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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
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.
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公开(公告)号：US11131250B2
公开(公告)日：2021-09-28
申请号：US16526221
申请日：2019-07-30
Applicant: ROLLS-ROYCE plc
Inventor： Craig W Bemment
Abstract: A gas turbine engine for an aircraft and a method of operating a gas turbine engine on an aircraft. Embodiments disclosed include a gas turbine engine for an aircraft including: an engine core has a turbine, a compressor, and a core shaft; a fan located upstream of the engine core, the fan has a plurality of fan blades; a nacelle surrounding the engine core and defining a bypass duct and bypass exhaust nozzle; and a gearbox that receives an input from the core shaft and outputs drive to the fan wherein the gas turbine engine is configured such that a jet velocity ratio of a first jet velocity exiting from the bypass exhaust nozzle to a second jet velocity exiting from an exhaust nozzle of the engine core at idle conditions is greater by a factor of 2 or more than the jet velocity ratio at maximum take-off conditions.
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公开(公告)号：US11002194B2
公开(公告)日：2021-05-11
申请号：US16443942
申请日：2019-06-18
Applicant: ROLLS-ROYCE plc
Inventor： Craig W Bemment
Abstract: The present disclosure relates to a geared gas turbine engine for an aircraft. Example embodiments include a gas turbine engine for an aircraft including: an engine core having a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan having a plurality of fan blades; and a gearbox that receives an input from the core shaft to drive the fan at a lower rotational speed than the core shaft, the gearbox having a gear ratio of around 3.4 or higher, wherein the gas turbine engine is configured such that a jet velocity ratio between a first jet velocity exiting from a bypass duct of the engine and a second jet velocity exiting from an exhaust nozzle of the engine core is within a range from around 0.75 to around 0.82.
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公开(公告)号：US10794294B1
公开(公告)日：2020-10-06
申请号：US16695472
申请日：2019-11-26
Applicant: ROLLS-ROYCE plc
Inventor： Pascal Dunning , Craig W Bemment
Abstract: 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.
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