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    TURBINE AIRFOIL COOLING SYSTEM WITH CHORDWISE EXTENDING SQUEALER TIP COOLING CHANNEL
    1.
    发明申请
    TURBINE AIRFOIL COOLING SYSTEM WITH CHORDWISE EXTENDING SQUEALER TIP COOLING CHANNEL 审中-公开

    公开(公告)号：US20170370232A1

    公开(公告)日：2017-12-28

    申请号：US15544112

    申请日：2015-01-22

    Applicant: Siemens Energy, Inc.

    Inventor： Ching-Pang Lee

    IPC: F01D5/18

    CPC classification number: F01D5/187 ,  F01D5/08 ,  F01D5/186 ,  F01D5/20 ,  F05D2220/32 ,  F05D2240/307 ,  F05D2260/202 ,  F05D2260/2212

    Abstract: An internal cooling system (10) for an airfoil (12) in a turbine engine (14) whereby the cooling system (10) includes a chordwise extending tip cooling channel (16) radially inward of a squealer tip (18) and formed at least in part by an inner wall (20) with a nonlinear outer surface (22) is disclosed. The nonlinear outer surface (22) of the inner wall (20) of the chordwise extending tip cooling channel (16) may be formed from pressure and suction side sections (24, 26) that intersect at a point (74) that is closer to the inner surface (30) of an outer wall (32) forming at least a portion of the squealer tip (18) than other aspects of the pressure side section (24) and the suction side section (26). The configurations of the pressure and suction side sections (24, 26) reduces the flow cross-sectional area, which accelerates the cooling fluid flow in a chordwise direction within the chordwise extending tip cooling channel (16) and directs cooling fluid toward the pressure and suction side outer walls (34, 36) for improved cooling efficiency.

    TURBINE AIRFOIL COOLING SYSTEM WITH PLATFORM COOLING CHANNELS
    4.
    发明申请
    TURBINE AIRFOIL COOLING SYSTEM WITH PLATFORM COOLING CHANNELS 审中-公开

    公开(公告)号：US20170081960A1

    公开(公告)日：2017-03-23

    申请号：US15303569

    申请日：2014-06-05

    Applicant: Siemens Energy, Inc.

    Inventor： Ching-Pang Lee ,  Ralph W. Matthews ,  Nan Jiang ,  Gm Salam Azad

    IPC: F01D5/18

    CPC classification number: F01D5/186 ,  F01D5/187 ,  F05D2260/202

    Abstract: A cooling system (10) positioned within a turbine airfoil (12) useable in a turbine engine and having cooling channels (16) positioned within a platform (18) of the turbine airfoil (12) with exhaust outlets (20) at the pressure and suction side edges (22, 24) to prevent hot gas ingestion under the platform (18) is disclosed. The cooling channels (16) may be formed from main channels (26) extending from cooling fluid supply channels (64) aligned with the airfoil (12) and branch channels (30) extending between the main channels (26) and the pressure or suction side edges (22, 24). The cooling system (10) reduces the cooling surface area adjacent to the airfoil fillet (32) at the intersection (34) of the platform (18) and airfoil (12) and increases cooling surface area adjacent to the pressure side and suction side mate faces (22, 24) as compared with conventional designs. Such configuration of the cooling system (10) yields a more uniform platform temperature distribution, colder and higher pressure cooling air for platform cooling and less manufacturing expense than conventional designs.

    Turbine abradable layer with asymmetric ridges or grooves
    5.
    发明授权
    Turbine abradable layer with asymmetric ridges or grooves 有权
    涡轮磨损层与不对称的脊或槽

    公开(公告)号：US09249680B2

    公开(公告)日：2016-02-02

    申请号：US14189035

    申请日：2014-02-25

    Applicant: Siemens Energy, Inc.

