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公开(公告)号：US20170370290A1
公开(公告)日：2017-12-28
申请号：US15604999
申请日：2017-05-25
Applicant: ROLLS-ROYCE plc
Inventor： Stephen J BRADBROOK
CPC classification number: F02C7/143 , F02C3/04 , F02K3/06 , F02K3/115 , F05D2220/323 , F05D2260/205 , F05D2260/211 , Y02T50/675
Abstract: An aircraft gas turbine engine includes a fan arranged to be driven by a gas turbine engine core. The core includes a first core module including a first compressor and a fan drive turbine interconnected by a first shaft, and a second core module including a second compressor and a second turbine interconnected by a second shaft, the first and second core modules being axially spaced. The gas turbine engine further includes an intercooler arrangement configured to cool core airflow between the first and second compressors, the intercooler arrangement including a cooling air duct provided in heat exchange relationship with a compressor duct provided between the first and second compressors, the cooling air duct including a fan air inlet configured to ingest fan air downstream of the fan, wherein the cooling air duct includes a flow modulation valve configured to modulate air mass flow through the fan air inlet.
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公开(公告)号：US20240102431A1
公开(公告)日：2024-03-28
申请号：US18215864
申请日：2023-06-29
Applicant: ROLLS-ROYCE plc
Inventor： Stephen J BRADBROOK , William J SMITH , Jonathan P BRADLEY
CPC classification number: F02K3/06 , F01D9/041 , F01D25/24 , F05D2240/35
Abstract: A novel configuration for an axial flow gas turbine engine for aircraft has an engine core having a core length and including a first turbine, an axial compressor, and a drive connecting the first turbine to the axial compressor, the engine core further including a second turbine, and a fan shaft connecting the second turbine to a fan located upstream of the engine core, the fan including a plurality of fan blades extending from a fan hub, the fan having a tip radius from 90 mm to 225 mm and wherein the ratio of fan tip radius to an engine length is 0.15 to 0.25. The engine may have 3, 4 or 5 compressor stages and combustor volume (in litres) which when divided by fan tip radius (in mm) is 0.015 to 0.083. The fan may be multistage fan configured to permit an electric motor to be located within fan hub diameter.
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公开(公告)号：US20200017226A1
公开(公告)日：2020-01-16
申请号：US16449527
申请日：2019-06-24
Applicant: ROLLS-ROYCE plc
Inventor： Stephen J BRADBROOK
Abstract: A propulsion system includes a main gas turbine engine adapted for generating propulsive thrust during subsonic and supersonic flight operations and a supplementary propulsion unit adapted for generating additional thrust. The supplementary propulsion unit has an air intake and an exhaust for gas accelerated by the supplementary propulsion unit to provide the additional thrust and is adapted to generate the additional thrust during a limited range of subsonic flight operations, and to be dormant during other flight operations. The propulsion system has housing for the supplementary propulsion unit, including intake and exhaust covers which are moveable between deployed and stowed configurations. During the limited range of subsonic flight operations the intake and exhaust cover are moved to the deployed configuration to open the intake and the exhaust. During other flight operations the intake and exhaust cover are moved to the stowed configuration to close the intake and the exhaust.
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公开(公告)号：US20240011432A1
公开(公告)日：2024-01-11
申请号：US18215860
申请日：2023-06-29
Applicant: ROLLS-ROYCE plc
Inventor： Stephen J BRADBROOK , William J SMITH , Jonathan P BRADLEY
IPC: F02C3/04
CPC classification number: F02C3/04 , F05D2220/36
Abstract: A novel configuration for an axial flow gas turbine engine for an aircraft has an engine core having a core length and including a first turbine, an axial compressor, and a drive connecting the first turbine to the axial compressor, the engine core further including a second turbine, and a fan shaft connecting the second turbine to a fan located upstream of engine core, the fan including a plurality of fan blades extending from a fan hub, the fan having a tip radius from 90 mm to 225 mm and wherein the ratio of fan tip radius to engine length is 0.15 to 0.25. The engine may have 3, 4 or 5 compressor stages and a combustor volume (in litres) which when divided by the fan tip radius (in mm) is 0.015 to 0.083. The fan may be multistage fan configured to permit electric motor to be located within fan hub diameter.
