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    Casting core for a cooling arrangement for a gas turbine component
    122.
    发明授权
    Casting core for a cooling arrangement for a gas turbine component 有权
    用于燃气轮机部件的冷却装置的铸芯

    公开(公告)号：US08936067B2

    公开(公告)日：2015-01-20

    申请号：US13658045

    申请日：2012-10-23

    Applicant: Ching-Pang Lee ,  Benjamin E. Heneveld

    Inventor： Ching-Pang Lee ,  Benjamin E. Heneveld

    IPC: B22C9/10

    CPC classification number: B22C9/10 ,  F01D5/186 ,  F01D5/187 ,  F05D2230/21 ,  F05D2230/211

    Abstract: A ceramic casting core, including: a plurality of rows (162, 166, 168) of gaps (164), each gap (164) defining an airfoil shape; interstitial core material (172) that defines and separates adjacent gaps (164) in each row (162, 166, 168); and connecting core material (178) that connects adjacent rows (170, 174, 176) of interstitial core material (172). Ends of interstitial core material (172) in one row (170, 174, 176) align with ends of interstitial core material (172) in an adjacent row (170, 174, 176) to form a plurality of continuous and serpentine shaped structures each including interstitial core material (172) from at least two adjacent rows (170, 174, 176) and connecting core material (178).

    Abstract translation: 一种陶瓷铸造芯，包括：多个间隔（164）的排（162,166,168），每个间隙（164）限定翼型形状; 间隙芯材料（172），其限定和分隔每排（162,166,168）中的相邻间隙（164）; 以及连接芯材料（178），其连接间隙芯材料（172）的相邻行（170,174,176）。 一排（170,174,176）中的间隙芯材（172）的端部与相邻排（170,174,176）中的间隙芯材（172）的端部对准，以形成多个连续和蛇形形状的结构 包括来自至少两个相邻行（170,174,176）和连接芯材（178）的间隙芯材料（172）。

    AFT OUTER RIM SEAL ARRANGEMENT
    123.
    发明申请
    AFT OUTER RIM SEAL ARRANGEMENT 有权
    AFT外部RIM密封安排

    公开(公告)号：US20150003973A1

    公开(公告)日：2015-01-01

    申请号：US13930482

    申请日：2013-06-28

    Applicant: Ching-Pang Lee ,  Kok-Mun Tham ,  Eric Schroeder ,  Jamie Meeroff ,  Samuel R. Miller, Jr. ,  John J. Marra ,  Christian X. Campbell

    Inventor： Ching-Pang Lee ,  Kok-Mun Tham ,  Eric Schroeder ,  Jamie Meeroff ,  Samuel R. Miller, Jr. ,  John J. Marra ,  Christian X. Campbell

    IPC: F01D11/02 ,  F01D5/18

    CPC classification number: F01D11/02 ,  F01D5/082 ,  F01D5/18 ,  F01D5/187 ,  F01D11/04 ,  F05D2260/14

    Abstract: An outer rim seal arrangement (10), including: an annular rim (70) centered about a longitudinal axis (30) of a rotor disc (31), extending fore and having a fore-end (72), an outward-facing surface (74), and an inward-facing surface (76); a lower angel wing (62) extending aft from a base of a turbine blade (22) and having an aft end (64) disposed radially inward of the rim inward-facing surface to define a lower angel wing seal gap (80); an upper angel wing (66) extending aft from the turbine blade base and having an aft end (68) disposed radially outward of the rim outward-facing surface to define a upper angel wing seal gap (80, 82); and guide vanes (100) disposed on the rim inward-facing surface in the lower angel wing seal gap. Pumping fins (102) may be disposed on the upper angel wing seal aft end in the upper angel wing seal gap.

    Abstract translation: 外缘密封装置（10），包括：以转子盘（31）的纵向轴线（30）为中心的环形边缘（70），其在前面延伸并具有前端（72），面向外的表面 （74）和面向内的表面（76）; 从天线叶片（22）的底部向后延伸的下部天使翼（62），并且具有设置在所述边缘向内相对表面的径向内侧的后端（64），以限定下部天使翼密封间隙（80）; 从所述涡轮叶片基部向后延伸的上天使翼（66），并且具有设置在所述轮缘朝外的表面的径向外侧的后端（68），以限定上天使翼密封间隙（80,82）; 以及设置在下天使翼密封间隙中的边缘向内表面上的导向叶片（100）。 抽吸翼（102）可以设置在上天使翼密封间隙中的上天使翼密封后端。

    SEAL ASSEMBLY INCLUDING GROOVES IN AN INNER SHROUD IN A GAS TURBINE ENGINE
    126.
    发明申请
    SEAL ASSEMBLY INCLUDING GROOVES IN AN INNER SHROUD IN A GAS TURBINE ENGINE 有权
    密封组件，包括在天然气涡轮发动机内部的油井

