公开(公告)号：US11149650B2

公开(公告)日：2021-10-19

申请号：US16025038

申请日：2018-07-02

Applicant: RAYTHEON TECHNOLOGIES CORPORATION

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

IPC: F02C7/36 ,  F02C3/107 ,  F02C9/18 ,  F01D11/12 ,  F01D5/06 ,  F01D25/24 ,  F02C3/04 ,  F02C7/20 ,  F04D19/02 ,  F02K3/075 ,  F02K3/06 ,  F04D29/32 ,  F04D29/54

Abstract: A turbofan engine according to an example of the present disclosure includes, among other things, a fan including a circumferential array of fan blades, a compressor in fluid communication with the fan, the compressor including a first compressor section and a second compressor, the second compressor section including a second compressor section inlet with a second compressor section inlet annulus area, a fan duct including a fan duct annulus area outboard of the second compressor section inlet, a shaft assembly having a first portion and a second portion, a turbine in fluid communication with the combustor, the turbine having a first turbine section coupled to the first portion of the shaft assembly to drive the first compressor section, and a second turbine section coupled to the second portion of the shaft assembly to drive the fan, an epicyclic transmission coupled to the fan and rotatable by the second turbine section through the second portion of the shaft assembly to allow the second turbine to turn faster than the fan, wherein the second turbine section includes a maximum gas path radius and the fan blades include a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the fan blades is less than 0.6.

公开(公告)号：US20210301731A1

公开(公告)日：2021-09-30

申请号：US17230326

申请日：2021-04-14

Applicant: RAYTHEON TECHNOLOGIES CORPORATION

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

IPC: F02C7/36 ,  F02C3/107 ,  F02C9/18 ,  F01D11/12 ,  F01D5/06 ,  F01D25/24 ,  F02C3/04 ,  F02C7/20 ,  F04D19/02 ,  F02K3/075 ,  F02K3/06 ,  F04D29/32

Abstract: A turbofan engine according to an example of the present disclosure includes, among other things, a fan including an array of fan blades rotatable about an engine axis, a compressor including a high pressure compressor section and a low pressure compressor section, the low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor inlet annulus area, a fan duct including a fan duct annulus area outboard of the low pressure compressor section inlet, and a turbine having a high pressure turbine section and a low pressure turbine section driving the fan through a speed reduction mechanism, wherein the low pressure turbine section defines a maximum gas path radius and the fan blades define a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the fan blades is less than 0.6.

公开(公告)号：US11091272B2

公开(公告)日：2021-08-17

申请号：US16039849

申请日：2018-07-19

Applicant: Raytheon Technologies Corporation

Inventor： Daniel Bernard Kupratis ,  Frederick M. Schwarz ,  William G. Sheridan ,  Glenn Levasseur

IPC: F02C6/14 ,  B64D27/24 ,  B64D27/10 ,  F01D15/10 ,  F02C7/36 ,  H02K7/18 ,  B64D27/02 ,  F02C6/02 ,  F02C3/107

Abstract: An aircraft propulsion system is disclosed and includes a first gas turbine engine including a first input shaft driving a first gear system, a first fan driven by the first gear system, a first generator supported on the first input shaft and a fan drive electric motor providing a drive input to the first fan, a second gas turbine engine including a second input shaft driving a second gear system, a second fan driven by the second gear system, a second generator supported on the second input shaft and a second fan drive electric motor providing a drive input to the second fan and a controller controlling power output from each of the first and second generators and directing the power output between each of the first and second fan drive electric motors.

公开(公告)号：US20210071588A1

公开(公告)日：2021-03-11

申请号：US17086715

申请日：2020-11-02

Applicant: Raytheon Technologies Corporation

Inventor： Frederick M. Schwarz ,  Karl L. Hasel

IPC: F02C7/36 ,  F02K3/06 ,  F02C3/06

Abstract: A gas turbine engine comprises a fan includes a plurality of fan blades rotatable about an axis. A compressor section includes at least a first compressor section and a second compressor section, wherein components of the second compressor section are configured to operate at an average exit temperature that is between about 1000° F. and about 1500° F. A combustor is in fluid communication with the compressor section. A turbine section is in fluid communication with the combustor. A geared architecture is driven by the turbine section for rotating the fan about the axis.

公开(公告)号：US20210017936A1

公开(公告)日：2021-01-21

申请号：US17062197

申请日：2020-10-02

Applicant: Raytheon Technologies Corporation

Inventor： Michael E. McCune ,  Lawrence E. Portlock ,  Frederick M. Schwarz

IPC: F02K3/06 ,  F02C7/36 ,  F16H57/04

Abstract: A gas turbine engine includes a bypass ratio greater than about ten (10). A fan is supported on a fan shaft and has a plurality of fan blades. A gear system is connected to the fan shaft. A plurality of planetary gears and a first set of opposed angled ring gear teeth are separated from a second set of opposed angled ring gear teeth. A lubricant flow path is located axially between the first set of opposed angled ring gear teeth and the second set of opposed angled ring gear teeth. A gear system support is relative to a fixed housing facilitating segregation of vibrations. An annular channel is axially aligned with the lubricant flow path. A low pressure turbine has an inlet, an outlet, and a low pressure turbine pressure ratio greater than 5:1.

