公开(公告)号：EP4390097A1

公开(公告)日：2024-06-26

申请号：EP23216529.0

申请日：2023-12-14

申请人： Rolls-Royce plc

发明人： Madden, Christopher ,  Bemment, Craig ,  Smith, Andrew ,  Swann, Peter

IPC分类号： F02C7/14 ,  F02C7/224 ,  F02C7/228

CPC分类号： F02C7/224 ,  F02C7/228 ,  F02C7/14 ,  F05D2260/21320130101

摘要： There is provided a method (400) of operating a gas turbine engine (10). The gas turbine engine (10) comprises a combustor (16). The combustor (16) comprises a combustion chamber (120) and a plurality of fuel spray nozzles (124) configured to inject fuel into the combustion chamber (120). The plurality of fuel spray nozzles (124) comprises a first subset (124A) of fuel spray nozzles (124) and a second subset (124B) of fuel spray nozzles (124). The combustor (16) is operable in a condition in which the first subset (124A) of fuel spray nozzles (124) are supplied with more fuel than the second subset (124B) of fuel spray nozzles (124). A ratio of the number of fuel spray nozzles (124) in the first subset (124A) of fuel spray nozzles (124) to the number of fuel spray nozzles (124) in the second subset (124B) of fuel spray nozzles (124) is in the range of 1:2 to 1:5. The method (200) comprises: providing (401) a fuel to the one or more fuel-oil heat exchangers (114); transferring (402) heat from oil to the fuel in the one or more fuel-oil heat exchangers (114); and providing (403) the fuel from the one or more fuel-oil heat exchangers (114) to the plurality of fuel spray nozzles (124). Transferring (402) heat from the oil to the fuel comprises transferring 200-600 kW/m3 of heat from oil to the fuel in the one or more fuel-oil heat exchangers (114) at cruise conditions before providing the fuel to the plurality of fuel spray nozzles (124). Also provided is a gas turbine engine (10) for an aircraft.

公开(公告)号：EP4261395A1

公开(公告)日：2023-10-18

申请号：EP23165284.3

申请日：2023-03-30

申请人： Rolls-Royce plc

发明人： Swann, Peter ,  Madden, Christopher ,  Bemment, Craig

IPC分类号： F02C3/20 ,  F02C7/22 ,  F02C9/40 ,  G06Q10/0631

摘要： The present application discloses a computer implemented method (4070) of determining a fuel allocation for an aircraft (1). The aircraft (1) comprises a first fuel source (302) adapted to contain a first fuel having a first fuel characteristic and a second fuel source (304) adapted to contain a second fuel having a second fuel characteristic, the second fuel characteristic being different from the first. The aircraft (1) further comprises one or more gas turbine engines (10) powered by fuel from the first and second fuel sources (302, 304). The gas turbine engines (10) each comprise a staged combustion system (64) having pilot fuel injectors (313) and main fuel injectors (314), the staged combustion system (64) being operable in a pilot-only range of operation and a pilot-and-main range of operation. The gas turbine engines (10) each comprise a fuel delivery regulator (306) arranged control delivery of fuel to the pilot and main fuel injectors (313, 314) from the first fuel source (302) and the second fuel source (304). The method (4070) comprises the following steps: obtaining (4072) a proposed mission description comprising a list of operating conditions for the gas turbine engines during the mission; obtaining (4074) nvPM impact parameters for the gas turbine engines (10), the impact parameters being associated with each operating condition of the proposed mission using compositions of fuel which include fuel from the first fuel source (302), fuel from the second fuel source (304), or a blend thereof; calculating (4076) an optimised set of one or more fuel characteristics for each flight condition of the proposed flight defined in the flight description based on the nvPM impact parameters; and determining (4078) a fuel allocation based on the optimised set of one or more fuel characteristics. Also disclosed is method (4080) of loading fuel onto an aircraft (1), a non-transitory computer readable medium, a fuel allocation determination system (5000), and an aircraft (1).

公开(公告)号：EP3855005A3

公开(公告)日：2021-10-20

申请号：EP21150296.8

申请日：2021-01-05

申请人： Rolls-Royce plc

发明人： Bemment, Craig ,  Dunning, Pascal

IPC分类号： F02C9/00 ,  F02K3/06 ,  B64D31/00

摘要： A gas turbine engine (10) comprises a fan (23), a compressor (14, 15), a low pressure turbine (19) and a high pressure turbine (17). The fan diameter is greater than 250 cm and less than 381 cm, and the gas turbine engine has a first thrust at sea level static conditions and a second thrust at end of runway conditions, and a thrust take-off ratio greater than 1.32, wherein the thrust take-off ratio is the ratio of the first thrust to the second thrust.