    公开(公告)号：US20140286760A1

    公开(公告)日：2014-09-25

    申请号：US13747868

    申请日：2013-01-23

    Applicant: Ching-Pang Lee

    Inventor： Ching-Pang Lee

    IPC: F02C7/28

    CPC classification number: F02C7/28 ,  F01D5/082 ,  F01D11/001 ,  F01D11/04 ,  F05D2250/71

    Abstract: A seal assembly between a disc cavity and a hot gas path in a gas turbine engine includes a rotating blade assembly having a plurality of blades that rotate with a turbine rotor during operation of the engine, and a stationary vane assembly having a plurality of vanes and an inner shroud. The inner shroud includes a radially outwardly facing first surface, a radially inwardly facing second surface, and a plurality of grooves extending into the second surface. The grooves are arranged such that a space having a component in a circumferential direction is defined between adjacent grooves. During operation of the engine, the grooves guide purge air out of the disc cavity toward the hot gas path such that the purge air flows in a desired direction with reference to a direction of hot gas flow through the hot gas path.

    Abstract translation: 燃气涡轮发动机中的盘腔和热气路之间的密封组件包括旋转叶片组件，其具有在发动机运转期间与涡轮转子一起旋转的多个叶片，以及具有多个叶片和 内罩。 内护罩包括径向向外的第一表面，径向向内的第二表面和延伸到第二表面中的多个凹槽。 凹槽被布置成使得在相邻凹槽之间限定具有圆周方向的部件的空间。 在发动机操作期间，凹槽将吹扫空气从盘腔朝向热气路径引导，使得净化空气相对于热气流通过热气路径的方向在所需方向上流动。

    Trailing edge cooling system in a turbine airfoil assembly
    127.
    发明授权
    Trailing edge cooling system in a turbine airfoil assembly 有权
    涡轮机翼组件中的后缘冷却系统

    公开(公告)号：US08840363B2

    公开(公告)日：2014-09-23

    申请号：US13228567

    申请日：2011-09-09

    Applicant: Ching-Pang Lee

    Inventor： Ching-Pang Lee

    IPC: F01D5/18

    CPC classification number: F01D5/186 ,  F05D2240/122 ,  F05D2240/304 ,  F05D2250/183 ,  F05D2260/202

    Abstract: An airfoil in a gas turbine engine includes an outer wall, a cooling fluid cavity, and a plurality of cooling fluid passages. The outer wall has a leading edge, a trailing edge, a pressure side, a suction side, and radially inner and outer ends. The cooling fluid cavity is defined in the outer wall, extends generally radially between the inner and outer ends of the outer wall, and receives cooling fluid for cooling the outer wall. The cooling fluid passages are in fluid communication with the cooling fluid cavity and include zigzagged passages that include alternating angled sections, each section having both a radial component and a chordal component. The cooling fluid passages extend from the cooling fluid cavity toward the trailing edge of the outer wall and receive cooling fluid from the cooling fluid cavity for cooling the outer wall near the trailing edge.

    Abstract translation: 燃气涡轮发动机中的翼型件包括外壁，冷却流体腔和多个冷却流体通道。 外壁具有前缘，后缘，压力侧，吸力侧以及径向内外端。 冷却流体腔被限定在外壁中，大致径向延伸在外壁的内端和外端之间，并接收用于冷却外壁的冷却流体。 冷却流体通道与冷却流体腔流体连通，并且包括包括交替成角度部分的锯齿形通道，每个部分都具有径向部件和弦部件。 冷却流体通道从冷却流体腔朝向外壁的后缘延伸，并接收来自冷却流体腔的冷却流体，用于冷却靠近后缘的外壁。

    TURBINE BLADE WITH INTEGRATED SERPENTINE AND AXIAL TIP COOLING CIRCUITS
    128.
    发明申请
    TURBINE BLADE WITH INTEGRATED SERPENTINE AND AXIAL TIP COOLING CIRCUITS 有权
    具有集成SERPENTINE和轴向冷却电路的涡轮叶片

    公开(公告)号：US20140169962A1

    公开(公告)日：2014-06-19

    申请号：US13714518

    申请日：2012-12-14

    Applicant: Ching-Pang Lee

    Inventor： Ching-Pang Lee

    IPC: F01D5/18

    CPC classification number: F01D5/187 ,  F01D5/186 ,  F01D5/20 ,  F05D2240/307 ,  F05D2250/185 ,  F05D2250/75 ,  F05D2260/201 ,  F05D2260/202 ,  F05D2260/2212 ,  F05D2260/2214

    Abstract: An air cooled turbine blade including leading and trailing edges, and pressure and suction side walls extending between the leading and trailing edges. Leading and trailing edge cooling circuits extend spanwise adjacent to the leading and trailing edges, respectively. A forward flow mid-section serpentine cooling circuit extends spanwise and is located between the leading and trailing edge cooling circuits. An axial tip cooling circuit extends in the chordal direction and is located between a tip cap of the blade and the serpentine cooling circuit at an outer end of the serpentine cooling circuit. The axial tip cooling circuit has a forward end receiving cooling air from a final channel of the serpentine cooling circuit and discharges the cooling air adjacent to the trailing edge.

