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公开(公告)号：US20190271269A1
公开(公告)日：2019-09-05
申请号：US16407505
申请日：2019-05-09
摘要： A propulsion system includes a fan, a gear, a turbine configured to drive the gear to, in turn, drive the fan. The turbine has an exit point, and a diameter (Dt) is defined at the exit point. A nacelle surrounds a core engine housing. The fan is configured to deliver air into a bypass duct defined between the nacelle and the core engine housing. A core engine exhaust nozzle is provided downstream of the exit point. A downstream most point of the core engine exhaust nozzle is defined at a distance from the exit point. A ratio of the distance to the diameter is greater than or equal to about 0.90.
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公开(公告)号：US10371092B2
公开(公告)日：2019-08-06
申请号：US14817611
申请日：2015-08-04
IPC分类号： F02K1/78 , B64D29/00 , F02K1/40 , F02K1/52 , F02K1/00 , F02K1/82 , F02K1/36 , F02K3/077 , F02C3/04 , F02K3/06 , B64D29/02 , B64D29/04
摘要： A nacelle for a gas turbine engine includes a ring shaped body defining a center axis and having a radially outward surface and a radially inward surface. An aft portion of the radially inward surface includes an axially extending convergent-divergent exit nozzle. An axially extending secondary duct passes through the nacelle in the convergent-divergent exit nozzle. The axially extending secondary duct includes an inlet at a convergent portion of the convergent-divergent exit nozzle and an outlet at a divergent portion of the convergent-divergent exit nozzle.
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3.发明授权公开(公告)号：US09920653B2
公开(公告)日：2018-03-20
申请号：US13721095
申请日：2012-12-20
CPC分类号： F01D25/24 , F01D5/02 , F01D5/141 , F01D15/12 , F01D17/105 , F02C7/04 , F02K3/06 , F05D2220/32 , F05D2250/00 , F05D2260/40311 , F05D2260/96 , Y02T50/671 , Y02T50/673
摘要： According to an example embodiment, a gas turbine engine assembly includes, among other things, a fan that has a plurality of fan blades. A diameter of the fan has a dimension D that is based on a dimension of the fan blades. Each fan blade has a leading edge. An inlet portion is situated forward of the fan. A length of the inlet portion has a dimension L between a location of the leading edge of at least some of the fan blades and a forward edge on the inlet portion. A dimensional relationship of L/D is between about 0.2 and 0.45.
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公开(公告)号：US20160195038A1
公开(公告)日：2016-07-07
申请号：US14817611
申请日：2015-08-04
CPC分类号： F02K1/78 , B64D29/00 , B64D29/02 , B64D29/04 , F02C3/04 , F02K1/00 , F02K1/36 , F02K1/40 , F02K1/52 , F02K1/82 , F02K3/06 , F02K3/077
摘要： A nacelle for a gas turbine engine includes a ring shaped body defining a center axis and having a radially outward surface and a radially inward surface. An aft portion of the radially inward surface includes an axially extending convergent-divergent exit nozzle. An axially extending secondary duct passes through the nacelle in the convergent-divergent exit nozzle. The axially extending secondary duct includes an inlet at a convergent portion of the convergent-divergent exit nozzle and an outlet at a divergent portion of the convergent-divergent exit nozzle.
摘要翻译： 用于燃气涡轮发动机的机舱包括限定中心轴线并具有径向向外表面和径向向内表面的环形主体。 径向向内表面的后部包括轴向延伸的收敛发散出口。 轴向延伸的二次管道穿过会聚发散出口喷嘴中的机舱。 轴向延伸的次级管道包括在收敛发散出口喷嘴的收敛部分处的入口和在会聚发散出口喷嘴的发散部分处的出口。
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公开(公告)号：US10479519B2
公开(公告)日：2019-11-19
申请号：US15378300
申请日：2016-12-14
摘要： A fan assembly for a gas turbine engine includes a fan including a plurality of fan blades extending radially outwardly from a fan hub to a blade tip defining a fan diameter. A nacelle surrounds the fan and defines a fan inlet upstream of the fan, relative to an airflow direction into the fan. The nacelle has a forwardmost edge defining an inlet length from the forwardmost edge to a leading edge of a fan blade. A ratio of inlet length to fan diameter is between 0.20 and 0.45. A nacelle inner surface defines a nacelle flowpath having a convex throat portion and a diffusion portion between the throat portion and the leading edge of the fan blade at a topmost portion of the nacelle. The throat portion has a minimum throat radius measured from a fan central axis greater than a blade tip radius.
