公开(公告)号：US20160076460A1

公开(公告)日：2016-03-17

申请号：US14824292

申请日：2015-08-12

Applicant: United Technologies Corporation

Inventor： Frederick M. Schwarz ,  Robert E. Malecki

IPC: F02C9/18 ,  F02C3/04 ,  F02C7/36 ,  F01D25/24

CPC classification number: F02C9/18 ,  B64D2033/022 ,  F01D25/24 ,  F02C3/04 ,  F02C7/05 ,  F02C7/36 ,  F02K3/06 ,  F05D2250/14 ,  F05D2250/232 ,  F05D2250/24

Abstract: A gas turbine engine includes a nacelle defining a centerline axis and an annular splitter radially inward from the nacelle. A spinner is radially inward of the nacelle forward of a compressor section. A fan blade extends from a fan blade platform. A distance X is the axial distance from a first point to a second point, wherein the first point is defined on a leading edge of the annular splitter and the second point is defined on a leading edge of the fan blade where the fan blade meets the fan blade platform. A distance H is the radial distance from the first point to the second point. The relative position of the first point and the second point is governed by the ratio of X/H≧1.5 for reducing foreign object debris (FOD) intake into the compressor section.

Abstract translation: 燃气涡轮发动机包括限定中心线轴线的机舱和从机舱径向向内的环形分配器。 旋转器在压缩机部分的前方的机舱的径向内侧。 风扇叶片从风扇叶片平台延伸。 距离X是从第一点到第二点的轴向距离，其中第一点被限定在环形分配器的前缘上，第二点限定在风扇叶片的前缘上，其中风扇叶片与 风扇叶片平台。 距离H是从第一点到第二点的径向距离。 第一点和第二点的相对位置由减少异物（FOD）进入压缩机段的X /H≥1.5的比率决定。

公开(公告)号：US20150247425A1

公开(公告)日：2015-09-03

申请号：US14676354

申请日：2015-04-01

Applicant: United Technologies Corporation

Inventor： Michael J. Murphy ,  Robert E. Malecki

IPC: F01D25/24 ,  F02C3/10

CPC classification number: F01D25/24 ,  B64D29/02 ,  F02C3/10

Abstract: A nacelle for a turbofan propulsion system that extends along a centerline includes a forward cowling and an aft cowling. To improve the fit of a turbofan propulsion system in the space between the wing and ground of a fixed-wing aircraft, the aft cowling of the nacelle is modified. The aft cowling has a non-circular cross-sectional geometry disposed in a plane substantially perpendicular to the centerline. The non-circular cross-sectional geometry includes a radially recessed section disposed between first and second curved sections. The first and the second curved sections each have a radius that is greater than a radial distance between the centerline and a center point of the radially recessed section.

Abstract translation: 一个沿着中心线延伸的涡轮风扇推进系统的机舱包括一个正向整流罩和一个后罩。 为了提高涡轮风扇推进系统在固定翼飞机机翼与地面之间的空间配合，改装了机舱的后罩。 后罩具有设置在基本上垂直于中心线的平面中的非圆形横截面几何形状。 非圆形横截面几何形状包括设置在第一和第二弯曲部分之间的径向凹陷部分。 第一和第二弯曲部分各自具有大于中心线和径向凹陷部分的中心点之间的径向距离的半径。

公开(公告)号：US11015550B2

公开(公告)日：2021-05-25

申请号：US15887183

申请日：2018-02-02

Applicant: United Technologies Corporation

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

IPC: F02K3/068 ,  F01D25/24 ,  F02C7/04

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 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.

公开(公告)号：US20200025036A1

公开(公告)日：2020-01-23

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.

公开(公告)号：US10330047B2

公开(公告)日：2019-06-25

申请号：US14768830

申请日：2013-12-18

Applicant: UNITED TECHNOLOGIES CORPORATION

Inventor： Wesley K. Lord ,  Robert E. Malecki

IPC: F02K3/02 ,  F02K3/06 ,  F02K1/78 ,  B64D33/02 ,  F01D25/24 ,  F02C3/04 ,  B64D33/04

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.

公开(公告)号：US10294871B2

公开(公告)日：2019-05-21

申请号：US14875750

申请日：2015-10-06

Applicant: United Technologies Corporation

Inventor： Frederick M. Schwarz ,  Robert E. Malecki

IPC: F02K3/06 ,  F02C7/36 ,  F02C3/10 ,  F02K3/068 ,  B64D33/02 ,  F02C3/107

Abstract: 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.

公开(公告)号：US20180163568A1

公开(公告)日：2018-06-14

申请号：US15887136

申请日：2018-02-02

Applicant: United Technologies Corporation

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

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

CPC classification number: 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

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 no more than 0.45.

公开(公告)号：US20160108854A1

公开(公告)日：2016-04-21

申请号：US14882760

申请日：2015-10-14

Applicant: United Technologies Corporation

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

IPC: F02K3/06 ,  F04D19/02 ,  F01D15/12 ,  F01D17/10 ,  F01D5/02 ,  F01D5/12

CPC classification number: F02K3/06 ,  F01D5/02 ,  F01D5/141 ,  F01D15/12 ,  F01D17/105 ,  F02C7/04 ,  F02C7/045 ,  F04D29/522 ,  F04D29/547 ,  F05D2220/32 ,  F05D2250/00 ,  Y02T50/671 ,  Y02T50/673

Abstract: A gas turbine engine assembly according to an example of the present disclosure includes, among other things, a 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, a geared architecture configured to drive the fan, a turbine section configured to drive the geared architecture, a compressor section including a first compressor and a second compressor, and an inlet portion 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 about 0.45.

Abstract translation: 根据本公开的示例的燃气涡轮发动机组件包括除了别的以外的包括多个风扇叶片的风扇，具有基于风扇叶片的尺寸的尺寸D的风扇的直径，每个风扇 具有前缘的叶片，配置成驱动风扇的齿轮架构，构造成驱动齿轮架构的涡轮部分，包括第一压缩机和第二压缩机的压缩机部分以及风扇前部的入口部分。 入口部分的长度在至少一些风扇叶片的前缘的位置与入口部分的前边缘之间具有尺寸L。 L / D的尺寸关系为约0.2至约0.45。

公开(公告)号：US20140286754A1

公开(公告)日：2014-09-25

申请号：US13792303

申请日：2013-03-11

Applicant: UNITED TECHNOLOGIES CORPORATION

Inventor： Frederick M. Schwarz ,  Robert E. Malecki

IPC: F02C7/36

CPC classification number: F02C7/36 ,  B64D33/02 ,  B64D2033/0286 ,  F02C3/10 ,  F02C3/107 ,  F02K3/06 ,  F02K3/068 ,  F05D2220/36 ,  Y02T50/672

Abstract: 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.

Abstract translation: 推进系统包括风扇，齿轮，被配置为驱动齿轮的涡轮，从而驱动风扇。 涡轮机具有出口点，并且在出口处限定直径（Dt）。 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。 在出口点的下游设有核心发动机排气喷嘴。 核心发动机排气喷嘴的下游最点定义在离出口一定距离处。 距离直径之比大于或等于约0.90。

公开(公告)号：US10724541B2

公开(公告)日：2020-07-28

申请号：US15378371

申请日：2016-12-14

Applicant: United Technologies Corporation

Inventor： Yuan J. Qiu ,  Robert E. Malecki ,  Wesley K. Lord

IPC: F04D29/54 ,  F02C7/04 ,  F02K3/06 ,  B64D33/02 ,  F04D29/32

Abstract: 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.