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公开(公告)号：US10823013B2
公开(公告)日：2020-11-03
申请号：US15282547
申请日：2016-09-30
发明人： Daniel Waslo
摘要： A core engine includes a first tierod and a compressor rotor assembly including a plurality of compressor rotor disks arranged in a face to face orientation and spaced along the first tierod. The core engine includes a second tierod and a turbine rotor assembly including a plurality of turbine rotor disks arranged in a face to face orientation and spaced along the second tierod. The compressor rotor assembly is aft of the turbine rotor assembly.
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公开(公告)号：US20190154055A1
公开(公告)日：2019-05-23
申请号：US15819215
申请日：2017-11-21
摘要： A turbofan engine comprising a compressor rotor, a plurality of main blades, the main blades are circumferentially mounted on the compressor rotor; a plurality of spinner blades, the spinner blades are circumferentially mounted on the compressor rotor, the spinner blades are located forward to the leading edge of the main blades; each spinner blade is circumferentially offset to a corresponding main blade.
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公开(公告)号：US10287977B2
公开(公告)日：2019-05-14
申请号：US15044211
申请日：2016-02-16
发明人： Mark Gregory Wotzak , Daniel Waslo
摘要： An in-line propeller gearbox of a turboprop gas turbine engine includes an epicyclic gearing arrangement that has a sun gear, a ring gear and at least one planet gear disposed between and meshing with both the sun gear and the ring gear. The propeller gearbox includes a disk brake that can be operated to slow down or stop altogether the rotation of the propeller. The disk brake has an axially extending shaft having at one end a disk and at the opposite end a gear having teeth that engage with one of the gears in the epicyclic gearing arrangement. The disk brake includes a hydraulic caliper or an electric caliper that can be actuated to slow down rotation of the disk to an eventual full stop and to hold the disk at full stop to thereby stop rotation of the propeller.
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公开(公告)号：US20190061965A1
公开(公告)日：2019-02-28
申请号：US15682834
申请日：2017-08-22
发明人： Kenneth Arthur Gould , Martin Miles D'Angelo , Daniel Waslo , Brian Lewis Devendorf , Mark Gregory Wotzak
摘要： An apparatus and method for mounting a turbine engine to an aircraft can include an engine core for the turbine engine including a compressor section, a combustor section, and a turbine section in flow arrangement. At least one strut couples to the engine core about a single mount plane. A structural wall at least partially defining a mainstream flow path couples to the at least one strut and passes through the compressor section and the turbine section.
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公开(公告)号：US10683809B2
公开(公告)日：2020-06-16
申请号：US15150628
申请日：2016-05-10
IPC分类号： F02C9/18 , F01D5/04 , F02C3/08 , F02C3/05 , F02C6/08 , F01D5/02 , F01D9/04 , F01D25/12 , F02C3/04
摘要： The present disclosure is directed to a system for bleeding air from a compressed gas path of a gas turbine engine. The system includes an impeller positioned at a downstream end of a compressor in the gas turbine engine. The impeller includes an impeller hub, an impeller arm coupled to the impeller hub, and a plurality of circumferentially spaced apart impeller vanes extending radially outwardly from the impeller arm. The impeller arm defines an impeller arm aperture extending therethrough. A vortex spoiler is positioned radially inwardly from the impeller arm and defines a vortex spoiler passage extending radially therethrough. Bleed air flows from the compressed gas path radially inwardly through both the impeller arm aperture and the vortex spoiler passage.
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公开(公告)号：US10670037B2
公开(公告)日：2020-06-02
申请号：US15819215
申请日：2017-11-21
摘要： A turbofan engine includes a compressor rotor, a plurality of main blades circumferentially mounted on the compressor rotor, and a plurality of spinner blades circumferentially mounted on the compressor rotor. Each of the spinner blades is located forward of a leading edge of a corresponding one of the main blades. Each of the spinner blades is circumferentially offset to the corresponding one of the main blades.
