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公开(公告)号:EP4306790A1
公开(公告)日:2024-01-17
申请号:EP23180891.6
申请日:2023-06-22
申请人: Rolls-Royce plc
摘要: A novel configuration for an axial flow gas turbine engine for an aircraft has an engine core having a core length and comprising a first turbine, an axial compressor, and a drive connecting the first turbine to the axial compressor, the engine core further comprising a second turbine, and a fan shaft connecting the second turbine to a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a fan hub, the fan having a tip radius in the range from 90mm to 225mm and wherein the ratio of fan tip radius to an engine length is of the order 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 in the range 0.015 to 0.083. The fan may be a multistage fan configured to permit an electric motor to be located within the fan hub diameter.
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公开(公告)号:EP4306781A1
公开(公告)日:2024-01-17
申请号:EP23180888.2
申请日:2023-06-22
申请人: Rolls-Royce plc
摘要: A novel configuration for an axial flow gas turbine engine for an aircraft has an engine core having a core length and comprising a first turbine, an axial compressor, and a drive connecting the first turbine to the axial compressor, the engine core further comprising a second turbine, and a fan shaft connecting the second turbine to a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a fan hub, the fan having a tip radius in the range from 90mm to 225mm and wherein the ratio of fan tip radius to an engine length is of the order 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 in the range 0.015 to 0.083. The fan may be a multistage fan configured to permit an electric motor to be located within the fan hub diameter.
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公开(公告)号:EP3144484B1
公开(公告)日:2018-01-17
申请号:EP16182194.7
申请日:2016-08-01
申请人: Rolls-Royce plc
发明人: Smith, William , Shorney, Andrew , Mawhinney, Paul
IPC分类号: F01D21/04
CPC分类号: F01D11/02 , F01D11/001 , F01D21/045 , F05D2220/32 , F05D2240/12 , F05D2240/24 , F05D2270/09
摘要: Described is a gas turbine engine, comprising: a first chamber (330) and a second chamber (332) separated by a partition wall (322); wherein the partition wall (322) includes a stationary element and a rotating element separated by a seal and either of the stationary element and the rotating element of the partition wall (322) is segmented by a levered joint (346), wherein the levered joint (346) includes a lever (348) having a trigger plate (350), a fixture portion (352) and a fulcrum portion (354); and, the other of the stationary element and rotating element includes a hammer 360 which is axially aligned and separated from the trigger plate (350) in a first position, and forcibly contacts the hammer (360) portion in a second position so as to create a moment on the lever (348) via the trigger plate (350), the moment forcing the levered joint apart.
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公开(公告)号:EP4306782A1
公开(公告)日:2024-01-17
申请号:EP23180889.0
申请日:2023-06-22
申请人: Rolls-Royce plc
摘要: A novel configuration for an axial flow gas turbine engine for an aircraft has an engine core having a core length and comprising a first turbine, an axial compressor, and a drive connecting the first turbine to the axial compressor, the engine core further comprising a second turbine, and a fan shaft connecting the second turbine to a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a fan hub, the fan having a tip radius in the range from 90mm to 225mm and wherein the ratio of fan tip radius to an engine length is of the order 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 in the range 0.015 to 0.083. The fan may be a multistage fan configured to permit an electric motor to be located within the fan hub diameter.
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公开(公告)号:EP3144484A1
公开(公告)日:2017-03-22
申请号:EP16182194.7
申请日:2016-08-01
申请人: Rolls-Royce plc
发明人: Smith, William , Shorney, Andrew , Mawhinney, Paul
IPC分类号: F01D21/04
CPC分类号: F01D11/02 , F01D11/001 , F01D21/045 , F05D2220/32 , F05D2240/12 , F05D2240/24 , F05D2270/09
摘要: Described is a gas turbine engine, comprising: a first chamber (330) and a second chamber (332) separated by a partition wall (322); wherein the partition wall (322) includes a stationary element and a rotating element separated by a seal and either of the stationary element and the rotating element of the partition wall (322) is segmented by a levered joint (346), wherein the levered joint (346) includes a lever (348) having a trigger plate (350), a fixture portion (352) and a fulcrum portion (354); and, the other of the stationary element and rotating element includes a hammer 360 which is axially aligned and separated from the trigger plate (350) in a first position, and forcibly contacts the hammer (360) portion in a second position so as to create a moment on the lever (348) via the trigger plate (350), the moment forcing the levered joint apart.
摘要翻译: 描述了一种燃气涡轮发动机,包括:由分隔壁(322)分隔的第一室(330)和第二室(332); 其中所述分隔壁(322)包括固定元件和由密封件隔开的旋转元件,并且所述固定元件和所述分隔壁(322)的所述旋转元件中的任一个由所述杠杆接头(346)分段,其中所述杠杆接头 (346)包括具有触发板(350),固定部分(352)和支点部分(354)的杠杆(348)。 并且固定元件和旋转元件中的另一个包括在第一位置轴向对准并与触发板(350)分离的锤子360,并且在第二位置强制地接触锤子(360)部分,以便产生 通过触发板(350)在杠杆(348)上的力矩,迫使杠杆接头分开。
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