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公开(公告)号：US10718265B2
公开(公告)日：2020-07-21
申请号：US15605164
申请日：2017-05-25
摘要： The present disclosure is directed to a gas turbine engine defining a longitudinal direction, a radial direction extended from an axial centerline, and a circumferential direction. The gas turbine engine includes a compressor section, a combustion section, and a turbine section in serial flow arrangement along the longitudinal direction. The gas turbine engine includes a low speed turbine rotor comprising a hub extended along the longitudinal direction and radially within the combustion section; a high speed turbine rotor comprising a high pressure (HP) shaft coupling the high speed turbine rotor to a HP compressor in the compressor section; a first turbine bearing disposed radially between the hub of the low speed turbine rotor and the HP shaft. The HP shaft extends along the longitudinal direction and radially within the hub of the low speed turbine rotor. The first turbine bearing defines an outer air bearing along an outer diameter of the first turbine bearing and adjacent to the hub of the low speed turbine rotor. The first turbine bearing defines an inner air bearing along an inner diameter of the first turbine bearing and adjacent to the HP shaft.
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公开(公告)号：US10654577B2
公开(公告)日：2020-05-19
申请号：US15439122
申请日：2017-02-22
发明人： Brandon Wayne Miller , Thomas Ory Moniz , Monty Lee Shelton , Joel Francis Kirk , Jeffrey Donald Clements
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 low pressure (LP) turbine defining an outer flowpath. The outer flowpath defines a first outer flowpath radius at an upstream-most end of the LP turbine, a last outer flowpath radius disposed at a downstream-most end of the LP turbine, a middle outer flowpath radius disposed therebetween along the longitudinal direction. The middle outer flowpath radius is greater than the last outer flowpath radius.
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公开(公告)号：US10358942B2
公开(公告)日：2019-07-23
申请号：US15052939
申请日：2016-02-25
摘要： A counter-rotating shaft assembly of a gas turbine engine includes an outer shaft rotatable in a first direction about a virtual rotational axis, an inner shaft counter-rotatable about the virtual rotational axis in a second direction that is opposite to the first direction, a differential bearing rotatably connecting the two shafts, a centering spring connecting the inner shaft to the differential bearing, and a squeeze film damper between the differential bearing and the inner shaft.
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94.发明申请公开(公告)号：US20180340423A1
公开(公告)日：2018-11-29
申请号：US15605351
申请日：2017-05-25
CPC分类号： F01D25/22 , F01D5/082 , F01D5/087 , F02C7/06 , F05D2240/53 , F05D2250/44
摘要： The present disclosure is directed to a gas turbine engine defining a longitudinal direction, an axial centerline extended along the longitudinal direction, an upstream end and a downstream end opposite of the upstream end along the longitudinal direction, a radial direction, and a circumferential direction. The gas turbine engine includes a high speed turbine rotor coupled to a high pressure (HP) shaft and HP compressor, a low speed turbine rotor comprising an axially extended hub, and a first turbine bearing disposed radially between the low speed turbine rotor and the high speed turbine rotor. The high speed turbine rotor defines a turbine cooling conduit through the high speed turbine rotor. The low speed turbine rotor includes a rotating nozzle adjacent to the turbine cooling conduit. The first turbine bearing defines an outer air bearing and an inner air bearing. The first turbine bearing defines a stationary nozzle adjacent to the rotating nozzle of the first turbine rotor.
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公开(公告)号：US20180328287A1
公开(公告)日：2018-11-15
申请号：US15593748
申请日：2017-05-12
发明人： Thomas Ory Moniz , Alan Roy Stuart , James William Simunek , Jeffrey Donald Clements , Brandon Wayne Miller , Sridhar Adibhatla
CPC分类号： F02C7/36 , F02C3/067 , F02C7/22 , F02C9/20 , F04D29/563 , F05D2220/32 , F05D2270/304 , F05D2270/31
摘要： The present disclosure is directed to a method of control of a gas turbine engine comprising a fan section coupled to a low turbine together defining a low spool, an intermediate compressor coupled to an intermediate turbine together defining an intermediate spool, and a high compressor coupled to a high turbine together defining a high spool. The method includes providing an intermediate spool speed to low spool speed characteristic curve to a controller; providing a commanded power output to the controller; providing one or more of an environmental condition to the controller; determining, via the controller, a commanded fuel flow rate; determining, via the controller, a commanded intermediate compressor loading; and generating an actual power output of the engine, wherein the actual power output is one or more of an actual low spool speed, an actual intermediate spool speed, an actual high spool speed, and an actual engine pressure ratio.
