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公开(公告)号：US12012901B1
公开(公告)日：2024-06-18
申请号：US18318604
申请日：2023-05-16
Applicant: General Electric Company
Inventor： Pranav R. Kamat , Bhaskar Nanda Mondal , Jeffrey D Clements
Abstract: A turbomachinery engine includes a fan assembly, a low-pressure turbine, and a gearbox. The fan assembly includes a plurality of fan blades. The low-pressure turbine includes 3-5 rotating stages. The low-pressure turbine includes an area ratio equal to the annular exit area of an aft-most rotating stage of the low-pressure turbine divided by the annular exit area of a forward-most rotating stage of the low-pressure turbine. In some instances, the area ratio is within a range of 2.0-6.5. Additionally (or alternatively) the low-pressure turbine includes an area-EGT ratio within a range of 1.3-1.6.
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公开(公告)号：US12000308B2
公开(公告)日：2024-06-04
申请号：US17893747
申请日：2022-08-23
Applicant: General Electric Company
Inventor： Vinod Shashikant Chaudhari , Bhaskar Nanda Mondal , Thomas O. Moniz
CPC classification number: F01D5/3053 , F01D5/225 , F05D2220/32 , F05D2240/30
Abstract: A rotor blade assembly for a turbine engine, including an airfoil blade including an inner diameter end and an outer diameter end, a lower blade carrier coupled to the inner diameter end of the airfoil blade and rigidly coupled to a disk via a pin, an upper blade carrier coupled to the outer diameter end of the airfoil blade, and an outer drum coupled to the upper blade carrier via a radial joint. The radial joint supports radial motion of the upper blade carrier relative to an axis extending through a center of the rotor blade assembly.
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公开(公告)号：US20230272744A1
公开(公告)日：2023-08-31
申请号：US17826321
申请日：2022-05-27
Applicant: GENERAL ELECTRIC COMPANY
Inventor： Joseph G. Rose , Atanu Saha , Bhaskar Nanda Mondal , Craig W. Higgins , Darek Zatorski , Vaibhav Deshmukh M
CPC classification number: F02C7/06 , F02C3/06 , F05D2240/52 , F05D2260/15 , F05D2260/213 , F05D2260/232 , F05D2260/605 , F05D2260/607
Abstract: A turbine engine having a drive shaft rotatable about an axis, a multi-stage compressor, a turbine section, a thrust bearing, and a balance cavity. The thrust bearing being provided between the drive shaft and at least a portion of the multi-stage compressor section and rotationally supporting the drive shaft. During operation of the turbine engine, a first axial force is applied to the thrust bearing by the drive shaft and a second axial force is applied to the thrust bearing in an opposite direction of the first axial force by the balance cavity.
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公开(公告)号：US11280219B2
公开(公告)日：2022-03-22
申请号：US16697350
申请日：2019-11-27
Applicant: General Electric Company
Inventor： Shashank Suresh Puranik , Bhaskar Nanda Mondal , Richard Schmidt , Peeyush Pankaj , Narayanan Payyoor
Abstract: A rotor support system for a gas turbine engine includes a rotatable drum rotor and a non-rotatable support casing. The rotor support system includes a bearing assembly configured for positioning between the rotatable drum rotor and the non-rotatable support casing. The bearing assembly includes, at least, a stationary support frame and a rotatable race. Further, the rotatable race is configured to engage the rotatable drum rotor at separate and discrete locations that are circumferentially spaced apart around the rotatable drum rotor.
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公开(公告)号：US11193720B2
公开(公告)日：2021-12-07
申请号：US16828186
申请日：2020-03-24
Applicant: General Electric Company
Inventor： Karthikeyan Sampath , Atanu Saha , Bhaskar Nanda Mondal , Sanjeev Sai Kumar Manepalli
Abstract: A gas turbine engine having a heat absorption device and an associated method are disclosed. The gas turbine engine includes a compressor having a compressor discharge nozzle, a combustor coupled to the compressor, a turbine coupled to the combustor and the compressor, a fluid flow passage fluidly coupling the compressor and the turbine and bypassing the combustor, and a heat absorption device disposed fluidly along the fluid flow passage at a first predefined location. The heat absorption device includes a casing having an inlet and an outlet, a flow path within the casing and extending between the inlet and the outlet, wherein the flow path directs an input bleed flow diverted from a fluid stream discharged from the compressor, and a phase change material hermetically sealed within the casing. The phase change material is separated from the flow path. The heat absorption device is configured to exchange heat between the phase change material and the input bleed flow to generate an output bleed flow of a different temperature than the input bleed flow and discharge the output bleed flow to a second predefined location different from the first predefined location.
