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    TURBINE BLADE WITH TIP SHROUD COOLING PASSAGE
    12.
    发明申请
    TURBINE BLADE WITH TIP SHROUD COOLING PASSAGE 有权

    公开(公告)号：US20210131291A1

    公开(公告)日：2021-05-06

    申请号：US16674059

    申请日：2019-11-05

    Applicant: General Electric Company

    Inventor： Melbourne James Myers ,  Michael Anthony Wondrasek ,  Mary Virginia Holloway ,  Luke C. Sponseller ,  Brandon Joseph Kerins

    IPC: F01D5/18

    Abstract: A turbine blade includes a root for fixing the turbine blade to a turbine rotor; and an airfoil coupled to the root, the airfoil including a suction side and a pressure side, and at least one internal wall defining a coolant chamber in the airfoil for delivering a coolant through the airfoil. A tip shroud is coupled to a radially outer end of the airfoil by a suction side fillet and a pressure side fillet. The tip shroud includes a shroud body defining a substantially trident-shaped shroud cooling passage including: a trunk cooling passage, a center cooling passage in fluid communication with the trunk cooling passage, a suction side cooling passage in fluid communication with the trunk cooling passage, and a pressure side cooling passage in fluid communication with the trunk cooling passage. The center cooling passage is fluidly coupled to the coolant chamber to receive a coolant.

    Airfoil shape for turbine rotor blades
    14.
    发明授权
    Airfoil shape for turbine rotor blades 有权

    公开(公告)号：US10808538B2

    公开(公告)日：2020-10-20

    申请号：US16176275

    申请日：2018-10-31

    Applicant: General Electric Company

    Inventor： Melbourne James Myers ,  Sylvain Pierre ,  Michael Gordon Sherman ,  Luke C Sponseller ,  Zhirui Dong ,  Le Trong Tran

    IPC: F01D5/14 ,  F02C3/04

    Abstract: A turbine rotor blade having an airfoil that includes a pressure side portion of a nominal airfoil profile substantially in accordance with Cartesian coordinate values of X, Y, and Z of a pressure side as set forth in Table I. The Cartesian coordinate values of X, Y, and Z are non-dimensional values from 0% to 100% convertible to dimensional distances by multiplying the Cartesian coordinate values of X, Y and Z by a height of the airfoil defined along the Z axis. The X and Y values of the pressure side are coordinate values that, when connected by smooth continuing arcs, define pressure side sections of the pressure side portion of the nominal airfoil profile at each Z coordinate value. The pressure side sections may be joined smoothly with one another to form the pressure side portion.

    AIRFOIL SHAPE FOR TURBINE ROTOR BLADES
    15.
    发明申请
    AIRFOIL SHAPE FOR TURBINE ROTOR BLADES 审中-公开

    公开(公告)号：US20200131912A1

    公开(公告)日：2020-04-30

    申请号：US16176275

    申请日：2018-10-31

    Applicant: General Electric Company

    Inventor： Melbourne James Myers ,  Sylvain Pierre ,  Michael Gordon Sherman ,  Luke C. Sponseller ,  Zhirui Dong ,  Le Trong Tran

    IPC: F01D5/14 ,  F02C3/04

    Abstract: A turbine rotor blade having an airfoil that includes a pressure side portion of a nominal airfoil profile substantially in accordance with Cartesian coordinate values of X, Y, and Z of a pressure side as set forth in Table I. The Cartesian coordinate values of X, Y, and Z are non-dimensional values from 0% to 100% convertible to dimensional distances by multiplying the Cartesian coordinate values of X, Y and Z by a height of the airfoil defined along the Z axis. The X and Y values of the pressure side are coordinate values that, when connected by smooth continuing arcs, define pressure side sections of the pressure side portion of the nominal airfoil profile at each Z coordinate value. The pressure side sections may be joined smoothly with one another to form the pressure side portion.

