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    Cooling circuit with shaped cooling pins
    11.
    发明授权
    Cooling circuit with shaped cooling pins 有权

    公开(公告)号：US10954801B2

    公开(公告)日：2021-03-23

    申请号：US16546055

    申请日：2019-08-20

    Applicant: HONEYWELL INTERNATIONAL INC.

    Inventor： Benjamin Dosland Kamrath ,  Jason Smoke ,  Daniel C. Crites

    IPC: F01D5/18

    Abstract: A cooling circuit to receive a cooling fluid includes at least one shaped cooling pin disposed in the cooling circuit. The at least one shaped cooling pin has a first end and a second end extending along an axis. The first end has a first curved surface defined by a minor diameter and the second end has a second curved surface defined by a major diameter. The first curved surface is to be upstream in the cooling fluid and the minor diameter is less than the major diameter.

    Method of producing an abrasive tip for a turbine blade
    12.
    发明授权
    Method of producing an abrasive tip for a turbine blade 有权

    公开(公告)号：US10857596B1

    公开(公告)日：2020-12-08

    申请号：US16127383

    申请日：2018-09-11

    Applicant: HONEYWELL INTERNATIONAL INC.

    Inventor： Don Mittendorf ,  Jason Smoke ,  Brent Ludwig

    IPC: B22F5/04 ,  F01D5/20 ,  B22F3/15 ,  B23P15/04 ,  B29C43/10

    Abstract: A method of producing an abrasive tip for a turbine blade includes producing or obtaining a metal powder that is mixed with an abrasive ceramic powder and producing or obtaining a metallic mold that is in the shape of an airfoil. The metallic mold includes a hollow interior portion. The method further includes sealing the metal and ceramic powder mixture within the hollow interior portion of the metallic mold under vacuum and subjecting the sealed mold to a hot isostatic pressing process. The hot isostatic pressing process compacts and binds the metal and ceramic powder mixture together into a solid article in the shape of the airfoil. Still further, the method includes slicing the solid article into a plurality of airfoil-shaped slices and bonding one slice of the plurality of airfoil-shaped slices to a tip portion of a turbine blade.

    METHOD OF PRODUCING AN ABRASIVE TIP FOR A TURBINE BLADE
    13.
    发明申请
    METHOD OF PRODUCING AN ABRASIVE TIP FOR A TURBINE BLADE 审中-公开

    公开(公告)号：US20200360998A1

    公开(公告)日：2020-11-19

    申请号：US16127383

    申请日：2018-09-11

    Applicant: HONEYWELL INTERNATIONAL INC.

    Inventor： Don Mittendorf ,  Jason Smoke ,  Brent Ludwig

    IPC: B22F5/04 ,  B22F7/08 ,  B22F3/15 ,  B22F3/24

    Abstract: A method of producing an abrasive tip for a turbine blade includes producing or obtaining a metal powder that is mixed with an abrasive ceramic powder and producing or obtaining a metallic mold that is in the shape of an airfoil. The metallic mold includes a hollow interior portion. The method further includes sealing the metal and ceramic powder mixture within the hollow interior portion of the metallic mold under vacuum and subjecting the sealed mold to a hot isostatic pressing process. The hot isostatic pressing process compacts and binds the metal and ceramic powder mixture together into a solid article in the shape of the airfoil. Still further, the method includes slicing the solid article into a plurality of airfoil-shaped slices and bonding one slice of the plurality of airfoil-shaped slices to a tip portion of a turbine blade.

    AIRFOIL WITH LEADING EDGE CONVECTIVE COOLING SYSTEM
    15.
    发明申请
    AIRFOIL WITH LEADING EDGE CONVECTIVE COOLING SYSTEM 审中-公开

    公开(公告)号：US20200018172A1

    公开(公告)日：2020-01-16

    申请号：US16035184

    申请日：2018-07-13

    Applicant: HONEYWELL INTERNATIONAL INC.

    Inventor： Luis Tapia ,  Jason Smoke

    IPC: F01D5/18 ,  F01D9/04 ,  F01D25/12

    Abstract: An airfoil includes a leading edge and an opposing trailing edge. The airfoil includes a pressure sidewall and an opposing suction sidewall. A leading edge cavity is defined between the pressure sidewall and the suction sidewall. The leading edge cavity has a first end opposite the leading edge and a second end defined at a rib. The airfoil includes at least one pin structure defined in the leading edge cavity between the first end and the second end. The at least one pin structure includes a main body and a first branch. The main body is coupled to the second end and extends toward the first end. The first branch extends from the main body toward the first end.

    GAS TURBINE ENGINE COMPONENTS HAVING SEALED STRESS RELIEF SLOTS AND METHODS FOR THE FABRICATION THEREOF
    20.
    发明申请
    GAS TURBINE ENGINE COMPONENTS HAVING SEALED STRESS RELIEF SLOTS AND METHODS FOR THE FABRICATION THEREOF 有权
    具有密封应力消除槽的气体涡轮发动机部件及其制造方法

    公开(公告)号：US20150300192A1

    公开(公告)日：2015-10-22

    申请号：US14257485

    申请日：2014-04-21

    Applicant: HONEYWELL INTERNATIONAL INC.

    Inventor： Jason Smoke ,  Bradley Reed Tucker ,  Ardeshir Riahi ,  Ed Zurmehly ,  Alexander MirzaMoghadam

    IPC: F01D11/00 ,  F01D9/04 ,  B22D19/00 ,  B22D25/02 ,  B22D25/06 ,  F01D1/06 ,  F01D25/12

    CPC classification number: F01D11/005 ,  B22D19/00 ,  B22D25/02 ,  B22D25/06 ,  F01D1/06 ,  F01D9/041 ,  F01D25/12 ,  F05D2220/32 ,  F05D2230/30 ,  F05D2240/55 ,  F05D2260/20 ,  F05D2260/941

    Abstract: Embodiments of a gas turbine engine component having sealed stress relief slots are provided, as are embodiments of a gas turbine engine containing such a component and embodiments of a method for fabricating such a component. In one embodiment, the gas turbine engine includes a core gas flow path, a secondary cooling flow path, and a turbine nozzle or other gas turbine engine component. The component includes, in turn, a component body through which the core gas flow path extends, a radially-extending wall projecting from the component body and into the secondary cooling flow path, and one or more stress relief slots formed in the radially-extending wall. The stress relief slots are filled with a high temperature sealing material, which impedes leakage between the second cooling and core gas flow paths and which fractures to alleviate thermomechanical stress within the radially-extending wall during operation of the gas turbine engine.

    Abstract translation: 提供了具有密封的应力消除槽的燃气涡轮发动机部件的实施例，其中包括含有这种部件的燃气涡轮发动机和用于制造这种部件的方法的实施例的实施例。 在一个实施例中，燃气涡轮发动机包括核心气体流动路径，二次冷却流动路径，以及涡轮喷嘴或其它燃气涡轮发动机部件。 该组件又包括核心气体流路延伸穿过的组件主体，从组件主体突出并进入二次冷却流动路径的径向延伸的壁，以及一个或多个形成在径向延伸的应力释放槽 壁。 应力释放槽填充有高温密封材料，其阻止第二冷却和核心气体流动路径之间的泄漏，并且其在燃气涡轮发动机的操作期间破裂以减轻径向延伸的壁内的热机械应力。

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