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    TWIN TIP TURBINE PROPULSORS POWERED BY A SINGLE GAS TURBINE GENERATOR
    1.
    发明申请
    TWIN TIP TURBINE PROPULSORS POWERED BY A SINGLE GAS TURBINE GENERATOR 审中-公开
    单燃气轮机发电机供电的双头涡轮推进器

    公开(公告)号：WO2014109811A2

    公开(公告)日：2014-07-17

    申请号：PCT/US2013066531

    申请日：2013-10-24

    Applicant: UNITED TECHNOLOGIES CORP

    Inventor： SUCIU GABRIEL L ,  CHANDLER JESSE M ,  STAUBACH JOSEPH B ,  SUYDAM ADAM JOSEPH

    CPC classification number: B64D27/14 ,  F02C7/36 ,  F02K3/06

    Abstract: A gas turbine engine has a core engine incorporating a turbine, and a manifold positioned downstream of the turbine. The manifold delivers gas downstream of the turbine into at least two nacelles, with each of the nacelles receiving a fan rotor. The fan rotor is fixed to rotate with a tip turbine mounted at a radially outer location of the fan rotor, with the tip turbine being in the path of gases from the manifold. An aircraft is also disclosed.

    Abstract translation: 燃气涡轮发动机具有并入涡轮的核心发动机和位于涡轮下游的歧管。 歧管将涡轮下游的气体输送到至少两个机舱中，每个机舱接收风扇转子。 风扇转子被固定以随着安装在风扇转子的径向外部位置处的尖端涡轮机旋转，其中尖端涡轮机处于来自歧管的气体的路径中。 还披露了一架飞机。

    GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT
    4.
    发明专利
    GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT 未知

    公开(公告)号：CA2850225A1

    公开(公告)日：2013-07-04

    申请号：CA2850225

    申请日：2012-12-26

    Applicant: UNITED TECHNOLOGIES CORP

    Inventor： HASEL KARL L ,  STAUBACH JOSEPH B ,  MERRY BRIAN D ,  SUCIU GABRIEL L ,  DYE CHRISTOPHER M

    IPC: F01D1/04

    Abstract: A gas turbine engine includes a fan section, a gear arrangement configured to drive the fan section, a compressor section and a turbine section. The compressor section includes a low pressure compressor section and a high pressure compressor section. The turbine section is configured to drive compressor section and the gear arrangement. An overall pressure ratio, which is provided by a combination of a pressure ratio across said low pressure compressor section and a pressure ratio across said high pressure compressor section, is greater than about 35. The pressure ratio across the low pressure compressor section is between about 3 and about 8 whereas the pressure ratio across the high pressure compressor section is between about 7 and about 15.

    HIGH COVERAGE COOLING HOLE SHAPE
    5.
    发明专利
    HIGH COVERAGE COOLING HOLE SHAPE 未知

    公开(公告)号：SG138539A1

    公开(公告)日：2008-01-28

    申请号：SG2007040470

    申请日：2007-06-06

    Applicant: UNITED TECHNOLOGIES CORP

    Inventor： KOHLI ATUL ,  STAUBACH JOSEPH B ,  DOLL STEPHEN D

    Abstract: A cooling hole for a gas turbine engine component has a bi-lobed shape to improve cooling effectiveness. The gas turbine engine component has a first outer surface and a second outer surface separated from each other by a thickness. The cooling hole extends through the thickness from a first opening at the first outer surface to a second opening at the second outer surface. The first and second openings are defined by shapes that are different from each other. One of the first and second openings has the bi- lobed shape.

    GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT
    6.
    发明专利
    GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT 未知

    公开(公告)号：CA2850225C

    公开(公告)日：2015-10-13

    申请号：CA2850225

    申请日：2012-12-26

    Applicant: UNITED TECHNOLOGIES CORP

    Inventor： HASEL KARL L ,  STAUBACH JOSEPH B ,  MERRY BRIAN D ,  SUCIU GABRIEL L ,  DYE CHRISTOPHER M

    IPC: F02K3/06 ,  F02C3/04

    Abstract: A gas turbine engine includes a fan section, a gear arrangement configured to drive the fan section, a compressor section and a turbine section. The compressor section includes a low pressure compressor section and a high pressure compressor section. The turbine section is configured to drive compressor section and the gear arrangement. An overall pressure ratio, which is provided by a combination of a pressure ratio across said low pressure compressor section and a pressure ratio across said high pressure compressor section, is greater than about 35. The pressure ratio across the low pressure compressor section is between about 3 and about 8 whereas the pressure ratio across the high pressure compressor section is between about 7 and about 15.

    AEROSPACE ENGINE WITH AUGMENTING TURBOJET
    9.
    发明公开
    AEROSPACE ENGINE WITH AUGMENTING TURBOJET 有权
    LUFT- UND RAUMFAHRTMOTOR MITVERSTÄRKTEMSTRAHLTRIEBWERK

    公开(公告)号：EP2825760A4

    公开(公告)日：2015-10-21

    申请号：EP13760947

    申请日：2013-01-24

    Applicant: UNITED TECHNOLOGIES CORP

    Inventor： KUPRATIS DANIEL BERNARD ,  STAUBACH JOSEPH B

    IPC: F02K3/12 ,  F02C3/04 ,  F02C3/13 ,  F02C6/02 ,  F02C9/18

    CPC classification number: F02C3/04 ,  F02C3/13 ,  F02C6/00 ,  F02C6/02 ,  F02C9/18 ,  F02K3/12 ,  F05D2250/30 ,  F05D2250/311 ,  F05D2270/07

    Abstract: A gas turbine engine system includes a fan section, a low pressure compressor section downstream of the fan section, a first engine core downstream from the low pressure compressor section, a second engine core downstream from the low pressure compressor section, and a flowpath control mechanism configured to selectively restrict fluid flow through the second engine core. The first engine core includes a first engine core compressor section, a first engine core combustor downstream of the first engine core compressor section, and a first engine core turbine section downstream of the first engine core combustor. The second engine core includes a second engine core compressor section, a second engine core combustor downstream of the second engine core compressor section, and a second engine core turbine section downstream of the second engine core combustor.

    Abstract translation: 燃气涡轮发动机系统包括风扇部分，风扇部分下游的低压压缩机部分，低压压缩机部分下游的第一发动机芯，低压压缩机部分下游的第二发动机芯，以及流路控制机构 被配置为选择性地限制流过第二发动机芯的流体。 第一发动机机芯包括第一发动机核心压缩机部分，第一发动机核心压缩机部分下游的第一发动机核心燃烧器和第一发动机核心燃烧器下游的第一发动机芯涡轮机部分。 第二发动机机芯包括第二发动机芯压缩机部分，第二发动机芯压缩机部分下游的第二发动机核心燃烧器和第二发动机核心燃烧器下游的第二发动机芯涡轮机部分。

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