    TURBINE BLADES HAVING SHANK FEATURES
    16.
    发明申请
    TURBINE BLADES HAVING SHANK FEATURES 审中-公开

    公开(公告)号：US20180230828A1

    公开(公告)日：2018-08-16

    申请号：US15431968

    申请日：2017-02-14

    Applicant: General Electric Company

    Inventor： William Scott Zemitis ,  Richard Ryan Pilson ,  Felipe Roman-Morales ,  Melbourne James Myers

    IPC: F01D5/30 ,  F01D5/14

    CPC classification number: F01D5/3007 ,  F05D2250/74 ,  F05D2260/96 ,  Y02T50/673

    Abstract: A turbine blade includes an airfoil that extends radially between a root end and a tip end, a platform coupled to the root end, and a shank that extends radially inwardly from the platform. The shank includes a downstream cover plate and a downstream angel wing that extends axially from the downstream cover plate. A portion of the downstream cover plate radially outward of the downstream angel wing defines an approximately linear profile. The turbine blade also includes a dovetail region that extends radially inwardly from the shank. A profile of a portion of the downstream cover plate radially inward of the downstream angel wing is recessed relative to the linear profile.

    METHOD FOR MODIFYING AN AIRFOIL SHROUD AND AIRFOIL
    18.
    发明申请
    METHOD FOR MODIFYING AN AIRFOIL SHROUD AND AIRFOIL 审中-公开
    用于修改航空飞机和飞机的方法

    公开(公告)号：US20160362988A1

    公开(公告)日：2016-12-15

    申请号：US15245985

    申请日：2016-08-24

    Applicant: General Electric Company

    Inventor： John David Ward ,  Kelvin Rono Aaron ,  James Ryan Connor ,  Melbourne James Myers ,  Brad Wilson VanTassel

    IPC: F01D5/22

    CPC classification number: F01D5/225 ,  F01D5/005 ,  F05D2220/32 ,  F05D2230/10 ,  F05D2230/50 ,  F05D2230/80 ,  F05D2260/94 ,  F05D2270/332 ,  Y10T29/49234 ,  Y10T83/04

    Abstract: A method is provided for modifying an airfoil shroud located at a tip of a blade of an airfoil, the airfoil shroud having a first end edge, a second end edge, a leading edge, and a trailing edge. The method includes locating a reference location in the first end edge of the airfoil shroud, the reference location including a portion of a seal rail extending circumferentially from a radially outer surface of the airfoil shroud, and a fillet extending from the radially outer surface and positioned directly adjacent to the seal rail. The method further includes forming a relief cut in the seal rail and fillet without performing a weld process on the airfoil shroud to remove the reference location, and installing the airfoil shroud in a turbomachine directly following forming the relief cut, wherein modifying the airfoil shroud is complete following forming the relief cut.

    Abstract translation: 提供了一种用于改变位于翼型件的叶片的尖端处的翼型护罩的方法，翼型护罩具有第一端边缘，第二端边缘，前缘和后缘。 该方法包括将参考位置定位在翼型护罩的第一端边缘中，参考位置包括从翼型护罩的径向外表面周向延伸的密封轨道的一部分，以及从径向外表面延伸的圆角， 直接靠近密封轨。 该方法还包括在密封导轨和圆角上形成浮雕切口，而不对翼型护罩进行焊接处理以去除参考位置，并且在形成浮雕切割之后将翼型件护罩直接安装在涡轮机中，其中修改翼型护罩是 完成以后形成救济切割。

    Turbomachine airfoil having a variable thickness thermal barrier coating
    19.
    发明授权
    Turbomachine airfoil having a variable thickness thermal barrier coating 有权

    公开(公告)号：US11629603B2

    公开(公告)日：2023-04-18

    申请号：US16835492

    申请日：2020-03-31

    Applicant: General Electric Company

    Inventor： Martin James Jasper ,  Jan Agudo ,  Melbourne James Myers ,  Richard Ryan Pilson ,  Christopher W. Kester

    IPC: F01D5/28

    Abstract: An airfoil includes a root and a tip, which define a span of the airfoil therebetween. The airfoil also includes a leading edge and a trailing edge downstream of the leading edge along a flow direction. The leading edge and the trailing edge each extend across the span of the airfoil from the root to the tip. The airfoil further includes a pressure side surface and a suction side surface. The airfoil also includes a thermal barrier coating on the pressure side surface and the suction side surface. The thermal barrier coating includes a base layer and a top coat. A thickness of the base layer varies across each of the pressure side surface and the suction side surface with a maximum thickness of the base layer at the leading edge.

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