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    PARTIAL CAVITY BAFFLES FOR AIRFOILS IN GAS TURBINE ENGINES
    181.
    发明申请
    PARTIAL CAVITY BAFFLES FOR AIRFOILS IN GAS TURBINE ENGINES 有权
    用于气体涡轮发动机中的空气的部分气囊

    公开(公告)号：US20170037732A1

    公开(公告)日：2017-02-09

    申请号：US14818379

    申请日：2015-08-05

    Applicant: United Technologies Corporation

    Inventor： Brandon W. Spangler

    IPC: F01D5/18

    CPC classification number: F01D5/187 ,  F01D5/188 ,  F01D5/189 ,  F05D2220/32 ,  F05D2230/60 ,  F05D2260/20 ,  F05D2260/201

    Abstract: An airfoil of a gas turbine engine having a hollow body defining at least one airfoil cavity therein, the hollow body defining an inner diameter and an outer diameter and a baffle positioned within the at least one airfoil cavity and extending over less than an entire length between the inner diameter and the outer diameter, the baffle configured to reduce the cross-sectional area within the at least one airfoil cavity. The at least one airfoil cavity includes a first portion having a length that is defined by an open cavity having a full cross-sectional area and a second portion having a length that is defined by a reduced cross-sectional area, the second portion being the length of the baffle within the at least one airfoil cavity.

    Abstract translation: 具有中空体的燃气涡轮发动机的翼型件，其中限定了至少一个翼型腔，所述中空体限定内径和外径，以及限定在所述至少一个翼型腔内的挡板，并且在所述至少一个翼型腔之间延伸超过整个长度， 内径和外径，挡板构造成减小至少一个翼型腔内的横截面积。 所述至少一个翼型腔包括具有由具有全横截面积的开放空腔限定的长度的第一部分和具有由减小的横截面面积限定的长度的第二部分，所述第二部分是 所述挡板在所述至少一个翼型腔内的长度。

    COOLING HOLE PATTERNED AIRFOIL
    185.
    发明申请
    COOLING HOLE PATTERNED AIRFOIL 审中-公开
    冷却孔图案空气

    公开(公告)号：US20160298464A1

    公开(公告)日：2016-10-13

    申请号：US14685410

    申请日：2015-04-13

    Applicant: United Technologies Corporation

    Inventor： Brandon W. Spangler ,  Gina Cavallo

    IPC: F01D5/18 ,  F01D9/04 ,  F01D25/12 ,  F01D5/14

    Abstract: An airfoil having a cooling hole pattern is disclosed. The cooling hole pattern may be an offset herringbone pattern. For instance, the airfoil may have rows of cooling holes arranged in filmrows, each filmrow divided into groups of cooling holes. A first group may be oriented to direct cooling air generally radially outward over a surface of the airfoil and a second group may be oriented to direct cooling air generally radially inward over a surface of the airfoil. Between the first group and the second group of cooling holes in each filmrow, a transition region exists. The adjacent filmrows are staggered to enhance the effectiveness of the convective cooling proximate to the transition regions by causing each filmrow to direct cooling air over the transition region of an adjacent filmrow.

    Abstract translation: 公开了具有冷却孔图案的翼型件。 冷却孔图案可以是偏移的人字形图案。 例如，翼型件可以具有排列在薄膜中的一排冷却孔，每个薄膜分成一组冷却孔。 第一组可以被定向成将冷却空气大致径向向外引导到翼型件的表面上，并且第二组可以被定向成将冷却空气大致径向向内引导到翼型件的表面上。 在每个成膜中的第一组和第二组冷却孔之间存在过渡区。 相邻的薄膜交错，以通过使每个薄膜引导冷却空气在相邻薄膜的过渡区域上来提高靠近过渡区域的对流冷却的有效性。

    GAS TURBINE ENGINE AIRFOIL HAVING SERPENTINE FED PLATFORM COOLING PASSAGE
    186.
    发明申请
    GAS TURBINE ENGINE AIRFOIL HAVING SERPENTINE FED PLATFORM COOLING PASSAGE 审中-公开
    具有SERPENTINE FED平台冷却通气的气体涡轮发动机

    公开(公告)号：US20160230567A1

    公开(公告)日：2016-08-11

    申请号：US15022622

    申请日：2014-09-12

    Applicant: UNITED TECHNOLOGIES CORPORATION

    Inventor： Scott W. Gayman ,  Brandon W. Spangler

    IPC: F01D5/18

    CPC classification number: F01D5/187 ,  F05D2240/81 ,  F05D2250/185

    Abstract: A gas turbine engine airfoil includes a platform, and spaced apart walls that provide an exterior airfoil surface that extends radially from the platform to an end opposite the platform. A serpentine cooling passage is arranged between the walls and has a first passageway that extends from the platform toward the end and a second passageway fluidly connecting to the first passageway and extending from the end toward the platform to an end. A platform cooling passageway is fluidly connected to the end and extends transversely into the platform. A cooling hole fluidly connects the platform cooling passageway to an exterior surface.