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公开(公告)号：US20180320633A1
公开(公告)日：2018-11-08
申请号：US15439167
申请日：2017-02-22
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：US20180274365A1
公开(公告)日：2018-09-27
申请号：US15412157
申请日：2017-01-23
摘要： The present disclosure is directed to a gas turbine engine defining a longitudinal direction, a radial direction, and a circumferential direction, and wherein the gas turbine engine defines an upstream end and a downstream end along the longitudinal direction. The gas turbine engine includes a turbine section that includes a first rotating component and a second rotating component. The first rotating component includes an inner shroud and an outer shroud outward of the inner shroud in the radial direction. The outer shroud defines a plurality of outer shroud airfoils extended inward of the outer shroud along the radial direction. The first rotating component further includes at least one connecting airfoil coupling the inner shroud and the outer shroud. The second rotating component is upstream of the one or more connecting airfoils of the first rotating component along the longitudinal direction. The second rotating component includes a plurality of second airfoils extended outward in the radial direction. The first rotating component defines at least one stage of the plurality of outer shroud airfoils upstream of the second rotating component.
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公开(公告)号：US10066630B2
公开(公告)日：2018-09-04
申请号：US15183385
申请日：2016-06-15
IPC分类号： G06F19/00 , G06G7/70 , F04D27/00 , F04D19/00 , F04D19/02 , F04D25/04 , F04D29/32 , F04D29/38 , F04D29/52 , F04D29/58 , F02C3/04 , F02K3/06
摘要： A fan assembly is provided. The fan assembly includes a fan, a fan casing circumscribing the fan, and a fan casing heating system in thermal communication with the fan casing. The fan includes a hub, and a plurality of fan blades extending from the hub. Each fan blade of the plurality of fan blades terminates at a respective blade tip. A clearance gap is defined between the fan casing and the blade tips. The fan casing heating system is configured to apply heat to the fan casing when the fan is operating in a first operational mode, and remove the applied heat when the fan transitions into a second operational mode.
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公开(公告)号：US10060290B2
公开(公告)日：2018-08-28
申请号：US14984372
申请日：2015-12-30
CPC分类号： F01D25/20 , F01D5/06 , F01D25/162 , F01D25/18 , F01D25/183 , F04D1/12 , F04D29/321 , F05D2220/32 , F05D2240/60 , F05D2260/98 , Y02T50/671
摘要： The pump includes a first rotatable member including a radially inward facing groove having an edge. The first rotatable member configured to receive a plurality of flows of fluid over the edge. The first rotatable member configured to rotate at a first angular velocity. The pump also includes a second rotatable member including a collector configured to rotate at a second angular velocity. The second rotatable member also includes a plurality of scoop tubes extending radially outwardly from the collector. Each scoop tube of the plurality of scoop tubes includes a first end coupled in flow communication to the collector and a second end including an inlet opening extending into the groove. The second end curved such that the inlet opening is open in a direction of rotation of the second rotatable member. The inlet opening configured to scoop a fluid collected in the groove.
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公开(公告)号：US20180216576A1
公开(公告)日：2018-08-02
申请号：US15293358
申请日：2016-10-14
发明人： Brandon Wayne Miller , Mark John Laricchiuta , Daniel Robert Dwyer , Jeffrey Donald Clements , Kenneth Scheffel , Thomas Ory Moniz
CPC分类号： F02K3/075 , B64D29/00 , F02K3/06 , F05D2220/80 , F05D2260/96
摘要： A supersonic turbofan engine includes a fan section having a single-stage fan defining a fan pressure ratio greater than 1.9. The supersonic turbofan engine also includes a core turbine engine defining a core air flowpath. A nacelle at least partially surrounds the fan of the fan section and the core turbine engine. The supersonic turbofan engine defines a bypass ratio, the bypass ratio being greater than or equal to three.
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