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    CONSTRUCTION OF MULTI-LAYERED REFRACTORY METAL CORE FOR INVESTMENT CASTING
    12.
    发明公开
    CONSTRUCTION OF MULTI-LAYERED REFRACTORY METAL CORE FOR INVESTMENT CASTING 审中-公开

    公开(公告)号：EP3527302A1

    公开(公告)日：2019-08-21

    申请号：EP19157502.6

    申请日：2019-02-15

    申请人： United Technologies Corporation

    发明人： AUXIER, James Tilsley

    IPC分类号： B22C3/00 ,  B22C9/10 ,  B22C9/12 ,  B22C9/22 ,  B22F7/08

    摘要： A refractory metal core laminate assembly, comprising: a first refractory metal core (64) having exterior surfaces (66) and a first side (68) and a second side (70) opposite the first side (68); a second refractory metal core (72) having exterior surfaces (74) and a first side (76) and a second side (78) opposite the first side (76), the second refractory metal core (72) being arranged above the first refractory metal core (64) with the second refractory metal core first side (76) facing the first refractory metal core second side (70); a layer (82) of a powder bed material (80) between the first refractory metal core second side (70) and the second refractory metal core first side (76); and a coating (92) of the powder bed material (80) coupled to the first refractory metal core exterior surfaces (66) and the second refractory metal core exterior surfaces (74).

    COMBUSTOR PANEL STANDOFF PIN
    13.
    发明公开
    COMBUSTOR PANEL STANDOFF PIN 审中-公开

    公开(公告)号：EP3524885A1

    公开(公告)日：2019-08-14

    申请号：EP19155605.9

    申请日：2019-02-05

    申请人： United Technologies Corporation

    发明人： PORTER, Steven D. ,  AUXIER, James Tilsley

    IPC分类号： F23R3/00 ,  F23R3/50 ,  F23R3/60

    摘要： A heat shield panel (126, 128; 402, 404) for use in a gas turbine engine combustor (56) is disclosed. The heat shield panel (126, 128; 402, 404) includes a hot side, a cold side and at least one attachment mechanism (275; 406) having a stud (250; 408) and a central axis extending through the stud (250; 408) and a plurality of standoff pins (154; 254; 300; 410) positioned circumferentially around the stud (250; 408), the standoff pins (154; 254; 300; 410) having a radial extent, a circumferential extent that is greater than the radial extent, a radially outer surface having a radially convex shape and a radially inner surface having a radially concave shape.

    DUAL-FED AIRFOIL TIP
    15.
    发明公开
    DUAL-FED AIRFOIL TIP 有权
    双翼飞机翼尖

    公开(公告)号：EP3196414A1

    公开(公告)日：2017-07-26

    申请号：EP17152351.7

    申请日：2017-01-20

    申请人： United Technologies Corporation

    发明人： AUXIER, James Tilsley ,  ZELESKY, Mark F. ,  TELLER, Bret M. ,  JARIWALA, Parth ,  MONGILLO, Dominic J.

    IPC分类号： F01D5/18 ,  F01D5/30 ,  F02C3/16

    CPC分类号： F01D5/187 ,  F01D9/02 ,  F01D25/12 ,  F04D29/324 ,  F04D29/542 ,  F04D29/582 ,  F05D2220/32 ,  F05D2250/185 ,  F05D2260/202 ,  F05D2260/22141 ,  F05D2260/606 ,  Y02T50/676

    摘要： An airfoil (201; 301; 401; 501) of a gas turbine engine is provided including a leading edge (212; 312) extending in a radial direction, a tip (232; 332; 432; 532) extending in an axial direction from the leading edge, a first rib (232; 332; 432; 532) extending radially within the airfoil, the leading edge and the first rib defining a leading edge cavity (220; 320; 420; 520) within the airfoil, a second rib (234; 334), the second rib and the first rib defining a serpentine cavity (228; 328; 428; 528) therein, a third rib (236; 336; 436; 536) extending axially within the tip, a flag tip cavity (226; 326; 426; 526) defined by the third rib, the leading edge, and the tip, the leading edge cavity fluidly connected to the flag tip cavity, and a bypass aperture (240; 340; 440; 540) formed between the first rib and the third rib, the bypass aperture configured to fluidly connect the serpentine cavity with the flag tip cavity.

