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    GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT
    13.
    发明申请
    GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT 审中-公开
    燃气涡轮发动机压缩机装置

    公开(公告)号：US20160290241A1

    公开(公告)日：2016-10-06

    申请号：US15184253

    申请日：2016-06-16

    Applicant: United Technologies Corporation

    Inventor： Karl L. Hasel ,  Joseph B. Staubach ,  Brian D. Merry ,  Gabriel L. Suciu ,  Christopher M. Dye

    IPC: F02C7/36 ,  F02C3/04 ,  F02K3/06 ,  F01D5/12 ,  F01D25/24 ,  F01D9/02 ,  F01D25/28 ,  F02C7/28 ,  F02C7/06 ,  F02K3/02

    CPC classification number: F02C7/36 ,  F01D5/06 ,  F01D5/12 ,  F01D9/02 ,  F01D21/003 ,  F01D25/24 ,  F01D25/28 ,  F02C3/04 ,  F02C3/06 ,  F02C3/107 ,  F02C7/06 ,  F02C7/28 ,  F02C9/18 ,  F02K3/025 ,  F02K3/06 ,  F05D2220/32 ,  F05D2220/323 ,  F05D2240/12 ,  F05D2240/35 ,  F05D2240/55 ,  F05D2260/40 ,  F05D2260/4031 ,  F05D2260/40311

    Abstract: A gas turbine engine includes, among other things, a fan section, a core engine, a bypass passage, and a bypass ratio defined as the volume of air passing into the bypass passage compared to the volume of air passing into the core engine, the bypass ratio being greater than or equal to about 8 at cruise power. A gear arrangement is configured to drive the fan section. A compressor section includes both a first compressor section and a second compressor section. A turbine section is configured to drive the gear arrangement, and may have a low pressure turbine with four stages and a low pressure turbine pressure ratio greater than about 5:1, and a high pressure turbine with two stages. An overall pressure ratio is provided by the combination of a pressure ratio across the first compressor section and a pressure ratio across the second compressor section, and greater than about 40, measured at sea level and at a static, full-rated takeoff power. The pressure ratio across the second compressor section is greater than about 7.

    Abstract translation: 除了别的以外，燃气涡轮发动机还包括风扇部分，核心发动机，旁通通道和旁路比，其被定义为与进入核心发动机的空气体积相比进入旁路通道的空气量， 旁路比在巡航功率时大于或等于约8。 齿轮装置被配置成驱动风扇部分。 压缩机部分包括第一压缩机部分和第二压缩机部分。 涡轮部分构造成驱动齿轮装置，并且可以具有四级的低压涡轮机和大于约5:1的低压涡轮机压力比，以及具有两级的高压涡轮机。 整个压力比通过跨第一压缩机段的压力比和跨越第二压缩机段的压力比的组合提供，并且在海平面和静态的额定额定起飞功率下测量大于约40。 跨越第二压缩机部分的压力比大于约7。

    GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT
    14.
    发明申请
    GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT 审中-公开
    燃气涡轮发动机压缩机装置

    公开(公告)号：US20140157756A1

    公开(公告)日：2014-06-12

    申请号：US14179771

    申请日：2014-02-13

    Applicant: United Technologies Corporation

    Inventor： Karl L. Hasel ,  Joseph B. Staubach ,  Brian D. Merry ,  Gabriel L. Suciu ,  Christopher M. Dye

    IPC: F02K3/06

    CPC classification number: F02K3/06 ,  F02C3/107 ,  F05D2260/4031

    Abstract: A gas turbine engine has a fan section, including a fan, a gear arrangement configured to drive the fan section, the geared arrangement defining a gear reduction ratio greater than or equal to about 2.6. A compressor section includes both a first compressor and a second compressor. A turbine section is configured to drive the compressor section and the gear arrangement. An overall pressure ratio is provided by the combination of the first compressor and the second compressor.

    Abstract translation: 燃气涡轮发动机具有风扇部分，其包括风扇，构造成驱动风扇部分的齿轮装置，齿轮装置限定大于或等于约2.6的齿轮减速比。 压缩机部分包括第一压缩机和第二压缩机。 涡轮部分构造成驱动压缩机部分和齿轮装置。 总压力比由第一压缩机和第二压缩机的组合提供。

    GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT
    16.
    发明申请
    GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT 审中-公开
    燃气涡轮发动机压缩机装置

    公开(公告)号：US20140157753A1

    公开(公告)日：2014-06-12

    申请号：US14179864

    申请日：2014-02-13

    Applicant: United Technologies Corporation

    Inventor： Karl L. Hasel ,  Joseph B. Staubach ,  Brian D. Merry ,  Gabriel L. Suciu ,  Christopher M. Dye

    IPC: F02K3/06

    CPC classification number: F02K3/06 ,  F02C3/107 ,  F05D2260/4031

    Abstract: A method of designing a gas turbine engine includes providing a fan section including a fan; driving the fan section via a gear arrangement; providing a compressor section, including both a first compressor and a second compressor; and driving the compressor section and the gear arrangement via a turbine section. An overall pressure ratio, being provided by the combination of a pressure ratio across the first compressor and a pressure ratio across the second compressor, is greater than or equal to about 35. The pressure ratio across the first compressor is greater than or equal to about 7.

