公开(公告)号：US20230021948A1

公开(公告)日：2023-01-26

申请号：US17951370

申请日：2022-09-23

Applicant: RAYTHEON TECHNOLOGIES CORPORATION

Inventor： Gabriel L. Suciu ,  William K. Ackermann ,  Harold W. Hipsky

IPC: F02C9/18 ,  F02C3/04 ,  F02C7/06 ,  F02C7/32 ,  B64D13/06 ,  F02C6/08

Abstract: An environmental control system for an aircraft includes a higher pressure tap associated with a higher compression location in a main compressor section. The higher pressure tap leads into a turbine section of a turbocompressor such that air in the higher pressure tap drives the turbine section to in turn drive a compressor section of the turbocompressor. A combined outlet receives airflow from a turbine outlet and a compressor outlet intermixing airflow and passing the mixed airflow downstream to be delivered to an aircraft system. A buffer air outlet communicates airflow to an engine buffer air system.

公开(公告)号：US11512651B2

公开(公告)日：2022-11-29

申请号：US17159525

申请日：2021-01-27

Applicant: Raytheon Technologies Corporation

Inventor： Brian D. Merry ,  Gabriel L. Suciu ,  Michael G. McCaffrey

IPC: F02C9/18 ,  F02C7/32 ,  F01D17/10 ,  F01D25/12 ,  F02C7/36 ,  F04D29/28 ,  F02C7/143 ,  F02C7/18

Abstract: A gas turbine engine includes a main compressor section with a downstream most location. A turbine section has a high pressure turbine. A tap line is connected to tap air from a location upstream of the downstream most location in the main compressor section. The tapped air is connected to a heat exchanger and then to a cooling compressor. The cooling compressor compresses air downstream of the heat exchanger, and is connected to deliver air into the high pressure turbine. A bypass valve is positioned downstream of the main compressor section, and upstream of the heat exchanger. The bypass valve selectively delivers air directly to the cooling compressor without passing through the heat exchanger under certain conditions.

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：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：US20220145807A1

公开(公告)日：2022-05-12

申请号：US17579938

申请日：2022-01-20

Applicant: RAYTHEON TECHNOLOGIES CORPORATION

Inventor： Glenn Levasseur ,  Gabriel L. Suciu ,  Gary M. Stetson

IPC: F02C7/36 ,  F02C9/20 ,  F01D17/16 ,  F02C3/06 ,  F02C7/32

Abstract: A gas turbine engine includes a propulsor having a plurality of blades, a compressor section including a first compressor and a second compressor aft of the first compressor. The first compressor includes at least one array of first variable guide vanes that control operation of the first compressor. The second compressor includes at least one array of second variable guide vanes that control operation of the second compressor. A turbine section includes a first turbine and a second turbine. A geared architecture is driven by the second turbine for rotating the propulsor.

公开(公告)号：US11215197B2

公开(公告)日：2022-01-04

申请号：US16251133

申请日：2019-01-18

Applicant: Raytheon Technologies Corporation

Inventor： Gabriel L. Suciu ,  Jesse M. Chandler ,  Joseph Brent Staubach ,  Brian D. Merry

IPC: F04D29/58 ,  F02C6/08 ,  F02C3/13 ,  F02C7/18 ,  F02C9/18 ,  F01D25/12 ,  F02C7/143

Abstract: A gas turbine engine comprises a main compressor section having a high pressure compressor with a downstream most end, and more upstream locations. A turbine section has a high pressure turbine. A first tap taps air from at least one of the more upstream locations in the main compressor section, passes the tapped air through a heat exchanger and then to a cooling compressor. The cooling compressor compresses air downstream of the heat exchanger. A second tap taps air from a location closer to the downstream most end than the location(s) of the first tap. The first and second tap mix together and are delivered into the high pressure turbine. An intercooling system for a gas turbine engine is also disclosed.

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：US11149689B2

公开(公告)日：2021-10-19

申请号：US16152502

申请日：2018-10-05

Applicant: Raytheon Technologies Corporation

Inventor： Brian D. Merry ,  Gabriel L. Suciu ,  Karl L. Hasel

IPC: F02C7/36 ,  F02K3/00 ,  F01D25/16 ,  F02K3/04 ,  F02C7/06 ,  F02C3/04

Abstract: A gas turbine engine includes a housing includes an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. A rotor is connected to the hub and supports a compressor section. The geared architecture includes an epicyclic gear train. A fan is rotationally driven by the geared architecture. First and second bearings support the shaft relative to the intermediate case and the inlet case, respectively. The radially inner boundary of the core inlet is at a location of a core inlet stator and the radially inner boundary of the compressor section inlet is at a location of the first stage low-pressure compressor rotor.

公开(公告)号：US11047337B2

公开(公告)日：2021-06-29

申请号：US15485512

申请日：2017-04-12

Applicant: RAYTHEON TECHNOLOGIES CORPORATION

Inventor： Michael E. McCune ,  Jason Husband ,  Frederick M. Schwarz ,  Daniel Bernard Kupratis ,  Gabriel L. Suciu ,  William K. Ackermann

IPC: F02K3/06 ,  F02C3/107 ,  F02C7/36 ,  F02C7/20 ,  F01D15/12 ,  F01D5/06 ,  F04D19/00 ,  F04D25/04 ,  F04D29/053 ,  F04D29/32 ,  F01D9/04

Abstract: A gas turbine engine includes a gear system that provides a speed reduction between a fan drive turbine and a fan rotor. Aspects of the gear system are provided with defined flexibility. The fan drive turbine has a first exit area and rotates at a first speed. A second turbine section has a second exit area and rotates at a second speed, which is faster than said first speed. A performance quantity can be defined for both turbine sections as the products of the respective areas and respective speeds squared. A performance quantity ratio of the performance quantity for the fan drive turbine to the performance quantity for the second turbine section is between 0.5 and 1.5.

公开(公告)号：US20210148289A1

公开(公告)日：2021-05-20

申请号：US17159525

申请日：2021-01-27

Applicant: Raytheon Technologies Corporation

Inventor： Brian D. Merry ,  Gabriel L. Suciu ,  Michael G. McCaffrey

IPC: F02C9/18 ,  F02C7/143 ,  F02C7/32 ,  F02C7/18 ,  F01D17/10 ,  F01D25/12 ,  F02C7/36 ,  F04D29/28

Abstract: A gas turbine engine includes a main compressor section with a downstream most location. A turbine section has a high pressure turbine. A tap line is connected to tap air from a location upstream of the downstream most location in the main compressor section. The tapped air is connected to a heat exchanger and then to a cooling compressor. The cooling compressor compresses air downstream of the heat exchanger, and is connected to deliver air into the high pressure turbine. A bypass valve is positioned downstream of the main compressor section, and upstream of the heat exchanger. The bypass valve selectively delivers air directly to the cooling compressor without passing through the heat exchanger under certain conditions.