公开(公告)号：US11754021B2

公开(公告)日：2023-09-12

申请号：US17407880

申请日：2021-08-20

Applicant: Raytheon Technologies Corporation

Inventor： Amanda Jean Learned Boucher ,  Joseph B. Staubach

IPC: F02K3/115 ,  F02K3/06

CPC classification number: F02K3/115 ,  F02K3/06 ,  F05D2220/323 ,  F05D2240/35 ,  F05D2260/213 ,  F05D2260/232 ,  F05D2260/606

Abstract: Aircraft propulsion systems including a closed loop-supercritical fluid system having a turbine, a cooler heat exchanger, a compressor, and a recovery heat exchanger arranged along a closed-loop flow path of a supercritical fluid. A shaft is operably coupled to the turbine and configured to be rotationally driven by the turbine. A fan is configured to generate thrust, the fan operably coupled to the shaft to be rotationally driven by the shaft. A burner is configured to combust a fuel and air from the fan to generate a combusted gas and supply said combusted gas to the recovery heat exchanger of the closed loop-supercritical fluid system and out an exhaust nozzle.

公开(公告)号：US11753993B1

公开(公告)日：2023-09-12

申请号：US17669632

申请日：2022-02-11

Applicant: Raytheon Technologies Corporation

Inventor： Neil J. Terwilliger ,  Joseph E. Turney ,  Joseph B. Staubach

IPC: F02C7/14 ,  F02C6/00 ,  F02C7/224 ,  F02C7/141 ,  F02C3/22 ,  F02C3/30 ,  F02C7/18

CPC classification number: F02C7/14 ,  F02C3/22 ,  F02C3/30 ,  F02C6/006 ,  F02C7/141 ,  F02C7/18 ,  F02C7/224 ,  F05D2220/323 ,  F05D2240/35 ,  F05D2260/207 ,  F05D2260/213 ,  F05D2260/232

Abstract: Aircraft engines include a turbine engine comprising a compressor section, a burner section, and a turbine section arranged along a shaft, with a core flow path through the turbine engine such that exhaust from the burner section passes through the turbine section, a condensing assembly arranged downstream of the turbine section of the turbine engine along the core flow path, and an exhaust compressor arranged downstream of the condensing assembly along the core flow path. The condensing assembly is configured to reduce a mass flow of the exhaust compressor by condensing water vapor from the core flow and removing liquid water from the core flow.

公开(公告)号：US20230258126A1

公开(公告)日：2023-08-17

申请号：US17670080

申请日：2022-02-11

Applicant: Raytheon Technologies Corporation

Inventor： Neil J. Terwilliger ,  Walter A. Ledwith, JR. ,  David L. Ma ,  Joseph B. Staubach

IPC: F02C6/04 ,  F02C6/20 ,  F02C9/40 ,  B64D27/10 ,  F02C7/22

CPC classification number: F02C6/04 ,  F02C6/20 ,  F02C9/40 ,  B64D27/10 ,  F02C7/22

Abstract: A powerplant is provided for an aircraft. This powerplant includes an engine and an energy recovery system. The engine includes an engine combustor, an engine turbine, a flowpath and a fluid delivery system. The flowpath extends out of the engine combustor and through the engine turbine. The fluid delivery system is configured to provide fluid hydrogen and fluid oxygen for combustion within the engine combustor to produce combustion products within the flowpath. The energy recovery system includes an energy recovery system condenser, an energy recovery system pump, an energy recovery system evaporator and an energy recovery system turbine. The energy recovery system pump is configured to pump liquid water from the energy recovery system condenser to the energy recovery system evaporator. The energy recovery system evaporator is configured to transfer heat from the combustion products into the liquid water to evaporate at least some of the liquid water into water vapor to drive the energy recovery system turbine.

公开(公告)号：US11725544B1

公开(公告)日：2023-08-15

申请号：US17670043

申请日：2022-02-11

Applicant: Raytheon Technologies Corporation

Inventor： Neil J. Terwilliger ,  Walter A. Ledwith, Jr. ,  Joseph B. Staubach ,  David L. Ma

IPC: F01K25/00 ,  F01K27/02 ,  F01K15/02 ,  F01K7/16 ,  F01K7/12

CPC classification number: F01K25/005 ,  F01K7/12 ,  F01K7/16 ,  F01K15/02 ,  F01K27/02

Abstract: A powerplant is provided that includes an engine and a water recovery system. The engine includes an engine combustor, an engine turbine, a flowpath and a fluid delivery system. The flowpath extends out of the engine combustor and through the engine turbine. The fluid delivery system includes a hydrogen reservoir and an oxygen reservoir. The hydrogen reservoir is configured to store fluid hydrogen as liquid hydrogen. The oxygen reservoir is configured to store fluid oxygen as liquid oxygen. The fluid delivery system is configured to provide the fluid hydrogen and the fluid oxygen for combustion within the engine combustor to produce combustion products within the flowpath. The water recovery system is configured to extract water from the combustion products. The water recovery system is configured to provide the water to a component of the powerplant.