    Abstract translation: 包括前缘和后缘的空气冷却涡轮叶片，以及在前缘和后缘之间延伸的压力和吸力侧壁。 前缘和后缘冷却回路分别沿着前缘和后缘分别延伸。 向前流动的中段蛇形冷却回路沿横向延伸并且位于前缘和后缘冷却回路之间。 轴向尖端冷却回路在弦向延伸，位于蛇形冷却回路的外端处，位于叶片的顶盖和蛇形冷却回路之间。 轴向尖端冷却回路具有从蛇形冷却回路的最终通道接收冷却空气的前端，并排出与后缘相邻的冷却空气。

    Ring segment with serpentine cooling passages
    129.
    发明授权
    Ring segment with serpentine cooling passages 有权
    带有蛇形冷却通道的环段

    公开(公告)号：US08727704B2

    公开(公告)日：2014-05-20

    申请号：US13213417

    申请日：2011-08-19

    Applicant: Ching-Pang Lee ,  Eric C. Berrong

    Inventor： Ching-Pang Lee ,  Eric C. Berrong

    IPC: F01D11/08

    CPC classification number: F01D11/08 ,  F05D2250/70 ,  F05D2260/20

    Abstract: A ring segment for a gas turbine engine includes a panel and a cooling system. The cooling system receives cooling fluid from an outer side of the panel for cooling the panel and includes at least one cooling fluid supply passage, at least one serpentine cooling passage, and at least one cooling fluid discharge passage. The cooling fluid supply passage(s) receive the cooling fluid from the outer side of the panel and deliver the cooling fluid to a first cooling fluid chamber within the panel. The serpentine cooling passage(s) receive the cooling fluid from the first cooling fluid chamber, wherein the cooling fluid provides convective cooling to the panel as it passes through the serpentine cooling passage(s). The cooling fluid discharge passage(s) discharge the cooling fluid from the cooling system.

    Abstract translation: 用于燃气涡轮发动机的环段包括面板和冷却系统。 冷却系统从面板的外侧接收冷却流体，用于冷却面板，并且包括至少一个冷却流体供应通道，至少一个蛇形冷却通道和至少一个冷却流体排出通道。 冷却流体供应通道从面板的外侧接收冷却流体，并将冷却流体输送到面板内的第一冷却流体室。 蛇形冷却通道从第一冷却流体室接收冷却流体，其中冷却流体在通过蛇形冷却通道时向面板提供对流冷却。 冷却流体排出通道从冷却系统排出冷却流体。

    COOLING ARRANGEMENT FOR A GAS TURBINE COMPONENT
    130.
    发明申请
    COOLING ARRANGEMENT FOR A GAS TURBINE COMPONENT 有权
    燃气涡轮机组件的冷却装置

    公开(公告)号：US20140112799A1

    公开(公告)日：2014-04-24

    申请号：US13657923

    申请日：2012-10-23

    Applicant: CHING-PANG LEE ,  BENJAMIN E. HENEVELD

    Inventor： CHING-PANG LEE ,  BENJAMIN E. HENEVELD

    IPC: F01D5/18

    CPC classification number: F01D5/187 ,  F05D2250/185 ,  F05D2260/22141

    Abstract: A cooling arrangement (82) for a gas turbine engine component, the cooling arrangement (82) having a plurality of rows (92, 94, 96) of airfoils (98), wherein adjacent airfoils (98) within a row (92, 94, 96) define segments (110, 130, 140) of cooling channels (90), and wherein outlets (114, 134) of the segments (110, 130) in one row (92, 94) align aerodynamically with inlets (132, 142) of segments (130, 140) in an adjacent row (94, 96) to define continuous cooling channels (90) with non continuous walls (116, 120), each cooling channel (90) comprising a serpentine shape.

    Abstract translation: 一种用于燃气涡轮发动机部件的冷却装置（82），所述冷却装置（82）具有多排（92,94,96）的翼型件（98），其中在一排（92,94）内相邻的翼型件 ，96）限定冷却通道（90）的段（110,130,140），并且其中一排（92,94）中的段（110,130）的出口（114,134）在空气动力学上与入口（132， 142）在相邻排（94,96）中的段（130,140）中以限定具有非连续壁（116,120）的连续冷却通道（90），每个冷却通道（90）包括蛇形形状。

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