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公开(公告)号：US20170122219A1
公开(公告)日：2017-05-04
申请号：US15411147
申请日：2017-01-20
Applicant: United Technologies Corporation
Inventor： Karl L. Hasel , Joseph B. Staubach , Brian D. Merry , Gabriel L. Suciu , Christopher M. Dye
IPC: F02C7/36 , F02C7/06 , F02K3/06 , F02C9/18 , F01D21/00 , F01D9/02 , F01D25/24 , F02C3/06 , F01D5/06
CPC classification number: F02C7/36 , F01D5/06 , F01D5/12 , F01D9/02 , F01D21/003 , F01D25/24 , F01D25/28 , F02C3/04 , F02C3/06 , F02C3/107 , F02C7/06 , F02C7/28 , F02C9/18 , F02K3/025 , F02K3/06 , F05D2220/32 , F05D2220/323 , F05D2240/12 , F05D2240/35 , F05D2240/55 , F05D2260/40 , F05D2260/4031 , F05D2260/40311
Abstract: A gas turbine engine includes, among other things, a fan section, a core engine, a bypass passage, and a bypass ratio defined as the volume of air passing into the bypass passage compared to the volume of air passing into the core engine, the bypass ratio being greater than 8 at cruise power. A gear arrangement drives the fan section. A guide vane includes a forward attachment, the forward attachment positioned aft of a plumbing connection area. A compressor section includes both a first compressor and a second compressor. A lubrication system and a compressed air system are in fluid communication with the gear arrangement. A turbine section drives the gear arrangement, and includes a low pressure turbine. An overall pressure ratio is provided by the combination of a pressure ratio across the first compressor and a pressure ratio across the second compressor, and greater than about 50, measured at sea level and at a static, full-rated takeoff power. The pressure ratio across the second compressor is greater than about 7.
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公开(公告)号：US20150345392A1
公开(公告)日：2015-12-03
申请号：US14821409
申请日：2015-08-07
Applicant: United Technologies Corporation
Inventor： Brian D. Merry , Gabriel L. Suciu , Karl L. Hasel
CPC classification number: F02C7/06 , F01D25/162 , F02C3/04 , F02C3/107 , F02C7/36 , F02K3/00 , F02K3/04 , Y02T50/671 , Y10T29/4932
Abstract: A gas turbine engine includes a housing includes an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. A rotor is connected to the hub and supports a compressor section. The geared architecture includes an epicyclic gear train. A fan is rotationally driven by the geared architecture. First and second bearings support the shaft relative to the intermediate case and the inlet case, respectively. The radially inner boundary of the core inlet is at a location of a core inlet stator and the radially inner boundary of the compressor section inlet is at a location of the first stage low-pressure compressor rotor.
Abstract translation: 燃气涡轮发动机包括壳体，其包括分别设置入口壳体流动路径和中间壳体流动路径的入口壳体和中间壳体。 转子连接到轮毂并支撑压缩机部分。 齿轮架构包括行星齿轮系。 风扇由齿轮架构旋转驱动。 第一和第二轴承分别相对于中间壳体和进气口支撑轴。 芯入口的径向内边界处于芯入口定子的位置，压缩机部分入口的径向内边界位于第一级低压压缩机转子的位置。
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公开(公告)号：US20150192298A1
公开(公告)日：2015-07-09
申请号：US14592043
申请日：2015-01-08
Applicant: United Technologies Corporation
Inventor： Karl L. Hasel , Peter G. Smith , Stuart S. Ochs
CPC classification number: F23R3/02 , F01D17/162 , F04D29/563 , F05D2220/36 , Y02T50/671
Abstract: A turbofan engine includes a fan driven by a low pressure turbine through a gear reduction. The gear reduction has a gear ratio of greater than or equal to about 2.4. The low pressure turbine has an expansion ratio greater than or equal to about 5. The fan has a bypass ratio greater than or equal to about 8. In other features, a turbofan engine includes a variable geometry fan exit guide vane (FEGV) system having a multiple of circumferentially spaced radially extending fan exit guide vanes. Rotation of the fan exit guide vanes between a nominal position and a rotated position selectively changes a fan bypass flow path to permit efficient operation at various flight conditions.