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Patent Agency Ranking
公开(公告)号：US20230167784A1
公开(公告)日：2023-06-01
申请号：US17940034
申请日：2022-09-08
Applicant: ROLLS-ROYCE plc
Inventor： Benjamin J SELLERS , Stephen J BRADBROOK , Robert J CORIN , Richard J HUNSLEY
CPC classification number: F02K3/06 , F01D15/10 , F05D2220/323 , F05D2220/36
Abstract: An aircraft turbofan gas turbine engine includes a fan assembly, a compressor module, and a turbine module. An electric machine is positioned downstream of the fan assembly and is connected to the turbine module. The fan assembly includes a highest pressure fan stage having a plurality of fan blades defining a fan diameter. The turbine module includes a lowest pressure turbine stage having a row of rotor blades. The gas turbine engine has a fan tip axis that joins a radially outer tip of the leading edge of one of the plurality of fan blades of the highest pressure fan stage, and the radially outer tip of the trailing edge of one of the rotor blades of the lowest pressure turbine stage. The fan tip axis lies in a longitudinal plane which contains a centreline of the gas turbine engine. The fan axis angle is between 11 and 20 degrees.
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公开(公告)号：US20230219694A1
公开(公告)日：2023-07-13
申请号：US17940030
申请日：2022-09-08
Applicant: ROLLS-ROYCE plc
Inventor： Benjamin J SELLERS , Andrew J NEWMAN , Gordon MARGARY , Paul R DAVIES , Stephen J BRADBROOK
IPC: B64D33/08
CPC classification number: B64D33/08 , B64D2027/026
Abstract: A cooling system for an aircraft comprises a gas turbine engine, an ancillary apparatus, and a heat exchanger. The gas turbine engine comprises, in axial flow sequence, a compressor module, a combustor module, and a turbine module, with a first electric machine being rotationally connected to the turbine module. The first electrical machine is configured to generate an electrical power PEM1 (W). The heat exchanger is configured to transfer a total waste heat energy Q (W) generated by the gas turbine engine and the ancillary apparatus, to an airflow passing through the heat exchanger, and a ratio S of:
S
=
(
Total
Electrical
Power
Generated
=
P
EM
1
)
(
Total
Heat
Energy
Rejected
to
Airflow
=
Q
)
is in a range of between 0.50 and 5.00.-
公开(公告)号：US20230184197A1
公开(公告)日：2023-06-15
申请号：US17940531
申请日：2022-09-08
Applicant: ROLLS-ROYCE PLC
Inventor： Benjamin J. SELLERS , Andrew J NEWMAN , David A JONES , Stephen J BRADBROOK
CPC classification number: F02K3/06 , B64D27/20 , F02K3/068 , F05D2200/14 , F05D2220/323
Abstract: An gas turbine engine for an aircraft includes, in axial flow sequence, a compressor module, a combustor module, and a turbine module, together with a first electrical machine rotationally connected to the turbine module. The combustor module has a combustor volume V (cm3). In use, at a full power condition, the gas turbine engine has a maximum corrected core flow Q (m3/sec), and a ratio T of:
T
=
(
Maximum
Corrected
Core
Flow
=
Q
)
(
Combustor
volume
=
V
)
is in a range of between 450 and 2,500.-
公开(公告)号：US20230167769A1
公开(公告)日：2023-06-01
申请号：US17940473
申请日：2022-09-08
Applicant: ROLLS-ROYCE PLC
Inventor： Andrew J NEWMAN , Martin N GOODHAND , Paul R DAVIES , David A JONES , Stephen J BRADBROOK
CPC classification number: F02C7/14 , F02C7/18 , F05D2220/323 , F05D2260/213 , F05D2260/232
Abstract: A cooling system for an aircraft comprises an apparatus, a heat exchanger, and a vapour cycle machine. The apparatus comprises a first fluid being circulated to provide cooling to the apparatus, with the heat exchanger being configured to transfer waste heat energy from the first fluid to a second fluid. The second fluid has a temperature T2 (°C), and the vapour cycle machine is configured to increase a temperature T1 (°C) of the first fluid, to a temperature greater than T2 (°C).
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公开(公告)号：US20220112813A1
公开(公告)日：2022-04-14
申请号：US17496193
申请日：2021-10-07
Applicant: ROLLS-ROYCE plc
Inventor： Stephen J BRADBROOK , Martin N GOODHAND , Paul M HIELD , Andrew PARSLEY , Natalie C WONG , Robert J CORIN , Thomas S BINNINGTON
Abstract: A turbofan gas turbine engine includes, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, a turbine module, and an exhaust module. The fan assembly includes fan blades defining a fan diameter. The heat exchanger module is in communication with the fan assembly by an inlet duct, and the heat exchanger module further includes radially-extending hollow vanes arranged in a circumferential array, with a channel extending axially between hollow vanes. Each hollow vane accommodates at least one heat transfer element to transfer heat from a first fluid contained within the or each heat transfer element to a corresponding vane airflow passing through the hollow vane and over a surface of the or each heat transfer element. Each hollow vane further includes a flow modulator configured to regulate airflow in proportion to total airflow entering the heat exchanger module in response to a user requirement.
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