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公开(公告)号：EP3526457A2
公开(公告)日：2019-08-21
申请号：EP17889699.9
申请日：2017-11-03
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公开(公告)号：EP4365428A1
公开(公告)日：2024-05-08
申请号：EP23202077.6
申请日：2023-10-06
发明人： NIERGARTH, Daniel Alan , CLEMENTS, Jeffrey Donald , SPRUILL, Jeffrey , OSGOOD, Daniel Endecott , KRAMMER, Erich Alois , MACDONALD, Matthew Kenneth , SCHIMMELS, Scott Alan
摘要： A gas turbine engine (100) is provided. The gas turbine engine includes a turbomachine (120) including a compressor section, a combustion section (130), and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor (128) defining a high pressure compressor exit area (AHPCExit) in square inches; wherein the gas turbine engine (100) defines a redline exhaust gas temperature (EGT) in degrees Celsius, a total sea level static thrust output (FnTotal) in pounds, and a corrected specific thrust, wherein the corrected specific thrust is greater than or equal to 42 and less than or equal to 90, the corrected specific determined as follows: FnTotal x EGT / (AHPCExit2 x 1000).
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公开(公告)号：EP3456916A1
公开(公告)日：2019-03-20
申请号：EP18192783.1
申请日：2018-09-05
发明人： VAN DER MERWE, Gert Johannes , ZATORSKI, Darek Tomasz , WESLING, Richard Alan , CLEMENTS, Jeffrey Donald
摘要： The present disclosure is directed to a gas turbine engine 10 including a torque frame 101. The torque frame 101 includes an inner shroud 112 defined circumferentially around the axial centerline 12, an outer shroud 114 surrounding the inner shroud 112 and defined circumferentially around the axial centerline 12, and a structural member 116 extended along the radial direction and coupled to the inner shroud 112 and the outer shroud 114. The torque frame 101 is configured to rotate around the axial centerline 12.
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公开(公告)号：EP3464823A1
公开(公告)日：2019-04-10
申请号：EP17733151.9
申请日：2017-04-19
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公开(公告)号：EP3460182A1
公开(公告)日：2019-03-27
申请号：EP18193679.0
申请日：2018-09-11
摘要： The present disclosure is directed to a gas turbine engine 10 including a turbine section 90 including a first rotating component 110 interdigitated along a longitudinal direction with a second rotating component 120. The first rotating component 110 and the second rotating component 120 are each coupled to a speed reduction assembly 45 in counter-rotating arrangement. The first rotating component 110 comprising an outer shroud 114 and a plurality of outer shroud airfoils 118 extended inward along a radial direction from the outer shroud (114). A connecting member 116 couples the outer shroud 114 to a radially extended first rotor 113. The second rotating component 120 comprising an inner shroud (112) and a plurality of inner shroud airfoils 119 extended outward along the radial direction from the inner shroud 112, the plurality of inner shroud airfoils 119 in alternating arrangement along the longitudinal direction with the plurality of outer shroud airfoils 118. The gas turbine engine 10 defines a radius per unit thrust defined by a maximum radius (RR) at the turbine section 90 over a maximum thrust output between approximately .0004 to approximately .0010 inches per pound thrust.
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公开(公告)号：EP3428424A1
公开(公告)日：2019-01-16
申请号：EP18182064.8
申请日：2018-07-05
发明人： CLEMENTS, Jeffrey Donald , ZATORSKI, Darek Tomasz , MILLER, Brandon Wayne , DICKMAN, Joseph Robert
摘要： The present disclosure is directed to a gas turbine engine 10 defining a longitudinal direction, a radial direction, and a circumferential direction. The gas turbine engine 10 includes a power turbine 100 including a first turbine rotor assembly 110 interdigitated with a second turbine rotor assembly 120 along the longitudinal direction; a gear assembly 300 coupled to the first turbine rotor assembly 110 and the second turbine rotor assembly 120, wherein the gear assembly 300 includes a first input interface 310 coupled to the first turbine rotor assembly 110, a second input interface 320 coupled to the second turbine rotor assembly 120, and one or more third gears 330 coupled to the first input interface 310 and the second input interface 320 therebetween; and a first output shaft 121 and a second output shaft 122, wherein each of the first output shaft 121 and the second output shaft 122 are configured to couple to an electrical load device.
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公开(公告)号：EP3354889A1
公开(公告)日：2018-08-01
申请号：EP18153271.4
申请日：2018-01-24
发明人： MILLER, Brandon Wayne , POLAKOWSKI, Matthew Ryan , MARRINAN, Patrick Michael , KIRK, Joel Francis , VONDRELL, Randy M. , CLEMENTS, Jeffrey Donald
CPC分类号： F02K3/06 , F01D25/26 , F02C3/04 , F02C7/18 , F02C7/36 , F05D2220/36 , F05D2300/20 , Y02T50/672
摘要： A turbofan engine 10 is provided including a fan 38 having a plurality of rotatable fan blades 40 and defining a fan pressure ratio during operation of the turbofan engine 10. The turbofan engine 10 also includes a turbomachine 16 operably coupled to the fan 38 for driving the fan 38, the turbomachine 16 including a compressor section, a combustion section 26, and a turbine section in serial flow order and together defining a core air flowpath 37. The turbofan 38 also includes an outer nacelle 50 at least partially surrounding the fan 38 and the turbomachine 16, the outer nacelle 50 defining a bypass passage 56 with the turbomachine 16. A bypass ratio of an amount of airflow through the bypass passage 56 to an amount of airflow through the core air flowpath 37 during operation of the turbofan 38 is less than or equal to about 11 and wherein the fan pressure ratio is less than or equal to about 1.5.
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公开(公告)号：EP3351725A1
公开(公告)日：2018-07-25
申请号：EP18151720.2
申请日：2018-01-15
CPC分类号： F01D5/027 , F01D3/00 , F01D11/00 , F01D11/08 , F01D25/16 , F01D25/24 , F04D25/045 , F04D29/053 , F04D29/325 , F05D2220/32 , F05D2270/051
摘要： The present disclosure is directed to a rotor thrust balanced turbine engine (10) that may increase engine performance and efficiency while managing thrust mismatch or imbalance in a low pressure (LP) spool between a fan assembly and a turbine rotor assembly (20). The gas turbine engine (10) defines a radial direction, a longitudinal direction, and a circumferential direction, an upstream end (99) and a downstream end (98) along the longitudinal direction, and an axial centerline (15) extended along the longitudinal direction. The gas turbine engine (10) includes a turbine rotor assembly (20) and a turbine frame (30). The turbine rotor assembly (20) defines a first flowpath radius (11) and a second flowpath radius (12) each extended from the axial centerline (15). The first flowpath radius (11) is disposed at the upstream end (99) of the turbine rotor assembly (20), and wherein the second flowpath radius (12) is disposed at the downstream end (98) of the turbine rotor assembly (20). The turbine frame (30) and the turbine rotor assembly (20) together define a seal interface radius (25) inward of the turbine rotor assembly (20) along the radial direction and concentric to the axial centerline (15), and wherein the turbine rotor assembly (20) defines a ratio of the first flowpath radius (11) to the seal interface radius (25) less than or equal to approximately 1.79.
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