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    AXIALLY DIVIDED INNER RING FOR A TURBOMACHINE AND GUIDE VANE RING
    25.
    发明申请
    AXIALLY DIVIDED INNER RING FOR A TURBOMACHINE AND GUIDE VANE RING 审中-公开
    用于涡轮机和导向环的轴向分配的内圈

    公开(公告)号：US20160237855A1

    公开(公告)日：2016-08-18

    申请号：US15017729

    申请日：2016-02-08

    Applicant: MTU Aero Engines AG

    Inventor： Werner Humhauser ,  Josef Wagner ,  Hermann Klingels

    IPC: F01D25/24 ,  F04D29/64 ,  F04D29/52 ,  F04D29/54 ,  F01D11/00 ,  F01D9/04

    CPC classification number: F01D9/042 ,  F01D9/04 ,  F01D11/001 ,  F01D17/162 ,  F01D25/246 ,  F04D29/522 ,  F04D29/542 ,  F04D29/563 ,  F04D29/644 ,  F05D2220/32 ,  F05D2240/11 ,  F05D2240/128 ,  F05D2260/30

    Abstract: The present invention relates to an axially divided inner ring (100) for a turbomachine, for fastening to guide vanes (13) of the turbomachine. The inner ring (100) comprises at least one first, solid ring segment (1) disposed upstream, and a second, solid ring segment (3) disposed downstream, wherein the first ring segment (1) is joined to the second ring segment (3) in a detachable manner by means of at least one fastening element. The first ring segment (1) and/or the second ring segment (3) is joined to at least one sealing segment. The inner ring comprises a securing element for securing the fastening element, wherein the securing element is joined to the first ring segment and/or to the second ring segment (3). In addition, the present invention relates to a guide vane ring of a turbomachine having guide vanes, which have an axially divided inner ring according to the invention.

    Abstract translation: 本发明涉及一种用于涡轮机的轴向分开的内环（100），用于紧固到涡轮机的引导叶片（13）。 内环（100）包括至少一个设置在上游的第一实心环段（1）和设置在下游的第二固体环段（3），其中第一环段（1）连接到第二环段 3）通过至少一个紧固元件以可拆卸的方式。 第一环段（1）和/或第二环段（3）连接到至少一个密封段。 所述内环包括用于固定所述紧固元件的固定元件，其中所述固定元件连接到所述第一环段和/或所述第二环段（3）。 此外，本发明涉及具有导向叶片的涡轮机的导向叶环，其具有根据本发明的轴向分开的内圈。

    TURBOFAN AIRCRAFT ENGINE
    26.
    发明申请
    TURBOFAN AIRCRAFT ENGINE 审中-公开
    涡轮飞机发动机

    公开(公告)号：US20160032826A1

    公开(公告)日：2016-02-04

    申请号：US14450882

    申请日：2014-08-04

    Applicant: MTU Aero Engines AG

    Inventor： Klaus Peter Rued ,  Werner Humhauser ,  Hermann Klingels ,  Rudolf Stanka ,  Eckart Heinrich ,  Hans-Peter Hackenberg ,  Claus Riegler ,  Erich Steinhardt ,  Jochen Gier ,  Manfred Feldmann ,  Norbert Huebner ,  Karl Maar ,  Stefan Weber

    IPC: F02C3/107 ,  F01D21/00

    CPC classification number: F02C3/107

    Abstract: A turbofan aircraft engine has at least one stage pressure ratio is at least 1.5, and a quotient of the total blade count divided by 110 is less than a difference ([(p1/p2)−1]) of the total pressure ratio minus one, and the total pressure ratio is greater than 4.5, and the turbine has at least two and no more than five turbine stages; and/or a product (An2) of an exit area (AL) of the second turbine and a square of a rotational speed of the second turbine at the design point is at least 4.5·1010 [in2·rpm2], and a blade tip velocity (uTIP) of at least one turbine stage of the second turbine at the design point is at least 400 meters per second. A jet and method are also provided.