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Patent Agency Ranking
公开(公告)号：US11098592B2
公开(公告)日：2021-08-24
申请号：US15710164
申请日：2017-09-20
Applicant: General Electric Company
Inventor： Peeyush Pankaj , Shashank Suresh Puranik , Darek Tomasz Zatorski , Christopher Charles Glynn , Richard Schmidt , Bhaskar Nanda Mondal
Abstract: A turbomachine includes a spool; and a turbine section including a turbine and a turbine center frame. The turbine includes a first plurality of turbine rotor blades and a second plurality of turbine rotor blades alternatingly spaced along an axial direction and rotatable with one another. The turbomachine also includes a first support member, the first plurality of turbine rotor blades coupled to the spool through the first support member; a second support member, the second plurality of turbine rotor blades supported by the second support member; and a bearing assembly including a first bearing and a second bearing, the first bearing and the second bearing each rotatably supporting the second support member and each being supported by the turbine center frame.
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公开(公告)号：US20210171211A1
公开(公告)日：2021-06-10
申请号：US16705384
申请日：2019-12-06
Applicant: General Electric Company
Inventor： Bhaskar Nanda Mondal , Arvind Namadevan , Shankar Jayaraman , Keith EJ Blodgett , Arnab Sen
Abstract: An aircraft defining a longitudinal centerline and extending between a forward end and an aft end is provided. The aircraft includes a fuselage extending longitudinally between the forward end of the aircraft and the aft end of the aircraft; a primary wing assembly extending laterally outwardly with respect to the longitudinal centerline from a portion of the fuselage; a first engine mounted to the primary wing assembly; and a first engine wing assembly extending outward from the first engine.
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公开(公告)号：US10781709B2
公开(公告)日：2020-09-22
申请号：US15865359
申请日：2018-01-09
Applicant: GENERAL ELECTRIC COMPANY
Inventor： Arnab Sen , Jeffrey Douglas Rambo , Rajesh Kumar , Nathan Evan McCurdy Gibson , Robert Proctor , Bhaskar Nanda Mondal , Steven Douglas Johnson
Abstract: A turbine engine includes an engine core defining a higher pressure region and a lower pressure region. A seal can fluidly separate the higher pressure region from the lower pressure region and be movably mounted to a component within the turbine engine, where a side of the seal can confront the component.
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公开(公告)号：US10724435B2
公开(公告)日：2020-07-28
申请号：US15625186
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
Abstract: A gas turbine engine includes a turbomachine and a fan rotatable by the turbomachine. The fan includes a plurality of fan blades. The gas turbine engine also includes an outer nacelle surrounding the plurality of fan blades and a plurality of part-span inlet guide vanes cantilevered from the outer nacelle at a location forward of the plurality of fan blades along the axial direction. Each of the plurality of inlet guide vanes defines an inner end along the radial direction and it is unconnected with an adjacent part-span inlet guide vane at the inner end.
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公开(公告)号：US10711797B2
公开(公告)日：2020-07-14
申请号：US15625212
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
Abstract: A gas turbine engine includes a turbomachine and a fan rotatable by the turbomachine. The fan includes a plurality of fan blades, each of the plurality of fan blades defining a fan blade span along a radial direction. The gas turbine engine further includes an outer nacelle surrounding the plurality of fan blades and a plurality of part-span inlet guide varies attached to the outer nacelle at a location forward of the plurality of fan blades along an axial direction. Each of the plurality of inlet guide vanes defines an IGV span along the radial direction, the IGV span being at least about five percent of the fan blade span and up to about fifty-five percent of the fan blade span.
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