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    Turbine airfoil
    1.
    发明专利
    Turbine airfoil 有权
    涡轮机空气

    公开(公告)号：JP2013007381A

    公开(公告)日：2013-01-10

    申请号：JP2012136478

    申请日：2012-06-18

    Applicant: United Technologies Corp ,  ユナイテッド テクノロジーズ コーポレイションUnited Technologies Corporation

    Inventor： SPANGLER BRANDON W ,  ABDEL-MESSEH WILLIAM

    IPC: F01D5/18 ,  F01D9/02 ,  F01D25/12 ,  F02C7/16

    CPC classification number: F01D5/187 ,  F05D2240/126 ,  F05D2240/127

    Abstract: PROBLEM TO BE SOLVED: To provide a cooling system for a turbine airfoil provided with pedestals having an ice-cream-cone-shape or teardrop-shape.SOLUTION: The turbine airfoil includes a wall portion, a cooling channel, a plurality of trip strips and a plurality of pedestals. The wall portion includes a leading edge, a trailing edge, a positive pressure side and a negative pressure side. The cooling channel is for receiving cooling air and extends radially through an interior of the wall portion between the pressure side and the suction side. The plurality of trip strips line the wall portion inside the cooling channel along the pressure side and the suction side. Each of the pedestals is an elongate, tapered pedestal having a curved leading edge. The plurality of pedestals is interposed within the trip strips and connects the pressure side with the suction side.

    Abstract translation: 要解决的问题：提供一种具有设置有具有冰淇淋锥形或泪珠形状的基座的涡轮机翼型的冷却系统。

    解决方案：涡轮机翼型件包括壁部分，冷却通道，多个跳闸条和多个基座。 壁部包括前缘，后缘，正压侧和负压侧。 冷却通道用于接收冷却空气并径向延伸穿过压力侧和吸力侧之间的壁部分的内部。 多个跳闸带沿着压力侧和吸力侧排列在冷却通道内部的壁部分。 每个基座是具有弯曲前缘的细长的锥形基座。 多个基座插入跳闸带内，并将压力侧与吸力侧连接。 版权所有（C）2013，JPO＆INPIT

    2.
    发明专利
    未知

    公开(公告)号：AT459442T

    公开(公告)日：2010-03-15

    申请号：AT04256360

    申请日：2004-10-15

    Applicant: UNITED TECHNOLOGIES CORP

    Inventor： BEALS JAMES T ,  LOPES JOSE ,  DRAPER SAMUEL D ,  MURRAY STEPHEN D ,  SPANGLER BRANDON W

    IPC: B22C9/10 ,  F01D5/18 ,  B22C1/00 ,  B22C7/02 ,  B22C9/04 ,  B22C21/14 ,  F01D5/28 ,  F01D9/02 ,  F01D25/00 ,  F02C7/00

    Abstract: In accordance with the present invention, a casting system is provided which broadly comprises a core (10) and a wax die (12) spaced from said core (10), a refractory metal core (14) having a first end seated within a slot (18) in the core (10) and a second end contacting the wax die (12) for positioning the core (10) relative to the wax die (12), and the refractory metal core having at least one of a mechanism for providing spring loading when closed in the wax die and a mechanism for mechanically locking the wax die to the core. The spring loading mechanism may comprise one or more spring tabs (20). The locking mechanism may comprise an end (32) of the refractory metal core (14') engaging a slot (34) in the wax die (12') (Fig 2).

    GAS TURBINE ENGINE TURBINE VANE AIRFOIL PROFILE
    4.
    发明申请
    GAS TURBINE ENGINE TURBINE VANE AIRFOIL PROFILE 审中-公开
    燃气涡轮发动机涡轮叶片翼型轮廓

    公开(公告)号：WO2014035516A3

    公开(公告)日：2014-05-22

    申请号：PCT/US2013044295

    申请日：2013-06-05

    Applicant: UNITED TECHNOLOGIES CORP

    Inventor： SPANGLER BRANDON W ,  BERGMAN RUSSELL J

    IPC: F01D5/14 ,  F02C7/00

    CPC classification number: F01D25/12 ,  F01D9/02 ,  F01D9/041 ,  F05D2240/81 ,  F05D2250/74 ,  F05D2260/202 ,  Y02T50/672 ,  Y02T50/673 ,  Y02T50/676

    Abstract: A turbine vane for a gas turbine engine includes inner and outer platforms joined by a radially extending airfoil. The airfoil includes leading and trailing edges joined by spaced apart pressure and suction sides to provide an exterior airfoil surface. The inner and outer platforms respectively include inner and outer sets of film cooling holes. One of the inner and outer sets of film cooling holes are formed in substantial conformance with platform cooling hole locations described by one of the sets of Cartesian coordinates set forth in Tables 1 and 2. The Cartesian coordinates are provided by an axial coordinate, a circumferential coordinate, and a radial coordinate, relative to a zero-coordinate. The cooling holes with Cartesian coordinates in Tables 1 and 2 have a diametrical surface tolerance relative to the specified coordinates of 0.200 inches (5.08 mm).

