公开(公告)号：US10982624B2

公开(公告)日：2021-04-20

申请号：US15943100

申请日：2018-04-02

Applicant: RAYTHEON TECHNOLOGIES CORPORATION

Inventor： Gabriel L. Suciu ,  Brian Merry ,  Christopher M. Dye

IPC: F02C7/20 ,  F02K3/06 ,  F02C7/36

Abstract: A turbine engine is disclosed that includes a fan case surrounding a fan rotatable about an axis. A core is supported relative to the fan case by support structure arranged downstream from the fan. The core includes a core housing having an inlet case arranged to receive airflow from the fan. A compressor case is arranged axially adjacent to the inlet case and surrounds a compressor stage having a rotor blade with a blade trailing edge. The support structure includes a support structure leading edge facing the fan and a support structure trailing edge on a side opposite the support structure leading edge. The support structure trailing edge is arranged axially forward of the blade trailing edge. In one example, a forward attachment extends from the support structure to the inlet case.

公开(公告)号：US20210071587A1

公开(公告)日：2021-03-11

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：US10815888B2

公开(公告)日：2020-10-27

申请号：US15943825

申请日：2018-04-03

Applicant: Raytheon Technologies Corporation

Inventor： Brian D. Merry ,  Gabriel L. Suciu

IPC: F01D25/16 ,  F02C7/06

Abstract: A geared turbofan gas turbine engine includes a fan section and a core section. The core section includes a compressor section, a combustor section and a turbine section. The fan section includes a gearbox and a fan. A low spool includes a low turbine within the turbine section and a forward connection to a gearbox for driving the fan. The low spool is supported for rotation about the axis at a forward most position by a forward roller bearing and at an aft position by a thrust bearing.

公开(公告)号：US20230323836A1

公开(公告)日：2023-10-12

申请号：US17716523

申请日：2022-04-08

Applicant: Raytheon Technologies Corporation

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

IPC: F02K3/06 ,  F02C7/36

CPC classification number: F02K3/06 ,  F02C7/36 ,  F01D5/02

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：US11731773B2

公开(公告)日：2023-08-22

申请号：US17395553

申请日：2021-08-06

Applicant: RAYTHEON TECHNOLOGIES CORPORATION

Inventor： Gabriel L. Suciu ,  Brian D. Merry ,  Christopher M. Dye ,  Steven B. Johnson ,  Frederick M. Schwarz

IPC: B64D27/26 ,  F01D25/28 ,  F02C7/20 ,  F02K3/06 ,  F02C9/20 ,  F02C9/18 ,  F02C7/36 ,  F02C3/107 ,  F01D5/06 ,  F01D9/02 ,  F01D15/12 ,  F01D25/24 ,  F02K1/15

CPC classification number: B64D27/26 ,  F01D5/06 ,  F01D9/02 ,  F01D15/12 ,  F01D25/24 ,  F01D25/28 ,  F02C3/107 ,  F02C7/20 ,  F02C7/36 ,  F02C9/18 ,  F02C9/20 ,  F02K1/15 ,  F02K3/06 ,  B64D2027/266 ,  B64D2027/268 ,  F05D2220/32 ,  F05D2220/36 ,  F05D2260/4031 ,  F05D2260/40311 ,  F05D2270/42

Abstract: A gas turbine engine includes, among other things, a propulsor section including a rotor, a gear train, a low spool and a high spool. A static structure includes a first case and a second case. A mount system includes a forward mount and an aft mount arranged in a plane containing an engine axis of rotation. The forward mount is secured to the first case. The aft mount is secured to the second case.

公开(公告)号：US20220403774A1

公开(公告)日：2022-12-22

申请号：US17716290

申请日：2022-04-08

Applicant: Raytheon Technologies Corporation

Inventor： Jesse M. Chandler ,  Gabriel L. Suciu

IPC: F02C3/10 ,  F02K3/062 ,  B64D27/10 ,  B64D33/02 ,  B64D33/04 ,  F01D25/30 ,  F02C7/36 ,  F02K3/12

Abstract: A gas turbine engine for an aircraft includes a core engine assembly including a compressor section communicating air to a combustor section where the air is mixed with fuel and ignited to generate a high-energy gas flow that is expanded through a turbine section. The turbine section is coupled to drive the compressor section. A free turbine is configured to be driven by gas flow from the core engine. A propulsor section aft of the core engine and is driven by the free turbine. An exhaust duct routes exhaust gases from the core engine to the free turbine. The free turbine is disposed aft of the propulsor section and the exhaust duct includes an outlet aft of the propulsor section communicating gas flow to drive the free turbine. An aircraft is also disclosed.

公开(公告)号：US11486314B2

公开(公告)日：2022-11-01

申请号：US17020034

申请日：2020-09-14

Applicant: Raytheon Technologies Corporation

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

IPC: F02C9/00 ,  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：US11408372B2

公开(公告)日：2022-08-09

申请号：US17217831

申请日：2021-03-30

Applicant: Raytheon Technologies Corporation

Inventor： Gabriel L. Suciu ,  Brian Merry ,  Christopher M. Dye

IPC: F02C7/20 ,  F02K3/06 ,  F02C7/36

Abstract: A turbine engine is disclosed that includes a fan case surrounding a fan rotatable about an axis. A core is supported relative to the fan case by support structure arranged downstream from the fan. The core includes a core housing having an inlet case arranged to receive airflow from the fan. A compressor case is arranged axially adjacent to the inlet case and surrounds a compressor stage having a rotor blade with a blade trailing edge. The support structure includes a support structure leading edge facing the fan and a support structure trailing edge on a side opposite the support structure leading edge. The support structure trailing edge is arranged axially forward of the blade trailing edge. In one example, a forward attachment extends from the support structure to the inlet case.

公开(公告)号：US20210380260A1

公开(公告)日：2021-12-09

申请号：US17411529

申请日：2021-08-25

Applicant: Raytheon Technologies Corporation

Inventor： Gabriel L. Suciu ,  Brian Merry ,  Stephen H. Taylor ,  Charles E. Lents

IPC: B64D13/08 ,  F02C6/08 ,  F02C9/18 ,  F02K3/115 ,  B64D27/10 ,  F01D15/10 ,  F02C3/04 ,  F02K3/06 ,  H02K7/18 ,  F02C7/18 ,  F02K3/04

Abstract: An engine driven environmental control system (ECS) air circuit includes a gas turbine engine having a compressor section. The compressor section includes a plurality of compressor bleeds. A selection valve selectively connects each of said bleeds to an input of an intercooler. A second valve is configured to selectively connect an output of said intercooler to at least one auxiliary compressor. The output of each of the at least one auxiliary compressors is connected to an ECS air input.

公开(公告)号：US20210301730A1

公开(公告)日：2021-09-30

申请号：US17230271

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 ,  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 low pressure compressor section and a high pressure compressor section, the low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area, a fan duct including a fan duct annulus area outboard of the a low pressure compressor section inlet, a turbine in fluid communication with the combustor, the turbine having a high pressure turbine section and a low pressure turbine that drives the fan, a speed reduction mechanism coupled to the fan and rotatable by the low pressure turbine section to allow the low pressure turbine section to turn faster than the fan, wherein the low pressure 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 between 0.50 and 0.55, or is greater than 0.55 and less than or equal to 0.65.