公开(公告)号：US20230159175A1

公开(公告)日：2023-05-25

申请号：US17535200

申请日：2021-11-24

Applicant: Raytheon Technologies Corporation

Inventor： Jon E. Sobanski ,  Paul R. Hanrahan ,  Joseph B. Staubach ,  Jill Klinowski

IPC: B64D27/24 ,  B64D27/26 ,  B64D29/02 ,  B64C39/10 ,  B64C11/28

CPC classification number: B64D27/24 ,  B64D27/26 ,  B64D29/02 ,  B64C39/10 ,  B64C11/28 ,  B64D2027/262

Abstract: A propulsor includes a propulsor body and a prop assembly in rotational communication with the propulsor body. The prop assembly includes a plurality of prop blades configured for rotation about an axial centerline of the propulsor. The plurality of prop blades are rotatable between a deployed position and a stowed position. The propulsor further includes at least one linkage having a first linkage end and a second linkage end. The first linkage end of the at least one linkage is rotatably mounted to the propulsor body and the second linkage end is configured to be rotatably mounted to an aircraft body. The propulsor further includes a first motor coupled to the at least one linkage and configured to rotate the at least one linkage relative to the propulsor body between a first rotational position and a second rotational position.

公开(公告)号：US11614036B2

公开(公告)日：2023-03-28

申请号：US17530544

申请日：2021-11-19

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 gas turbine engine according to an example of the present disclosure includes, among other things, a propulsor including a circumferential array of blades, a low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area and a low pressure turbine section. The low pressure turbine section includes a maximum gas path radius, the blades include a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the blades is equal to or greater than 0.35, and is less than 0.55.

公开(公告)号：US20230056536A1

公开(公告)日：2023-02-23

申请号：US17407880

申请日：2021-08-20

Applicant: Raytheon Technologies Corporation

Inventor： Amanda Jean Learned Boucher ,  Joseph B. Staubach

IPC: F02K3/115 ,  F02K3/06

Abstract: Aircraft propulsion systems including a closed loop-supercritical fluid system having a turbine, a cooler heat exchanger, a compressor, and a recovery heat exchanger arranged along a closed-loop flow path of a supercritical fluid. A shaft is operably coupled to the turbine and configured to be rotationally driven by the turbine. A fan is configured to generate thrust, the fan operably coupled to the shaft to be rotationally driven by the shaft. A burner is configured to combust a fuel and air from the fan to generate a combusted gas and supply said combusted gas to the recovery heat exchanger of the closed loop-supercritical fluid system and out an exhaust nozzle.

公开(公告)号：US20230010158A1

公开(公告)日：2023-01-12

申请号：US17860742

申请日：2022-07-08

Applicant: Raytheon Technologies Corporation

Inventor： Marc J. Muldoon ,  Joseph B. Staubach ,  Charles E. Lents

IPC: F02C9/40 ,  B64D37/30 ,  B64D37/34 ,  F02C3/22

Abstract: A turbine engine system includes aircraft systems including at least one hydrogen fuel tank, engine systems comprising a compressor section, a combustor section having a burner, and a turbine section, and a hydrogen fuel flow supply line configured to supply hydrogen fuel from the at least one hydrogen fuel tank into the burner for combustion. The turbine engine system has a bypass ratio between 5 to 20.

公开(公告)号：US20220364513A1

公开(公告)日：2022-11-17

申请号：US17321052

申请日：2021-05-14

Applicant: Raytheon Technologies Corporation

Inventor： Marc J. Muldoon ,  Joseph B. Staubach ,  Charles E. Lents

IPC: F02C7/224

Abstract: An aircraft propulsion systems and aircraft having the same are described. The aircraft propulsion systems have one or more aircraft systems including at least one hydrogen tank and a first heat exchanger and one or more engine systems including at least a main engine core, a second heat exchanger, and a third heat exchanger. The main engine core comprises a compressor section, a combustor section having a burner, and a turbine section. Hydrogen is configured to be supplied from the at least one hydrogen tank through a hydrogen flow path, passing through the first heat exchanger of the aircraft systems, the second heat exchanger of the engine systems, and the third heat exchanger of the engine systems, and then supplied into the burner for combustion.

公开(公告)号：US11480108B2

公开(公告)日：2022-10-25

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.