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    Turbomachine thrust reverser
    23.
    发明授权
    Turbomachine thrust reverser 有权

    公开(公告)号：US10145335B2

    公开(公告)日：2018-12-04

    申请号：US13712240

    申请日：2012-12-12

    Applicant: United Technologies Corporation

    Inventor： Robert L. Gukeisen ,  Claude Mercier ,  Robert E. Malecki

    IPC: F02K1/72 ,  F02K1/58 ,  F02K1/64 ,  F02C7/20 ,  B64D29/06

    Abstract: An example thrust reverser of a gas turbine engine is configured to connect to an aircraft wing via a pylon via one or more thrust reverser mounts located adjacent to a top circumferential apex of the engine according to an exemplary aspect of the present disclosure includes, among other things, a first cowl moveable between a stowed position and a deployed position relative to a second cowl. The first cowl in the deployed position configured to permit thrust to be redirected from an engine to slow the engine. The first cowl forming a portion of a substantially annular encasement of the engine. The first cowl directly interfaces with second cowl of the encasement at a cowl interface position that is more than 18 degrees circumferentially offset from the top circumferential apex when the first cowl is in the stowed position.

    BYPASS DUCT HEAT EXCHANGER PLACEMENT
    25.
    发明申请
    BYPASS DUCT HEAT EXCHANGER PLACEMENT 审中-公开
    旁路导管热交换器放置

    公开(公告)号：US20160215732A1

    公开(公告)日：2016-07-28

    申请号：US15024254

    申请日：2014-07-30

    Applicant: UNITED TECHNOLOGIES CORPORATION

    Inventor： Robert E. Malecki

    IPC: F02K3/115 ,  F02C7/14

    CPC classification number: F02K3/115 ,  F02C7/14 ,  F02C7/185 ,  F05D2250/52 ,  F05D2260/213

    Abstract: A bypass duct for a gas turbine engine includes as an inner surface, an intermediate case, an inner fixed structure (IFS), and a heat exchanger outlet mounted to the outer surface of the intermediate case or the forward portion of the IFS. The heat exchanger outlet is oriented to bathe the surface of the bypass duct case downstream of the heat exchanger outlet with low momentum heat exchanger exhaust to reduce skin friction losses for the bypass duct.

    Abstract translation: 用于燃气涡轮发动机的旁通管包括作为内表面的中间壳体，内部固定结构（IFS）和安装到中间壳体的外表面或IFS的前部的热交换器出口。 热交换器出口定向成以低动量热交换器排气将热交换器出口下游的旁通管道壳体的表面洗涤，以减少旁路管道的表面摩擦损失。

    Asymmetric Fan Nozzle in High-BPR Separate-Flow Nacelle
    27.
    发明申请
    Asymmetric Fan Nozzle in High-BPR Separate-Flow Nacelle 审中-公开
    高BPR分流机舱不对称风机喷嘴

    公开(公告)号：US20160003194A1

    公开(公告)日：2016-01-07

    申请号：US14768830

    申请日：2013-12-18

    Applicant: UNITED TECHNOLOGIES CORPORATION

    Inventor： Wesley K. Lord ,  Robert E. Malecki

    IPC: F02K3/06 ,  F01D25/24 ,  F02C3/04

    CPC classification number: F02K3/06 ,  B64D33/04 ,  B64D2033/0286 ,  F01D25/24 ,  F02C3/04 ,  F02K1/78 ,  F05D2220/32 ,  F05D2240/24 ,  F05D2240/35 ,  F05D2250/73

    Abstract: A fan nozzle for an aircraft gas turbine engine is comprised of a core engine cowl that is disposed within a fan cowl so that an air flow area is defined therebetween. The core engine cowl and fan cowl are disposed around a horizontal central plane. The fan cowl has a substantially circular shape and is formed of an upper substantially semi-circular portion having a first radius and a lower substantially semi-circular portion having a second radius. The core engine cowl has a substantially circular shape and is formed of an upper substantially semi-circular portion having a third radius and a lower substantially semi-circular portion having a third radius. The upper substantially semi-circular portion of the core engine cowl includes a left arcuate member and a right arcuate member. The second radius is less than the first radius and the third radius is less than the fourth radius.