    Inventor： Ching-Pang Lee ,  Gm Salam Azad ,  Zhihong Gao ,  Neil Hitchman ,  Nicholas F. Martin, Jr. ,  David G. Sansom ,  Ramesh Subramanian

    IPC: F01D11/12

    CPC classification number: F01D11/122 ,  F01D11/08 ,  F05D2230/31 ,  F05D2250/28 ,  F05D2300/611

    Abstract: Turbine and compressor casing abradable component embodiments for turbine engines, with composite grooves and vertically projecting asymmetric non-parallel walls or trapezoidal cross section ridges that reduce, redirect and/or block blade tip airflow leakage downstream into the grooves rather than from turbine blade airfoil high to low pressure sides. In some embodiments at least one angularly oriented first groove formed in the ridge plateau is adapted for angular orientation upstream a turbine blade rotation direction to resist blade tip airflow leakage and the ridges are separated by second grooves that are skewed relative to the respective ridge plateaus and the substrate that are also adapted for orientation upstream the turbine blade rotation direction to resist blade tip airflow leakage.

    Abstract translation: 用于涡轮发动机的涡轮机和压缩机壳体可磨损部件实施例，其具有复合凹槽和垂直突出的非对称非平行壁或梯形横截面脊，其减小，重定向和/或阻挡下游进入凹槽的叶片尖端气流泄漏，而不是涡轮叶片翼型高 到低压侧。 在一些实施例中，形成在脊台平台中的至少一个角度定向的第一凹槽适于在涡轮机叶片旋转方向上游的角度定向以抵抗叶片尖端气流泄漏，并且脊部由相对于相应的脊部高原倾斜的第二凹槽分开， 该衬底还适于在涡轮机叶片旋转方向上游定向以抵抗叶片尖端气流泄漏。

    INTERNAL COOLING SYSTEM WITH CORRUGATED INSERT FORMING NEARWALL COOLING CHANNELS FOR AIRFOIL USABLE IN A GAS TURBINE ENGINE
    6.
    发明申请
    INTERNAL COOLING SYSTEM WITH CORRUGATED INSERT FORMING NEARWALL COOLING CHANNELS FOR AIRFOIL USABLE IN A GAS TURBINE ENGINE 审中-公开
    内部冷却系统，带有腐蚀插件，形成可用于气体涡轮发动机的空气涡轮机的近壁冷却通道

    公开(公告)号：US20150198050A1

    公开(公告)日：2015-07-16

    申请号：US14155422

    申请日：2014-01-15

    Applicant: Siemens Energy, Inc.

    Inventor： Ching-Pang Lee ,  Jae Y. Um ,  Gerald L. Hillier ,  Eric Schroeder ,  Mohamed Abdullah

    IPC: F01D5/18 ,  F01D9/02

    CPC classification number: F01D5/188 ,  F01D5/186 ,  F01D5/189 ,  F01D9/02 ,  F01D9/041 ,  F05D2240/127 ,  F05D2250/61 ,  F05D2260/201

    Abstract: An airfoil for a gas turbine engine in which the airfoil includes an internal cooling system formed from one or more midchord cooling channels with a corrugated insert positioned therein and creating nearwall leading edge, pressure side and suction side nearwall cooling channels. The corrugated insert may be formed from a wall that oscillates in a repeating pattern between peaks and valleys, such that the peaks are closer to an inner surface of the outer wall forming the generally elongated hollow airfoil The corrugated insert may work in concert with the rows of partition walls to create periodic impingement on the inner surface of the outer wall Such cooling system provides adequate cooling for use in environments in which few, if any, cooling holes are desired, such as in crude oil engine applications

    Abstract translation: 一种用于燃气涡轮发动机的翼型件，其中所述翼型件包括由一个或多个中间冷却通道形成的内部冷却系统，其中定位有波纹嵌入件，并且形成近壁前缘，压力侧和吸力侧近壁冷却通道。 波纹插入件可以由在峰和谷之间以重复图案振荡的壁形成，使得峰更接近形成大致细长的中空翼型的外壁的内表面。波纹插入件可以与行一起工作 的隔壁以在外壁的内表面上产生周期性的冲击。这种冷却系统提供足够的冷却，用于在诸如原油引擎应用中几乎没有（如果有的话）冷却孔的环境中使用

    OUTER RIM SEAL ASSEMBLY IN A TURBINE ENGINE
    7.
    发明申请
    OUTER RIM SEAL ASSEMBLY IN A TURBINE ENGINE 有权
    涡轮发动机中的外部密封组件

    公开(公告)号：US20150071763A1

    公开(公告)日：2015-03-12

    申请号：US14546309

    申请日：2014-11-18

    Applicant: Siemens Energy, Inc.

