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公开(公告)号：US11898518B2
公开(公告)日：2024-02-13
申请号：US17892549
申请日：2022-08-22
IPC分类号： F02K3/075 , F02K3/06 , F02C7/36 , F02C3/13 , F02C3/107 , F02C7/20 , F01D21/04 , B64D29/02 , B64D27/12 , B64D33/04
CPC分类号： F02K3/075 , B64D27/12 , B64D29/02 , F01D21/045 , F02C3/107 , F02C3/13 , F02C7/20 , F02C7/36 , F02K3/06 , B64D33/04 , F05D2260/40 , F05D2300/433 , F05D2300/601 , F05D2300/6033
摘要： The present disclosure is directed to a gas turbine engine defining a radial direction, a longitudinal direction, and a circumferential direction, an upstream end and a downstream end along the longitudinal direction, and an axial centerline extended along the longitudinal direction. The gas turbine engine includes a fan assembly including a plurality of fan blades rotatably coupled to a fan rotor in which the fan blades define a maximum fan diameter and a fan pressure ratio. The gas turbine engine further includes a low pressure (LP) turbine defining a core flowpath therethrough generally along the longitudinal direction. The core flowpath defines a maximum outer flowpath diameter relative to the axial centerline. The gas turbine engine defines a fan to turbine diameter ratio of the maximum fan diameter to the maximum outer flowpath diameter. The fan to turbine diameter ratio over the fan pressure ratio is approximately 0.90 or greater.
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公开(公告)号：US11753939B2
公开(公告)日：2023-09-12
申请号：US16280531
申请日：2019-02-20
发明人： Alan Roy Stuart , Tsuguji Nakano , Thomas Ory Moniz , Andrew Breeze-Stringfellow , Richard Schmidt
IPC分类号： F01D5/06
CPC分类号： F01D5/06 , F05D2220/32
摘要： A gas turbine engine is provided including a turbine section including a turbine having a plurality of first speed turbine rotor blades; a compressor section including a compressor having a plurality of first speed compressor rotor blades and a plurality of second speed compressor rotor blades; a gearbox; and a first spool rotatable by the plurality of first speed turbine rotor blades, the first spool coupled to the plurality of first speed compressor rotor blades for driving the plurality of first speed compressor rotor blades in a first direction and to the plurality of second speed compressor rotor blades across the gearbox for driving the plurality of second speed compressor rotor blades in a second direction, opposite the first direction.
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公开(公告)号：US20220403799A1
公开(公告)日：2022-12-22
申请号：US17892549
申请日：2022-08-22
IPC分类号： F02K3/075 , F02K3/06 , F02C7/36 , F02C3/13 , F02C3/107 , F02C7/20 , F01D21/04 , B64D29/02 , B64D27/12
摘要： The present disclosure is directed to a gas turbine engine defining a radial direction, a longitudinal direction, and a circumferential direction, an upstream end and a downstream end along the longitudinal direction, and an axial centerline extended along the longitudinal direction. The gas turbine engine includes a fan assembly including a plurality of fan blades rotatably coupled to a fan rotor in which the fan blades define a maximum fan diameter and a fan pressure ratio. The gas turbine engine further includes a low pressure (LP) turbine defining a core flowpath therethrough generally along the longitudinal direction. The core flowpath defines a maximum outer flowpath diameter relative to the axial centerline. The gas turbine engine defines a fan to turbine diameter ratio of the maximum fan diameter to the maximum outer flowpath diameter. The fan to turbine diameter ratio over the fan pressure ratio is approximately 0.90 or greater.
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公开(公告)号：US11421627B2
公开(公告)日：2022-08-23
申请号：US15439167
申请日：2017-02-22
IPC分类号： F02K3/075 , F02K3/06 , F02C7/36 , F02C3/13 , F02C3/107 , F02C7/20 , F01D21/04 , B64D29/02 , B64D27/12 , B64D33/04
摘要： The present disclosure is directed to a gas turbine engine defining a radial direction, a longitudinal direction, and a circumferential direction, an upstream end and a downstream end along the longitudinal direction, and an axial centerline extended along the longitudinal direction. The gas turbine engine includes a fan assembly including a plurality of fan blades rotatably coupled to a fan rotor in which the fan blades define a maximum fan diameter and a fan pressure ratio. The gas turbine engine further includes a low pressure (LP) turbine defining a core flowpath therethrough generally along the longitudinal direction. The core flowpath defines a maximum outer flowpath diameter relative to the axial centerline. The gas turbine engine defines a fan to turbine diameter ratio of the maximum fan diameter to the maximum outer flowpath diameter. The fan to turbine diameter ratio over the fan pressure ratio is approximately 0.90 or greater.
