted into the guide vane holder in such a way that an outer lateral surface of the axle element faces an inner lateral surface of the guide vane holder. A protective coating is applied to at least parts of the guide vane airfoil. A protective coating is applied to at least one of the lateral surfaces. The invention also relates to a compressor module, a turbomachine, and a method for producing a guide vane assembly.

Abstract: The present invention relates to a method for testing a component by a penetrant test, in which a penetrant is introduced onto a surface of the component; at least one portion of the penetrant is removed from the surface by blasting the surface with a solid blasting material; the surface is optically detected after the blasting.

Abstract: The present invention relates to an aircraft comprising at least one wing, at least one flight propulsion drive, and a retainer, particularly an engine pylon, which interconnects the wing and the flight propulsion drive. The aircraft comprises at least one heat exchanger for cooling exhaust gas of the fight propulsion drive and/or at least one water removal channel having at least one removal apparatus for removing water from exhaust gas of the flight propulsion drive, especially after the exhaust gas has flowed through the heat exchanger. The removal apparatus is disposed on, more particularly in, the retainer or is connected to the wing by means of the retainer, and/or the flight propulsion drive is fastened to the retainer by means of at least one flight propulsion drive suspension means, and the heat exchanger is fastened, independently thereof, to the retainer by means of at least one heat exchanger suspension means.

Abstract: A gas turbine has a rotor blade. The rotor blade has a blade root connected to an airfoil. The blade root has a root contour with respect to a cross-sectional view. From a lower end of the blade root, the blade root contour has convex contour portions and concave contour portions. From the lower end along the blade root contour between a convex contour portion and an adjoining concave contour portion, there is a contour portion as a flank portion that is load-bearing. From the lower end along the blade root contour between a concave contour portion and an adjoining convex contour portion, there is a contour portion as a flank portion that is not load-bearing in operation. At least one of the concave contour portions has a first arc portion, a second arc portion, and a straight portion disposed between the two arc portions.

Abstract: The invention relates to a blisk segment for a gas turbine, comprising at least one first blade having an airfoil, a leading edge, a trailing edge, a blade root, a suction side and a pressure side, a pedestal, and a first fillet having a device for influencing crack growth. The robustness of the gas turbine is improved in accordance with the invention in that, at the leading edge of the first blade, at least on the first fillet, a first surface structure is arranged and interacts with the crack-influencing device for influencing crack growth and for flow influencing.

Abstract: In general, a generator set system may include a primary module including: an engine; a generator coupled to the engine and configured to convert mechanical energy of the engine to electrical energy; a primary frame configured to support the engine and the generator; and a first coupling. The generator set system also may include a secondary module including a secondary module including: at least one of a heat recovery module, an exhaust module, a cooling system module, a fuel system module, a solar power module, a battery module, an electrical system module, an air intake module, or an air outlet module; a secondary frame; and a second coupling. The first coupling and the second coupling are configured to automatically electrically or fluidically couple the primary module to the secondary module upon assembling the primary frame and the secondary frame.

Abstract: The invention relates to a method for determining individual vectors for open-loop and/or closed-loop control of at least one energy beam of a layering apparatus, comprising at least the steps of: providing layer data characterizing at least one component layer of a component to be additively manufactured, on the basis of the layer data, determining individual vectors, according to which at least one energy beam is to be moved relative to a construction and joining zone of the layering apparatus in order to solidify a material powder selectively to the component layer, determining at least one node point of a plurality of individual vectors, and adapting at least one property of at least one individual vector of the at least one node point, the at least one property being selected from a group comprising spatial orientation, radiation sequence in relation to at least one other individual vector, and vector length.

Abstract: The present invention relates to a lever for adjusting an adjustable vane of a turbomachine, wherein the lever has a fulcrum for the rotatable mounting of the lever about an axis of rotation, a first load arm with a first adjusting connection point for at least indirect connection to an adjustable vane and a first force arm with a first actuating connection point for connection to an adjusting device, wherein the lever further has a first strut, which connects the first adjusting connection point and the first actuating connection point to each other and thereby extends in an arc-shaped manner about the axis of rotation.

Abstract: An example hybrid power system includes a mobile platform comprising: a generator set comprising: a reciprocating engine configured to convert natural gas into rotational mechanical energy; and a generator configured to convert rotational mechanical energy sourced from the reciprocating engine into electrical energy; an electrical energy storage system (ESS) configured to store electrical energy; an electrical motor configured to convert electrical energy sourced from a combination of the generator set and the ESS into rotational mechanical energy; and a pump configured to operate using rotational mechanical energy sourced from the electrical motor.

Abstract: A method for machining a blade and a blade for a turbomachine comprising a shroud which is positioned on a tip side of the blade. The shroud has an outer surface with at least one circumferential fin arranged thereon, whereby at least one section of the outer surface beside the at least one fin is processed in at least two manufacturing steps. At least one first section of the outer surface is processed to have a first shape and at least one second section of the outer surface is processed to have a second shape.

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a propulsor section, a geared architecture, a high spool and a low spool. The high spool includes a high pressure compressor and a high pressure turbine. The low spool includes a low pressure compressor and a low pressure turbine. 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 1.55. A mechanical tip rotational Mach number of the blades is greater than or equal to 0.5 at an approach speed.