公开(公告)号：EP3808962A1

公开(公告)日：2021-04-21

申请号：EP20189729.5

申请日：2020-08-06

申请人： Rolls-Royce plc

发明人： Dunning, Pascal ,  Bemment, Craig

IPC分类号： F02K3/06

摘要： A gas turbine engine (10) for an aircraft comprising: an engine core (11) comprising a turbine system comprising one or more turbines (17, 19), a compressor system comprising one or more compressors (14,15), and a core shaft (26) connecting the turbine system to the compressor system, wherein a compressor exit temperature (T30) is defined as an average temperature of airflow at the exit of the highest pressure compressor of the compressor system at cruise conditions and a compressor exit pressure (P30) is defined as an average pressure of airflow at the exit of the highest pressure compressor of the compressor system at cruise conditions, the engine core (11) further comprising an annular splitter (70) at which flow is divided between a core flow (A) that flows through the engine core, and a bypass flow (B) that flows along a bypass duct (22), wherein stagnation streamlines (110) around the circumference of the engine (10), stagnating on a leading edge of the annular splitter (70), form a streamsurface (110) forming a radially outer boundary of a streamtube that contains all of the core flow (A); a fan (23) located upstream of the engine core (11), the fan comprising a plurality of fan blades (64) extending from a hub, each fan blade (64) having a leading edge (64a) and a trailing edge (64b), each fan blade (64) having a radially inner portion (65a) lying within the streamtube that contains the core flow (A), and wherein a fan root entry temperature (T20) is defined as an average temperature of airflow across the leading edge (64a) of the radially inner portion of each fan blade (64) at cruise conditions and wherein a fan root entry pressure (P20) is defined as an average pressure of airflow across the leading edge (64a) of the radially inner portion of each fan blade (64) at cruise conditions; and a nacelle (21) surrounding the engine core (11), the nacelle (21) defining the bypass duct (22) and a bypass exhaust nozzle (18). An overall pressure ratio is defined as the compressor exit pressure (P30) divided by the fan root entry pressure (P20). A bypass nozzle pressure ratio is defined as the nozzle pressure ratio of the bypass exhaust nozzle at cruise conditions. A core temperature rise is defined as the compressor exit temperature (T30) in Kelvin divided by the fan root entry temperature (T20) in Kelvin. A temperature-pressure ratio defined as: the core temperature rise the bypass nozzle pressure ratio is in a range between 1.52 and 1.8, and the overall pressure ratio is in a range between 42.5 and 70. A method of operating the gas turbine engine is also disclosed.

公开(公告)号：EP3741974A1

公开(公告)日：2020-11-25

申请号：EP20172015.8

申请日：2020-04-29

申请人： Rolls-Royce plc

发明人： Bemment, Craig

IPC分类号： F02C7/36 ,  F02K3/06

摘要： The disclosure relates to a gas turbine engine (10) for an aircraft, comprising: an engine core (11) comprising a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor, the engine core having an inlet (29) upstream of the compressor and an outlet (20) downstream of the turbine (19); a fan (23) located upstream of the engine core (11), the fan comprising a plurality of fan blades; a gearbox (30) that receives an input from the core shaft (26) and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft; and a nacelle (21) surrounding the engine core (11), the nacelle (21) defining a bypass duct (22) and a bypass exhaust nozzle (18), wherein the gas turbine engine is configured such that an axial Mach number at the engine core inlet multiplied by an axial Mach number of an exhaust airflow from the bypass exhaust nozzle (18) is within a range from around 0.30 to around 0.56 at maximum take-off conditions, where the axial Mach number at the engine core inlet is less than around 0.7 at maximum take-off conditions.

公开(公告)号：EP3450726A1

公开(公告)日：2019-03-06

申请号：EP18189079.9

申请日：2018-08-15

申请人： Rolls-Royce plc

发明人： Phelps, Benedict ,  Baralon, Stephane ,  Bemment, Craig ,  Wilson, Mark ,  Gonzalez-Gutierrez, Gabriel ,  Smith, Nigel ,  Barale, Marco ,  Johal, Kashmir

IPC分类号： F02C7/36 ,  F01D5/16

摘要： A gas turbine engine 10 is provided in a fan root to tip pressure ratio, defined as the ratio of the mean total pressure of the flow at the fan exit that subsequently flows through the engine core (P 102 ) to the mean total pressure of the flow at the fan exit that subsequently flows through the bypass duct (P 104 ), is no greater than a certain value. The gas turbine engine 10 may provide improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.