Abstract translation: 涡轮风扇发动机包括由低压涡轮机通过齿轮减速器驱动的风扇。 齿轮减速器具有大于或等于约2.4的齿轮比。 低压涡轮机具有大于或等于约5的膨胀比。风扇具有大于或等于约8的旁通比。在其它特征中，涡轮风扇发动机包括可变几何风扇出口导叶（FEGV）系统，其具有 周向间隔开的径向延伸的风扇出口导叶的多个。 风扇出口导向叶片在标称位置和旋转位置之间的旋转选择性地改变风扇旁路流动路径以允许在各种飞行条件下的有效操作。
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14.发明授权Method for setting a gear ratio of a fan drive gear system of a gas turbine engine 有权用于设定燃气轮机的风扇驱动齿轮系的齿轮比的方法公开(公告)号：US08814494B1
公开(公告)日：2014-08-26
申请号：US14203863
申请日：2014-03-11
Applicant: United Technologies Corporation
Inventor： William G. Sheridan , Karl L. Hasel
CPC classification number: F02C7/36 , F05D2230/60 , F05D2260/40311 , F05D2260/941
Abstract: A gas turbine engine includes a fan section including a fan rotatable about an axis. A speed reduction device is connected to the fan. The speed reduction device includes a planetary fan drive gear system with a planet gear ratio of at least 2.5. A bypass ratio is greater than about 11.0.
Abstract translation: 燃气涡轮发动机包括风扇部分，风扇部分包括可围绕轴线旋转的风扇。 减速装置连接到风扇。 减速装置包括具有至少2.5的行星齿轮比的行星风扇驱动齿轮系统。 旁路比大于约11.0。
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公开(公告)号：US20140157755A1
公开(公告)日：2014-06-12
申请号：US14179743
申请日：2014-02-13
Applicant: United Technologies Corporation
Inventor： Karl L. Hasel , Joseph B. Staubach , Brian D. Merry , Gabriel L. Suciu , Christopher M. Dye
IPC: F02K3/06
CPC classification number: F02K3/06 , F02C3/107 , F05D2260/4031
Abstract: A gas turbine engine has a fan section, including a fan, a gear arrangement configured to drive the fan section, and a compressor section, including both a first compressor and a second compressor. A turbine section is configured to drive the compressor section and the gear arrangement. An overall pressure ratio is provided by the combination of the first compressor and the second compressor. The fan is configured to deliver a portion of air into the compressor section, and a portion of air into a bypass duct. A bypass ratio which is defined as a volume of air passing to the bypass duct compared to a volume of air passing into the compressor section is greater than or equal to about 8.
Abstract translation: 一种燃气涡轮发动机具有风扇部分，包括风扇，构造成驱动风扇部分的齿轮装置和包括第一压缩机和第二压缩机的压缩机部分。 涡轮部分构造成驱动压缩机部分和齿轮装置。 总压力比由第一压缩机和第二压缩机的组合提供。 风扇被配置为将一部分空气输送到压缩机部分中，并且一部分空气进入旁通管道。 被定义为与通过压缩机部分的空气体积相比通过旁通管道的空气体积的旁路比大于或等于约8。
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Patent Agency Ranking
公开(公告)号：US10215094B2
公开(公告)日：2019-02-26
申请号：US14821409
申请日：2015-08-07
Applicant: United Technologies Corporation
Inventor： Brian D. Merry , Gabriel L. Suciu , Karl L. Hasel
Abstract: A gas turbine engine includes a housing includes an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. A rotor is connected to the hub and supports a compressor section. The geared architecture includes an epicyclic gear train. A fan is rotationally driven by the geared architecture. First and second bearings support the shaft relative to the intermediate case and the inlet case, respectively. The radially inner boundary of the core inlet is at a location of a core inlet stator and the radially inner boundary of the compressor section inlet is at a location of the first stage low-pressure compressor rotor.
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17.发明申请公开(公告)号：US20180073439A9
公开(公告)日：2018-03-15
申请号：US14705577
申请日：2015-05-06
Applicant: United Technologies Corporation
Inventor： William G. Sheridan , Karl L. Hasel
CPC classification number: F02C7/36 , F02C3/04 , F02C3/06 , F02C3/107 , F02C3/113 , F02C7/00 , F02C7/32 , F05D2210/12 , F05D2220/32 , F05D2220/36 , F05D2240/40 , F05D2260/40311 , F16H1/28 , F16H1/36
Abstract: A gas turbine engine includes a fan section that includes a fan rotatable about an axis of rotation of the gas turbine engine. A speed reduction device is connected to the fan. The speed reduction device includes a star drive gear system with a star gear ratio of at least 1.5. A bypass ratio is greater than about 11.0.