公开(公告)号：US20210010426A1

公开(公告)日：2021-01-14

申请号：US16919620

申请日：2020-07-02

Applicant: Raytheon Technologies Corporation

Inventor： Frederick M. Schwarz ,  William G. Sheridan

IPC: F02C7/36 ,  F02K3/06 ,  F02C3/107 ,  F02C3/04

Abstract: A gas turbine engine comprises a fan rotor having a hub and a plurality of fan blades extending radially outwardly of the hub. A compressor is positioned downstream of the fan rotor, and has a first compressor blade row defined along a rotational axis of the fan rotor and the compressor rotor. A gear reduction is positioned axially between the first compressor blade row and the fan rotor, and includes a ring gear and a carrier. The carrier has an axial length and the ring gear has an outer diameter. A ratio of the axial length to the outer diameter may be greater than or equal to about 0.20 and less than or equal to about 0.40. The gear reduction is connected to drive the hub to rotate. A method of designing a gas turbine engine is also disclosed.

公开(公告)号：US10830153B2

公开(公告)日：2020-11-10

申请号：US16186811

申请日：2018-11-12

Applicant: RAYTHEON TECHNOLOGIES CORPORATION

Inventor： Frederick M. Schwarz ,  Daniel Bernard Kupratis

IPC: F02C7/36 ,  F02C3/107 ,  F02K3/06 ,  F01D1/26 ,  F01D15/12

Abstract: A gas turbine engine turbine has a compressor section including a low pressure compressor and a high pressure compressor, a turbine section including a low pressure turbine and a high pressure turbine, the high pressure turbine rotatable with the high pressure compressor as a high pressure spool in a first direction about an engine longitudinal axis, the low pressure turbine rotatable as a low pressure spool in the first direction about the engine longitudinal axis, a power density of greater than or equal to 4.75 and less than or equal to 5.5 lbf/in3, a fan section including a fan having a plurality of fan blades, the fan connected to the low pressure spool via a geared architecture such that the fan is rotatable about the engine longitudinal axis in a second direction opposed to the first direction, and a bypass ratio greater than 6.

公开(公告)号：US10823052B2

公开(公告)日：2020-11-03

申请号：US16531704

申请日：2019-08-05

Applicant: Raytheon Technologies Corporation

Inventor： Frederick M. Schwarz

IPC: F02C3/107 ,  F01D5/14 ,  F02K3/06

Abstract: A turbofan engine includes a fan section including a fan blade having a leading edge and hub to tip ratio of less than about 0.34 and greater than about 0.020 measured at the leading edge and a speed change mechanism with gear ratio greater than about 2.6 to 1. A first compression section includes a last blade trailing edge radial tip length that is greater than about 67% of the radial tip length of a leading edge of a first stage of the first compression section. A second compression section includes a last blade trailing edge radial tip length that is greater than about 57% of a radial tip length of a leading edge of a first stage of the first compression section.

公开(公告)号：US10823051B2

公开(公告)日：2020-11-03

申请号：US16157224

申请日：2018-10-11

Applicant: Raytheon Technologies Corporation

Inventor： Frederick M. Schwarz ,  Karl L. Hasel

IPC: F02C3/06 ,  F02C7/36 ,  F02K3/06

Abstract: A gas turbine engine comprises a fan includes a plurality of fan blades rotatable about an axis. A compressor section includes at least a first compressor section and a second compressor section, wherein components of the second compressor section are configured to operate at an average exit temperature that is between about 1000° F. and about 1500° F. A combustor is in fluid communication with the compressor section. A turbine section is in fluid communication with the combustor. A geared architecture is driven by the turbine section for rotating the fan about the axis.

公开(公告)号：US20230303254A1

公开(公告)日：2023-09-28

申请号：US18082848

申请日：2022-12-16

Applicant: Raytheon Technologies Corporation

Inventor： Frederick M. Schwarz ,  Michael Winter ,  Charles E. Lents ,  Nathan Snape ,  Alan H. Epstein

IPC: B64D13/08 ,  F02C7/141 ,  F02C7/16 ,  F02C9/18 ,  F25B9/14 ,  F25J3/04 ,  B01D45/08 ,  B01D53/00 ,  B03C1/28 ,  B64D27/12 ,  B64D33/08 ,  F01D15/00 ,  F01D25/12 ,  F02C6/08 ,  F02C7/143 ,  F02C7/18 ,  F02K3/06 ,  B64D13/06 ,  F17C7/02 ,  F17C7/04

CPC classification number: B64D13/08 ,  F02C7/141 ,  F02C7/16 ,  F02C9/18 ,  F25B9/14 ,  F25J3/04254 ,  B01D45/08 ,  B01D53/002 ,  B03C1/288 ,  B64D27/12 ,  B64D33/08 ,  F01D15/00 ,  F01D25/12 ,  F02C6/08 ,  F02C7/143 ,  F02C7/185 ,  F02K3/06 ,  F25J3/04975 ,  B64D13/06 ,  F17C7/02 ,  F17C7/04 ,  B64D37/32

Abstract: A gas turbine engine includes a compressor section and a turbine section operably coupled to the compressor section. The gas turbine engine further includes a means for selectively releasing a cooling fluid flow produced at a cryogenic temperature and a plumbing system in fluid communication with the means for selectively releasing the cooling fluid flow. The plumbing system is configured to route the cooling fluid flow to one or more of the compressor section and the turbine section.