Abstract: A method is described for reducing contrails during operation of an aircraft (100, 100?) having a heat engine 10. The method includes inducing a condensation of moisture contained in the exhaust gas (A) of the heat engine by mixing at least a portion of the exhaust gas with ambient air (U) of the aircraft as well as separating the condensed-out water on the aircraft. Also described is an aircraft (100, 100?) having a heat engine (10). The aircraft includes at least one nozzle (30) which is adapted for conducting exhaust gas (A) from the heat engine of the aircraft at least partially into ambient air (U) of the aircraft and thus to produce a gas mixture, and at least one separator device (40) for separating condensed-out water from the gas mixture.

Abstract: An intermediate element is for a blade/rotor disk connection in a rotor of a fluid flow machine. The intermediate element is adapted to a shape of a blade root of a blade and to a blade root slot in a rotor disk for receiving the blade root such that, when arranged between the blade root and rotor disk, the intermediate element prevents contact between the blade root and rotor disk. The intermediate element has, on an outer surface that faces the rotor disk, at least one protrusion to reduce an air flow parallel to an axis of rotation of the rotor between the rotor disk and the intermediate element; and on an inner surface that faces the blade root, a recess that corresponds to the at least one protrusion.

Abstract: An arrangement reduces oscillation (vibration) of an oscillatory structure. The arrangement has a structure having at least one mode in at least one direction; and an oscillation-reducing device (vibration-reducing device). The oscillation-reducing devices has a housing formed by or provided on the structure, a cavity, and a body configured for making impact contacts with the housing and disposed in the cavity in such a manner that the body is configured to make impact contacts with the housing at least temporarily for as long as the structure is excited in the at least one mode in the at least on direction.

Abstract: An arrangement reduces oscillation of an oscillatory structure. The arrangement has: a structure having at least one mode in at least one direction; and an oscillation-reducing device. The oscillation-reducing device includes a housing on the structure; a cavity; and a body. The body is configured to make impact contact with a first surface portion and a second surface portion of an inner wall of the housing and disposed is in the cavity such that the body can make impact contact with the first surface portion and the second surface portion of the inner wall of the housing at least temporarily for as long as the structure is excited in the at least one mode in the at least one direction. The first surface portion or the second surface portion of the inner wall of the housing has a curved profile.

Abstract: An adjustable guide vane for a compressor, in particular a high-pressure compressor, of a gas turbine, in particular an aircraft gas turbine is described, the vane comprising a radially outer bearing section, a radially inner bearing section, and a vane section, which extends in the radial direction between the outer bearing section and the inner bearing section, wherein the outer and the inner bearing sections are designed in such a way that the adjustable guide vane can be taken up rotatably about a vane axis in the compressor, and wherein the radially inner bearing section is configured like a journal (cone-shaped) and has a lateral surface that is formed circumferentially about the vane axis, the lateral surface being of convex shape. In this way, it is provided that the radius of curvature of the convex lateral surface is at least double the maximum diameter of the bearing section.

Abstract: The present invention relates to an adjustment assembly for the adjustment of adjustable blades or vanes of a turbomachine, having an adjustment ring for coupling to the adjustable blades or vanes of a blade or vane ring and having an adjusting mechanism, which has a coupling rod for coupling to a further blade or vane ring as well as a lever and a push rod, wherein the lever is rotatably mounted at a pivot and has a load arm as well as a force arm, wherein the load arm of the lever is coupled to the push rod and its force arm is coupled to the coupling rod, on different sides of the lever, so that an offset of the coupling rod via the lever and the push rod is converted to a rotation of the adjustment ring around the ring axis thereof.

Abstract: The present invention is directed to a method for coating a component, wherein the component has a first and a second surface, and wherein the first and the second surface adjoin each other at an edge, in which method i) first of all, the edge between the first and the second surface is rounded, and ii) subsequently, a coating is applied to the first surface.

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: The invention is directed to an engine that has a fan, a compressor with a high-pressure compressor, and a combustion chamber. The high-pressure compressor of the engine has a mean stage pressure ratio and an overall pressure ratio formed between the fan and the combustion chamber.

Abstract: An integrally bladed rotor for a turbomachine, in particular a compressor or turbine stage of a gas turbine, to which at least one separately formed impulse element housing (40; 40?) is fastened by at least one fastening element (30; 30?) which engages for this purpose into an opening (41) of the impulse element housing and into an opening (11) of the rotor, the impulse element housing having at least one cavity (44) in which at least one impulse element (5) is accommodated with play.

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 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: 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: 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.