    Abstract translation: 涡轮风扇飞机发动机的至少一级压力比至少为1.5，并且总叶片数除以110的商小于总压力比减1的差（[（p1 / p2）-1]） 并且总压力比大于4.5，并且涡轮机具有至少两个且不超过五个涡轮级; 和/或第二涡轮机的出口区域（AL）的产品（An2）和在设计点处的第二涡轮机的转速的平方值为至少4.5×10 10 [in 2·rpm 2]，并且叶片尖端 在设计点的第二涡轮机的至少一个涡轮机级的速度（uTIP）至少为400米/秒。 还提供喷射和方法。

    Blade-disk assembly, method and turbomachine
    27.
    发明申请
    Blade-disk assembly, method and turbomachine 审中-公开
    刀片盘组件，方法和涡轮机

    公开(公告)号：US20150139811A1

    公开(公告)日：2015-05-21

    申请号：US14546675

    申请日：2014-11-18

    Applicant: MTU Aero Engines AG

    Inventor： Werner Humhauser ,  Herman Klingels ,  Alexander Kloetzer

    IPC: F01D5/30 ,  F01D5/32 ,  F01D5/14

    CPC classification number: F01D5/3038 ,  F01D5/147 ,  F01D5/32 ,  F05D2230/60 ,  Y10T29/49321

    Abstract: Disclosed is a blade-disk assembly of a turbomachine, the blade-disk assembly having a plurality of adjacent rotor blades and a closure blade which are tilted into an anchoring groove, and at least one circumferential retention element which interlockingly cooperates with at least one blade, as well as a plurality of tilt-out prevention elements which are disposed between the groove base and the root portions and which, in the rest state, space the blades from the groove base when in the upper position. The blade-disk assembly further has a locking element, a portion of which is located between the groove base and the root portions and which, in the rest state, spaces the closure blade from the groove base when in the upper position. Also disclosed are a method for assembling such a blade-disk assembly, as well as a turbomachine.

    Abstract translation: 公开了一种涡轮机的叶片盘组件，叶片盘组件具有多个相邻的转子叶片和倾斜到锚定槽中的封闭叶片，以及至少一个周向保持元件，其与至少一个叶片 以及设置在槽基部和根部之间的多个防倾倒元件，并且在静止状态下，当处于上部位置时，叶片从槽底部间隔开。 叶片盘组件还具有锁定元件，其一部分位于槽基部和根部之间，并且在静止状态下，当处于上部位置时，封闭刀片与槽基座间隔开。 还公开了一种用于组装这种叶片盘组件以及涡轮机的方法。

    Unknown
    28.
    发明申请
    Unknown 有权

    公开(公告)号：US20250122807A1

    公开(公告)日：2025-04-17

    申请号：US18915017

    申请日：2024-10-14

    Applicant: MTU Aero Engines AG

    Inventor： Markus Schlemmer ,  Werner Humhauser ,  Andreas Hartung ,  Adam Koscso ,  Dennis Von Brauneck

    IPC: F01D9/04 ,  F01D25/04 ,  F01D25/24

    Abstract: The invention relates to a guide vane for a gas turbine, comprising an airfoil, a platform arranged at a radial end of the airfoil, an upstream flange extending radially from the platform, and a downstream flange extending radially from the platform, wherein the flanges, together with a section of the platform lying between the flanges, bound a groove extending in the circumferential direction of the gas turbine for the arrangement of a damping element. A surface of the section of the platform bounding the groove is arched radially at least in regions thereof in the direction of an opening of the groove. The invention further relates to a guide vane cluster, a housing for a gas turbine, as well as to a gas turbine.

    COMPRESSOR FOR AN ENGINE
    30.
    发明公开
    COMPRESSOR FOR AN ENGINE 审中-公开

    公开(公告)号：US20240003353A1

    公开(公告)日：2024-01-04

    申请号：US18345319

    申请日：2023-06-30

    Applicant: MTU Aero Engines AG

    Inventor： Sergio Elorza Gomez ,  Werner Humhauser ,  Roland Matzgeller

    IPC: F04D19/02 ,  F02C3/06

    CPC classification number: F04D19/02 ,  F02C3/06

    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.

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