    Abstract translation: 用于飞机燃气涡轮发动机的风扇喷嘴由设置在风扇罩内的核心发动机罩组成，从而在其间限定空气流动区域。 核心发动机罩和风扇罩布置在水平中心平面周围。 风扇罩具有大致圆形形状，并且由具有第一半径的上部基本上半圆形部分和具有第二半径的下部基本上半圆形部分形成。 芯发动机整流罩具有大致圆形形状，并且由具有第三半径的上部大致半圆形部分和具有第三半径的下部基本上半圆形部分形成。 核心发动机罩的上部基本上半圆形的部分包括左弓形件和右弓形件。 第二半径小于第一半径，第三半径小于第四半径。

    TURBOMACHINE THRUST REVERSER
    28.
    发明申请
    TURBOMACHINE THRUST REVERSER 审中-公开

    公开(公告)号：US20140246514A1

    公开(公告)日：2014-09-04

    申请号：US13712240

    申请日：2012-12-12

    Applicant: United Technologies Corporation

    Inventor： Robert L. Gukeisen ,  Claude Mercier ,  Robert E. Malecki

    IPC: F02K1/58

    CPC classification number: F02K1/58 ,  B64D29/06 ,  F02C7/20 ,  F02K1/64 ,  F02K1/72 ,  F05D2250/34

    Abstract: An example thrust reverser of a gas turbine engine is configured to connect to an aircraft wing via a pylon via one or more thrust reverser mounts located adjacent to a top circumferential apex of the engine according to an exemplary aspect of the present disclosure includes, among other things, a first cowl moveable between a stowed position and a deployed position relative to a second cowl. The first cowl in the deployed position configured to permit thrust to be redirected from an engine to slow the engine. The first cowl forming a portion of a substantially annular encasement of the engine. The first cowl directly interfaces with second cowl of the encasement at a cowl interface position that is more than 18 degrees circumferentially offset from the top circumferential apex when the first cowl is in the stowed position.

    Abstract translation: 燃气涡轮发动机的示例推力反向器被构造成根据本公开的示例性方面经由经由一个或多个相对于发动机的顶部圆周顶点定位的一个或多个推力反向器安装座经由塔架连接到飞机机翼， 事物，第一整流罩可相对于第二整流罩在收起位置和展开位置之间移动。 在展开位置的第一个整流罩配置成允许从发动机重新引导推力以减慢发动机。 第一整流罩形成发动机的大致环形的壳体的一部分。 当第一整流罩处于收起位置时，第一整流罩直接与外壳的第二整流罩接合在整流罩接口位置，该顶盖接口位置与顶部圆周顶点周向偏移大于18度。

    Low pressure ratio fan engine having a dimensional relationship between inlet and fan size
    29.
    发明授权
    Low pressure ratio fan engine having a dimensional relationship between inlet and fan size 有权

    公开(公告)号：US11286811B2

    公开(公告)日：2022-03-29

    申请号：US16391476

    申请日：2019-04-23

    Applicant: United Technologies Corporation

    Inventor： Wesley K. Lord ,  Robert E. Malecki ,  Yuan J. Qiu ,  Becky E. Rose ,  Jonathan Gilson

    IPC: F01D25/24 ,  F02K3/06 ,  F01D5/14 ,  F02C7/04 ,  F01D5/02 ,  F01D15/12 ,  F01D17/10

    Abstract: According to an example embodiment, a gas turbine engine assembly includes, among other things, a fan section including a fan, the fan including a plurality of fan blades, a diameter of the fan having a dimension D that is based on a dimension of the fan blades, each fan blade having a leading edge, and a forward most portion on the leading edges of the fan blades in a first reference plane, a geared architecture, a turbine section including a high pressure turbine and a low pressure turbine, the low pressure turbine driving the fan through the geared architecture, a nacelle surrounding the fan, the nacelle including an inlet portion forward of the fan, a forward edge on the inlet portion in a second reference plane, and a length of the inlet portion having a dimension L measured along an engine axis between the first reference plane and the second reference plane. A dimensional relationship of L/D is between 0.20 and 0.40.

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