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公开(公告)号：US12092037B2
公开(公告)日：2024-09-17
申请号：US17940419
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
CPC classification number: F02C7/32 , F02C6/00 , F02C6/20 , F02C7/12 , F05D2220/323 , F05D2220/76 , F05D2260/213
Abstract: A gas turbine engine for an aircraft includes, 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 a total electrical power PEM1 (W), and the gas turbine engine is configured to generate a total shaft power PSHAFT (W); and a ratio R of:
R
=
(
Total
Shaft
Power
=
P
SHAFT
)
(
Total
Electrical
Power
Generated
=
P
EM
1
)
is in a range of between 0.005 and 0.020.-
公开(公告)号：US11920540B2
公开(公告)日：2024-03-05
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：US11181042B2
公开(公告)日：2021-11-23
申请号：US16411347
申请日：2019-05-14
Applicant: ROLLS-ROYCE plc
Inventor： Michael O Hales , Craig W Bemment , Stephane M M Baralon , Benjamin J Sellers , Christopher Benson , Benedict R Phelps , Mark J Wilson
Abstract: A gas turbine engine has a cycle operability parameter β in a defined range to achieve improved overall performance, taking into account fan operability and/or bird strike requirements as well as engine efficiency. The defined range of cycle operability parameter β may be particularly beneficial for gas turbine engines in which the fan is driven by a turbine through a gearbox.
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Patent Agency Ranking
公开(公告)号：US12104530B2
公开(公告)日：2024-10-01
申请号：US17940035
申请日：2022-09-08
Applicant: ROLLS-ROYCE plc
Inventor： Benjamin J Sellers , Jonathan A Cherry , Craig Town
CPC classification number: F02C7/12 , F02C3/06 , F05D2220/323 , F05D2260/213 , F05D2260/40
Abstract: An aircraft turbofan gas turbine engine includes a fan assembly, a compressor module, and a turbine module. An electric machine positioned downstream of the fan assembly is rotationally connected to the turbine module. The fan assembly is in fluid communication with the compressor module by an intermediate duct and includes a highest pressure fan stage having a plurality of fan blades defining a fan diameter. The compressor module includes a lowest pressure compressor stage having a row of rotor blades. An intermediate flow axis is defined joining a radially outer tip of a trailing edge of one of the fan blades of the highest pressure fan stage, and a radially outer tip of a leading edge of one of the rotor blades of a leading edge of a lowest-pressure compressor blade. An intermediate flow axis angle and the intermediate flow axis angle is from −20 to −30 degrees.
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公开(公告)号：US12092030B2
公开(公告)日：2024-09-17
申请号：US17940055
申请日：2022-09-08
Applicant: ROLLS-ROYCE plc
Inventor： Natalie C Wong , Jonathan A Cherry , Paul R Davies , David A Jones , Andrew J Newman , Benjamin J Sellers , Stephen J Bradbrook
CPC classification number: F02C7/12 , F02C6/20 , F02C7/32 , F05D2220/323 , F05D2220/76 , F05D2260/213
Abstract: An aircraft gas turbine engine includes a heat exchanger module, and a core engine including an intermediate-pressure compressor, a high-pressure compressor, a high pressure turbine, and a low-pressure turbine. The high-pressure compressor is connected to the high-pressure turbine by a first shaft, and the intermediate-pressure compressor is connected to the low-pressure turbine by a second shaft. The heat exchanger module includes a central hub and heat transfer elements extending radially from the central hub and spaced in a circumferential array, for transferring heat energy from a fluid within the heat transfer elements to an inlet airflow passing over the heat transfer elements prior to entry of the airflow into an inlet to the core engine. The gas turbine engine further includes a first electric machine connected to the first shaft and positioned downstream of the heat exchanger module, and a second electric machines connected to the second shaft.
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公开(公告)号：US12006055B2
公开(公告)日：2024-06-11
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
CPC classification number: B64D33/08 , B64D27/026
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：US11994067B2
公开(公告)日：2024-05-28
申请号：US17940055
申请日：2022-09-08
Applicant: ROLLS-ROYCE plc
Inventor： Natalie C Wong , Jonathan A Cherry , Paul R Davies , David A Jones , Andrew J Newman , Benjamin J Sellers , Stephen J Bradbrook
CPC classification number: F02C7/12 , F02C6/20 , F02C7/32 , F05D2220/323 , F05D2220/76 , F05D2260/213
Abstract: An aircraft gas turbine engine includes a heat exchanger module, and a core engine including an intermediate-pressure compressor, a high-pressure compressor, a high pressure turbine, and a low-pressure turbine. The high-pressure compressor is connected to the high-pressure turbine by a first shaft, and the intermediate-pressure compressor is connected to the low-pressure turbine by a second shaft. The heat exchanger module includes a central hub and heat transfer elements extending radially from the central hub and spaced in a circumferential array, for transferring heat energy from a fluid within the heat transfer elements to an inlet airflow passing over the heat transfer elements prior to entry of the airflow into an inlet to the core engine. The gas turbine engine further includes a first electric machine connected to the first shaft and positioned downstream of the heat exchanger module, and a second electric machines connected to the second shaft.
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公开(公告)号：US11519363B2
公开(公告)日：2022-12-06
申请号：US17196382
申请日：2021-03-09
Applicant: ROLLS-ROYCE plc
Inventor： Michael O Hales , Craig W Bemment , Benjamin J Sellers , Ian J Bousfield , Amarveer S Mann
Abstract: A gas turbine engine (10) comprising: a high pressure turbine (17); a low pressure turbine (19); a high pressure compressor (15) coupled to the high pressure turbine (17) by a high pressure shaft (27); a propulsor (23) and a low pressure compressor (14) coupled to the low pressure turbine (19) via a low pressure shaft (26) and a reduction gearbox (30); wherein the low pressure compressor (14) consists of four compressor stages (14) and defines a cruise pressure ratio of between 2.4:1 and 3.3:1; the high pressure compressor (15) defines a cruise pressure ratio of less than 17:1; and the high pressure compressor (15) and low pressure compressor (14) together define a cruise core overall pressure ratio of greater than 36:1.
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