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 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: 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 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 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: The present invention relates to an aircraft having at least one wing, on which at least one propulsion unit is arranged, comprising at least one heat engine, especially a gas turbine, as well as an exhaust gas passage for conducting exhaust gas of the heat engine into and inside the wing.

Abstract: A turbine module (2) for a turbomachine (1). The turbine module (2) includes a main channel (26) to guide a main flow (36) through the turbine module (2), a rotor blade (21) and a stator vane (22), the stator vane (22) including a stator airfoil (22) and a platform (23), with the stator airfoil (22) arranged downstream of the rotor blade (21) in the main channel (26), and a cavity (30) including an inlet (31) for injecting a part (36.2) of the main flow (36) into the cavity (30), an outlet (32) for a reinjection of the part (36.2) of the main flow (36) from the cavity (30) into the main channel (26), wherein the cavity (30) is arranged at an axial position of the stator vane (20) and is radially offset from the stator airfoil (22).

Abstract: The present invention relates to a method for manufacturing a plate of a turbomachine having a plurality of receiving grooves, which are formed on the cylindrical peripheral surface of the plate in order to receive blade roots of blades of the turbomachine. The receiving grooves extend in a straight line from one face of the plate to the other face of the plate, such that a straight line delimits the receiving groove from each point on the cross section of the receiving groove on one of the faces to the corresponding point on the cross section on the other face. A circular plate is provided and a plurality of receiving grooves are introduced into the plate simultaneously on opposing regions relative to a central axis of rotation of the plate.

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, a radially outer shroud, and an airfoil extending between the blade root and the shroud. It is provided that the outer shroud have only a single sealing element, which projects radially from the shroud, in particular only a single sealing fin.

Abstract: A stator vane (3) for a turbine (50c) of a turbomachine (50), the stator vane having a stator vane airfoil (3c), an inner shroud (3a) and an outer shroud (3b), the inner shroud (3a) and the outer shroud (3b) bounding an annular space (2), in which working gas (51) is conveyed during operation, radially with respect to a longitudinal axis (52) of the turbomachine (50), and the stator vane airfoil (3c) having a stator vane airfoil channel (3d) extending through its interior between a radially inner inlet (6) and a radially outer outlet (7). A characteristic features is that the inlet (6) is disposed in such a manner that a gas (8) flowing through the stator vane airfoil channel (3d) during operation is at least partially formed of the working gas (51) conveyed in the annular space (2), and thus the working gas is redistributed from radially inward to radially outward.

Abstract: The invention relates to a turbine blade which comprises at least one cavity that is defined by a wall with one or more surface discontinuities which preferably are selected from elevations, depressions and undercuts and preferably change at least one eigenmode of the blade. The blade may suitably be produced by a generative production method such as selective laser melting (SLM).

Abstract: An exhaust gas treatment device, an aircraft propulsion system and method for treating an exhaust gas stream are provided. The exhaust gas treatment has a condenser, which condenses at least a portion of water contained in the exhaust gas stream from the turbomachine and thereby releases a first energy; an evaporator, which evaporates at least a portion of the water condensed in the condenser and thereby absorbs a second energy, which is extracted from the exhaust gas stream from the turbomachine; a turbine, which is driven by steam output by the evaporator and expands the steam; a fan, which can be driven by the turbine and feeds the condenser ambient air in order that it absorb the first energy; and at least one exhaust apparatus, out of which the ambient air given off by the condenser and/or a dehumidified exhaust gas stream is exhausted from the condenser.

Abstract: The invention relates to a blade or vane, a blade or vane segment, and an assembly for a turbomachine, wherein the blade or vane has a blade or vane element with a blade or vane element profile and a radially inner end of the blade or vane element and a radially outer end of the blade or vane element, and wherein the blade or vane has, in addition to the blade or vane element profile, at least one first guide profile in the region of the radially inner end of the blade or vane element (EI) and/or of the radially outer end of the blade or vane element and is spaced apart from the associated blade or vane element end in the radial direction, and extends at least partly in the axial direction and at least partly in the peripheral direction and/or at least partly in the tangential direction.

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: 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: 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, a radially outer shroud, and an airfoil extending between the blade root and the shroud. It is provided that the outer shroud have only a single sealing element, which projects radially from the shroud, in particular only a single sealing fin.

Abstract: The invention is directed to a guide vane ring for a turbomachine, in which a vane bearing is produced in the inner ring via platform plates of the vanes and reliability against disintegration is produced via journals that extend radially inward from the platform plates, as well as a turbomachine.

Abstract: A seal system for a moving blade system of a gas turbine includes a seal carrier including a ring section extending in the axial direction and the circumferential direction, on which a radially inner seal element is situated. A first flange section and a second flange section extend outwardly from the ring section in the circumferential direction and the radial direction and are situated at a distance from each other in the axial direction. An inner shroud section of at least one guide blade or of at least one guide blade segment includes a sealing section extending inwardly in the radial direction and in the circumferential direction, the sealing section being accommodated between the first flange section and the second flange section of the seal carrier and being connected thereto. A disk-like first seal is situated upstream from the first flange section with respect to the main flow direction of a working medium and partially abuts the first flange section radially inwardly.

Abstract: A turbine center frame for a gas turbine, in particular an aircraft gas turbine, having a hot gas-conveying, segmented annular duct (6), a radial outer casing (4) that is disposed at a distance around the annular duct (6), a hub element (7) that is disposed radially inwardly at a distance from the annular duct (6), a plurality of struts (5) that essentially extend in the radial direction (RR) through the annular duct (6) and that are coupled radially inwardly to the hub element (7) and radially outwardly to the outer casing (4), and having at least one connecting element (30), that is formed obliquely to the struts (5), that extends through the annular duct (6), and is coupled radially inwardly to a seal carrier element (20), the seal carrier element (20) being supported axially forwardly on the hub element (7). At least one connecting element (30) is coupled radially outwardly to one of the struts (5).