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    GEARED TURBOFAN ENGINE WITH HIGH COMPRESSOR EXIT TEMPERATURE
    22.
    发明申请
    GEARED TURBOFAN ENGINE WITH HIGH COMPRESSOR EXIT TEMPERATURE 审中-公开
    具有高压缩机出口温度的齿轮涡轮发动机

    公开(公告)号：US20140260326A1

    公开(公告)日：2014-09-18

    申请号：US14354175

    申请日：2013-03-12

    Applicant: UNITED TECHNOLOGIES CORPORATION

    Inventor： Frederick M. Schwarz ,  Karl L. Hasel

    IPC: F02C3/13

    CPC classification number: F02C3/13 ,  F01D5/34 ,  F02K3/06 ,  F05D2220/327 ,  F05D2220/36 ,  F05D2260/40311 ,  Y02T50/671

    Abstract: A gas turbine engine includes a fan with a plurality of fan blades rotatable about an axis, and a compressor section that includes at least first and second compressor sections. An average exit temperature of the compressor section is between about 1000° F. and about 1500° F. The engine also includes a combustor that is in fluid communication with the compressor section, and a turbine section that is in fluid communication with the combustor. A geared architecture is driven by the turbine section for rotating the fan about the axis.

    Abstract translation: 燃气涡轮发动机包括具有可围绕轴线旋转的多个风扇叶片的风扇和至少包括第一和第二压缩机部分的压缩机部分。 压缩机部分的平均出口温度在约1000°F和约1500°F之间。发动机还包括与压缩机部分流体连通的燃烧器和与燃烧器流体连通的涡轮机部分。 齿轮架构由涡轮部分驱动，用于围绕轴线旋转风扇。

    GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT
    24.
    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：US20140157752A1

    公开(公告)日：2014-06-12

    申请号：US14179827

    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. 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。 跨越第一压缩机的压力比大于或等于约7。

    Geared Turbofan Engine With Increased Bypass Ratio and Compressor Ratio ...
    25.
    发明申请
    Geared Turbofan Engine With Increased Bypass Ratio and Compressor Ratio ... 审中-公开
    具有增加旁路比和压缩比的齿轮涡轮风扇发动机

    公开(公告)号：US20140096509A1

    公开(公告)日：2014-04-10

    申请号：US13716253

    申请日：2012-12-17

    Applicant: UNITED TECHNOLOGIES CORPORATION

    Inventor： Karl L. Hasel

    IPC: F02K3/075

    CPC classification number: F02K3/075 ,  F02K3/06

    Abstract: A gas turbine engine is typically comprised of a fan stage, multiple compressor stages, and multiple turbine stages. These stages are made up of alternating rotating blade rows and static vane rows. The total number of blades and vanes is the airfoil count. An overall pressure ratio is greater than 30. A bypass ratio is greater than 8. A stage ratio is the product of the bypass ratio and the overall pressure ratio divided by the number of stages. An airfoil ratio is that product divided by the airfoil count. The stage ratio is greater than or equal to 22 and/or the airfoil ratio is greater than or equal to 0.12.

    Abstract translation: 燃气涡轮发动机通常包括风扇台，多个压缩机级和多个涡轮级。 这些阶段由交替的旋转刀片行和静态叶片行组成。 叶片和叶片的总数是机翼数量。 总压力比大于30.旁路比大于8.级比是旁路比和总压力比除以级数的乘积。 翼型比是产品除以翼型数。 台阶比大于或等于22，和/或翼型比大于或等于0.12。

    Gas turbine engine compressor arrangement
    27.
    发明授权
    Gas turbine engine compressor arrangement 有权

    公开(公告)号：US10830152B2

    公开(公告)日：2020-11-10

    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 ,  F02K3/06 ,  F02C3/107 ,  F02C3/04 ,  F02K3/02 ,  F02C3/06

    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.

    GAS TURBINE ENGINE SHAFT BEARING CONFIGURATION
    28.
    发明申请
    GAS TURBINE ENGINE SHAFT BEARING CONFIGURATION 审中-公开

    公开(公告)号：US20190107051A1

    公开(公告)日：2019-04-11

    申请号：US16152502

    申请日：2018-10-05

    Applicant: United Technologies Corporation

    Inventor： Brian D. Merry ,  Gabriel L. Suciu ,  Karl L. Hasel

    IPC: F02C7/06 ,  F02C7/36 ,  F02C3/04

    Abstract: A gas turbine engine includes a housing includes an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. A rotor is connected to the hub and supports a compressor section. The geared architecture includes an epicyclic gear train. A fan is rotationally driven by the geared architecture. First and second bearings support the shaft relative to the intermediate case and the inlet case, respectively. The radially inner boundary of the core inlet is at a location of a core inlet stator and the radially inner boundary of the compressor section inlet is at a location of the first stage low-pressure compressor rotor.

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