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公开(公告)号：US10436121B2
公开(公告)日：2019-10-08
申请号：US14038886
申请日：2013-09-27
摘要： A propulsion system includes a fan, a gear, a turbine configured to drive the gear to, in turn, drive the fan. The turbine has an exit point, and a diameter (Dt) is defined at the exit point. A nacelle surrounds a core engine housing. The fan is configured to deliver air into a bypass duct defined between the nacelle and the core engine housing. A core engine exhaust nozzle is provided downstream of the exit point. A downstream most point of the core engine exhaust nozzle is defined at a distance from the exit point. A ratio of the distance to the diameter is greater than or equal to about 0.90.
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公开(公告)号：US10316757B2
公开(公告)日：2019-06-11
申请号：US14875762
申请日：2015-10-06
摘要： A propulsion system according to an example of the present disclosure includes, among other things, a geared architecture configured to drive a fan section including a fan, and a turbine configured to drive the geared architecture. The turbine has an exit point, and a diameter (Dt) defined as the radially outer diameter of a last blade airfoil stage in the turbine at the exit point. A nacelle at least partially surrounds a core engine housing. The fan configured to deliver air into a bypass duct is defined between the nacelle and the core engine housing. A core engine exhaust nozzle is downstream of the exit point, with a downstream most point of the core engine exhaust nozzle being defined at a distance (Lc or Ln) from the exit point.
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8.发明申请公开(公告)号：US20180209379A1
公开(公告)日：2018-07-26
申请号：US15887183
申请日：2018-02-02
CPC分类号： F02K3/068 , F01D25/24 , F02C7/04 , F05D2220/36 , F05D2240/303 , F05D2260/40311 , F05D2260/96 , Y02T50/671
摘要： 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 turbine section including a high pressure turbine and a low pressure turbine, the low pressure turbine driving the fan, a 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 no more than 0.45.
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公开(公告)号：US20170298954A1
公开(公告)日：2017-10-19
申请号：US15378371
申请日：2016-12-14
发明人： Yuan J. Qiu , Robert E. Malecki , Wesley K. Lord
摘要： A fan assembly for a gas turbine engine includes a fan including a plurality of fan blades. Each fan blade extends radially outwardly from a fan hub to a blade tip. The plurality of blade tips define a fan diameter. A nacelle surrounds the fan and defines a fan inlet upstream of the fan, relative to an airflow direction into the fan. The nacelle has a forwardmost edge defining an inlet length from the forwardmost edge to a leading edge of a fan blade of the plurality of fan blades. A ratio of inlet length to fan diameter is between 0.20 and 0.45. A nacelle inner surface defines a nacelle flowpath. The nacelle flowpath has a convex throat portion and a concave diffusion portion between the throat portion and the leading edge of the fan blade at a bottommost portion of the nacelle.
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公开(公告)号：US20160201569A1
公开(公告)日：2016-07-14
申请号：US14875750
申请日：2015-10-06
CPC分类号： F02C7/36 , B64D33/02 , B64D2033/0286 , F02C3/10 , F02C3/107 , F02K3/06 , F02K3/068 , F05D2220/36 , Y02T50/672
摘要： A propulsion system according to an example of the present disclosure includes, among other things, a geared architecture configured to drive a fan section including a fan, and a turbine section configured to drive the geared architecture. The turbine section has an exit point, and a diameter (Dt) defined as the outer diameter of a last blade airfoil stage in the turbine section at the exit point. A nacelle surrounds a core engine housing. The fan is configured to deliver air into a bypass duct defined between the nacelle and the core engine housing. A core engine exhaust nozzle is downstream of the exit point. A downstream most point of the core engine exhaust nozzle is defined at a distance (Lc or Ln) from the exit point.
摘要翻译： 根据本公开的示例的推进系统包括构造成驱动包括风扇的风扇部分和被配置为驱动齿轮架构的涡轮机部分的齿轮架构。 涡轮部分具有出口点，并且直径（Dt）被定义为在出口处涡轮部分中的最后叶片翼型段的外径。 机舱围绕核心发动机外壳。 风扇被配置为将空气输送到限定在机舱和核心发动机壳体之间的旁路管道中。 核心发动机排气喷嘴位于出口点的下游。 核心发动机排气喷嘴的下游最点位于出口点的距离（Lc或Ln）处。
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