    Inventor： Ching-Pang Lee ,  Kok-Mun Tham ,  Manjit Shivanand ,  Vincent Paul Laurello ,  Gm Salam Azad ,  Nicholas F. Martin, JR.

    IPC: F01D25/12 ,  F01D5/08 ,  F01D11/12 ,  F01D11/04

    CPC classification number: F01D25/12 ,  F01D5/081 ,  F01D11/001 ,  F01D11/04 ,  F01D11/122 ,  F05D2260/202

    Abstract: A seal assembly between a hot gas path and a disc cavity in a turbine engine includes a non-rotatable vane assembly including a row of vanes and an inner shroud, a rotatable blade assembly axially adjacent to the vane assembly and including a row of blades and a turbine disc that forms a part of a turbine rotor, and an annular wing member located radially between the hot gas path and the disc cavity. The wing member extends generally axially from the blade assembly toward the vane assembly and includes a plurality of circumferentially spaced apart flow passages extending therethrough from a radially inner surface thereof to a radially outer surface thereof. The flow passages each include a portion that is curved as the passage extends radially outwardly to effect a scooping of cooling fluid from the disc cavity into the flow passages and toward the hot gas path.

    Abstract translation: 在涡轮发动机中的热气路径和盘腔之间的密封组件包括不可旋转的叶片组件，其包括一排叶片和内护罩，轴向相邻于叶片组件的可旋转刀片组件，并且包括一排叶片和 形成涡轮转子的一部分的涡轮盘，以及径向位于热气路径和盘腔之间的环形翼件。 翼构件大致轴向地从叶片组件延伸到叶片组件，并且包括从其径向内表面延伸到其径向外表面的多个周向间隔开的流动通道。 流动通道各自包括当通道径向向外延伸时弯曲的部分，以实现将冷却流体从盘腔抽入流动通道并朝向热气体通道。

    Turbine abradable layer with progressive wear zone terraced ridges
    8.
    发明授权
    Turbine abradable layer with progressive wear zone terraced ridges 有权
    涡轮磨损层与渐进磨损区梯田

    公开(公告)号：US08939707B1

    公开(公告)日：2015-01-27

    申请号：US14188992

    申请日：2014-02-25

    Applicant: Siemens Energy, Inc.

    Inventor： Ching-Pang Lee ,  Kok-Mun Tham ,  Gm Salam Azad ,  Zhihong Gao ,  Neil Hitchman ,  David G. Sansom ,  Barrry L. Allmon

    IPC: F01D11/08 ,  F01D11/12

    CPC classification number: F01D11/00 ,  F01D11/08 ,  F01D11/12 ,  F01D11/122 ,  F01D25/24 ,  F05D2220/32 ,  F05D2240/11 ,  F05D2250/181 ,  F05D2250/185 ,  F05D2300/6033

    Abstract: Turbine and compressor casing abradable component embodiments for turbine engines, with composite grooves and vertically projecting rows of stepped first ridges in planform patterns, to reduce, redirect and/or block blade tip airflow leakage downstream into the grooves rather than from turbine blade airfoil high to low pressure sides. Each stepped first ridge has a first portion proximal the substrate surface with a pair of first opposed lateral walls terminating in a plateau, and a second portion terminating in a ridge tip. These ridge or rib embodiments have first lower and second upper wear zones. The lower zone, which at and below first portion height, optimizes engine airflow characteristics, while the upper zone, between the plateau and the second portion ridge is optimized to minimize blade tip gap and wear by being more easily abradable than the lower zone.

    Abstract translation: 用于涡轮发动机的涡轮机和压缩机壳体可磨损部件实施例，其具有平面状图案中的复合凹槽和垂直突出的阶梯状第一脊的行，以减少，重定向和/或阻挡下游进入凹槽的叶片尖端气流泄漏，而不是从涡轮叶片翼型件高到 低压侧。 每个阶梯形第一脊具有靠近基板表面的第一部分，一对终止于平台的第一相对的侧壁和终止于脊尖的第二部分。 这些脊或肋实施例具有第一下和第二上磨损区。 在第一部分高度处和下方的下部区域优化发动机气流特性，同时优化平台和第二部分脊之间的上部区域，以通过比下部区域更容易磨损来最小化叶片尖端间隙和磨损。

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