    摘要翻译： 提供了一种燃气涡轮发动机的翼型件（201; 301; 401; 501），该翼型件包括在径向方向上延伸的前缘（212; 312），沿轴向方向延伸的尖端（232; 332; 432; 532） 所述前缘，在所述翼型中径向延伸的第一肋（232; 332; 432; 532），所述前缘和所述第一肋在所述翼型内限定出前缘腔（220; 320; 420; 520）;第二肋 （234; 334），所述第二肋和所述第一肋在其中限定了蛇形腔（228; 328; 428; 528），在所述尖端内轴向延伸的第三肋（236; 336; 436; 536） 由所述第三肋，所述前缘和所述尖端限定的所述前缘腔（226; 326; 426; 526），所述前缘腔流体地连接至所述旗尖端腔;以及旁通孔（240; 340; 440; 540） 所述第一肋和所述第三肋，所述旁路孔配置成将所述蛇形腔与所述旗尖端腔流体连接。

    METHOD OF MANUFACTURE OF AIRFOIL CASTINGS USING AUTONOMOUS ADAPTIVE MACHINING
    17.
    发明公开
    METHOD OF MANUFACTURE OF AIRFOIL CASTINGS USING AUTONOMOUS ADAPTIVE MACHINING 审中-公开

    公开(公告)号：EP3682984A1

    公开(公告)日：2020-07-22

    申请号：EP20152313.1

    申请日：2020-01-16

    申请人： United Technologies Corporation

    发明人： AUXIER, James Tilsley ,  BARRON, Alan C. ,  BALZER, Wolfgang

    IPC分类号： B22C9/18 ,  B22D31/00 ,  F01D5/18 ,  G05B19/4097

    摘要： A method (100) of forming an airfoil (28) includes casting the airfoil (28) with an internal cooling circuit and an exterior surface (30) with a positive feature (32, 34). The exterior surface (30) of the airfoil (28) is scanned with a first probe (20). A size and a location of the positive feature (32, 34) are identified based on the scan of the exterior surface (30). A transformation matrix is created with a controller (12) such that the transformation matrix includes toolpath transformation instructions. A transformed set of machine toolpath instructions is created by applying the transformation matrix using the controller (12) to a first set of machine toolpath instructions to align the first set of machine toolpath instructions relative to the positive feature (32, 34). A contour is then machined into the exterior surface (30) of the airfoil (28) based on the transformed set of machine toolpath instructions.

    GAS TURBINE ENGINE BLADE, CORRESPONDING GAS TURBINE ENGINE AND METHOD FOR A GAS TURBINE ENGINE BLADE
    19.
    发明公开
    GAS TURBINE ENGINE BLADE, CORRESPONDING GAS TURBINE ENGINE AND METHOD FOR A GAS TURBINE ENGINE BLADE 审中-公开
    燃气轮机发动机叶片，相应的燃气涡轮发动机及用于燃气涡轮发动机叶片的方法

    公开(公告)号：EP3296511A2

    公开(公告)日：2018-03-21

    申请号：EP17187494.4

    申请日：2017-08-23

    申请人： United Technologies Corporation

    发明人： JENNINGS, Timothy J. ,  NIEZELSKI, David A. ,  AUXIER, James Tilsley

    IPC分类号： F01D5/18

    CPC分类号： F01D5/187 ,  F01D5/186 ,  F01D11/08 ,  F02C3/04 ,  F05D2220/32 ,  F05D2240/307 ,  F05D2260/202 ,  Y02T50/676

    摘要： A gas turbine engine (20) blade (64) includes a platform (68) that has an inner side (68A) and an outer side (68B), a root (70) that extends outwardly from the inner side (68A), and an airfoil (72) that extends outwardly from a base (72A) at the outer side (68B) to a tip end (72B). The airfoil (72) includes a leading edge and a trailing edge and a first side wall (72C) and a second side wall (72D). The first side wall (72C) and the second side wall (72D) join the leading edge and the trailing edge and at least partially define one or more cavities (74) in the airfoil (72). The airfoil (72) has a span from the base (72A) to the tip end (72B), with the base (72A) being at 0% of the span and the tip end (72B) being at 100% of the span. The first side wall (72C) includes an axial row (76) of cooling holes (78) at 90% or greater of the span.

    摘要翻译： 燃气涡轮发动机（20）叶片（64）包括具有内侧（68A）和外侧（68B），从内侧（68A）向外延伸的根部（70）的平台（68），以及 在外侧（68B）处从基部（72A）向外延伸至尖端（72B）的翼型件（72）。 翼型件（72）包括前缘和后缘以及第一侧壁（72C）和第二侧壁（72D）。 第一侧壁（72C）和第二侧壁（72D）连接前缘和后缘并且至少部分地限定翼型件（72）中的一个或多个空腔（74）。 翼型件72具有从基座72A到尖端72B的跨度，其中基座72A占跨距的0％并且末端72B处于跨距的100％。 第一侧壁（72C）包括在跨度的90％或更大处的轴向排（76）的冷却孔（78）。