专利快速检索-快速检索全球专利，免费商用专利数据库-IPRDB

1. 发明授权

US11136922B2 Gas turbine engine transfer efficiency 有权
公开(公告)号：US11136922B2
公开(公告)日：2021-10-05
申请号：US16545003
申请日：2019-08-20
申请人： ROLLS-ROYCE plc
发明人： Craig W Bemment , Pascal Dunning
IPC分类号： F02C9/28 , F01D11/14 , F01D13/00 , F01D25/24
摘要： 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.

2. 发明授权

US10794294B1 Efficient jet 有权
公开(公告)号：US10794294B1
公开(公告)日：2020-10-06
申请号：US16695472
申请日：2019-11-26
申请人： ROLLS-ROYCE plc
发明人： Pascal Dunning , Craig W Bemment
IPC分类号： F02C7/36 , F02K3/06
摘要： A gas turbine engine includes: an engine core, compressor system, and core shaft. A compressor exit pressure is defined as an average airflow pressure at the exit of the highest pressure compressor at cruise conditions. The core has an annular splitter and bypass flow. Stagnation streamlines around the engine circumference form a streamsurface. A fan is upstream the core with blades having leading and trailing edges, and a radially inner portion within the streamtube. A fan root entry pressure is an average airflow pressure across the radially inner portion leading edge of each fan blade at cruise conditions. An overall pressure ratio, OPR, is defined as the compressor exit pressure divided by the fan root entry pressure. A bypass jet velocity is defined as the jet velocity of air flow exiting the bypass exhaust nozzle at cruise conditions. A jet velocity to OPR ratio is in a range between 4.7 m/s and 7.7 m/s.

3. 发明授权

US11668246B2 Gas turbine engine 有权
公开(公告)号：US11668246B2
公开(公告)日：2023-06-06
申请号：US17209989
申请日：2021-03-23
申请人： ROLLS-ROYCE PLC
发明人： Pascal Dunning , Roderick M Townes , Michael J Whittle
IPC分类号： F02C7/36 , F02C3/073 , F01D5/28 , F01D5/30
CPC分类号： F02C7/36 , F01D5/284 , F01D5/3084 , F02C3/073 , F05D2240/12 , F05D2240/30
摘要： 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.

4. 发明授权

US11635021B2 Compression in a gas turbine engine 有权
公开(公告)号：US11635021B2
公开(公告)日：2023-04-25
申请号：US17697630
申请日：2022-03-17
申请人： ROLLS-ROYCE PLC
发明人： Craig W Bemment , Pascal Dunning
IPC分类号： F02C3/04 , F02C7/36
摘要： 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.

5. 发明授权

US11560853B2 Gas turbine engine transfer efficiency 有权
公开(公告)号：US11560853B2
公开(公告)日：2023-01-24
申请号：US17466086
申请日：2021-09-03
申请人： ROLLS-ROYCE PLC
发明人： Craig W Bemment , Pascal Dunning
IPC分类号： F02C9/28 , F01D11/14 , F01D13/00 , F01D25/24
摘要： 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.

6. 发明授权

US11466617B2 Gas turbine engine with efficient thrust generation 有权
公开(公告)号：US11466617B2
公开(公告)日：2022-10-11
申请号：US17345328
申请日：2021-06-11
申请人： ROLLS-ROYCE plc
发明人： Pascal Dunning , Michael J Whittle , Roderick M Townes
IPC分类号： F02C3/107 , F02C3/073 , F02C7/36 , F01D5/28
摘要： 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.

7. 发明授权

US11053842B2 Compression in a gas turbine engine 有权
公开(公告)号：US11053842B2
公开(公告)日：2021-07-06
申请号：US16558417
申请日：2019-09-03
申请人： ROLLS-ROYCE plc
发明人： Craig W Bemment , Pascal Dunning
IPC分类号： F02C3/04 , F02C7/36
摘要： 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.

8. 发明授权

US11326512B2 Compression in a gas turbine engine 有权
公开(公告)号：US11326512B2
公开(公告)日：2022-05-10
申请号：US17345588
申请日：2021-06-11
申请人： ROLLS-ROYCE PLC
发明人： Craig W Bemment , Pascal Dunning
IPC分类号： F02C7/36 , F02C3/04
摘要： 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.

9. 发明授权

US11047301B2 Gas turbine engine with efficient thrust generation 有权
公开(公告)号：US11047301B2
公开(公告)日：2021-06-29
申请号：US16411224
申请日：2019-05-14
申请人： ROLLS-ROYCE plc
发明人： Pascal Dunning , Michael J Whittle , Roderick M Townes
IPC分类号： F02C3/107 , F02C3/073 , F02C7/36 , F01D5/28
摘要： 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.

10. 发明授权

US10989112B2 Gas turbine engine 有权
公开(公告)号：US10989112B2
公开(公告)日：2021-04-27
申请号：US16411341
申请日：2019-05-14
申请人： ROLLS-ROYCE plc
发明人： Pascal Dunning , Roderick M Townes , Michael J Whittle
IPC分类号： F02C3/073 , F02C7/36 , F01D5/28 , F01D5/30
摘要： 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.

你已经成功收藏专利！

检索式保存成功!

IPRDB

热门服务

关于我们

友情链接

联系方式