公开(公告)号：US12078102B2

公开(公告)日：2024-09-03

申请号：US17951972

申请日：2022-09-23

Applicant: RAYTHEON TECHNOLOGIES CORPORATION

Inventor： Stephen H. Taylor ,  Alan Retersdorf ,  Oliver V. Atassi ,  Nathan A. Snape

IPC: F02C6/08 ,  F02C3/13 ,  F02C7/08 ,  F02C7/10 ,  F02C7/32 ,  F02C9/18 ,  F02C9/52 ,  F02G5/00 ,  F02K3/02

CPC classification number: F02C6/08 ,  F02C3/13 ,  F02C7/08 ,  F02C7/10 ,  F02C7/32 ,  F02C9/18 ,  F02C9/52 ,  F02G5/00 ,  F02K3/02

Abstract: A gas turbine engine assembly including a tap that is at a location up stream of the combustor section for drawing a bleed airflow. An exhaust heat exchanger is configured to transfer thermal energy from the exhaust gas flow into the bleed airflow and communicate the heated bleed airflow into the turbine section where it is expanded to drive the turbine section.

公开(公告)号：US12031478B2

公开(公告)日：2024-07-09

申请号：US17951991

申请日：2022-09-23

Applicant: RAYTHEON TECHNOLOGIES CORPORATION

Inventor： Stephen H. Taylor ,  Alan Retersdorf ,  Oliver V. Atassi

IPC: F02C6/08 ,  F02C3/10 ,  F02C7/18 ,  F02C7/32 ,  F02C7/36 ,  F02C9/18

CPC classification number: F02C6/08 ,  F02C3/10 ,  F02C7/18 ,  F02C7/32 ,  F02C7/36 ,  F02C9/18 ,  F05D2220/323 ,  F05D2220/36

Abstract: An aircraft propulsion system includes a core engine that includes a core flow path where a core airflow is compressed in a compressor section, communicated to a combustor section, mixed with fuel and ignited to generate a gas flow that is expanded through a turbine section for powering a primary propulsor. The aircraft propulsion system further includes a tap that is at a location upstream of the combustor section where a bleed airflow is drawn, a heat exchanger where the bleed airflow is heated by the gas flow, a power turbine through which heated bleed airflow is expanded to generate a work output, and a secondary propulsor that is driven by the work output that is generated by the power turbine.

公开(公告)号：US20240117764A1

公开(公告)日：2024-04-11

申请号：US17951991

申请日：2022-09-23

Applicant: RAYTHEON TECHNOLOGIES CORPORATION

Inventor： Stephen H. Taylor ,  Alan Retersdorf ,  Oliver V. Atassi

IPC: F02C9/18 ,  F02C7/32 ,  F02C7/36

CPC classification number: F02C9/18 ,  F02C7/32 ,  F02C7/36 ,  F05D2220/323 ,  F05D2220/36

Abstract: An aircraft propulsion system includes a core engine that includes a core flow path where a core airflow is compressed in a compressor section, communicated to a combustor section, mixed with fuel and ignited to generate a gas flow that is expanded through a turbine section for powering a primary propulsor. The aircraft propulsion system further includes a tap that is at a location upstream of the combustor section where a bleed airflow is drawn, a heat exchanger where the bleed airflow is heated by the gas flow, a power turbine through which heated bleed airflow is expanded to generate a work output, and a secondary propulsor that is driven by the work output that is generated by the power turbine.

公开(公告)号：US11773778B1

公开(公告)日：2023-10-03

申请号：US17951936

申请日：2022-09-23

Applicant: RAYTHEON TECHNOLOGIES CORPORATION

Inventor： Alan Retersdorf ,  Stephen H. Taylor ,  Oliver V. Atassi

IPC: F02C6/18 ,  F02C7/32 ,  F02C9/18

CPC classification number: F02C6/18 ,  F02C7/32 ,  F02C9/18 ,  F05D2220/62 ,  F05D2260/10 ,  F05D2260/213 ,  F05D2260/232

Abstract: A turbine engine assembly includes a tap is at a location upstream of the combustor section where a bleed airflow is drawn. The bleed air is pressurized in an auxiliary compressor section, heated in an exhaust heat exchanger and expanded through a power turbine that is coupled to drive the auxiliary compressor section.

公开(公告)号：US11702981B1

公开(公告)日：2023-07-18

申请号：US17724904

申请日：2022-04-20

Applicant: RAYTHEON TECHNOLOGIES CORPORATION

Inventor： Stephen H. Taylor ,  Alan Retersdorf

IPC: F02C6/08 ,  F02C9/18 ,  F02C3/10 ,  F02C7/14 ,  F02C7/18 ,  F02C7/36

CPC classification number: F02C6/08 ,  F02C3/10 ,  F02C7/14 ,  F02C7/18 ,  F02C9/18 ,  F02C7/36

Abstract: A gas turbine engine assembly includes a core engine that includes a core flow path where a core airflow is compressed in a compressor section, communicated to a combustor section, mixed with fuel and ignited to generate a high energy combusted gas flow that is expanded through a turbine section, a first tap at a location up stream of the combustor section for communicating a portion of the core airflow as a bleed airflow downstream of the combustor section, a heat exchanger that places the bleed airflow that is communicated from the first tap in thermal communication with the high energy combusted gas flow downstream of the combustor section, and a power turbine that is configured to generate shaft power from expansion of the heated bleed airflow, the power turbine includes an inlet that is configured to receive the heated bleed airflow from the heat exchanger.

公开(公告)号：US20240102417A1

公开(公告)日：2024-03-28

申请号：US17951972

申请日：2022-09-23

Applicant: RAYTHEON TECHNOLOGIES CORPORATION

Inventor： Stephen H. Taylor ,  Alan Retersdorf ,  Oliver V. Atassi ,  Nathan A. Snape

IPC: F02C6/08 ,  F02C9/18 ,  F02G5/00 ,  F02K3/02

CPC classification number: F02C6/08 ,  F02C9/18 ,  F02G5/00 ,  F02K3/02

Abstract: A gas turbine engine assembly including a tap that is at a location up stream of the combustor section for drawing a bleed airflow. An exhaust heat exchanger is configured to transfer thermal energy from the exhaust gas flow into the bleed airflow and communicate the heated bleed airflow into the turbine section where it is expanded to drive the turbine section.