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公开(公告)号：US20190136698A1
公开(公告)日：2019-05-09
申请号：US15806871
申请日：2017-11-08
Applicant: General Electric Company
Inventor： David William Crall
Abstract: An airfoil defining a chordwise dimension, a spanwise dimension, a leading edge, a trailing edge, a root, and a tip, is generally provided. The airfoil includes a first material substrate defining a pressure side and a suction side. The first material substrate defines a plurality of discrete volumes extended from at least one of the pressure side or the suction side into the first material substrate. The plurality of discrete volumes is arranged at least partially along the chordwise dimension and a second material substrate different from the first material substrate is defined at least partially within the volume.
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公开(公告)号：US20180363554A1
公开(公告)日：2018-12-20
申请号：US15625101
申请日：2017-06-16
Applicant: General Electric Company
Inventor： Christopher James Kroger , Brandon Wayne Miller , Trevor Wayne Goerig , David William Crall , Tsuguji Nakano , Jeffrey Donald Clements , Bhaskar Nanda Mondal
CPC classification number: F02C7/057 , F02C3/04 , F02C7/042 , F02K3/06 , F04D29/522 , F04D29/541 , F04D29/681
Abstract: A gas turbine engine includes a compressor section and a turbine section. The turbine section includes a drive turbine and is located downstream of the compressor section. The gas turbine engine also includes a fan mechanically coupled to and rotatable with the drive turbine such that the fan is rotatable by the drive turbine at the same rotational speed as the drive turbine, the fan defining a fan pressure ratio and including a plurality of fan blades, each fan blade defining a fan tip speed. During operation of the gas turbine engine at a rated speed, the fan pressure ratio of the fan is less than 1.5 and the fan tip speed of each of the fan blades is greater than 1,250 feet per second.
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公开(公告)号：US11187156B2
公开(公告)日：2021-11-30
申请号：US15819327
申请日：2017-11-21
Applicant: General Electric Company
Abstract: A combustion engine includes a combustion section; a fuel delivery system for providing a fuel flow to the combustion section, the fuel delivery system including an oxygen reduction unit for reducing an oxygen content of the fuel flow; a thermal management system including a heat sink heat exchanger, the heat sink heat exchanger in thermal communication with the fuel delivery system at a location downstream of the oxygen reduction unit; and a control system including a sensor operable with the fuel delivery system for sensing data indicative of an operability of the oxygen reduction unit and a controller operable with the sensor, the controller configured to initiate a corrective action based on the data sensed by the sensor indicative of the operability of the oxygen reduction unit.
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公开(公告)号：US20210221491A1
公开(公告)日：2021-07-22
申请号：US17071114
申请日：2020-10-15
Applicant: General Electric Company
Abstract: In an aspect of the present disclosure, an aircraft defining a longitudinal direction and a lateral direction is provided. The aircraft includes: a fuselage; a single unducted rotor engine mounted at a location spaced from the fuselage of the aircraft, the single unducted rotor engine comprising an unducted rotor assembly having a single stage of rotor blades; and a fuselage shield attached to or formed integrally with the fuselage at a location in alignment with the single stage of rotor blades of the unducted rotor assembly along the lateral direction.
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公开(公告)号：US10815886B2
公开(公告)日：2020-10-27
申请号：US15625101
申请日：2017-06-16
Applicant: General Electric Company
Inventor： Christopher James Kroger , Brandon Wayne Miller , Trevor Wayne Goerig , David William Crall , Tsuguji Nakano , Jeffrey Donald Clements , Bhaskar Nanda Mondal
IPC: F02C7/05 , F02C3/04 , F02K3/06 , F04D29/52 , F04D29/68 , F02C7/04 , F04D29/54 , F02C7/057 , F02C7/042
Abstract: A gas turbine engine includes a compressor section and a turbine section. The turbine section includes a drive turbine and is located downstream of the compressor section. The gas turbine engine also includes a fan mechanically coupled to and rotatable with the drive turbine such that the fan is rotatable by the drive turbine at the same rotational speed as the drive turbine, the fan defining a fan pressure ratio and including a plurality of fan blades, each fan blade defining a fan tip speed. During operation of the gas turbine engine at a rated speed, the fan pressure ratio of the fan is less than 1.5 and the fan tip speed of each of the fan blades is greater than 1,250 feet per second.