    Abstract translation: 用于燃气涡轮发动机的涡轮叶片包括通过径向延伸的翼型连接的内部平台和外部平台。 翼型包括由间隔开的压力侧和吸力侧连接的前缘和后缘以提供外部翼型表面。 内部和外部平台分别包括内部和外部组膜冷却孔。 内部和外部组膜冷却孔中的一个基本上与由表1和2中所述的笛卡尔坐标系中的一个描述的平台冷却孔位置一致地形成。笛卡尔坐标由轴向坐标，圆周 坐标和相对于零坐标的径向坐标。 表1和表2中具有笛卡尔坐标的冷却孔相对于0.200英寸（5.08毫米）的指定坐标具有直径表面公差。

    GAS TURBINE ENGINE TURBINE VANE PLATFORM CORE
    5.
    发明申请
    GAS TURBINE ENGINE TURBINE VANE PLATFORM CORE 审中-公开
    气体涡轮发动机涡轮叶片平台核心

    公开(公告)号：WO2014035517A2

    公开(公告)日：2014-03-06

    申请号：PCT/US2013044419

    申请日：2013-06-06

    Applicant: UNITED TECHNOLOGIES CORP

    Inventor： BERGMAN RUSSELL J ,  BACH LEONARD A ,  SPANGLER BRANDON W

    CPC classification number: F01D9/02 ,  B22C9/04 ,  B22C9/10 ,  F01D5/187 ,  F01D9/041 ,  F01D25/12 ,  F05D2230/21 ,  F05D2230/211 ,  F05D2240/81 ,  F05D2250/185 ,  F05D2260/2214 ,  Y02T50/672 ,  Y02T50/676

    Abstract: A turbine vane for a gas turbine engine includes inner and outer platforms joined by a radially extending airfoil. The airfoil includes leading and trailing edges joined by spaced apart pressure and suction sides to provide an exterior airfoil surface. The airfoil includes an airfoil cooling passage. A platform cooling passage is arranged within at least one of the inner and outer platforms. The platform cooling passage includes multiple cooling regions with one of the cooling regions arranged beneath the airfoil cooling passage.

    Abstract translation: 用于燃气涡轮发动机的涡轮叶片包括通过径向延伸的翼型件连接的内平台和外平台。 翼型件包括通过间隔开的压力和吸力侧连接的前缘和后缘，以提供外部翼面表面。 翼型件包括翼型件冷却通道。 平台冷却通道布置在内平台和外平台中的至少一个中。 平台冷却通道包括多个冷却区域，其中一个冷却区域布置在翼型冷却通道下方。

    GAS TURBINE ENGINE COOLING HOLE WITH CIRCULAR EXIT GEOMETRY
    7.
    发明申请
    GAS TURBINE ENGINE COOLING HOLE WITH CIRCULAR EXIT GEOMETRY 审中-公开
    带循环出口几何的燃气轮机冷却孔

    公开(公告)号：WO2014109801A2

    公开(公告)日：2014-07-17

    申请号：PCT/US2013061101

    申请日：2013-09-22

    Applicant: UNITED TECHNOLOGIES CORP ,  MCBRIEN JOHN ,  SPANGLER BRANDON W ,  DEVORE MATTHEW A

    Inventor： MCBRIEN JOHN ,  SPANGLER BRANDON W ,  DEVORE MATTHEW A

    CPC classification number: F01D5/186 ,  F05D2250/14 ,  F05D2250/141 ,  F05D2250/314 ,  F23R3/002 ,  F23R3/06 ,  F23R2900/00018 ,  F23R2900/03042 ,  Y02T50/676

    Abstract: A gas turbine engine component includes a structure having an exterior surface. A cooling hole extends from a cooling passage to the exterior surface to provide an exit area on the exterior surface that is substantially circular in shape. A gas turbine engine includes a compressor section and a turbine section. A combustor is provided between the compressor and turbine sections. A component in at least one of the compressor and turbine sections has an exterior surface. A film cooling hole extends from a cooling passage to the exterior surface to provide an exit area that is substantially circular in shape. A method of machining a film cooling hole includes providing a component having an internal cooling passage and an exterior surface, machining a film cooling hole from the exterior surface to the internal cooling passage to provide a substantially circular exit area on the exterior surface.

