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    AIRCRAFT PROPULSION SYSTEM GEARTRAIN
    3.
    发明公开
    AIRCRAFT PROPULSION SYSTEM GEARTRAIN 审中-公开

    公开(公告)号：EP4365429A1

    公开(公告)日：2024-05-08

    申请号：EP23208062.2

    申请日：2023-11-06

    申请人： RTX Corporation

    发明人： HANRAHAN, Paul R. ,  MYLREA, Benjamin T. ,  STACHOWIAK, David A.

    IPC分类号： F02C7/36 ,  F02C3/107 ,  F02K3/072

    CPC分类号： F02C3/107 ,  F02C7/36 ,  F05D2260/4031120130101 ,  F05D2260/90220130101 ,  F05D2260/402320130101 ,  F02K3/072

    摘要： An aircraft assembly includes a geartrain (72) with a first gear system (94) and a second gear system (96). The first gear system (94) includes a first sun gear (98), a first ring gear (100), a plurality of first intermediate gears (102) and a first carrier (104). The first sun gear (98) and the first ring gear (100) are each rotatable about an axis (28,40,106). The first intermediate gears (102) are between and meshed with the first sun gear (98) and the first ring gear (100). Each of the first intermediate gears (102) is rotatably mounted to the first carrier (104). The second gear system (96) is interconnected with the first gear system (94). The second gear system (96) includes a second sun gear (108), a second ring gear (110), a plurality of second intermediate gears (112) and a second carrier (114). The second intermediate gears (112) are between and meshed with the second sun gear (108) and the second ring gear (110). Each of the second intermediate gears (112) is rotatably mounted to the second carrier (114).

    GAS TURBINE ENGINE WITH HIGHER LOW SPOOL TORQUE-TO-THRUST RATIO
    8.
    发明公开
    GAS TURBINE ENGINE WITH HIGHER LOW SPOOL TORQUE-TO-THRUST RATIO 审中-公开

    公开(公告)号：EP4123142A1

    公开(公告)日：2023-01-25

    申请号：EP22185719.6

    申请日：2022-07-19

    申请人： Raytheon Technologies Corporation

    发明人： PIXTON, Stephen G. ,  HASEL, Karl L.

    IPC分类号： F02C3/107 ,  F02C7/36 ,  F02K3/06

    摘要： A gas turbine engine (20) includes a fan drive turbine driving a low pressure compressor (44), and driving a gear reduction to in turn drive a fan rotor (42) at a speed slower than the fan drive turbine. The turbine section (28) further includes a high pressure turbine (54) driving the high pressure compressor (52). The fan drive turbine and low pressure compressor (44) are connected by a shaft (40) and the fan drive turbine, the shaft (40) and the low pressure compressor (44) define a low pressure spool. The gas turbine engine (20) is rated to provide an amount of thrust at maximum takeoff, and a low spool thrust ratio defined as a ratio of a torque on the low pressure spool at maximum takeoff in ft-lbs and the maximum takeoff thrust being defined in lbf, with the low spool torque ratio being greater than or equal to 0.70 ft-lb/lbf (0.21 Nm/N), and less than or equal to 1.2 ft-lb/lbf (0.37 Nm/N).

    GAS TURBINE ENGINE WITH HIGH LOW SPOOL POWER EXTRACTION RATIO
    9.
    发明公开
    GAS TURBINE ENGINE WITH HIGH LOW SPOOL POWER EXTRACTION RATIO 审中-公开

    公开(公告)号：EP4123141A1

    公开(公告)日：2023-01-25

    申请号：EP22185711.3

    申请日：2022-07-19

    申请人： Raytheon Technologies Corporation

    发明人： PIXTON, Stephen G. ,  HASEL, Karl L.

    IPC分类号： F02C3/107 ,  F02C7/36 ,  F02K3/06

    摘要： A gear reduction (48) drives a fan rotor (42) at a speed slower than a fan drive turbine. The turbine section further includes a high pressure turbine (54) driving a high pressure compressor (52). The fan drive turbine and low pressure compressor (44) are connected by a shaft (40). The low pressure turbine, the shaft (40) and the low pressure compressor (44) define a low pressure spool (30). The low pressure spool (30) has a torque at maximum takeoff defined in ft-lbs and also has a low pressure spool power defined in horsepower and at maximum takeoff. A ratio of the low pressure spool torque to the low pressure spool power is defined, and the ratio of the low pressure spool torque to the low pressure spool power is greater than or equal to 0.6 ft-lb/hp (0.0011 Nm/Nm.s -1 ) and less than or equal to 1.2 ft-lb/hp (0.022 Nm/Nm.s -1 ).

    GAS TURBINE ENGINES HAVING CRYOGENIC FUEL SYSTEMS
    10.
    发明公开
    GAS TURBINE ENGINES HAVING CRYOGENIC FUEL SYSTEMS 审中-公开

    公开(公告)号：EP4019752A1

    公开(公告)日：2022-06-29

    申请号：EP21216172.3

    申请日：2021-12-20

    申请人： Raytheon Technologies Corporation

    发明人： JOHNSON, Steven B. ,  MURPHY, Andrew J.

    IPC分类号： F02C3/20 ,  F02C3/22 ,  F02C6/20 ,  F02K3/06 ,  F02C3/107

    摘要： Turbine engine systems (200; 300; 400) and aircraft having such systems are described. The turbine engine systems (200; 300; 400) include a combustor (210; 306) arranged along a core flow path of the turbine engine (200; 300; 400), a drive shaft (220; 312) having at least a compressor section (206, 208; 304) and a turbine section (212, 214; 308) coupled thereto, a cryogenic fuel tank (224; 318; 404) configured to supply a fuel to the combustor (210; 306), and an expansion turbine (234; 326; 412) mechanically coupled to the drive shaft (220; 312), the expansion turbine (234; 326; 412) configured to receive fuel from the cryogenic fuel tank (224; 318; 404) and expand said fuel, wherein expansion of said fuel by the expansion turbine (234; 326; 412) drives rotation of the expansion turbine (234; 326; 412) to provide power input to the drive shaft (220; 312).