公开(公告)号：US20230011956A1

公开(公告)日：2023-01-12

申请号：US17860215

申请日：2022-07-08

Applicant: Raytheon Technologies Corporation

Inventor： James Wiedenhoefer ,  Joseph B. Staubach ,  Marc J. Muldoon ,  Charles E. Lents ,  Brian M. Holley

IPC: F02C7/224 ,  F02C7/22 ,  F02C3/22

Abstract: A turbine engine system includes at least one hydrogen fuel tank, a core flow path heat exchanger in a core flow path; and engine systems located in the core flow path. The engine system including at least a compressor section, a combustor section having a burner, and a turbine section. The core flow path heat exchanger is arranged in the core flow path downstream of the combustor section. The hydrogen fuel is supplied from the at least one hydrogen fuel tank through a hydrogen fuel supply line, passing through the core flow path heat exchanger and then supplied into the burner for combustion.

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：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：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：US12180893B2

公开(公告)日：2024-12-31

申请号：US18329044

申请日：2023-06-05

Applicant: Raytheon Technologies Corporation

Inventor： Neil J. Terwilliger ,  Joseph B. Staubach

IPC: F02C7/141 ,  F02C3/22 ,  F02C3/30

Abstract: A propulsion system for an aircraft includes a core engine that includes a core flow path where air is compressed in a compressor section, communicated to a combustor section, mixed with a gaseous fuel and ignited to generate an exhaust gas flow that is expanded through a turbine section. A fuel system supplies a fuel to the combustor through a fuel flow path, a first heat exchanger thermally communicates a first heat load into a cooling flow, a turboexpander where a heated cooling flow from the first heat exchanger is expanded to generate shaft power and cooled to provide a cooled cooling flow, and a second heat exchanger thermally communicates a second heat load to the cooled cooling flow that is communicated from the turboexpander, cooling flow from the second heat exchanger is communicated to the combustor section.

公开(公告)号：US12129774B2

公开(公告)日：2024-10-29

申请号：US18316505

申请日：2023-05-12

Applicant: RAYTHEON TECHNOLOGIES CORPORATION

Inventor： Joseph B. Staubach

IPC: F01K15/02 ,  B01D45/16 ,  B01D53/26 ,  F02C7/16

CPC classification number: F01K15/02 ,  B01D45/16 ,  B01D53/265 ,  F02C7/16 ,  B01D2257/80 ,  B01D2259/4575 ,  F05D2260/211

Abstract: A propulsion system for an aircraft includes a gas generating core engine that generates an exhaust gas flow that is expanded through a turbine section. A power turbine engine is forward of the core engine and is coupled to drive a propulsor. A hydrogen fuel system supplies hydrogen fuel to the combustor through a fuel flow path. A condenser extracts water from the exhaust gas flow. A water separator is in communication with the condenser and directs the extracted water to a water storage tank. An evaporator receives a portion of the water that is extracted by the condenser and generates a steam flow. The steam flow is injected into the core flow path upstream of the turbine section.

公开(公告)号：US12098645B2

公开(公告)日：2024-09-24

申请号：US18314958

申请日：2023-05-10

Applicant: RAYTHEON TECHNOLOGIES CORPORATION

Inventor： Neil J. Terwilliger ,  Joseph B. Staubach

IPC: F01D15/02 ,  F01D15/10 ,  F01K3/26 ,  F01K7/16 ,  F01K7/38

CPC classification number: F01D15/02 ,  F01D15/10 ,  F01K3/26 ,  F01K7/16 ,  F01K7/38

Abstract: A turbine engine assembly includes a core engine generating a high energy gas flow that is expanded through a turbine section, a hydrogen fuel system supplying hydrogen fuel to a combustor through a fuel flow path, a condenser extracting water from the high energy gas flow, an evaporator inputting thermal energy into the water extracted by the condenser to generate a steam flow, and at least one superheater receiving the steam flow from the evaporator and input thermal energy for heating the steam flow. The steam flow from the at least one superheater is injected into the core flow path upstream of the turbine section.

公开(公告)号：US20240254913A1

公开(公告)日：2024-08-01

申请号：US18102356

申请日：2023-01-27

Applicant: Raytheon Technologies Corporation

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

IPC: F02C3/30

CPC classification number: F02C3/30 ,  F05D2220/323 ,  F05D2240/35

Abstract: A turbine engine assembly an inter-turbine burner that is disposed between at least two of a plurality of turbine stages where an additional amount of fuel is mixed with an exhaust gas flow to generate a reheated gas flow. A condenser extracts water from the reheated gas flow and an evaporator system generates a steam flow from at least a portion of water that is extracted by the condenser for injection into the core airflow.

公开(公告)号：US20240052792A1

公开(公告)日：2024-02-15

申请号：US17886950

申请日：2022-08-12

Applicant: RAYTHEON TECHNOLOGIES CORPORATION

Inventor： Neil J. Terwilliger ,  Nathan A. Snape ,  Joseph B. Staubach

IPC: F02C9/40 ,  F02C7/22 ,  F02C3/06

CPC classification number: F02C9/40 ,  F02C7/22 ,  F02C3/06

Abstract: A propulsion system for an aircraft includes a core engine that includes a core flow path where a core flow is compressed in a compressor section, communicated to a combustor section, mixed with a hydrogen-based fuel and ignited to generate a gas flow that is expanded through a turbine section. A fuel system is configured to supply a hydrogen based fuel to the combustor through a fuel flow path. A condenser is arranged along the core flow path and configured to extract water from the gas flow. An intercooling system receives a portion of water from the condenser for cooling a portion of the core flow at a first location within the compressor section. Heated water from the intercooling system is exhausted to a second location within the core flow path downstream of the first location.