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Patent Agency Ranking
公开(公告)号：US10508547B2
公开(公告)日：2019-12-17
申请号：US15039751
申请日：2014-11-25
Applicant: GENERAL ELECTRIC COMPANY
Abstract: According to some embodiments, a tie-bolt support assembly is provided which includes a support spring for engagement with both the tie-bolt and a rotor assembly to maintain a load path between the tie-bolt and the rotor assembly while also allowing for axial movement of the tie-bolt.
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公开(公告)号：US20190242399A1
公开(公告)日：2019-08-08
申请号：US15891423
申请日：2018-02-08
Applicant: GENERAL ELECTRIC COMPANY
Inventor： Nicholas Joseph Kray , Gregory Carl Gemeinhardt , Douglas Duane Ward , David William Crall , Wendy Wen-Ling Lin
CPC classification number: F04D29/388 , F01D5/147 , F01D21/045 , F04D29/023 , F04D29/325
Abstract: A blade for a turbine engine comprising a composite core defining a pressure side and a suction side extending axially between a leading edge and a trailing edge defining a chord-wise direction and extending radially between a root and a tip defining a span-wise direction. The composite core is formed from two materials with different compositions.
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公开(公告)号：US20190242260A1
公开(公告)日：2019-08-08
申请号：US15891415
申请日：2018-02-08
Applicant: GENERAL ELECTRIC COMPANY
Inventor： Nicholas Joseph Kray , Gregory Carl Gemeinhardt , Douglas Duane Ward , David William Crall
Abstract: A blade for the fan section of a turbine engine comprising a composite core defining a pressure side and a suction side extending axially between a core leading edge and a core trailing edge defining a chord-wise direction and extending radially between a core root and a core tip defining a span-wise direction and a leading edge strip mounted to the core leading edge.
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公开(公告)号：US20190153952A1
公开(公告)日：2019-05-23
申请号：US15819327
申请日：2017-11-21
Applicant: General Electric Company
Abstract: A combustion engine includes a combustion section; a fuel delivery system for providing a fuel flow to the combustion section, the fuel delivery system including an oxygen reduction unit for reducing an oxygen content of the fuel flow; a thermal management system including a heat sink heat exchanger, the heat sink heat exchanger in thermal communication with the fuel delivery system at a location downstream of the oxygen reduction unit; and a control system including a sensor operable with the fuel delivery system for sensing data indicative of an operability of the oxygen reduction unit and a controller operable with the sensor, the controller configured to initiate a corrective action based on the data sensed by the sensor indicative of the operability of the oxygen reduction unit.
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公开(公告)号：US20180237120A1
公开(公告)日：2018-08-23
申请号：US15439082
申请日：2017-02-22
Applicant: General Electric Company
Inventor： Brandon Wayne Miller , Richard Byron Stewart , Jeffrey Donald Clements , Richard David Cedar , David William Crall
CPC classification number: B64C3/32 , B64D27/18 , B64D2027/262 , F01D21/045 , F02K3/06 , F05D2300/603 , F05D2300/6033 , Y02T50/672
Abstract: The present disclosure is directed to an aircraft including a fuselage to which a pair or more of wings attaches. The aircraft defines a transverse direction, a longitudinal direction, and a latitudinal direction. The aircraft includes a wing extended from the fuselage along the transverse direction in which the wing defines a leading edge, and a gas turbine engine coupled to the wing. The engine defines an axial centerline therethrough along the longitudinal direction. The engine includes a nacelle including an outer wall extended around the axial centerline. The nacelle defines a radial reference plane extended perpendicular from the axial centerline. The outer wall defines an outer wall point closest to the fuselage. The radial reference plane extends through a reference line defined along the latitudinal direction from the outer wall point to the leading edge of the wing. The engine further includes a low pressure (LP) turbine rotor that includes an upstream-most first turbine rotor concentric to the axial centerline. The first turbine rotor is disposed downstream along the longitudinal direction of the radial reference plane.
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