公开(公告)号：US10472081B2

公开(公告)日：2019-11-12

申请号：US15874335

申请日：2018-01-18

Applicant: United Technologies Corporation

Inventor： Gabriel L. Suciu ,  Jesse M. Chandler ,  Wesley K. Lord

IPC: B64D35/08 ,  B64C11/00 ,  B64C21/06 ,  B64C39/00 ,  B64C1/06 ,  B64D35/02 ,  B64C23/02 ,  B64C5/02 ,  F04D17/04 ,  B64D27/02 ,  B64D41/00

Abstract: A cross flow fan to be incorporated into an aircraft fuselage comprises an ingestion fan rotor to be positioned in a tail section of an aircraft fuselage to reduce boundary layer air from a top surface of the fuselage and to drive the air away from the top surface, and a drive arrangement for the ingestion fan rotor. An aircraft is also disclosed.

公开(公告)号：US09951633B2

公开(公告)日：2018-04-24

申请号：US14518009

申请日：2014-10-20

Applicant: UNITED TECHNOLOGIES CORPORATION

Inventor： Wesley K. Lord

IPC: F01D9/02 ,  F02C3/107 ,  F01D9/04

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.

公开(公告)号：US09920653B2

公开(公告)日：2018-03-20

申请号：US13721095

申请日：2012-12-20

Applicant: United Technologies Corporation

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

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

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 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.

公开(公告)号：US09771863B2

公开(公告)日：2017-09-26

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: B64D27/12 ,  F02C3/10 ,  F02K1/54 ,  F02C6/20 ,  F02C7/36 ,  B64D35/04

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.

公开(公告)号：US20160207631A1

公开(公告)日：2016-07-21

申请号：US14600946

申请日：2015-01-20

Applicant: United Technologies Corporation

Inventor： Gabriel L. Suciu ,  Wesley K. Lord ,  Jesse M. Chandler

IPC: B64D27/18 ,  F02C7/20 ,  B64D29/02 ,  B64D33/02 ,  B64C3/32

Abstract: An aircraft is provided that includes a wing and a pusher fan engine. The pusher fan engine is configured in the wing.

Abstract translation: 提供一种包括机翼和推进器风扇发动机的飞机。 推进器风扇发动机配置在机翼中。

公开(公告)号：US20160131044A1

公开(公告)日：2016-05-12

申请号：US15000077

申请日：2016-01-19

Applicant: UNITED TECHNOLOGIES CORPORATION

Inventor： Gabriel L. Suciu ,  Alan H. Epstein ,  Wesley K. Lord ,  Michael E. McCune ,  Brian D. Merry

IPC: F02C7/36 ,  F02C3/06

CPC classification number: F02C7/36 ,  F01D15/12 ,  F02C3/06 ,  F02C3/107 ,  F05D2220/32 ,  F05D2220/36 ,  F05D2240/24 ,  F05D2240/35 ,  F05D2240/40 ,  F05D2240/60 ,  F05D2260/4031 ,  F05D2260/40311

Abstract: A gas turbine engine includes a core engine with a compressor section, a combustor and a turbine. The turbine drives an output shaft, and the output shaft drives at least four gears. Each of the at least four gears extends through a drive shaft to drive an associated fan rotor.

公开(公告)号：US20150377123A1

公开(公告)日：2015-12-31

申请号：US14793785

申请日：2015-07-08

Applicant: United Technologies Corporation

Inventor： Paul R. Adams ,  Shankar S. Magge ,  Joseph Brent Staubach ,  Wesley K. Lord ,  Frederick M. Schwarz ,  Gabriel L. Suciu

IPC: F02C3/107 ,  F02K3/02 ,  F02C9/18

CPC classification number: F02C7/36 ,  F01D5/06 ,  F01D11/122 ,  F01D25/24 ,  F02C3/04 ,  F02C3/107 ,  F02C7/20 ,  F02C9/18 ,  F02K3/06 ,  F02K3/075 ,  F04D19/02 ,  F05B2250/283 ,  F05D2220/32 ,  F05D2220/323 ,  F05D2240/35 ,  F05D2240/60 ,  F05D2260/40311

Abstract: A turbofan engine comprises a fan having fan blades. A compressor is in communication with the fan section. The fan is configured to communicate a portion of air into a bypass path defining a bypass area outwardly of the compressor and a portion into the compressor. A bypass ratio is defined as air communicated through the bypass path relative to air communicated to the compressor being greater than about 6.0. A combustor is in fluid communication with the compressor. A turbine is in communication with the combustor. The turbine has a first turbine section that includes two or more stages and a second turbine section that includes at least two stages. A ratio of airfoils in the first turbine section to the bypass ratio is less than about 170. The first turbine section includes a maximum gas path radius. A ratio of the maximum gas path radius to a maximum radius of the fan blades is less than about 0.50. A speed reduction mechanism is coupled to the fan and rotatable by the turbine.

