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    TURBINE SECTION OF HIGH BYPASS TURBOFAN
    15.
    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：US20150377124A1

    公开(公告)日：2015-12-31

    申请号：US14793795

    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 ,  F01D5/06 ,  F01D9/04 ,  F01D25/24

    CPC classification number: F02C3/107 ,  F01D9/041 ,  F01D25/24 ,  F02C7/36 ,  F02K3/06 ,  F05D2260/40311

    Abstract: A turbofan engine includes an engine case, a gaspath through the engine case, a fan having a circumferential 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 bypass area ratio is between about 8.0 and about 20.0. A ratio of maximum gaspath radius along the fan drive turbine section to maximum radius of the fan is less than about 0.50. A ratio of a turbine section airfoil count to the bypass area ratio is between about 10 and about 170. The fan drive turbine section airfoil count being the total number of blade airfoils and vane airfoils of the fan drive 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，具有圆周阵列的风扇叶片的风扇，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。 旁路面积比在约8.0和约20.0之间。 风扇驱动涡轮机部分的最大气路径半径与风扇的最大半径的比值小于约0.50。 涡轮机翼型计数与旁路面积比的比率在约10和约170之间。风扇驱动涡轮机翼型计数是风扇驱动涡轮机部分的叶片翼型和叶片翼型的总数。

    Gas turbine engine mid turbine frame with flow turning features
    17.
    发明授权
    Gas turbine engine mid turbine frame with flow turning features 有权
    燃气涡轮发动机中涡轮机架具有流动转向特征

    公开(公告)号：US08561414B1

    公开(公告)日：2013-10-22

    申请号：US13778202

    申请日：2013-02-27

    Applicant: United Technologies Corporation

    Inventor： Thomas J. Praisner ,  Shankar S. Magge ,  Matthew B. Estes

    IPC: F02D9/02

    CPC classification number: F01D9/041 ,  F01D5/141 ,  F01D9/04 ,  F01D9/065 ,  F01D25/28 ,  F02K3/06 ,  F05D2240/12 ,  Y02T50/673

    Abstract: A gas turbine engine includes first and second stages having a rotational axis. A circumferential array of airfoils is arranged axially between the first stage and the second stage. At least one of the airfoils have a curvature provided equidistantly between pressure and suction sides. The airfoils extend from a leading edge to a trailing edge at a midspan plane along the airfoil. An angle is defined between first and second lines respectively tangent to the intersection of the midspan plane and the curvature at airfoil leading and trailing edges. The angle is equal to or greater than about 10°.

    Abstract translation: 燃气涡轮发动机包括具有旋转轴线的第一和第二级。 在第一阶段和第二阶段之间轴向布置有圆周阵列的翼型件。 至少一个翼型件具有在压力侧和吸力侧之间等距离设置的曲率。 翼型件从翼缘沿着翼型在中跨平面处从前缘延伸到后缘。 第一和第二线之间的角度分别与翼型前缘和后缘的中跨平面和曲率的交点相切。 该角度等于或大于约10°。

    TURBINE SECTION OF HIGH BYPASS TURBOFAN
    19.
    发明申请
    TURBINE SECTION OF HIGH BYPASS TURBOFAN 审中-公开

    公开(公告)号：US20190017446A1

    公开(公告)日：2019-01-17

    申请号：US16025038

    申请日：2018-07-02

    Applicant: United Technologies Corporation

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

    IPC: F02C7/36 ,  F02K3/06 ,  F02C9/18 ,  F01D11/12 ,  F01D5/06 ,  F01D25/24 ,  F02C3/04 ,  F02C7/20 ,  F04D19/02 ,  F02K3/075 ,  F02C3/107

    Abstract: A turbofan engine according to an example of the present disclosure includes, among other things, a fan including a circumferential array of fan blades, a compressor in fluid communication with the fan, the compressor including a first compressor section and a second compressor, the second compressor section including a second compressor section inlet with a second compressor section inlet annulus area, a fan duct including a fan duct annulus area outboard of the second compressor section inlet, a shaft assembly having a first portion and a second portion, a turbine in fluid communication with the combustor, the turbine having a first turbine section coupled to the first portion of the shaft assembly to drive the first compressor section, and a second turbine section coupled to the second portion of the shaft assembly to drive the fan, an epicyclic transmission coupled to the fan and rotatable by the second turbine section through the second portion of the shaft assembly to allow the second turbine to turn faster than the fan, wherein the second turbine section includes a maximum gas path radius and the fan blades include a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the fan blades is less than 0.6.

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