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    SEAL ASSEMBLY OF GAS TURBINE ENGINE, ROTOR ASSEMBLY, BLADE FOR ROTOR ASSEMBLY AND INTER-STAGE CAVITY SEAL
    1.
    发明专利
    SEAL ASSEMBLY OF GAS TURBINE ENGINE, ROTOR ASSEMBLY, BLADE FOR ROTOR ASSEMBLY AND INTER-STAGE CAVITY SEAL 审中-公开

    公开(公告)号：JP2006342797A

    公开(公告)日：2006-12-21

    申请号：JP2006154370

    申请日：2006-06-02

    Applicant: UNITED TECHNOLOGIES CORP

    Inventor： ALVANOS IOANNIS ,  AGRAWAL RAJENDRA ,  VIRTUE JOHN P JR ,  HILL JAMES D ,  PINERO HECTOR M ,  JOHNSON ERIKA

    IPC: F01D11/02 ,  F02C7/28

    Abstract: PROBLEM TO BE SOLVED: To provide a seal assembly for restricting leaking-out of cooling air in constant. SOLUTION: An inside cavity 50 is positioned between a vane support part 52 and a rotary assembly 54 on the inside in the radial direction of an annular duct 30. The cooling air 40 is introduced into the cavity 50 by a duct 70, and cools a blade 58 by flowing in a conduit 72. The cavity 50 is separated from the duct 30 by a seal 80, and air pressure is properly maintained over the whole engine conditions. A segment-like ring 86 extends by crossing the cavity 50 from a divided rim area 110, and a runner 170 is aligned with land 84. One complete segment-like ring 86 is formed by aligning a plurality of adjacent ring segments 164 and 166 in a state of inserting the blade 58 into a disc. Leaking-out restriction is improved by extending the runner 170 outside or inside in the radial direction from the ring 86 by inclining to a shaft 11. COPYRIGHT: (C)2007,JPO&INPIT

    4.
    发明专利
    未知

    公开(公告)号：DE602005012009D1

    公开(公告)日：2009-02-12

    申请号：DE602005012009

    申请日：2005-03-30

    Applicant: UNITED TECHNOLOGIES CORP

    Inventor： ALVANOS IOANNIS ,  PINERO HECTOR M ,  AGRAWAL RAJENDRA ,  HU JIN

    IPC: F01D5/08 ,  F01D5/18 ,  F01D25/12 ,  F02C7/18

    Abstract: A gas turbine engine (20) having a turbine section cooling system is disclosed which is able to cool the turbine section (26) of the g as turbine engine (20) to an optimal operating temperature range, while at the same time not substantially degrading engine performance. The disclosure provides a number of different structures for doing so including the provision of turning foils (44), turning holes (54), and turning grooves within a secondary air flow cavity which turns or directs cooling and parasitic leakage air into the turbine section (26) in a direction in substantial alignment with a gas flow path through the turbine section (26).

    FLASH THERMOGRAPHY DOUBLE WALL THICKNESS MEASUREMENT
    6.
    发明公开
    FLASH THERMOGRAPHY DOUBLE WALL THICKNESS MEASUREMENT 有权
    BLITZTHERMOGRAFIEMESSUNG MIT DOPPELTER WANDDICKE

    公开(公告)号：EP2962061A4

    公开(公告)日：2016-11-30

    申请号：EP14757641

    申请日：2014-02-27

    Applicant: UNITED TECHNOLOGIES CORP

    Inventor： OUYANG ZHONG ,  RAULERSON DAVID A ,  SMITH KEVIN D ,  PINERO HECTOR M ,  TARASKEVICH JAIMIE ,  BOYER JESSE R

    IPC: G01B11/06 ,  G01B21/08 ,  G01J5/00 ,  G01N25/00

    CPC classification number: G01B21/085 ,  G01J5/0088 ,  G01J5/0896 ,  G01J2005/0081 ,  G01N25/00

    Abstract: A method of determining the thickness of an internal wall in a gas turbine engine component includes the steps of utilizing flash thermography to measure a complete thickness of a component between an outer wall and at least one enlarged cooling channel at a location where an outer cooling channel is positioned between the outer wall and the at least one enlarged cooling channel and where at least one member spans the cooling channel, such that the thickness is through the member which spans the outer cooling channel. An outer thickness of the component is measured from the outer wall to an outer wall of the outer cooling channel. A thickness is determined from an inner wall of the outer cooling channel to the at least one enlarged cooling channel by subtracting the measured outer thickness from the complete thickness, and also subtracting a known thickness of the outer cooling channel.

    Abstract translation: 一种确定燃气涡轮发动机部件中的内壁厚度的方法包括以下步骤：利用闪蒸热成像测量在外壁与至少一个扩大的冷却通道之间的部件的完整厚度，其中外冷却通道 定位在外壁和至少一个扩大的冷却通道之间，并且其中至少一个构件跨越冷却通道，使得厚度穿过跨过外部冷却通道的构件。 从外壁到外部冷却通道的外壁测量部件的外部厚度。 通过从完全厚度减去所测量的外部厚度，并且还减去外部冷却通道的已知厚度，从外部冷却通道的内壁到至少一个扩大的冷却通道确定厚度。

    HIGH PRESSURE TURBINE COOLANT SUPPLY SYSTEM
    7.
    发明公开
    HIGH PRESSURE TURBINE COOLANT SUPPLY SYSTEM 审中-公开
    冷却液供给系统高压涡轮

    公开(公告)号：EP2855884A4

    公开(公告)日：2016-05-11

    申请号：EP13797827

    申请日：2013-05-15

    Applicant: UNITED TECHNOLOGIES CORP

    Inventor： MOSLEY JOHN H ,  ALVANOS IOANNIS ,  STRIPINIS PHILIP S ,  FREIBERG DOUGLAS PAUL ,  PINERO HECTOR M ,  O'CONNOR JOHN J ,  PIETROBON JON

    IPC: F02C7/12 ,  F01D5/08 ,  F01D25/08 ,  F01D25/12

    CPC classification number: F01D5/082 ,  F01D25/12

    Abstract: A gas turbine engine configured to rotate in a circumferential direction about an axis extending through a center of the gas turbine engine comprises a turbine stage. The turbine stage comprises a disk, a plurality of blades and a mini-disk. The disk comprises an outer diameter edge having slots, an inner diameter bore surrounding the axis, a forward face, and an aft face. The plurality of blades is coupled to the slots. The mini-disk is coupled to the aft face of the rotor to define a cooling plenum therebetween in order to direct cooling air to the slots. In one embodiment of the invention, the cooling plenum is connected to a radially inner compressor bleed air inlet through all rotating components so that cooling air passes against the inner diameter bore.

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