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公开(公告)号：US09816443B2
公开(公告)日：2017-11-14
申请号：US15337171
申请日：2016-10-28
Applicant: United Technologies Corporation
Inventor： William G. Sheridan , Karl L. Hasel
CPC classification number: F02C7/36 , F01D1/04 , F01D9/041 , F01D25/162 , F02C3/04 , F02C7/00 , F02K3/06 , F05D2220/36 , F05D2240/12 , F05D2240/24 , F05D2240/50 , F05D2260/40311
Abstract: A gas turbine engine includes an engine centerline longitudinal axis and a fan section including a fan with fan blades and rotatable about the engine centerline longitudinal axis. A low corrected fan tip speed less than about 1400 ft/sec and the low corrected fan tip speed is an actual fan tip speed determined at an ambient temperature divided by [(Tram ° R)/(518.7 ° R)]0.5, where T represents the ambient temperature in degrees Rankine. A bypass ratio greater than about 11 and a speed reduction device having a gear system with a gear ratio. A low and high pressure turbine in communication with a first and second shaft, respectively. The first and second shafts are concentric and mounted via at least one of the bearing systems for rotation about the engine centerline longitudinal axis and the first shaft is in communication with the fan through the speed reduction device and the low pressure turbine includes four stages.
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19.发明申请METHOD FOR SETTING A GEAR RATIO OF A FAN DRIVE GEAR SYSTEM OF A GAS TURBINE ENGINE 审中-公开用于设定气体涡轮发动机的风扇驱动齿轮系统的齿轮比的方法公开(公告)号：US20170051677A1
公开(公告)日：2017-02-23
申请号：US15337171
申请日：2016-10-28
Applicant: United Technologies Corporation
Inventor： William G. Sheridan , Karl L. Hasel
CPC classification number: F02C7/36 , F01D1/04 , F01D9/041 , F01D25/162 , F02C3/04 , F02C7/00 , F02K3/06 , F05D2220/36 , F05D2240/12 , F05D2240/24 , F05D2240/50 , F05D2260/40311
Abstract: A gas turbine engine includes an engine centerline longitudinal axis and a fan section including a fan with fan blades and rotatable about the engine centerline longitudinal axis. A low corrected fan tip speed less than about 1400 ft/sec and the low corrected fan tip speed is an actual fan tip speed determined at an ambient temperature divided by [(Tram ° R)/(518.7° R)]0.5, where T represents the ambient temperature in degrees Rankine. A bypass ratio greater than about 11 and a speed reduction device having a gear system with a gear ratio. A low and high pressure turbine in communication with a first and second shaft, respectively. The first and second shafts are concentric and mounted via at least one of the bearing systems for rotation about the engine centerline longitudinal axis and the first shaft is in communication with the fan through the speed reduction device and the low pressure turbine includes four stages.
Abstract translation: 燃气涡轮发动机包括发动机中心线纵向轴线和风扇部分，风扇部分包括具有风扇叶片的风扇并可围绕发动机中心线纵向轴线旋转。 小于约1400ft / sec的低校正风扇叶尖速度和低校正风扇叶尖速度是在环境温度除以[（Tram°R）/（518.7°R）] 0.5时确定的实际风扇叶尖速度，其中T 表示环境温度，以度兰为单位。 旁路比大于约11，减速装置具有齿轮比齿轮系。 分别与第一和第二轴连通的低压和高压涡轮机。 第一和第二轴是同心的，并且经由至少一个轴承系统安装，用于围绕发动机中心线纵向轴线旋转，并且第一轴通过减速装置与风扇连通，低压涡轮机包括四个阶段。
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20.发明申请METHOD FOR SETTING A GEAR RATIO OF A FAN DRIVE GEAR SYSTEM OF A GAS TURBINE ENGINE 审中-公开用于设定气体涡轮发动机的风扇驱动齿轮系的齿轮比的方法公开(公告)号：US20150308335A1
公开(公告)日：2015-10-29
申请号：US14746910
申请日：2015-06-23
Applicant: United Technologies Corporation
Inventor： William G. Sheridan , Karl L. Hasel
CPC classification number: F02C7/32 , F02C3/04 , F02C3/06 , F02C3/107 , F02C3/113 , F02C7/00 , F02C7/36 , F05D2210/12 , F05D2220/32 , F05D2220/36 , F05D2240/40 , F05D2260/40311 , F16H1/28 , F16H1/36
Abstract: A gas turbine engine has a fan section including a fan rotatable about an axis. A speed reduction device is in communication with the fan. The speed reduction device includes a planetary fan drive gear system with a planet gear ratio of at least 2.6. A fan blade tip speed of the fan is less than 1400 fps. A low pressure turbine section is in communication with the speed reduction device. The low pressure turbine section includes three or four stages. A bypass ratio is between 11.0 and about 22.0. A method of improving performance of a gas turbine engine comprising
Abstract translation: 燃气涡轮发动机具有风扇部分，风扇部分包括可围绕轴线旋转的风扇。 减速装置与风扇通信。 减速装置包括具有至少2.6的行星齿轮比的行星风扇驱动齿轮系统。 风扇的风扇叶片尖端速度小于1400 fps。 低压涡轮部分与减速装置连通。 低压涡轮机部分包括三级或四级。 旁路比在11.0到22.0之间。 一种改善燃气涡轮发动机性能的方法，包括：
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