公开(公告)号：US12110830B2

公开(公告)日：2024-10-08

申请号：US18137123

申请日：2023-04-20

Applicant: ROLLS-ROYCE plc

Inventor： Pascal Dunning ,  Roderick M. Townes ,  Michael J. Whittle

IPC: F02C7/36 ,  F01D5/28 ,  F01D5/30 ,  F02C3/073

CPC classification number: F02C7/36 ,  F01D5/284 ,  F01D5/3084 ,  F02C3/073 ,  F05D2240/12 ,  F05D2240/30

Abstract: A highly efficient gas turbine engine is provided. The fan of the gas turbine engine is driven from a turbine via a gearbox, such that the fan has a lower rotational speed than the driving turbine, thereby providing efficiency gains. The efficient fan system is mated to a core that has low cooling flow requirements and/or high temperature capability, and which may have particularly low mass for a given power.

公开(公告)号：US11560853B2

公开(公告)日：2023-01-24

申请号：US17466086

申请日：2021-09-03

Applicant: ROLLS-ROYCE PLC

Inventor： Craig W Bemment ,  Pascal Dunning

IPC: F02C9/28 ,  F01D11/14 ,  F01D13/00 ,  F01D25/24

Abstract: A gas turbine engine for an aircraft includes an engine core including a first, lower pressure, turbine, a first compressor, and a first core shaft connecting the first turbine to the first compressor; and a second, higher pressure, turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor, and a fan located upstream of the engine core and including a plurality of fan blades extending from a hub. A turbine to fan tip temperature change ratio of a low pressure turbine temperature change to a fan tip temperature rise is in the range from 1.46 to 2.0.

公开(公告)号：US11466617B2

公开(公告)日：2022-10-11

申请号：US17345328

申请日：2021-06-11

Applicant: ROLLS-ROYCE plc

Inventor： Pascal Dunning ,  Michael J Whittle ,  Roderick M Townes

IPC: F02C3/107 ,  F02C3/073 ,  F02C7/36 ,  F01D5/28

Abstract: A highly efficient gas turbine engine includes the fan of the gas turbine engine driven from a turbine via a gearbox, such that the fan has a lower rotational speed than the driving turbine, thereby providing efficiency gains. The efficient fan system is mated to a core that has low cooling flow requirements and/or high temperature capability, and which may have particularly low mass for a given power.

公开(公告)号：US11053842B2

公开(公告)日：2021-07-06

申请号：US16558417

申请日：2019-09-03

Applicant: ROLLS-ROYCE plc

Inventor： Craig W Bemment ,  Pascal Dunning

IPC: F02C3/04 ,  F02C7/36

Abstract: An engine core including turbine, compressor, and core shaft connecting the turbine to the compressor, wherein a compressor exit temperature has an average airflow; and a fan upstream including a plurality of fan blades extending from a hub, each fan blade having a leading and trailing edge, wherein a fan rotor entry temperature has an average airflow across the leading edge of each blade at cruise conditions and fan tip rotor exit temperature has an average temperature of airflow across a radially outer portion of each blade at the trailing edge cruise conditions. A fan tip temperature rise as: the ⁢ ⁢ fan ⁢ ⁢ tip ⁢ ⁢ rotor ⁢ ⁢ exit ⁢ ⁢ temperature the ⁢ ⁢ fan ⁢ ⁢ rotor ⁢ ⁢ entry ⁢ ⁢ temperature . A core temperature rise as: the ⁢ ⁢ compressor ⁢ ⁢ exit ⁢ ⁢ temperature the ⁢ ⁢ fan ⁢ ⁢ rotor ⁢ ⁢ entry ⁢ ⁢ temperature , A core to fan tip temperature rise ratio of: the ⁢ ⁢ core ⁢ ⁢ temperature ⁢ ⁢ rise the ⁢ ⁢ fan ⁢ ⁢ tip ⁢ ⁢ temperature ⁢ ⁢ rise is in the range from 2.845-3.8.

公开(公告)号：US12146440B2

公开(公告)日：2024-11-19

申请号：US18583395

申请日：2024-02-21

Applicant: ROLLS-ROYCE PLC

Inventor： Roderick M Townes ,  Michael J Whittle ,  Pascal Dunning

IPC: F02C7/08 ,  F01D5/28 ,  F02C7/36 ,  F02C9/16 ,  F02K3/06

Abstract: A highly efficient gas turbine engine is a system wherein the fan of the gas turbine engine is driven from a turbine via a gearbox, such that the fan has a lower rotational speed than the driving turbine, thereby providing efficiency gains. The efficient fan system is mated to a core that has low cooling flow requirements and/or high temperature capability, and which may have particularly low mass for a given power.