    Abstract translation: 燃气涡轮发动机翼型件包括平台和间隔开的壁，其提供从平台径向延伸到与平台相对的端部的外部翼面表面。 蛇形冷却通道布置在壁之间并且具有从平台朝向端部延伸的第一通道和流体连接到第一通道并从端部朝向平台延伸到端部的第二通道。 平台冷却通道流体地连接到端部并横向延伸到平台中。 冷却孔将平台冷却通道流畅地连接到外表面。

    PARTIAL TIP FLAG
    187.
    发明申请
    PARTIAL TIP FLAG 审中-公开
    部分提示标志

    公开(公告)号：US20160194965A1

    公开(公告)日：2016-07-07

    申请号：US14853182

    申请日：2015-09-14

    Applicant: United Technologies Corporation

    Inventor： Brandon W. Spangler

    IPC: F01D5/18 ,  F01D9/04

    CPC classification number: F01D5/187 ,  F01D5/186 ,  F01D9/041 ,  F05D2220/32 ,  F05D2230/00 ,  F05D2240/12 ,  F05D2240/30 ,  F05D2240/303 ,  F05D2240/304 ,  F05D2250/185 ,  F05D2260/202 ,  Y02T50/673 ,  Y02T50/676

    Abstract: A rotor blade of a turbine engine may have internal passages to permit the travel of cooling air through the blade. These passages may include a tip flag, a serpentine channel, and a trailing edge channel. The tip flag may extend radially outward along the leading edge of the rotor blade and may turn axially aftward along the tip of the rotor blade. The tip flag may terminate forward of a portion of the serpentine channel and the trailing edge channel. Thus the tip flag may be a “partial tip flag.” The internal passages may be arranged to ameliorate the effect of ambient pressure variations, such as between the leading edge and the trailing edge of the rotor blade, on the flow travel of cooling air through the rotor blade.

    Abstract translation: 涡轮发动机的转子叶片可以具有允许冷却空气通过叶片的行进的内部通道。 这些通道可以包括尖端标记，蛇形通道和后缘通道。 尖端标志可沿着转子叶片的前缘径向向外延伸，并且可沿转子叶片的尖端轴向向后转动。 尖端标志可以在蛇形通道和后缘通道的一部分的前方终止。 因此，尖端标记可以是“部分尖端标志”。内部通道可以被布置成改善环境压力变化（例如在转子叶片的前缘和后缘之间）对冷却空气的流动的影响 通过转子叶片。

    GAS TURBINE ENGINE TURBINE BLADE TIP COOLING
    189.
    发明申请
    GAS TURBINE ENGINE TURBINE BLADE TIP COOLING 审中-公开
    气体涡轮发动机涡轮叶片冷却

    公开(公告)号：US20160102561A1

    公开(公告)日：2016-04-14

    申请号：US14853286

    申请日：2015-09-14

    Applicant: United Technologies Corporation

    Inventor： Brandon W. Spangler ,  Laurie L. Burstynski

    IPC: F01D5/18 ,  F01D5/14

    CPC classification number: F01D5/186 ,  F01D5/141 ,  F01D5/20 ,  F05D2220/32 ,  F05D2240/307 ,  F05D2260/202 ,  Y02T50/676

    Abstract: An airfoil for a gas turbine engine includes pressure and suction walls that are spaced apart from one another and joined at leading and trailing edges to provide an airfoil that has an exterior surface that extends in a radial direction to a tip. A film cooling hole is provided in the tip and extends at an angle relative to the radial direction. The film cooling hole includes a diffuser.

    Abstract translation: 用于燃气涡轮发动机的翼型件包括彼此间隔开并在前缘和后缘连接的压力和抽吸壁，以提供具有沿径向方向延伸到尖端的外表面的翼型件。 薄片冷却孔设置在尖端中并相对于径向成一定角度延伸。 薄膜冷却孔包括扩散器。

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