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公开(公告)号：US20240210037A1
公开(公告)日：2024-06-27
申请号：US18211624
申请日：2023-06-20
Applicant: ROLLS-ROYCE plc
Inventor： Christopher P MADDEN , Craig W BEMMENT , Andrew T SMITH , Peter SWANN
CPC classification number: F02C9/40 , F02C3/20 , F23R3/28 , F05D2240/35 , F05D2270/08
Abstract: A method of operating a gas turbine engine including a combustor. The combustor includes a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the chamber. The nozzles include a first and second subset of fuel spray nozzles. The combustor is operable in a condition wherein the first subset is supplied with more fuel than the second. A ratio of the number of nozzles in the first subset to the second is 1:2 to 1:5. The method includes operating the engine so a reduction of 20-80% in an average of particles/kg of nvPM in the exhaust when the engine is operating at 85% available thrust for given operating conditions and when the engine is operating at 30% is obtained when fuel provided to the nozzles is a sustainable aviation fuel instead of a fossil-based hydrocarbon fuel. Also, a gas turbine engine for an aircraft.
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公开(公告)号：US20210231060A1
公开(公告)日：2021-07-29
申请号：US17231676
申请日：2021-04-15
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：US20250003592A1
公开(公告)日：2025-01-02
申请号：US18882197
申请日：2024-09-11
Applicant: ROLLS-ROYCE plc
Inventor： Christopher P. MADDEN , Craig W BEMMENT , Andrew T SMITH , Peter SWANN
Abstract: A method of operating a gas turbine engine; the gas turbine engine includes a combustor. The combustor includes a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber. The plurality of fuel spray nozzles includes a first subset of fuel spray nozzles and a second subset of fuel spray nozzles. The combustor is operable in a condition in which the first subset of fuel spray nozzles are supplied with more fuel than the second subset of fuel spray nozzles. A ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5. The method includes: providing fuel to the one or more fuel-oil heat exchangers. Also provided is a gas turbine engine for an aircraft.
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公开(公告)号：US20240401531A1
公开(公告)日：2024-12-05
申请号：US18420204
申请日：2024-01-23
Applicant: ROLLS-ROYCE plc
Inventor： Craig W BEMMENT , Benjamin J KEELER , Kevin R MCNALLY , Andrea MINELLI
Abstract: A method of operating a gas turbine engine and a gas turbine engine includes a fuel delivery system arranged to provide fuel, a combustor arranged to combust at least a proportion of the fuel, a primary fuel-oil heat exchanger arranged to have up to 100% of the fuel provided by the fuel delivery system flow therethrough, and a secondary fuel-oil heat exchanger arranged to have a proportion of the fuel from the primary fuel-oil heat exchanger flow therethrough. Fuel is arranged to flow from the primary fuel-oil heat exchanger to the secondary fuel-oil heat exchanger whereas oil is arranged to flow from the secondary fuel-oil heat exchanger to the primary fuel-oil heat exchanger. A fuel viscosity is adjusted to a maximum of 0.58 mm2/s on entry to the combustor at cruise conditions.
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公开(公告)号：US20230323824A1
公开(公告)日：2023-10-12
申请号：US18098433
申请日：2023-01-18
Applicant: ROLLS-ROYCE PLC
Inventor： Peter SWANN , Christopher P MADDEN , Craig W BEMMENT
CPC classification number: F02C9/40 , F02C9/28 , F05D2270/31 , F05D2220/323 , F05D2270/306 , F05D2220/80
Abstract: A aircraft gas turbine engine and operation method, the engine including: a staged combustion system having pilot and main fuel injectors, and operates in a pilot-only range wherein fuel delivers to pilot fuel injectors, and a pilot-and-main operation range wherein fuel is delivered to at least the main fuel injectors. The engine further includes a fuel delivery regulator to pilot and main fuel injectors, which receives fuel from a first and second source containing fuels each with different characteristics. The staged combustion system switches between pilot-only and pilot-and-main range operation when in steady cruise mode, the mode defining a boundary between first and second engine cruise operation range. The fuel delivery regulator delivers fuel to pilot fuel injectors during at least part of the first engine cruise operation with different fuel characteristics from fuel delivered to one or both pilot and main fuel injectors the second engine cruise operation range.
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Patent Agency Ranking
公开(公告)号：US20230228232A1
公开(公告)日：2023-07-20
申请号：US18123091
申请日：2023-03-17
Applicant: ROLLS-ROYCE plc
Inventor： Craig W BEMMENT , Pascal DUNNING
IPC: F02K3/06
CPC classification number: F02K3/06 , F05D2220/36 , F05D2200/14 , F05D2260/4031 , F05D2240/307 , F05D2220/323
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.
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公开(公告)号：US20220389875A1
公开(公告)日：2022-12-08
申请号：US17889158
申请日：2022-08-16
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：US20220099035A1
公开(公告)日：2022-03-31
申请号：US17466086
申请日：2021-09-03
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 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.
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公开(公告)号：US20210310407A1
公开(公告)日：2021-10-07
申请号：US17345588
申请日：2021-06-11
Applicant: ROLLS-ROYCE PLC
Inventor： Craig W BEMMENT , Pascal DUNNING
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.
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公开(公告)号：US20240301828A1
公开(公告)日：2024-09-12
申请号：US18663558
申请日：2024-05-14
Applicant: ROLLS-ROYCE plc
Inventor： Craig W BEMMENT , Benjamin J. KEELER , Christopher P. MADDEN , Andrea MINELLI , Peter SWANN , Martin K. YATES
CPC classification number: F02C7/14 , F02C7/06 , F02C7/10 , F02C7/224 , F02C7/36 , F05D2220/323 , F05D2260/205 , F05D2260/213 , F05D2260/98
Abstract: A gas turbine engine includes an engine core with a combustor; a turbine including turbine blades; a compressor as a source of cooling air for the turbine blades; and an inducer to accelerate and direct the cooling air onto the turbine blades, and a modulating valve to allow or block cooling air flow into a subset of airflow passageways of the inducer; and a fuel management system to provide fuel to the combustor. The fuel management system includes a primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger through which oil and the fuel flow, to transfer heat between the oil and the fuel. A method of operating the engine includes using the modulating valve to adjust the cooling air flow based on turbine inlet temperature; and transferring 200-600 kJ/m3 of heat to the fuel from the oil in the primary fuel-oil heat exchanger at cruise conditions.
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