公开(公告)号：US11668246B2

公开(公告)日：2023-06-06

申请号：US17209989

申请日：2021-03-23

Applicant: ROLLS-ROYCE PLC

Inventor： Pascal Dunning ,  Roderick M Townes ,  Michael J Whittle

IPC: F02C7/36 ,  F02C3/073 ,  F01D5/28 ,  F01D5/30

CPC classification number: F02C7/36 ,  F01D5/284 ,  F01D5/3084 ,  F02C3/073 ,  F05D2240/12 ,  F05D2240/30

Abstract: A highly efficient gas turbine engine is provided. The fan of the gas turbine engine is driven from a turbine via a gearbox, such that the fan has a lower rotational speed than the driving turbine, thereby providing efficiency gains. The efficient fan system is mated to a core that has low cooling flow requirements and/or high temperature capability, and which may have particularly low mass for a given power.

公开(公告)号：US11635021B2

公开(公告)日：2023-04-25

申请号：US17697630

申请日：2022-03-17

Applicant: ROLLS-ROYCE PLC

Inventor： Craig W Bemment ,  Pascal Dunning

IPC: F02C3/04 ,  F02C7/36

Abstract: A gas turbine engine for an aircraft comprises an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, wherein a compressor exit temperature is defined as an average temperature of airflow at the exit from the compressor; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, each fan blade having a leading edge and a trailing edge, wherein a fan rotor entry temperature is defined as an average temperature of airflow across the leading edge of each fan blade at cruise conditions and a fan tip rotor exit temperature is defined as an average temperature of airflow across a radially outer portion of each fan blade at the trailing edge at cruise conditions. A core to fan tip temperature rise ratio is in the range from 2.845 to 3.8.

公开(公告)号：US12297770B2

公开(公告)日：2025-05-13

申请号：US18405485

申请日：2024-01-05

Applicant: ROLLS-ROYCE plc

Inventor： Craig W Bemment ,  Pascal Dunning

IPC: F02C3/04 ,  F02C7/36 ,  F02K3/06

Abstract: A gas turbine engine for an aircraft comprises an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, wherein a compressor exit temperature is defined as an average temperature of airflow at the exit from the compressor; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, each fan blade having a leading edge and a trailing edge, wherein a fan rotor entry temperature is defined as an average temperature of airflow across the leading edge of each fan blade at cruise conditions and a fan tip rotor exit temperature is defined as an average temperature of airflow across a radially outer portion of each fan blade at the trailing edge at cruise conditions. A core to fan tip temperature rise ratio is in the range from 2.845 to 3.8.

公开(公告)号：US11898489B2

公开(公告)日：2024-02-13

申请号：US18123091

申请日：2023-03-17

Applicant: ROLLS-ROYCE plc

Inventor： Craig W Bemment ,  Pascal Dunning

IPC: F02K3/06 ,  F02C3/04 ,  F02C7/36

CPC classification number: F02C3/04 ,  F02C7/36 ,  F02K3/06 ,  F05D2200/14 ,  F05D2220/323 ,  F05D2220/36 ,  F05D2240/307 ,  F05D2260/4031

Abstract: A gas turbine engine for an aircraft comprises an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, wherein a compressor exit temperature is defined as an average temperature of airflow at the exit from the compressor; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, each fan blade having a leading edge and a trailing edge, wherein a fan rotor entry temperature is defined as an average temperature of airflow across the leading edge of each fan blade at cruise conditions and a fan tip rotor exit temperature is defined as an average temperature of airflow across a radially outer portion of each fan blade at the trailing edge at cruise conditions. A core to fan tip temperature rise ratio is in the range from 2.845 to 3.8.

公开(公告)号：US11136922B2

公开(公告)日：2021-10-05

申请号：US16545003

申请日：2019-08-20

Applicant: ROLLS-ROYCE plc

Inventor： Craig W Bemment ,  Pascal Dunning

IPC: F02C9/28 ,  F01D11/14 ,  F01D13/00 ,  F01D25/24

Abstract: A gas turbine engine for an aircraft includes an engine core including a first, lower pressure, turbine, a first compressor, and a first core shaft connecting the first turbine to the first compressor; and a second, higher pressure, turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor, and a fan located upstream of the engine core and including a plurality of fan blades extending from a hub. A low pressure turbine temperature change is defined as: the ⁢ ⁢ first ⁢ ⁢ turbine ⁢ ⁢ entrance ⁢ ⁢ temperature the ⁢ ⁢ first ⁢ ⁢ turbine ⁢ ⁢ exit ⁢ ⁢ temperature . A fan tip temperature rise is defined as: the ⁢ ⁢ fan ⁢ ⁢ tip ⁢ ⁢ rotor ⁢ ⁢ exit ⁢ ⁢ temperature the ⁢ ⁢ fan ⁢ ⁢ rotor ⁢ ⁢ entry ⁢ ⁢ temperature . A turbine to fan tip temperature change ratio of: the ⁢ ⁢ low ⁢ ⁢ pressure ⁢ ⁢ turbine ⁢ ⁢ temperature ⁢ ⁢ change the ⁢ ⁢ fan ⁢ ⁢ tip ⁢ ⁢ temperature ⁢ ⁢ rise is in the range from 1.46 to 2.0.