    Abstract translation: 一种设计燃气涡轮发动机的方法包括：提供包括风扇的风扇部分; 通过齿轮装置驱动风扇部分; 提供包括第一压缩机和第二压缩机的压缩机部分; 并且经由涡轮机部分驱动压缩机部分和齿轮装置。 通过跨第一压缩机的压力比和跨第二压缩机的压力比的组合提供的总压力比大于或等于约35.跨第一压缩机的压力比大于或等于约 7。

    Twin Tip Turbine Propulsors Powered by a Single Gas Turbine Generator
    17.
    发明申请
    Twin Tip Turbine Propulsors Powered by a Single Gas Turbine Generator 审中-公开
    由单个燃气轮机发电机提供的双尖头涡轮推进器

    公开(公告)号：US20140117152A1

    公开(公告)日：2014-05-01

    申请号：US13662879

    申请日：2012-10-29

    Applicant: UNITED TECHNOLOGIES CORPORATION

    Inventor： Gabriel L. Suciu ,  Jesse M. Chandler ,  Joseph B. Staubach ,  Adam Joseph Suydam

    IPC: B64D27/14 ,  F02K3/02

    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机的核心发动机和位于涡轮机下游的歧管。 歧管将涡轮下游的气体输送到至少两个机舱中，其中每个机舱都接收风扇转子。 风扇转子被固定以与安装在风扇转子的径向外部位置处的尖端涡轮一起旋转，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。

    Supercritical CO2 cycle for gas turbine engines having supplemental cooling
    18.
    发明授权
    Supercritical CO2 cycle for gas turbine engines having supplemental cooling 有权

    公开(公告)号：US11506124B2

    公开(公告)日：2022-11-22

    申请号：US16832535

    申请日：2020-03-27

    Applicant: United Technologies Corporation

    Inventor： Joseph B. Staubach ,  Amanda J. L. Boucher

    IPC: F02C7/141

    Abstract: Gas turbine engines are described. The gas turbine engines include a compressor section, a combustor section, a turbine section, and a nozzle, wherein the compressor section, the combustor section, the turbine section, and the nozzle define a core flow path that expels through the nozzle. A waste heat recovery system is operably connected to the gas turbine engine, the waste heat recovery system having a working fluid. An auxiliary cooling system is configured to provide cooling to a working fluid of the waste heat recovery system.

    Descent operation for an aircraft parallel hybrid gas turbine electric propulsion system
    19.
    发明授权
    Descent operation for an aircraft parallel hybrid gas turbine electric propulsion system 有权

    公开(公告)号：US11428170B2

    公开(公告)日：2022-08-30

    申请号：US15200149

    申请日：2016-07-01

    Applicant: United Technologies Corporation

    Inventor： Charles E. Lents ,  Larry W. Hardin ,  Jonathan Rheaume ,  Joseph B. Staubach

    IPC: F02C7/36 ,  F01D15/10 ,  B64D31/06 ,  F02C6/14 ,  F02C9/50 ,  B64D27/24 ,  B64D27/10 ,  B64D27/02

    Abstract: A gas turbine engine includes a core having a compressor section with a first compressor and a second compressor, a turbine section with a first turbine and a second turbine, and a primary flowpath fluidly connecting the compressor section and the turbine section. The first compressor is connected to the first turbine via a first shaft, the second compressor is connected to the second turbine via a second shaft, and a motor is connected to the first shaft such that rotational energy generated by the motor is translated to the first shaft. The gas turbine engine includes a takeoff mode of operation, a top of climb mode of operation, and at least one additional mode of operation. The gas turbine engine is undersized relative to a thrust requirement in at least one of the takeoff mode of operation and the top of climb mode of operation, and a controller is configured to control the mode of operation of the gas turbine engine.

    Gas Turbine Engine Heat Exchanger for Annular Flowpaths
    20.
    发明申请
    Gas Turbine Engine Heat Exchanger for Annular Flowpaths 有权

    公开(公告)号：US20210207534A1

    公开(公告)日：2021-07-08

    申请号：US16733646

    申请日：2020-01-03

    Applicant: United Technologies Corporation

    Inventor： Joseph B. Staubach ,  Amanda Jean Boucher ,  Jesse M. Chandler

    IPC: F02C7/143 ,  F02C7/16 ,  F28D7/02

    Abstract: A heat exchanger has arcuate inlet and outlet manifolds and a plurality of tube banks, each tube bank coupling one of the inlet manifold outlets to an associated one of the outlet manifold inlets. Each tube bank partially nests with one or more others of the tube banks and has: a first header coupled to the associated inlet manifold outlet and the associated the outlet manifold inlet; a second header; and a plurality of tube bundles each having a first end coupled to the associated first header and a second end coupled to the associated second header. A flowpath from the each inlet manifold outlet passes sequentially through flowpath legs formed by each of the tube bundles in the associated tube bank to exit the tube bank to the associated outlet manifold inlet.

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