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    TURBINE SECTION OF HIGH BYPASS TURBOFAN
    1.
    发明申请
    TURBINE SECTION OF HIGH BYPASS TURBOFAN 审中-公开
    高旁路涡轮机涡轮部分

    公开(公告)号：US20150377123A1

    公开(公告)日：2015-12-31

    申请号：US14793785

    申请日：2015-07-08

    Applicant: United Technologies Corporation

    Inventor： Paul R. Adams ,  Shankar S. Magge ,  Joseph Brent Staubach ,  Wesley K. Lord ,  Frederick M. Schwarz ,  Gabriel L. Suciu

    IPC: F02C3/107 ,  F02K3/02 ,  F02C9/18

    CPC classification number: F02C7/36 ,  F01D5/06 ,  F01D11/122 ,  F01D25/24 ,  F02C3/04 ,  F02C3/107 ,  F02C7/20 ,  F02C9/18 ,  F02K3/06 ,  F02K3/075 ,  F04D19/02 ,  F05B2250/283 ,  F05D2220/32 ,  F05D2220/323 ,  F05D2240/35 ,  F05D2240/60 ,  F05D2260/40311

    Abstract: A turbofan engine comprises a fan having fan blades. A compressor is in communication with the fan section. The fan is configured to communicate a portion of air into a bypass path defining a bypass area outwardly of the compressor and a portion into the compressor. A bypass ratio is defined as air communicated through the bypass path relative to air communicated to the compressor being greater than about 6.0. A combustor is in fluid communication with the compressor. A turbine is in communication with the combustor. The turbine has a first turbine section that includes two or more stages and a second turbine section that includes at least two stages. A ratio of airfoils in the first turbine section to the bypass ratio is less than about 170. The first turbine section includes a maximum gas path radius. A ratio of the maximum gas path radius to a maximum radius of the fan blades is less than about 0.50. A speed reduction mechanism is coupled to the fan and rotatable by the turbine.

    Abstract translation: 涡轮风扇发动机包括具有风扇叶片的风扇。 压缩机与风扇部分连通。 风扇被构造成将一部分空气传送到限定压缩机外部的旁路区域的旁路，以及一部分进入压缩机。 旁路比定义为相对于连通到压缩机的空气通过旁路通路的空气大于约6.0。 燃烧器与压缩机流体连通。 涡轮机与燃烧器连通。 涡轮机具有包括两级或更多级的第一涡轮部分和包括至少两级的第二涡轮部分。 第一涡轮机部分中的翼型件与旁路比率的比率小于约170.第一涡轮机部分包括最大气体路径半径。 风扇叶片的最大气路半径与最大半径的比值小于约0.50。 减速机构联接到风扇并由涡轮旋转。

    TURBINE SECTION OF HIGH BYPASS TURBOFAN
    2.
    发明申请
    TURBINE SECTION OF HIGH BYPASS TURBOFAN 审中-公开
    高旁路涡轮机涡轮部分

    公开(公告)号：US20150377122A1

    公开(公告)日：2015-12-31

    申请号：US14793770

    申请日：2015-07-08

    Applicant: United Technologies Corporation

    Inventor： Paul R. Adams ,  Shankar S. Magge ,  Joseph Brent Staubach ,  Wesley K. Lord ,  Frederick M. Schwarz ,  Gabriel L. Suciu

    IPC: F02C3/10 ,  F02K3/06 ,  F02C7/36

    CPC classification number: F02K3/06 ,  F02C3/107 ,  F02C7/36 ,  F02K3/075 ,  F05D2260/40311

    Abstract: A turbofan engine includes an engine case, a gaspath through the engine case, a fan having an array of fan blades, a compressor in fluid communication with the fan, a combustor in fluid communication with the compressor, and a turbine in fluid communication with the combustor. The turbine has a fan drive turbine section having 3 to 6 blade stages. A speed reduction mechanism couples the fan drive turbine section to the fan. A ratio of maximum gaspath radius along the low pressure turbine section to maximum radius of the fan blades is less than about 0.55. A bypass area ratio is greater than about 6.0. A ratio of a fan drive turbine section airfoil count to the bypass area ratio is less than about 170 and a second turbine section.

