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    METHODS AND SYSTEMS FOR DETECTING LEAN BLOWOUT IN GAS TURBINE SYSTEMS
    4.
    发明申请
    METHODS AND SYSTEMS FOR DETECTING LEAN BLOWOUT IN GAS TURBINE SYSTEMS 有权
    用于检测气体涡轮机系统中的亮度喷雾的方法和系统

    公开(公告)号：US20150219019A1

    公开(公告)日：2015-08-06

    申请号：US14171001

    申请日：2014-02-03

    申请人： General Electric Company

    发明人： Sarah Lori Crothers ,  Scott Arthur Day ,  Scott Alan Kopcho

    IPC分类号： F02C9/28

    CPC分类号： F02C9/28 ,  F05D2220/32 ,  F05D2260/83 ,  F05D2270/083 ,  F05D2270/092 ,  F05D2270/54

    摘要： A system includes a control system. The control system includes a coherence derivation system configured to derive a coherence between respective outputs of each of a plurality of combustors coupled to a gas turbine system, and a phase derivation system configured to derive a phase difference between the respective outputs of each of the plurality of combustors coupled to the gas turbine system. The control system is configured to derive an indication of an impending lean blowout (LBO) or an actual LBO of at least one of the plurality of combustors based at least in part on the coherence derivation, the phase derivation, or a combination thereof.

    摘要翻译： 系统包括控制系统。 控制系统包括相干导出系统，被配置为导出耦合到燃气涡轮机系统的多个燃烧器中的每一个的相应输出之间的相干性;以及相位导出系统，被配置为导出多个燃烧器中的每一个的相应输出之间的相位差 的燃烧器连接到燃气轮机系统。 控制系统被配置为至少部分地基于相干导出，相位推导或其组合来导出即将到来的贫气井喷（LBO）或至少一个燃烧器的实际LBO的指示。

    Method and device for monitoring a turbine engine of an aircraft
    5.
    发明授权
    Method and device for monitoring a turbine engine of an aircraft 有权
    用于监测飞行器的涡轮发动机的方法和装置

    公开(公告)号：US08918264B2

    公开(公告)日：2014-12-23

    申请号：US13488574

    申请日：2012-06-05

    申请人： Patrick Jegu ,  Manfred Birnfeld

    发明人： Patrick Jegu ,  Manfred Birnfeld

    IPC分类号： G06F19/00 ,  F01D21/00 ,  F02C7/262 ,  F02C9/00

    CPC分类号： F01D21/003 ,  F02C7/262 ,  F02C9/00 ,  F05D2220/323 ,  F05D2260/80 ,  F05D2270/092 ,  F05D2270/303 ,  Y02T50/671

    摘要： A method and device aids in monitoring at least one turbine engine in an aircraft. The method includes determining a value illustrating a thermomechanical state of the turbine engine, and comparing the determined value with at least one threshold. The method also includes displaying in the cockpit of the aircraft, at least one indication related to the operation of the turbine engine, according to this comparison. As a result, conditions that could result in mechanical degradation or failure of the turbine engine are identified for a crew of the aircraft.

    摘要翻译： 方法和装置有助于监测飞行器中的至少一个涡轮发动机。 该方法包括确定示出涡轮发动机的热机械状态的值，并将确定的值与至少一个阈值进行比较。 根据该比较，该方法还包括在飞行器的驾驶舱中显示与涡轮发动机的操作有关的至少一个指示。 因此，可以为飞机的船员确定可能导致涡轮发动机机械降解或故障的条件。

    GAS TURBINE ENGINE HAVING SLIM-LINE NACELLE
    6.
    发明申请
    GAS TURBINE ENGINE HAVING SLIM-LINE NACELLE 有权
    气体涡轮发动机具有滑动线

    公开(公告)号：US20140216002A1

    公开(公告)日：2014-08-07

    申请号：US14248444

    申请日：2014-04-09

    申请人： UNITED TECHNOLOGIES CORPORATION

    发明人： Stephen A. Morford ,  Michael J. Larkin

    IPC分类号： F02K1/16 ,  F02K1/06

    CPC分类号： F02K1/165 ,  F02K1/06 ,  F02K1/15 ,  F02K3/06 ,  F05D2260/80 ,  F05D2270/09 ,  F05D2270/092 ,  Y02T50/671 ,  Y10T29/4932 ,  Y10T137/0536

    摘要： A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a nacelle assembly that includes an inlet lip section and an inlet internal diffuser section downstream of the inlet lip section. A variable area fan nozzle is positioned near an aft segment of the nacelle assembly, the variable area fan nozzle adaptable to move between a first position having a first discharge airflow area and a second position having a second discharge airflow area greater than the first discharge airflow area. At least one boundary layer control device is positioned near one of the inlet lip section and the inlet internal diffuser section. A controller is configured to move the variable area fan nozzle from the first position to the second position and to actuate the at least one boundary layer control device to introduce an airflow in response to an operability condition.