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7.发明申请公开(公告)号：US20180195395A1
公开(公告)日：2018-07-12
申请号：US15401461
申请日：2017-01-09
发明人： Brian David Johnson , Daniel Waslo
CPC分类号： F01D5/3015 , F01D5/082 , F05D2260/205 , Y02T50/671 , Y02T50/676
摘要： The turbine includes a first turbine rotor disc, a second turbine rotor disc, a first source of cooling fluid, and a first cooling fluid conduit. The first turbine rotor disc has a first outer circumference and a first plurality of blades spaced along the first outer circumference. The second turbine rotor disc has a second outer circumference and a second plurality of blades spaced along the second outer circumference. The first cooling fluid conduit is configured to channel a first flow of cooling fluid from the first source of cooling fluid through the first turbine rotor disc to a blade of the second plurality of blades.
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公开(公告)号：US20170234217A1
公开(公告)日：2017-08-17
申请号：US15044211
申请日：2016-02-16
发明人： Mark Gregory Wotzak , Daniel Waslo
CPC分类号： F02C3/107 , B64D27/10 , B64D35/00 , F02C3/04 , F02C6/20 , F02C6/206 , F02C7/36 , F05D2220/323 , F05D2240/60 , F05D2260/40311 , F05D2260/902 , F16H1/28 , F16H57/10
摘要： An in-line propeller gearbox of a turboprop gas turbine engine includes an epicyclic gearing arrangement that has a sun gear, a ring gear and at least one planet gear disposed between and meshing with both the sun gear and the ring gear. The propeller gearbox includes a disk brake that can be operated to slow down or stop altogether the rotation of the propeller. The disk brake has an axially extending shaft having at one end a disk and at the opposite end a gear having teeth that engage with one of the gears in the epicyclic gearing arrangement. The disk brake includes an hydraulic caliper or an electric caliper that can be actuated to slow down rotation of the disk to an eventual full stop and to hold the disk at full stop to thereby stop rotation of the propeller.
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公开(公告)号：US11143041B2
公开(公告)日：2021-10-12
申请号：US15401461
申请日：2017-01-09
发明人： Brian David Johnson , Daniel Waslo
摘要： The turbine includes a first turbine rotor disc, a second turbine rotor disc, a first source of cooling fluid, and a first cooling fluid conduit. The first turbine rotor disc has a first outer circumference and a first plurality of blades spaced along the first outer circumference. The second turbine rotor disc has a second outer circumference and a second plurality of blades spaced along the second outer circumference. The first cooling fluid conduit is configured to channel a first flow of cooling fluid from the first source of cooling fluid through the first turbine rotor disc to a blade of the second plurality of blades.
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公开(公告)号：US20180230805A1
公开(公告)日：2018-08-16
申请号：US15434658
申请日：2017-02-16
CPC分类号： F01D5/08 , F01D1/26 , F01D5/03 , F01D5/06 , F01D5/186 , F01D5/187 , F01D11/08 , F02C3/04 , F05D2220/32 , F05D2240/14 , F05D2240/24 , F05D2240/55 , F05D2250/185 , F05D2260/202
摘要： The present disclosure is directed to a gas turbine engine defining a radial direction, a circumferential direction, an axial centerline along a longitudinal direction, and an upstream end and a downstream end along the longitudinal direction. The gas turbine engine defines a core flowpath extended generally along the longitudinal direction. The gas turbine engine includes a first turbine rotor. The first turbine rotor includes an annular outer band disposed outward of the core flowpath along the radial direction. The first turbine rotor further includes a plurality of airfoils coupled to an inner diameter of the outer band in which the plurality of airfoils are extended generally inward along the radial direction. The outer band defines a plurality of airfoil cooling passages in which the plurality of airfoil cooling passages are extended at least partially in the radial direction in fluid communication with the plurality of airfoils.
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