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    AIRFOIL HAVING IMPROVED LEADING EDGE COOLING SCHEME AND DAMAGE RESISTANCE
    4.
    发明公开
    AIRFOIL HAVING IMPROVED LEADING EDGE COOLING SCHEME AND DAMAGE RESISTANCE 审中-公开

    公开(公告)号：EP3567219A1

    公开(公告)日：2019-11-13

    申请号：EP19172133.1

    申请日：2019-05-01

    Applicant: United Technologies Corporation

    Inventor： JENNINGS, Timothy J. ,  PROPHETER-HINCKLEY, Tracy A. ,  LANA, Kyle C.

    IPC: F01D5/18

    Abstract: Airfoil including an airfoil body (402) extending between a leading edge (412) and a trailing edge (414) in axial direction, between a pressure side (416) and a suction side (418) in circumferential direction, and between a root (406) and a tip (408) in radial direction, a first transitioning leading edge cavity (422) located proximate the leading edge proximate the root of the airfoil body and transitioning axially toward the trailing edge as the first transitioning leading edge cavity extends radially toward the tip, and a second transitioning leading edge cavity (424) located aft of the first transitioning leading edge cavity proximate the root of the airfoil body and transitioning axially toward the leading edge as the second transitioning leading edge cavity extends radially toward the tip. The second transitioning leading edge cavity includes an impingement sub-cavity (424a) and a film sub-cavity (424b) along the leading edge and proximate the tip.

    AIRFOIL FOR A GAS TURBINE ENGINE AND CORRESPONDING CORE STRUCUTURE FOR MANUFACTURING AN AIRFOIL
    6.
    发明公开
    AIRFOIL FOR A GAS TURBINE ENGINE AND CORRESPONDING CORE STRUCUTURE FOR MANUFACTURING AN AIRFOIL 审中-公开

    公开(公告)号：EP3467266A1

    公开(公告)日：2019-04-10

    申请号：EP18198482.4

    申请日：2018-10-03

    Applicant: United Technologies Corporation

    Inventor： PROPHETER-HINCKLEY, Tracy A. ,  CLUM, Carey ,  JENNINGS, Timothy J.

    IPC: F01D5/18 ,  F01D9/06 ,  B22C13/10

    Abstract: An airfoil for a gas turbine engine comprises an airfoil body with a first core cavity (702) and a second core cavity (704) located within the airfoil body that is adjacent the first core cavity. The second core cavity is defined by a first cavity wall (712), a second cavity wall (714), a first exterior wall (716), and a second exterior wall (718), wherein the first cavity wall is located between the second core cavity and the first core cavity and the first and second exterior walls are exterior walls of the airfoil body. The first cavity wall includes a first surface (724) angled toward the first exterior wall and a second surface (726) angled toward the second exterior wall. At least one first cavity impingement hole (720) is formed within the first surface and a central ridge (742, 744) extends into the second core cavity from at least one of the first cavity wall and the second wall and divides the second core cavity into a two-vortex chamber. A core structure for manufacturing an airfoil for a gas turbine engine comprises ajdacent first and second core cavity cores to form first and second core cavities. At least one first cavity impingement stem extends between the core cavity cores to form at least one first cavity impingement hole in a first cavity wall and a central channel extends into the second core cavity core to form a central ridge which divides the second core cavity into a two-vortex chamber.

    FILM HOLE WITH PROTRUDING FLOW ACCUMULATOR
    7.
    发明公开
    FILM HOLE WITH PROTRUDING FLOW ACCUMULATOR 审中-公开
    FILMKHHLLOCH MIT VORSTEHENDEMSTRÖMUNGSAKKUMULATOR

    公开(公告)号：EP3012407A1

    公开(公告)日：2016-04-27

    申请号：EP15190346.5

    申请日：2015-10-19

    Applicant: United Technologies Corporation

    Inventor： MUSTO, Clifford J. ,  LoRICCO, Nicholas M. ,  SLAVENS, Thomas N. ,  JENNINGS, Timothy J. ,  CLUM, Carey ,  MOORE, Sasha M. ,  McBRIEN, John ,  QUACH, San ,  WHITFIELD, Christopher

