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    SYSTEM AND METHOD FOR CONTROLLING AN ENVIRONMENTAL CONDITION OF AN ENGINE ELECTRONIC COMPONENT
    11.
    发明申请
    SYSTEM AND METHOD FOR CONTROLLING AN ENVIRONMENTAL CONDITION OF AN ENGINE ELECTRONIC COMPONENT 审中-公开
    用于控制发动机电子部件的环境条件的系统和方法

    公开(公告)号：US20160177781A1

    公开(公告)日：2016-06-23

    申请号：US14974808

    申请日：2015-12-18

    Applicant: United Technologies Corporation

    Inventor： Gabriel L. Suciu ,  Kurt J. Sobanski ,  Wesley K. Lord ,  William K. Ackermann

    IPC: F01D25/10 ,  F02C3/04

    CPC classification number: F01D25/10 ,  F01D25/12 ,  F02C3/04 ,  F02C7/18 ,  F02C7/32 ,  F02K3/115 ,  F05D2260/202 ,  Y02T50/671 ,  Y02T50/675

    Abstract: The present disclosure relates generally to a gas turbine engine that includes a fan configured to generate a fanstream and a fanstream duct configured to receive the fanstream flowing therethrough. An engine electronic component is positioned in flow communication with the fanstream. A heating element is positioned in the fanstream upstream from the engine electronic component and is operative to heat at least a portion of the fanstream in flow communication with the engine electronic component. The position of the engine electronic component passively thermally conditions the engine electronic component and the heating element actively thermally conditions the engine electronic component.

    Abstract translation: 本公开一般涉及一种燃气涡轮发动机，其包括被配置为产生风扇流的风扇和被配置为接收流过其中的风扇流的风扇流管。 发动机电子部件被定位成与风扇流动地流动连通。 加热元件位于发动机电子部件上游的风扇流中，并且可操作地加热与发动机电子部件流动连通的风扇流的至少一部分。 发动机电子部件的位置被动地热调节发动机电子部件和加热元件，使发动机电子部件发热。

    Asymmetric Fan Nozzle in High-BPR Separate-Flow Nacelle
    14.
    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。

    GAS TURBINE ENGINE WITH EMBEDDED DISTRIBUTED FANS
    15.
    发明申请
    GAS TURBINE ENGINE WITH EMBEDDED DISTRIBUTED FANS 审中-公开
    具有嵌入式分布式风机的气体涡轮发动机

    公开(公告)号：US20150361879A1

    公开(公告)日：2015-12-17

    申请号：US14600304

    申请日：2015-01-20

    Applicant: United Technologies Corporation

    Inventor： Gabriel L. Suciu ,  Alan H. Epstein ,  Wesley K. Lord ,  Jesse M. Chandler ,  Stephen G. Pixton

    IPC: F02C3/10 ,  B64D33/02 ,  B64D33/04 ,  F02K1/54 ,  B64D27/12

    CPC classification number: F02C3/10 ,  B64D27/12 ,  B64D35/04 ,  F02C6/206 ,  F02C7/36 ,  F02K1/54

    Abstract: An aircraft body comprises a support structure. A gas turbine engine comprises a gas generator having at least one compressor rotor, at least one gas generator turbine rotor, and a combustion section. A fan drive turbine is positioned downstream of at least one gas generator turbine rotor, and is configured to drive a shaft. The shaft engages gears to drive a plurality of fan rotors. The gas turbine engine is embedded into the support structure such that there is an inlet leading through the support structure to the fan rotors to deliver air to the fan rotors. A gas turbine engine is also disclosed.

    Abstract translation: 飞机主体包括支撑结构。 燃气涡轮发动机包括具有至少一个压缩机转子，至少一个气体发生器涡轮机转子和燃烧部分的气体发生器。 风扇驱动涡轮机位于至少一个气体发生器涡轮转子的下游，并且构造成驱动轴。 轴接合齿轮以驱动多个风扇转子。 燃气涡轮发动机被嵌入到支撑结构中，使得有一个入口通过支撑结构通向风扇转子，以将空气输送到风扇转子。 还公开了一种燃气涡轮发动机。

    Gas Turbine Engine with Low Fan Noise
    16.
    发明申请
    Gas Turbine Engine with Low Fan Noise 审中-公开
    具有低风扇噪音的燃气轮机发动机

    公开(公告)号：US20150361824A1

    公开(公告)日：2015-12-17

    申请号：US14761689

    申请日：2014-03-11

    Applicant: UNITED TECHNOLOGIES CORPORATION

    Inventor： Bruce L. Morin ,  Wesley K. Lord

    IPC: F01D25/04 ,  F01D17/14 ,  F01D9/02 ,  F01D5/02 ,  F01D5/14

    Abstract: In accordance with one aspect of the disclosure, a gas turbine engine, method of using and designing such is disclosed. The gas turbine engine may comprise a fan including a plurality of blades, and a variable area fan nozzle. The fan may be configured to have a design point fan tip leading edge relative flow angle βADP, and may be further configured to have an off-design point fan tip leading edge relative flow angle β at an off-design fan operating point. The variable area fan nozzle may be configured to manipulate the amount of air flowing through the fan so that the absolute value of a difference between the design point fan tip leading edge relative flow angle βADP and the off-design point fan tip leading edge relative flow angle β is in a specified range.