    摘要翻译： 根据本公开的示例性方面的燃气涡轮发动机尤其包括机舱组件，其包括入口唇部和入口唇部的下游的入口内部扩散器部分。 可变区域风扇喷嘴位于机舱组件的后段附近，可变区域风扇喷嘴适于在具有第一排放气流区域的第一位置和具有大于第一排出气流的第二排放气流区域的第二位置之间移动 区。 至少一个边界层控制装置位于入口唇部和入口内部扩散器部分中的一个附近。 控制器构造成将可变区域风扇喷嘴从第一位置移动到第二位置，并且致动至少一个边界层控制装置以响应于可操作性条件引入气流。

    Gas turbine engine with variable area fan nozzle positioned for starting
    7.
    发明授权
    Gas turbine engine with variable area fan nozzle positioned for starting 有权
    具有可变面积风扇喷嘴的燃气涡轮发动机定位用于启动

    公开(公告)号：US08596072B2

    公开(公告)日：2013-12-03

    申请号：US13617506

    申请日：2012-09-14

    申请人： Eric J. Wehmeier

    发明人： Eric J. Wehmeier

    IPC分类号： F02K3/02

    CPC分类号： F02C9/20 ,  F02K1/1207 ,  F02K1/165 ,  F05D2260/85 ,  F05D2270/092

    摘要： A gas turbine engine has a compressor section, a low spool, and a fan. The fan delivers air into the compressor section and into a bypass duct having a variable area nozzle. The compressor section compresses air and delivers it into a combustion section. The combustion section mixes air with fuel, igniting the fuel, and driving the products of the combustion across a turbine. The turbine drives the low spool. A control for the gas turbine engine is programmed to position the nozzle at startup of the engine to increase airflow across the fan. A variable inlet guide vane is positioned upstream of the compressor section. The control also positions the variable inlet guide vane at start-up to increase air flow across the compressor section.

    摘要翻译： 燃气涡轮发动机具有压缩机部，低阀芯和风扇。 风扇将空气输送到压缩机部分并进入具有可变区域喷嘴的旁路管道。 压缩机部分压缩空气并将其输送到燃烧部分。 燃烧部分将空气与燃料混合，点燃燃料，并将燃烧产物驱动穿过涡轮机。 涡轮驱动低阀芯。 燃气涡轮发动机的控制被编程为在发动机起动时定位喷嘴以增加风扇的气流。 可变入口引导叶片位于压缩机部分的上游。 该控制器还可以将起动时的可变入口导向叶片定位，以增加整个压缩机部分的空气流量。

    Gas turbine engine with variable area fan nozzle positioned for starting
    9.
    发明授权
    Gas turbine engine with variable area fan nozzle positioned for starting 有权
    具有可变面积风扇喷嘴的燃气涡轮发动机定位用于启动

    公开(公告)号：US08291690B1

    公开(公告)日：2012-10-23

    申请号：US13367579

    申请日：2012-02-07

    申请人： Eric J. Wehmeier

    发明人： Eric J. Wehmeier

    IPC分类号： F02K3/02

    CPC分类号： F02C9/20 ,  F02K1/1207 ,  F02K1/165 ,  F05D2260/85 ,  F05D2270/092

    摘要： A gas turbine engine includes a variable inlet guide vane positioned forwardly of a low pressure compressor. The angle of the inlet guide vane is controlled at startup to increase airflow into the compressor. This is particularly useful when the gas turbine engine is being restarted while an associated aircraft is in the air, and is relied upon to increase windmill speed of the compressor and turbine rotors. A method and nozzle are also disclosed.

    摘要翻译： 燃气涡轮发动机包括位于低压压缩机前方的可变入口引导叶片。 入口导叶的角度在启动时被控制，以增加进入压缩机的气流。 当燃气涡轮发动机在相关联的飞行器处于空气中时重新启动，并且依赖于增加压缩机和涡轮转子的风车速度时，这是特别有用的。 还公开了一种方法和喷嘴。

    FUEL SUPPLY SYSTEM WITH MULTIPLE PUMPING MEANS
    10.
    发明申请
    FUEL SUPPLY SYSTEM WITH MULTIPLE PUMPING MEANS 审中-公开
    具有多种泵送装置的燃油供应系统

    公开(公告)号：US20120204532A1

    公开(公告)日：2012-08-16

    申请号：US13369628

    申请日：2012-02-09

    申请人： Nicolas POTEL ,  Regis Michel Paul DELDALLE

    发明人： Nicolas POTEL ,  Regis Michel Paul DELDALLE

    IPC分类号： F02G3/00

    CPC分类号： F02C7/236 ,  F02C9/48 ,  F05D2260/85 ,  F05D2270/092 ,  F05D2270/64

    摘要： The invention proposes a system for supplying (10) fuel to an aircraft turbomachine, which includes a low-pressure circuit (14) in which a low-pressure pump (12) is installed, and a high-pressure circuit (44) supplying an injection device (20) of the turbomachine and actuators (22) of components of the turbomachine, which includes several high-pressure pumps (30, 32), characterised in that each of the said high-pressure pumps (30, 32) consists of a stage of a high-pressure pump with several stages, each stage of the said pump with several stages being associated with each of the said segments (24, 26).

    摘要翻译： 本发明提出了一种用于向飞机涡轮机供给（10）燃料的系统，该系统包括安装有低压泵（12）的低压回路（14）和提供低压泵（12）的高压回路（44） 涡轮机的喷射装置（20）和涡轮机的部件的致动器（22），其包括几个高压泵（30,32），其特征在于，每个所述高压泵（30,32）由 具有几级的高压泵的级，所述泵的每个级与所述段（24,26）中的每一个相关联。