公开(公告)号：US20230042974A1

公开(公告)日：2023-02-09

申请号：US17969179

申请日：2022-10-19

Applicant: RAYTHEON TECHNOLOGIES CORPORATION

Inventor： Frederick M. Schwarz ,  William G. Sheridan

IPC: F01D25/24 ,  F02K3/06 ,  F16H1/28 ,  F02C3/107 ,  F01D5/02 ,  F01D5/12 ,  F01D17/10 ,  F04D29/32

Abstract: A gas turbine engine includes, among other things, a propulsor section including a propulsor hub, the hub including a hub diameter supporting a plurality of propulsor blades. A compressor section includes a first compressor and a second compressor. A turbine section includes a first turbine and a second turbine. A geared architecture interconnects the first turbine and the propulsor hub. The geared architecture includes a gear volume. A compressor inlet passage is disposed annularly about the geared architecture.

公开(公告)号：US11486311B2

公开(公告)日：2022-11-01

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：US11377221B2

公开(公告)日：2022-07-05

申请号：US17028578

申请日：2020-09-22

Applicant: RAYTHEON TECHNOLOGIES CORPORATION

Inventor： Nigel David Sawyers-Abbott ,  Frederick M. Schwarz

IPC: B64D27/26 ,  F02C7/20 ,  F01D25/28 ,  F02K3/06

Abstract: A turbofan engine may comprise an inlet and a fan case coupled to the inlet. An engine case may be coupled to the fan case via a vane extending between the fan case and the engine case. A strut apparatus may extend from the fan case and limit deflection of the fan case. The strut apparatus may comprise a first end proximate the fan case, and a second end coupled to at least one of the engine case or a structure for mounting the turbofan engine to an aircraft.

公开(公告)号：US11346289B2

公开(公告)日：2022-05-31

申请号：US16881567

申请日：2020-05-22

Applicant: Raytheon Technologies Corporation

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

IPC: F02C9/18 ,  F02C3/04 ,  F02C7/20 ,  F01D1/02 ,  F01D5/02 ,  F01D9/04 ,  F01D25/24 ,  F02C7/36 ,  F02C3/107 ,  F02K3/06 ,  F02C3/113 ,  F02K3/02 ,  F02K3/075 ,  F02K7/06

Abstract: A turbofan engine has an engine case and a gaspath through the engine case. A fan has a circumferential array of fan blades. The engine further has a compressor, a combustor, a gas generating turbine, and a low pressure turbine section. A speed reduction mechanism couples the low pressure turbine section to the fan. A bypass area ratio is greater than about 6.0. The low pressure turbine section airfoil count to bypass area ratio is below about 170.

公开(公告)号：US11280267B2

公开(公告)日：2022-03-22

申请号：US17007166

申请日：2020-08-31

Applicant: Raytheon Technologies Corporation

Inventor： Frederick M. Schwarz ,  William G. Sheridan

IPC: F02C7/36 ,  F02C7/06 ,  F02K1/09 ,  F02C3/107 ,  F02K3/04 ,  F01D5/02 ,  F01D9/06 ,  F01D17/10 ,  F02C3/06 ,  F02C3/10 ,  F04D29/32 ,  F01D5/06 ,  F02C3/04 ,  F02C7/28 ,  F02C9/18 ,  F02K1/06 ,  F02K3/06 ,  F04D29/38 ,  F16H1/28

Abstract: A gas turbine engine according to the present disclosure includes, among other things, a fan section having a plurality of fan blades, the plurality of fan blades having a peak tip radius Rt and an inboard leading edge radius Rh at a first inboard boundary of a first flowpath, and a core engine including a first turbine configured to drive a first compressor, and a fan drive turbine configured to drive the fan section.

公开(公告)号：US20220082065A1

公开(公告)日：2022-03-17

申请号：US17538048

申请日：2021-11-30

Applicant: RAYTHEON TECHNOLOGIES CORPORATION

Inventor： Daniel Bernard Kupratis ,  Frederick M. Schwarz

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

Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a propulsor section, a compressor section including a low pressure compressor and a second compressor section, and a turbine section including a low pressure turbine and a high pressure turbine. The low pressure turbine drives the low pressure compressor and the gear arrangement to drive the propulsor. A core split power ratio is provided by power input to the high pressure compressor divided by a power input to the low pressure compressor measured in horsepower.

公开(公告)号：US20220056847A1

公开(公告)日：2022-02-24

申请号：US17519887

申请日：2021-11-05

Applicant: Raytheon Technologies Corporation

Inventor： James D. Hill ,  Frederick M. Schwarz

IPC: F02C7/18 ,  F02C3/107 ,  F01D25/12 ,  F02C6/08

Abstract: A gas turbine engine comprises a compressor section and a turbine section, the compressor section having a last compressor stage. High pressure cooling air is tapped from a location downstream of the last compressor stage and passed through a heat exchanger. Lower pressure air passes across the heat exchanger to cool the high pressure cooling air. A housing surrounds the compressor section and the turbine section and there being a space radially outwardly of the housing, and a mixing chamber received in the space radially outwardly of the housing, the mixing chamber receiving the high pressure cooling air downstream of the heat exchanger, and further receiving air at a temperature higher than a temperature of the high pressure cooling air downstream of the heat exchanger. Mixed air from the mixing chamber is returned into the housing and utilized to cool at least the turbine section.

公开(公告)号：US11215143B2

公开(公告)日：2022-01-04

申请号：US16684968

申请日：2019-11-15

Applicant: RAYTHEON TECHNOLOGIES CORPORATION

Inventor： Daniel Bernard Kupratis ,  Frederick M. Schwarz

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

Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section, and a compressor section including a low pressure compressor and a second compressor section, and a turbine section including a fan drive turbine and a high pressure turbine. The fan drive turbine drives the low pressure compressor and a gear arrangement to drive the fan section. A core split power ratio is provided by power input to the high pressure compressor divided by a power input to the low pressure compressor measured in horsepower.

公开(公告)号：US11215123B2

公开(公告)日：2022-01-04

申请号：US17062756

申请日：2020-10-05

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 ,  F02K3/06 ,  F02K3/075 ,  F01D5/06 ,  F01D25/24 ,  F02C3/04 ,  F02C7/20 ,  F04D19/02 ,  F01D11/12

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 low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area, and a fan duct annulus area outboard of the low pressure compressor section inlet, and a fan drive turbine section. The fan drive 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 equal to or greater than 0.35, and is less than 0.55.

公开(公告)号：US20210339880A1

公开(公告)日：2021-11-04

申请号：US17376675

申请日：2021-07-15

Applicant: Raytheon Technologies Corporation

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

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

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.