Patents by Inventor Alexander Halcoussis
Alexander Halcoussis has filed for patents to protect the following inventions. This listing includes patent applications that are pending as well as patents that have already been granted by the United States Patent and Trademark Office (USPTO).
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Publication number: 20230279777Abstract: 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.Type: ApplicationFiled: February 28, 2023Publication date: September 7, 2023Applicant: MTU Aero Engines AGInventors: Benjamin Hanschke, Alexander Halcoussis, Amadeusz Rekosiewicz, Moritz Thon, Benjamin Linke
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Publication number: 20230265766Abstract: A guide vane for arrangement in a gas duct of a turbomachine, including a guide vane blade and an outer platform which, relative to a longitudinal axis of the turbomachine, is arranged radially on an outside on the guide vane blade. The outer platform delimits the gas duct with an inner wall surface radially to the outside. The inner wall surface has a curved profile with an inflection point when viewed in an axial section, the inner wall surface having a concave curvature in a first axial section which is upstream from the inflection point with respect to a flow through the gas duct, and a convex curvature in a second axial section which is arranged downstream from the inflection point with respect to the flow through the gas duct.Type: ApplicationFiled: July 29, 2021Publication date: August 24, 2023Inventors: Alexander Halcoussis, Harsimar Sahota
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Patent number: 11085453Abstract: An engine, having a first, second, third and fourth quasi-stage, each with a stator and a downstream rotor, wherein a particular blade/vane solidity in the center section is in the first quasi-stage, at least 1.04 and at most 1.16 for the stator and at least 1.46 and at most 1.67 for the rotor; in the second quasi-stage, at least 1.13 and at most 1.32 for the stator and at least 1.32 and at most 1.61 for the rotor; in the third quasi-stage, at least 1.20 and at most 1.39 for the stator and at least 1.16 and at most 1.41 for the rotor; in the fourth quasi-stage, at least 1.37 and at most 1.63 for the stator and at least 1.15 and at most 1.41 for the rotor.Type: GrantFiled: July 22, 2019Date of Patent: August 10, 2021Assignee: MTU Aero Engines AGInventors: Harsimar Sahota, Alexander Halcoussis, Sergio Elorza Gomez
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Patent number: 11066949Abstract: The present invention relates to a guide vane ring element for a turbomachine, such as a gas turbine, with at least one drilled hole for the at least partial uptake of a flange of a rotatable guide vane, wherein an edge of the drilled hole on the vane element side is designed convexly along the edge in a first circumferential portion and, in particular, has a radius.Type: GrantFiled: July 29, 2016Date of Patent: July 20, 2021Assignee: MTU Aero Engines AGInventors: Vitalis Mairhanser, Birgit Effner, Alexander Halcoussis, Hannes Wolf
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Patent number: 10844726Abstract: The invention relates to a blade for a turbomachine, comprising a blade element with a suction side and a pressure side, which extend between a leading edge and a trailing edge of the blade element, as well as a blade root for connection of the blade at a main rotor body, wherein the blade comprises a crack-affecting device, which, in the radial direction, has an altered cross-sectional geometry in comparison to an aerodynamically optimized blade profile. The invention further relates to a rotor for a turbomachine having at least one such blade, a turbomachine having at least one such blade, and/or with such a rotor as well as a method for producing a blade.Type: GrantFiled: October 18, 2018Date of Patent: November 24, 2020Assignee: MTU Aero Engines AGInventors: Wolfgang Horn, Daniel Theurich, Martin Becker, Axel Stettner, Wilfried Schuette, Michael Junge, Hans-Peter Hackenberg, Alexander Halcoussis, Hannes Wolf
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Patent number: 10711798Abstract: A variable vane cascade for a turbomachine, in particular for a compressor stage or turbine stage of a gas turbine, having at least one first vane, in particular guide vane that has a first distance from a circumferentially adjacent vane, at least one second vane, in particular guide vane that has at least one second distance from at least one circumferentially adjacent vane that is smaller than the first distance, and an actuating device, in particular for jointly and/or reversibly adjusting the first and second vane from a first position where at least one airfoil cross section of the first vane and an airfoil cross section of the second vane each have a first stagger angle, into a second position where these airfoil cross sections have second stagger angles, the second stagger angle of the first vane differing from the second stagger angle of the second vane, in particular being larger than the second stagger angle of the second vane.Type: GrantFiled: July 11, 2017Date of Patent: July 14, 2020Assignee: MTU Aero Engines AGInventor: Alexander Halcoussis
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Patent number: 10677076Abstract: A guide vane ring for a turbomachine minimizes a leakage flow that passes through a recess into which a guide vane disk is inserted. It has a guide vane row having a plurality of guide vanes, each having a vane airfoil and a vane disk, as well as an inner ring having an inner ring surface facing the plurality of guide vanes. Viewed in the direction of a designated primary flow streaming through the turbomachine, the vane disks have a front and a rear surface region. In a nominal and/or a maximum open position of the guide vanes, the front and/or the rear surface region of at least one of the vane disks has an offset from the inner ring surface that is radially disposed (relative to a central axis of the inner ring).Type: GrantFiled: April 26, 2017Date of Patent: June 9, 2020Assignee: MTU AERO ENGINES AGInventors: Hannes Wolf, Alexander Halcoussis