    IPC: F01D5/18

    CPC classification number: F01D5/186 ,  B22C9/10 ,  B33Y10/00 ,  B33Y80/00 ,  F01D5/187 ,  F05D2230/21 ,  F05D2230/22 ,  F05D2230/31 ,  F05D2240/127 ,  F05D2260/202 ,  F05D2260/2212 ,  Y02T50/676

    Abstract: A wall (204; 302; 402; 502) of a gas turbine engine component/part is provided. The wall (204; 302; 402; 502) may comprise an external surface (207; 301; 403; 503) adjacent a gas path (222; 322; 416; 518) and an internal surface (205; 303; 401; 501) adjacent an internal flow path (220; 320; 412; 512). A film hole (213; 304; 404; 504) may have an inlet (214; 308) at the internal surface (205; 303; 401; 501) and an outlet (212; 306) at the external surface (207; 301; 403; 503). A flow accumulator (206; 312; 406; 506) adjacent the inlet (214; 308) may protrude from the internal surface (205; 303; 401; 501). A method of making an engine component is also provided and comprises the step of forming a component wall (204; 302; 402; 502) comprising an accumulator (206; 312; 406; 506) on an internal surface (205; 303; 401; 501) and a film hole (213; 304; 404; 504) defined by the component wall (204; 302; 402; 502). The film hole (213; 304; 404; 504) may include an opening adjacent the accumulator (206; 312; 406; 506) and defined by the internal surface (205; 303; 401; 501).

    Abstract translation: 提供了一种燃气涡轮发动机部件/部件的壁（204; 302; 402; 502）。 所述壁（204; 302; 402; 502）可包括邻近气体路径（222; 322; 416; 518）和内表面（205; 303; 401; 501）的外表面（207; 301; 403; ）相邻于内部流动路径（220; 320; 412; 512）。 薄膜孔（213; 304; 404; 504）可以在内表面（205; 303; 401; 501）处具有入口（214; 308）和外表面（207,301）处的出口（212; 306） ; 403; 503）。 邻近入口（214; 308）的流动蓄能器（206; 312; 406; 506）可从内表面（205; 303; 401; 501）突出。 还提供了制造发动机部件的方法，包括在内表面（205; 303; 401）上形成包括蓄能器（206; 312; 406; 506）的部件壁（204; 302; 402; 502） ; 501）和由所述部件壁（204; 302; 402; 502）限定的薄膜孔（213; 304; 404; 504）。 薄膜孔（213; 304; 404; 504）可以包括邻近储存器（206; 312; 406; 506）并由内表面（205; 303; 401; 501）限定的开口。

    GAS TURBINE ENGINE COMPONENT WITH FILM COOLING HOLE FEATURE
    8.
    发明公开
    GAS TURBINE ENGINE COMPONENT WITH FILM COOLING HOLE FEATURE 审中-公开
    GASTURBINENBAUTEIL MITFILMKÜHLUNGSÖFFNUNG。

    公开(公告)号：EP3009599A1

    公开(公告)日：2016-04-20

    申请号：EP15190282.2

    申请日：2015-10-16

    Applicant: United Technologies Corporation

    Inventor： SLAVENS, Thomas N. ,  JENNINGS, Timothy J. ,  PEREZ, Roberto J.