Abstract translation: 涡轮风扇发动机包括具有风扇叶片的风扇。 压缩机与风扇部分连通。 风扇被构造成将一部分空气传送到限定压缩机外部的旁路区域的旁路，以及一部分进入压缩机。 旁路比定义为相对于连通到压缩机的空气通过旁路通路的空气大于约6.0。 燃烧器与压缩机流体连通。 涡轮机与燃烧器连通。 涡轮机具有包括两级或更多级的第一涡轮部分和包括至少两级的第二涡轮部分。 第一涡轮机部分中的翼型件与旁路比率的比率小于约170.第一涡轮机部分包括最大气体路径半径。 风扇叶片的最大气路半径与最大半径的比值小于约0.50。 减速机构联接到风扇并由涡轮旋转。

公开(公告)号：US20150377122A1

公开(公告)日：2015-12-31

申请号：US14793770

申请日：2015-07-08

Applicant: United Technologies Corporation

Inventor： Paul R. Adams ,  Shankar S. Magge ,  Joseph Brent Staubach ,  Wesley K. Lord ,  Frederick M. Schwarz ,  Gabriel L. Suciu

IPC: F02C3/10 ,  F02K3/06 ,  F02C7/36

CPC classification number: F02K3/06 ,  F02C3/107 ,  F02C7/36 ,  F02K3/075 ,  F05D2260/40311

Abstract: A turbofan engine includes an engine case, a gaspath through the engine case, a fan having an array of fan blades, a compressor in fluid communication with the fan, a combustor in fluid communication with the compressor, and a turbine in fluid communication with the combustor. The turbine has a fan drive turbine section having 3 to 6 blade stages. A speed reduction mechanism couples the fan drive turbine section to the fan. A ratio of maximum gaspath radius along the low pressure turbine section to maximum radius of the fan blades is less than about 0.55. A bypass area ratio is greater than about 6.0. A ratio of a fan drive turbine section airfoil count to the bypass area ratio is less than about 170 and a second turbine section.

Abstract translation: 涡轮风扇发动机包括发动机壳体，通过发动机壳体的气路，具有风扇叶片阵列的风扇，与风扇流体连通的压缩机，与压缩机流体连通的燃烧器，以及与压缩机流体连通的涡轮机 燃烧器。 涡轮机具有风扇驱动涡轮机部分，具有3到6个叶片级。 减速机构将风扇驱动涡轮部分连接到风扇。 沿着低压涡轮机段的最大气路径与风扇叶片的最大半径之比小于约0.55。 旁路面积比大于约6.0。 风扇驱动涡轮机翼型计数与旁路面积比的比率小于约170，而第二涡轮部分。

公开(公告)号：US11480104B2

公开(公告)日：2022-10-25

申请号：US14767396

申请日：2014-02-26

Applicant: UNITED TECHNOLOGIES CORPORATION

Inventor： Yuan J. Qiu ,  Robert M. Miller ,  Wesley K. Lord ,  Gavin Hendricks

IPC: F02C7/04 ,  F02K3/06 ,  F01D5/02 ,  F01D17/10 ,  F01D25/24 ,  F02C3/04 ,  B64D33/02

Abstract: A gas turbine engine that includes a fan nacelle and a core nacelle that provide a bypass flow path radially between. The fan nacelle has an inlet including a throat. The inlet has an inlet forward-most point. A fan is arranged in the bypass flow path and rotatable about an axis. The fan has a leading edge recessed from the inlet forward-most point an inlet length in an axial direction. A spinner has a spinner length from a spinner forward-most point to the leading edge. A ratio of the spinner length to inlet length is equal to or greater than about 0.5.

公开(公告)号：US10830129B2

公开(公告)日：2020-11-10

申请号：US15104114

申请日：2014-12-12

Applicant: United Technologies Corporation

Inventor： Wesley K. Lord

IPC: F02C3/04 ,  F02C7/36 ,  F02K3/072 ,  B64D27/14 ,  B64D29/04 ,  F02C3/107 ,  F02C3/10 ,  F02K3/02 ,  B64D27/00

Abstract: The present disclosure relates generally to an aircraft with counter-rotating pusher props powered by a gas turbine engine having a power turbine disposed substantially perpendicular to the compressor, combustor and turbine gas generator power core axis, as well as to the aircraft longitudinal axis.