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Publication number: 20200024948Abstract: The present invention relates to a high pressure compressor for an engine, having a first, second, third and fourth quasi-stage, each with a stator and a downstream rotor, wherein a particular blade/vane solidity in the center section is in the first quasi-stage, at least 1.04 and at most 1.16 for the stator and at least 1.46 and at most 1.67 for the rotor; in the second quasi-stage, at least 1.13 and at most 1.32 for the stator and at least 1.32 and at most 1.61 for the rotor; in the third quasi-stage, at least 1.20 and at most 1.39 for the stator and at least 1.16 and at most 1.41 for the rotor; in the fourth quasi-stage, at least 1.37 and at most 1.63 for the stator and at least 1.15 and at most 1.41 for the rotor.Type: ApplicationFiled: July 22, 2019Publication date: January 23, 2020Applicant: MTU Aero Engines AGInventors: Harsimar SAHOTA, Alexander HALCOUSSIS, Sergio ELORZA GOMEZ
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Publication number: 20190120061Abstract: The invention relates to a blade for a turbomachine, comprising a blade element with a suction side and a pressure side, which extend between a leading edge and a trailing edge of the blade element, as well as a blade root for connection of the blade at a main rotor body, wherein the blade comprises a crack-affecting device, which, in the radial direction, has an altered cross-sectional geometry in comparison to an aerodynamically optimized blade profile. The invention further relates to a rotor for a turbomachine having at least one such blade, a turbomachine having at least one such blade, and/or with such a rotor as well as a method for producing a blade.Type: ApplicationFiled: October 18, 2018Publication date: April 25, 2019Applicant: MTU Aero Engines AGInventors: Wolfgang Horn, Daniel Theurich, Martin Becker, Axel Stettner, Wilfried Schuette, Michael Junge, Hans-Peter Hackenberg, Alexander Halcoussis, Hannes Wolf
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Publication number: 20180017081Abstract: A variable vane cascade for a turbomachine, in particular for a compressor stage or turbine stage of a gas turbine, having at least one first vane, in particular guide vane that has a first distance from a circumferentially adjacent vane, at least one second vane, in particular guide vane that has at least one second distance from at least one circumferentially adjacent vane that is smaller than the first distance, and an actuating device, in particular for jointly and/or reversibly adjusting the first and second vane from a first position where at least one airfoil cross section of the first vane and an airfoil cross section of the second vane each have a first stagger angle, into a second position where these airfoil cross sections have second stagger angles, the second stagger angle of the first vane differing from the second stagger angle of the second vane, in particular being larger than the second stagger angle of the second vane.Type: ApplicationFiled: July 11, 2017Publication date: January 18, 2018Inventor: Alexander HALCOUSSIS
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Publication number: 20170314406Abstract: A guide vane ring for a turbomachine minimizes a leakage flow that passes through a recess (11) into which a guide vane disk is inserted. It has a guide vane row having a plurality of guide vanes (20), each having a vane airfoil (24) and a vane disk (21), as well as an inner ring (10) having an inner ring surface (12) facing the plurality of guide vanes. Viewed in the direction (R) of a designated primary flow streaming through the turbomachine, the vane disks (21) have a front and a rear surface region (22a, 22b). In a nominal and/or a maximum open position of the guide vanes, the front and/or the rear surface region (22a, 22b) of at least one of the vane disks has an offset (31a, 31b) from the inner ring surface (12) that is radially disposed (relative to a central axis (X) of the inner ring (10)).Type: ApplicationFiled: April 26, 2017Publication date: November 2, 2017Inventors: Hannes Wolf, Alexander Halcoussis
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Publication number: 20170037741Abstract: The present invention relates to a guide vane ring element (10) for a turbomachine, such as a gas turbine, with at least one drilled hole (30) for the at least partial uptake of a flange (21) of a rotatable guide vane (20), wherein an edge of the drilled hole on the vane element side is designed convexly along the edge in a first circumferential portion (31A, 31B) and, in particular, has a radius (R).Type: ApplicationFiled: July 29, 2016Publication date: February 9, 2017Inventors: Vitalis Mairhanser, Birgit Effner, Alexander Halcoussis, Hannes Wolf
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Publication number: 20110164967Abstract: An axial flow machine having an asymmetrical air inlet and, downstream therefrom, a compressor having an inlet guide baffle composed of guide vanes is characterized in that at least some of the guide vanes of the inlet guide baffle have a vane profile and/or an angle of attack that deviate(s) from the remaining guide vanes. The inlet flow angle of the first compressor stage is hereby evened out circumferentially symmetrically. This is accomplished in that the different inlet angles resulting at various circumferential positions of the inlet guide baffle due to the asymmetry of the air inlet, are influenced by selective profiling and/or by selectively modifying the angle of incidence of individual guide vanes in such a way that a circumferentially symmetrical outflow angle from the inlet vane ring results.Type: ApplicationFiled: September 9, 2009Publication date: July 7, 2011Applicant: MTU AERO ENGINES GMBHInventors: Sergio Elorza Gomez, Alexander Halcoussis