    Abstract translation: 燃气涡轮发动机部件包括具有外表面的结构。 冷却孔从冷却通道延伸到外表面以在外表面上提供基本上圆形形状的出口区域。 燃气涡轮发动机包括压缩机部分和涡轮部分。 燃烧器设置在压缩机和涡轮机部分之间。 压缩机和涡轮机部分中的至少一个中的部件具有外表面。 膜冷却孔从冷却通道延伸到外表面以提供基本上圆形形状的出口区域。 一种加工膜冷却孔的方法包括提供具有内部冷却通道和外部表面的部件，从外部表面到内部冷却通道加工膜冷却孔以在外部表面上提供基本圆形的出口区域。

    TURBINE ENGINE VARIABLE AREA VANE
    8.
    发明申请
    TURBINE ENGINE VARIABLE AREA VANE 审中-公开
    涡轮发动机可变区域VANE

    公开(公告)号：WO2014011304A3

    公开(公告)日：2014-04-10

    申请号：PCT/US2013037618

    申请日：2013-04-22

    Applicant: UNITED TECHNOLOGIES CORP

    Inventor： SPANGLER BRANDON W ,  PROPHETER-HINCKLEY TRACY A

    IPC: F02C9/20 ,  F01D17/16 ,  F02K3/00

    CPC classification number: F01D17/162 ,  F05D2260/202

    Abstract: A turbine engine stator vane is provided that rotates about an axis, and includes an airfoil, a flange and a shaft. The airfoil extends axially between a first airfoil end and a second airfoil end. The airfoil includes a concave side surface, a convex side surface and a cavity. The concave and the convex side surfaces extend between an airfoil leading edge and an airfoil trailing edge. The cavity extends axially into the airfoil from a cavity inlet in an end surface at the second airfoil end. The flange is connected to the second airfoil end. The flange extends circumferentially around at least a portion of the cavity inlet, and radially away from the concave and the convex side surfaces to a distal flange edge. The shaft extends along the axis, and is connected to the second airfoil end.

    Abstract translation: 提供涡轮发动机定子叶片，其绕轴线旋转，并且包括翼型件，法兰和轴。 机翼在第一翼型端和第二翼型端之间轴向延伸。 翼型件包括凹面侧面，凸面侧面和空腔。 凹面和凸面侧面在翼型前缘和翼型后缘之间延伸。 空腔从第二翼型端的端面中的腔入口轴向延伸到翼型中。 法兰连接到第二翼型端。 凸缘围绕空腔入口的至少一部分周向延伸，并且从凹入和凸出的侧表面径向地远离到远侧凸缘边缘。 轴沿轴线延伸，并连接到第二翼型端。

    9.
    发明专利
    未知

    公开(公告)号：DE602004025779D1

    公开(公告)日：2010-04-15

    申请号：DE602004025779

    申请日：2004-10-15

    Applicant: UNITED TECHNOLOGIES CORP

    Inventor： BEALS JAMES T ,  LOPES JOSE ,  DRAPER SAMUEL D ,  MURRAY STEPHEN D ,  SPANGLER BRANDON W

    IPC: B22C9/10 ,  F01D5/18 ,  B22C1/00 ,  B22C7/02 ,  B22C9/04 ,  B22C21/14 ,  F01D5/28 ,  F01D9/02 ,  F01D25/00 ,  F02C7/00

    Abstract: In accordance with the present invention, a casting system is provided which broadly comprises a core (10) and a wax die (12) spaced from said core (10), a refractory metal core (14) having a first end seated within a slot (18) in the core (10) and a second end contacting the wax die (12) for positioning the core (10) relative to the wax die (12), and the refractory metal core having at least one of a mechanism for providing spring loading when closed in the wax die and a mechanism for mechanically locking the wax die to the core. The spring loading mechanism may comprise one or more spring tabs (20). The locking mechanism may comprise an end (32) of the refractory metal core (14') engaging a slot (34) in the wax die (12') (Fig 2).

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