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公开(公告)号：US20240280056A1
公开(公告)日：2024-08-22
申请号：US18430283
申请日：2024-02-01
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 3.1-5.1.
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公开(公告)号：US12006834B1
公开(公告)日：2024-06-11
申请号：US18187724
申请日：2023-03-22
Applicant: General Electric Company
Inventor： Ramesh Balla , Bhaskar Nanda Mondal
Abstract: A gas turbine engine includes a plurality of annular discs arranged in an axial stack and defining axial passage, with at least one of the annular discs having an annular engagement member. A tie rod extends through the axial passage and secures the axial discs in the axial stack. The tie rod includes an annular collar defining an annular groove therein. The annular engagement member is received in the annular groove of the tie rod.
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83.发明公开公开(公告)号：US20240110504A1
公开(公告)日：2024-04-04
申请号：US17955713
申请日：2022-09-29
Applicant: General Electric Company , GE Avio S.r.l.
Inventor： Ranganayakulu Alapati , Peeyush Pankaj , Sanjeev Sai Kumar Manepalli , Bhaskar Nanda Mondal , Thomas Moniz , N V Sai Krishna Emani , Shishir Paresh Shah , Anil Soni , Praveen Sharma , Randy T. Antelo , Antonio Giuseppe D'Ettole
CPC classification number: F02C3/067 , F02C3/10 , F05D2220/323 , F05D2240/24
Abstract: A gas turbine engine includes a fan located at a forward portion of the gas turbine engine, and a compressor section and a turbine section arranged in serial flow order. The compressor section and the turbine section together define a core airflow path. A rotary member is rotatable with the fan and with a low pressure turbine of the turbine section. The low pressure turbine includes a rotating drum to which a first airfoil structure is connected and extends radially inward toward the rotary member. A torque frame connects the rotating drum to the rotary member and transfers torque from the first airfoil structure mounted to the rotating drum to the rotary member. The torque frame includes an inner disk mounted to the rotary member, an outer ring and a second airfoil structure formed separately from the outer ring and connected thereto by a releasable connecting structure. The second airfoil structure extends radially inward from the outer ring toward the inner disk.
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公开(公告)号：US20240068372A1
公开(公告)日：2024-02-29
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：US20230211885A1
公开(公告)日：2023-07-06
申请号：US18058040
申请日：2022-11-22
Applicant: General Electric Company
Inventor： Narayanan Payyoor , Bhaskar Nanda Mondal , Arda Unsal
CPC classification number: B64D27/10 , F02C7/36 , F02C7/06 , F02C3/04 , F05D2260/40 , F05D2240/60 , F05D2300/614 , F05D2300/211 , F05D2300/2112 , F05D2300/2261 , F05D2300/6033 , F05D2260/96
Abstract: A turbomachine engine including a high-pressure compressor, a high-pressure turbine, a combustion chamber in flow communication with the high-pressure compressor and the high-pressure turbine, and a power turbine in flow communication with the high-pressure turbine. At least one of the high-pressure compressor, the high-pressure turbine, and the power turbine comprises a ceramic matrix composite (CMC) material. The turbomachine engine includes a low-pressure shaft coupled to the power turbine and characterized by a midshaft rating (MSR) between two hundred (ft/sec)1/2 and three hundred (ft/sec)1/2. The low-pressure shaft has a redline speed between fifty and two hundred fifty feet per second (ft/sec). The turbomachine engine is configured to operate up to the redline speed without passing through a critical speed associated with a first-order bending mode of the low-pressure shaft.
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Patent Agency Ranking
公开(公告)号：US11118506B2
公开(公告)日：2021-09-14
申请号：US16229533
申请日：2018-12-21
Applicant: General Electric Company
Inventor： Bhaskar Nanda Mondal
Abstract: A gear assembly for a turbo machine is provided. The gear assembly includes a first input shaft, a second input shaft, an output shaft, a fixed member, and a spindle. The spindle is extended along a spindle centerline axis. The first input shaft is drivingly connected to the spindle at a first interface. The second input shaft is drivingly connected to the spindle at a second interface. The spindle is connected to the fixed member providing a reactive force at a third interface. The spindle is connected to the output shaft at a fourth interface.
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公开(公告)号：US20210262416A1
公开(公告)日：2021-08-26
申请号：US16796256
申请日：2020-02-20
Applicant: General Electric Company
Inventor： Kishanjit Pal , Bhaskar Nanda Mondal , Lajith Vijayan
Abstract: A gas turbine engine, the engine including a core turbine engine forming a core flowpath, a rotatable first stage blade assembly in which a bypass airflow passage is formed downstream of the first stage blade assembly, and a shroud positioned at the bypass airflow passage radially outward of the core turbine engine, wherein a first flowpath is formed outward of the shroud at which a first portion of air is flowed, and wherein the shroud and the core turbine engine form a second flowpath therebetween, the core flowpath in fluid communication with the second flowpath to flow a mixture of a second portion of air and combustion gases in the second flowpath.
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公开(公告)号：US10941706B2
公开(公告)日：2021-03-09
申请号：US15895336
申请日：2018-02-13
Applicant: General Electric Company
Inventor： Arnab Sen , Brandon Wayne Miller , Bhaskar Nanda Mondal , Kishanjit Pal , Daniel Alan Niergarth
Abstract: An aeronautical gas turbine engine includes a turbomachine including a compressor section, a combustion section, a turbine section, and an exhaust section in serial flow order. The aeronautical gas turbine engine additionally includes a closed cycle heat engine including a compressor configured to compress a working fluid; a primary heat exchanger in thermal communication with the turbomachine and the working fluid, the primary heat exchanger configured to transfer heat from the turbomachine to the working fluid; an expander coupled to the compressor for expanding the working fluid; and an output shaft driven by the expander.
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公开(公告)号：US10914194B2
公开(公告)日：2021-02-09
申请号：US15710176
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
IPC: F01D25/16 , F01D5/30 , F02C3/04 , F02K3/04 , F01D5/03 , F02C7/06 , F02C3/067 , F01D1/26 , F02C7/36
Abstract: A turbomachine includes a spool and a turbine section including a turbine rear frame, a turbine, and a support member. The turbine includes a first plurality of turbine rotor blades and a second plurality of turbine rotor blades, the first plurality of turbine rotor blades and second plurality of turbine rotor blades alternatingly spaced along the axial direction and rotatable with one another. The support member extends between the first plurality of turbine rotor blades and the spool and couples the first plurality of turbine rotor blades to the spool. The turbomachine also includes a bearing assembly including a bearing, the bearing rotatably supporting the support member and supported by the turbine rear frame, the bearing spaced apart from the spool along the radial direction.
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公开(公告)号：US20200240720A1
公开(公告)日：2020-07-30
申请号：US16828186
申请日：2020-03-24
Applicant: General Electric Company
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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