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公开(公告)号：US11365629B1
公开(公告)日：2022-06-21
申请号：US17230826
申请日：2021-04-14
摘要： A turbine assembly including a first rotor assembly with a rotatable outer drum from which one or more stages of a plurality of outer drum airfoils is extended radially inward is provided. An outer casing surrounds the outer drum of the first rotor assembly. A seal assembly is coupled to the outer casing and positioned radially outward from an upstream-most stage of the plurality of outer drum airfoils. The seal assembly is positioned in axial alignment with the upstream-most stage of the plurality of outer drum airfoils. The seal assembly separates a first plenum from a second plenum. The second plenum is formed axially aft of the first plenum and is formed by the seal assembly, the outer casing, and the outer drum of the first rotor assembly. The first plenum is positioned radially outward from the upstream-most stage of the plurality of outer drum airfoils.
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公开(公告)号：US11053797B2
公开(公告)日：2021-07-06
申请号：US15412216
申请日：2017-01-23
IPC分类号： F01D3/00 , F01D5/02 , F01D25/16 , F02C7/28 , F01D11/00 , F01D11/08 , F01D25/24 , F04D25/04 , F04D29/053 , F04D29/32
摘要： The present disclosure is directed to a rotor thrust balanced turbine engine 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. The gas turbine engine defines a radial direction, a longitudinal direction, and a circumferential direction, an upstream end and a downstream end along the longitudinal direction, and an axial centerline extended along the longitudinal direction. The gas turbine engine includes a turbine rotor assembly and a turbine frame. The turbine rotor assembly defines a first flowpath radius and a second flowpath radius each extended from the axial centerline. The first flowpath radius is disposed at the upstream end of the turbine rotor assembly, and wherein the second flowpath radius is disposed at the downstream end of the turbine rotor assembly. The turbine frame and the turbine rotor assembly together define a seal interface radius inward of the turbine rotor assembly along the radial direction and concentric to the axial centerline, and wherein the turbine rotor assembly defines a ratio of the first flowpath radius to the seal interface radius less than or equal to approximately 1.79.
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公开(公告)号：US10823064B2
公开(公告)日：2020-11-03
申请号：US15286621
申请日：2016-10-06
摘要： A gas turbine engine includes a low pressure compressor and a low pressure turbine, with a low pressure shaft coupled to the low pressure turbine through a turbine extension and coupled to the low pressure compressor through a compressor extension. A forward bearing assembly supports the low pressure shaft within a compressor section and at a location aft of the compressor extension and an aft LP bearing assembly supports the low pressure shaft within a turbine section at a location aft of the turbine extension.
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公开(公告)号：US10738791B2
公开(公告)日：2020-08-11
申请号：US14970968
申请日：2015-12-16
IPC分类号： F04D27/00 , F01D11/24 , F01D25/12 , F02C7/12 , F02C6/08 , F04D29/32 , F04D29/52 , F04D29/54 , F04D29/58
摘要： In one embodiment, a compressor clearance control system is provided. The compressor clearance control system includes a cooling air source, a compressor casing, and a cooling air flowpath defined through the compressor casing. The flowpath includes an inlet configured to receive cooling air from the cooling air source and an outlet configured to exhaust the cooling air to the cooling air source such that the flowpath defines a closed loop within the compressor.
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公开(公告)号：US10683806B2
公开(公告)日：2020-06-16
申请号：US15398820
申请日：2017-01-05
摘要： A gas turbine engine defines a radial direction and an axial centerline. The gas turbine engine includes a core turbine engine that defines a core inlet. The core inlet is oriented with respect to the axial centerline and positioned along the radial direction such that the area available to capture foreign object debris is minimized. In one aspect, the gas turbine engine defines a capture ratio less than about 35%, wherein the capture ratio is a ratio of an area between a splitter radius and a tangency radius to an area encompassed by the splitter radius. The splitter radius is defined as a radial distance between the axial centerline and an outer lip of a splitter of the core turbine engine. The tangency radius is defined as a radial distance between the axial centerline and a tangency point, which can be defined at an inner lip of the core inlet.
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公开(公告)号：US10677264B2
公开(公告)日：2020-06-09
申请号：US15293368
申请日：2016-10-14
发明人： Thomas Ory Moniz , Daniel Robert Dwyer , Mark John Laricchiuta , Jeffrey Donald Clements , Brandon Wayne Miller
摘要： A gas turbine engine includes a fan section having a single-stage fan, a core turbine engine, and a nacelle at least partially surrounding the fan of the fan section and the core turbine engine. The gas turbine engine also includes an outlet guide vane extending between a casing of the core turbine engine and the nacelle. The gas turbine engine is configured to define an acoustic ratio greater than or equal to 2.3, with the acoustic ratio being a ratio of an axial spacing between the trailing edge of the fan blade and the leading edge of the outlet guide vane at a radial location seventy-five percent (75%) along the span of the fan blade to the axial width of the fan blade also at the radial location seventy-five percent (75%) along the span of the fan blade.
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