    Abstract translation: 根据本公开的一个方面，公开了一种燃气涡轮发动机，其使用和设计方法。 燃气涡轮发动机可以包括包括多个叶片的风扇和可变区域风扇喷嘴。 风扇可以被配置成具有设计点风扇顶端前缘相对流动角度ADP，并且还可以被配置为具有偏离设计点的风扇尖端前缘相对流动角度＆bgr; 在设计外的风扇工作点。 可变区域风扇喷嘴可以被配置为操纵流过风扇的空气量，使得设计点风扇叶片前缘相对流动角度与ADP和偏离设计点之间的差异的绝对值 相对流角 在指定范围内。

    TURBINE SECTION OF HIGH BYPASS TURBOFAN
    17.
    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：US20150345404A1

    公开(公告)日：2015-12-03

    申请号：US14692090

    申请日：2015-04-21

    Applicant: United Technologies Corporation

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

    IPC: F02C9/18 ,  F02C7/36 ,  F01D25/24 ,  F01D1/02 ,  F01D5/02 ,  F01D9/04 ,  F02C3/04 ,  F02C7/20

    Abstract: A turbofan engine has an engine case and a gaspath through the engine case. A fan has a circumferential array of fan blades. The engine further has a compressor, a combustor, a gas generating turbine, and a low pressure turbine section. A speed reduction mechanism couples the low pressure turbine section to the fan. A bypass area ratio is greater than about 6.0. The low pressure turbine section airfoil count to bypass area ratio is below about 170.

    Abstract translation: 涡轮风扇发动机具有发动机壳体和通过发动机壳体的气路。 风扇具有圆周阵列的风扇叶片。 发动机还具有压缩机，燃烧器，气体发生涡轮机和低压涡轮部分。 减速机构将低压涡轮部分连接到风扇。 旁路面积比大于约6.0。 低压涡轮机翼型计数到旁路面积比低于约170。

    Reduced Length Transition Ducts
    18.
    发明申请
    Reduced Length Transition Ducts 有权
    减少长度过渡管道

    公开(公告)号：US20150300253A1

    公开(公告)日：2015-10-22

    申请号：US14518009

    申请日：2014-10-20

    Applicant: United Technologies Corporation

    Inventor： Wesley K. Lord

    IPC: F02C7/04 ,  F01D25/24 ,  F02C3/04

    CPC classification number: F01D9/02 ,  F01D9/023 ,  F01D9/047 ,  F02C3/107 ,  F05D2240/12 ,  F05D2250/38 ,  Y02T50/671

    Abstract: A transition duct defining an airflow pathway between a low pressure compressor and a high pressure compressor of a gas turbine engine is disclosed. The transition duct may comprise an inner wall and an outer wall located radially outward of the inner wall with respect to a central axis of the gas turbine engine. It may further comprise a first bend configured to turn the airflow radially inward with respect to the central axis, and a turning vane located at the first bend between the inner wall and the outer wall. The turning vane may be configured to assist the first bend in turning the airflow radially inward.

    Abstract translation: 公开了一种限定燃气轮机的低压压缩机和高压压缩机之间的气流路径的过渡管道。 过渡管道可以包括相对于燃气涡轮发动机的中心轴线位于内壁的径向外侧的内壁和外壁。 其还可以包括构造成相对于中心轴径向向内转动气流的第一弯曲部，以及位于内壁和外壁之间的第一弯曲处的转动叶片。 转动叶片可以构造成辅助第一弯曲使气流径向向内转动。

    Gas turbine engine with low fan noise
    20.
    发明授权
    Gas turbine engine with low fan noise 有权

    公开(公告)号：US10704415B2

    公开(公告)日：2020-07-07

    申请号：US14761689

    申请日：2014-03-11

    Applicant: UNITED TECHNOLOGIES CORPORATION

    Inventor： Bruce L. Morin ,  Wesley K. Lord

    IPC: F01D25/04 ,  F01D5/02 ,  F01D5/14 ,  F01D9/02 ,  F02K1/15 ,  F02K1/16 ,  F01D17/14

    Abstract: In accordance with one aspect of the disclosure, a gas turbine engine, method of using and designing such is disclosed. The gas turbine engine may comprise a fan including a plurality of blades, and a variable area fan nozzle. The fan may be configured to have a design point fan tip leading edge relative flow angle βADP, and may be further configured to have an off-design point fan tip leading edge relative flow angle β at an off-design fan operating point. The variable area fan nozzle may be configured to manipulate the amount of air flowing through the fan so that the absolute value of a difference between the design point fan tip leading edge relative flow angle βADP and the off-design point fan tip leading edge relative flow angle β is in a specified range.

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