    Abstract translation: 涡轮风扇发动机包括发动机壳体，通过发动机壳体的气路，具有风扇叶片阵列的风扇，与风扇流体连通的压缩机，与压缩机流体连通的燃烧器，以及与压缩机流体连通的涡轮机 燃烧器。 涡轮机具有风扇驱动涡轮机部分，具有3到6个叶片级。 减速机构将风扇驱动涡轮部分连接到风扇。 沿着低压涡轮机段的最大气路径与风扇叶片的最大半径之比小于约0.55。 旁路面积比大于约6.0。 风扇驱动涡轮机翼型计数与旁路面积比的比率小于约170，而第二涡轮部分。

    INTERCOOLED COOLING AIR
    3.
    发明申请
    INTERCOOLED COOLING AIR 审中-公开

    公开(公告)号：US20190323789A1

    公开(公告)日：2019-10-24

    申请号：US16502518

    申请日：2019-07-03

    Applicant: United Technologies Corporation

    Inventor： Gabriel L. Suciu ,  Jesse M. Chandler ,  Joseph Brent Staubach ,  Brian D. Merry

    IPC: F41A9/63 ,  F02C3/04 ,  F02C7/18 ,  F02C7/32 ,  F41A9/65

    Abstract: A gas turbine engine comprises a main compressor section having a high pressure compressor with a downstream discharge, and more upstream locations. A turbine section has a high pressure turbine. A tap taps air from at least one of the more upstream locations in the compressor section, passes the tapped air through a heat exchanger and then to a cooling compressor. The cooling compressor compresses air downstream of the heat exchanger, and delivers air into the high pressure turbine. An intercooling system for a gas turbine engine is also disclosed.

    GAS TURBINE ENGINE WITH TRANSCRITICAL VAPOR CYCLE COOLING
    5.
    发明申请
    GAS TURBINE ENGINE WITH TRANSCRITICAL VAPOR CYCLE COOLING 审中-公开

    公开(公告)号：US20180080688A1

    公开(公告)日：2018-03-22

    申请号：US15269217

    申请日：2016-09-19

    Applicant: UNITED TECHNOLOGIES CORPORATION

    Inventor： Nathan Snape ,  Joseph Brent Staubach

    IPC: F25B9/00 ,  F02C3/04 ,  F02C7/18 ,  F02C7/224 ,  F02C7/22 ,  F25B9/06 ,  F25B40/02

    Abstract: A gas turbine engine has a compressor section, a combustor, and a turbine section. An associated fluid is to be cooled and an associated fluid is to be heated. A transcritical vapor cycle heats the fluid to be heated, and cools the fluid to be cooled. The transcritical vapor cycle includes a gas cooler in which the fluid to be heated is heated by a refrigerant in the transcritical vapor cycle. An evaporator heat exchanger at which the fluid to be cooled is cooled by the refrigerant in the transcritical vapor cycle. A compressor upstream of the gas cooler compresses the refrigerant to a pressure above a critical point for the refrigerant. An expansion device expands the refrigerant downstream of the gas cooler, with the evaporator heat exchanger being downstream of the expansion device, and such that the refrigerant passing through the gas cooler to heat the fluid to be heated is generally above the critical point.

    COMPLEX AIR SUPPLY SYSTEM FOR GAS TURBINE ENGINE AND ASSOCIATED AIRCRAFT
    10.
    发明申请
    COMPLEX AIR SUPPLY SYSTEM FOR GAS TURBINE ENGINE AND ASSOCIATED AIRCRAFT 审中-公开

    公开(公告)号：US20190162121A1

    公开(公告)日：2019-05-30

    申请号：US15823639

    申请日：2017-11-28

    Applicant: United Technologies Corporation

    Inventor： Joseph Brent Staubach ,  Nathan Snape ,  Frederick M. Schwarz ,  Charles E. Lents ,  Michael K. Ikeda

    IPC: F02C9/18 ,  F02C7/18 ,  B64D33/02 ,  B64D13/00

    Abstract: A lower pressure tap is connected to a first heat exchanger to be cooled by cooling air, and then to a selection valve. The selection valve selectively delivers the lower pressure tap air to a boost compressor. The lower pressure tap air downstream of the boost compressor is connected to cool the at least one turbine. The selection valve also selectively delivers a portion of the lower pressure tap air across a first cooling turbine, and to a line associated with an air delivery system for a cabin on an associated aircraft. A portion of the air downstream of the first cooling turbine is connected to a second cooling turbine, and air downstream of the second cooling turbine is connected for use in a cold loop A method of operating an air supply system is also disclosed.

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