    IPC: F01D5/18

    CPC classification number: F01D5/187 ,  F01D5/186 ,  F01D9/041 ,  F01D11/08 ,  F01D25/12 ,  F02C3/04 ,  F02C7/18 ,  F04D29/083 ,  F04D29/324 ,  F04D29/542 ,  F04D29/582 ,  F05D2220/32 ,  F05D2240/127 ,  F05D2240/35 ,  F05D2240/81 ,  F05D2260/202 ,  F05D2260/2212 ,  Y02T50/676

    Abstract: A gas turbine engine component (64) includes a wall (94) that provides an exterior surface (79) and an interior flow path surface (96). A film cooling hole (92) extends through the wall (94) and is configured to fluidly connect the interior flow path surface (96) to the exterior surface (79). The film cooling hole (92) has a diffuser (100) that is arranged downstream from a metering hole (98). The diffuser (100) includes inner and outer diffuser surfaces (104,102) opposite one another and respectively arranged on sides near the interior flow path surface (96)and the exterior surface (79). A protrusion (106) is arranged in the diffuser on the outer diffuser surface (102).

    Abstract translation: 燃气涡轮发动机部件（64）包括提供外表面（79）和内部流路表面（96）的壁（94）。 膜冷却孔（92）延伸穿过壁（94）并且构造成将内部流路表面（96）流体连接到外表面（79）。 膜冷却孔（92）具有布置在计量孔（98）下游的扩散器（100）。 扩散器（100）包括彼此相对的分别布置在内部流路表面（96）和外表面（79）附近的侧面的内部和外部扩散器表面（104,102）。 在外扩散器表面（102）上的扩散器中布置有突起（106）。

    AIRFOIL HAVING IMPROVED LEADING EDGE COOLING SCHEME AND DAMAGE RESISTANCE
    9.
    发明公开
    AIRFOIL HAVING IMPROVED LEADING EDGE COOLING SCHEME AND DAMAGE RESISTANCE 审中-公开

    公开(公告)号：EP3567218A1

    公开(公告)日：2019-11-13

    申请号：EP19172115.8

    申请日：2019-05-01

    Applicant: United Technologies Corporation

    Inventor： JENNINGS, Timothy J. ,  PROPHETER-HINCKLEY, Tracy A. ,  LANA, Kyle C.

    IPC: F01D5/18

    Abstract: Airfoil for a gas turbine including a body (402) extending between leading (412) and trailing edges (414) in an axial direction, between pressure (416) and suction sides (418) in a circumferential direction, and between a root (406) and tip (408) in a radial direction. A first transitioning leading edge cavity (422) is located adjacent one of the sides proximate the root of the body and transitions axially toward the leading edge as the first transitioning leading edge cavity extends radially toward the tip. A second transitioning leading edge cavity (424) is adjacent the other side and adjacent the leading edge proximate the root of the body and transitions axially toward the trailing edge as the second transitioning leading edge cavity extends radially toward the tip. A portion of the second transitioning leading edge cavity shields a portion of the first transitioning leading edge cavity proximate the root of the body.

    GAS TURBINE AIRFOIL COOLING CIRCUIT
    10.
    发明公开
    GAS TURBINE AIRFOIL COOLING CIRCUIT 审中-公开

    公开(公告)号：EP3477053A1

    公开(公告)日：2019-05-01

    申请号：EP18201749.1

    申请日：2018-10-22

    Applicant: United Technologies Corporation

    Inventor： JENNINGS, Timothy J. ,  CLUM, Carey

    IPC: F01D5/14 ,  F01D5/18

    Abstract: An airfoil (100) may include an airfoil body (110) that defines a central chamber (114), a skin chamber (112), and an impingement hole (113) extending between the central chamber (114) and the skin chamber (112). The central chamber (114) may be in fluidic communication with the skin chamber (112) via the impingement hole (113). In various embodiments, a first cross-sectional area of the impingement hole (113) is greater than about 25% of a second cross-sectional area of the skin chamber (112). In various embodiments, the impingement hole (113) is positioned and configured to deliver cooling circuit air to a predicted position of a hotspot on a surface of the airfoil (100). In various embodiments, the airfoil body (110) further defines at least one structural hole formed from at least one structural core tie, wherein a first cross-sectional area of the impingement hole (113) is at least twice a second cross-sectional area of the at least one structural hole.

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