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公开(公告)号：US20220349316A1
公开(公告)日：2022-11-03
申请号：US17864534
申请日：2022-07-14
Applicant: General Electric Company
Inventor： Bhaskar Nanda Mondal , John Joseph Rahaim , Thomas Moniz , Steven A. Ross , Joel Kirk , Scott Hunter , Daniel Fusinato
Abstract: A passive clearance control limits thermal expansion between stator components relative to rotor components. A control ring controls clearance in a passive manner and is located on or adjacent to stationary components which thermally expand during engine operation. The control ring is formed of material having low coefficient of thermal expansion such as CMCs (Ceramic Matrix Composites) and therefore limits, inhibits or restrains expansion of the adjacent stator components as temperatures increase. Limiting expansion of the stator component reduces rotor/stator clearances and limits parasitic leakage of fluid along the flow path through the engine core.
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公开(公告)号：US11428160B2
公开(公告)日：2022-08-30
申请号：US17139138
申请日：2020-12-31
Applicant: General Electric Company
Abstract: A gas turbine engine having an interdigitated turbine assembly including a first turbine rotor and a second turbine rotor, wherein a total number of stages at the interdigitated turbine assembly is between 3 and 8, and an average stage pressure ratio at the interdigitated turbine assembly is between 1.3 and 1.9. A gear assembly is configured to receive power from the interdigitated turbine assembly, and a fan assembly is configured to receive power from the gear assembly. The interdigitated turbine assembly and the gear assembly are together configured to allow the second turbine rotor to rotate at a second rotational speed greater than a first rotational speed at the first turbine rotor. The fan assembly and the gear assembly are together configured to allow the fan assembly to rotate at a third rotational speed less than the first rotational speed and the second rotational speed. The interdigitated turbine assembly, the gear assembly, and the fan assembly together have a maximum AN2 at the second turbine rotor between 30 and 90.
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公开(公告)号：US11231043B2
公开(公告)日：2022-01-25
申请号：US15900891
申请日：2018-02-21
Applicant: General Electric Company
Inventor： Veeraraju Vanapalli , Bhaskar Nanda Mondal , Jagata Laxmi Narasimharao , Tsuguji Nakano , Subramanian Narayanan
Abstract: The present disclosure is directed to a gas turbine engine including a compressor rotor. The compressor rotor includes a first stage compressor airfoil defined at an upstream-most stage of the compressor rotor. The first stage compressor airfoil defines a first stage pressure ratio of at least approximately 1.7 during operation of the gas turbine engine at a tip speed of at least approximately 472 meters per second.
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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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公开(公告)号：US10316681B2
公开(公告)日：2019-06-11
申请号：US15168581
申请日：2016-05-31
Applicant: General Electric Company
Inventor： Jagata Laxmi Narasimharao , Thomas Ory Moniz , Mohamed Musthafa Thoppil , Bhaskar Nanda Mondal , Atanu Saha
Abstract: A turbine assembly includes a rotor assembly including a shaft coupled to a plurality of rotor stages including a plurality of turbine blades. The shaft and the plurality of turbine blades define a wheelspace therein. The turbine assembly further includes a plurality of seals in series, at least one seal of the plurality of seals is coupled between a static support member and a respective rotor stage such that a plurality of turbine cavities in series are defined within the wheelspace. Each turbine cavity of the plurality of turbine cavities defined by the plurality of seals receives a pressurized fluid flow that applies an axially aft force to the respective rotor stage of the plurality of rotor stages that at least partially reduces net rotor thrust generated by the rotor assembly during operation, the pressurized fluid flow further provides turbine purge within the wheel space.
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Patent Agency Ranking
公开(公告)号：US20170268535A1
公开(公告)日：2017-09-21
申请号：US15447565
申请日：2017-03-02
Applicant: General Electric Company
Inventor： Bhaskar Nanda Mondal , Kenneth Edward Seitzer , Monty Lee Shelton , Thomas Ory Moniz , Atanu Saha
CPC classification number: F04D29/642 , F01D11/24 , F02K3/06 , F04D19/02 , F04D29/5846 , F04D29/5853 , F05D2220/32 , F05D2260/201 , Y02T50/671
Abstract: The turbomachine includes a compressor, an inner annular casing, and an outer annular casing. The inner annular casing and the outer annular casing define at least one cavity therebetween. The clearance control system includes a manifold system including at least one conduit disposed within the cavities and configured to channel a flow of cooling fluid between the cavities. The clearance control system also includes an impingement system including a header and at least one plenum configured to channel the flow of cooling fluid to the inner annular casing. The conduits configured to channel the flow of cooling fluid to the impingement system. The clearance control system further includes a channel system including at least one channels configured to channel the flow of cooling fluid to the turbomachine. The channels are configured to control the flow of cooling fluid to the manifold system.
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公开(公告)号：US20240392690A1
公开(公告)日：2024-11-28
申请号：US18202506
申请日：2023-05-26
Applicant: General Electric Company
Inventor： Vaishnav Raghuvaran , Bhaskar Nanda Mondal , Gary W. Bryant
Abstract: A fan assembly for an engine includes a fan disk, fan disk inserts, and fan blades. The fan disk includes disk posts extending in a radial direction from a central region of the fan disk. The disk posts include disk post pressure surfaces that define a first array of dovetail recesses. The fan disk inserts are retained within the first array of dovetail recesses. The fan disk inserts include fan disk insert pressure surfaces that define a second array of dovetail recesses. The fan blades are retained within the second array of dovetail recesses.
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公开(公告)号：US12152537B2
公开(公告)日：2024-11-26
申请号：US18430317
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：US12049833B1
公开(公告)日：2024-07-30
申请号：US18202503
申请日：2023-05-26
Applicant: General Electric Company
Inventor： Vaishnav Raghuvaran , Bhaskar Nanda Mondal , Gary W. Bryant
IPC: F01D5/30
CPC classification number: F01D5/30 , F05D2220/32 , F05D2230/60 , F05D2240/30
Abstract: A rotor disk for a gas turbine engine includes disk posts extending in a radial direction from a disk body. Each of the disk posts includes an insert receiving slot extending axially through the disk post and a disk post pressure surface. Disk post inserts are assembled within the insert receiving slots of each of the disk posts. A blade is retained by disk post pressure surfaces of each of the disk posts.
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公开(公告)号：US20230417152A1
公开(公告)日：2023-12-28
申请号：US18189740
申请日：2023-03-24
Applicant: General Electric Company
Inventor： Bhaskar Nanda Mondal , Narayanan Payyoor , Pranav Kamat
IPC: F01D15/00
CPC classification number: F01D15/00 , F05D2240/60 , F05D2300/6033
Abstract: A turbomachine engine includes an engine core including a high-pressure compressor, a high-pressure turbine, and a combustion chamber. The engine core has a length (LCORE), and the high-pressure compressor has an exit stage diameter (DCORE). A power turbine is in flow communication with the high-pressure turbine. A low-pressure shaft is coupled to the power turbine and characterized by a midshaft rating from one hundred fifty (ft/sec)1/2 to three hundred thirty (ft/sec)1/2. The low-pressure shaft has a redline speed between fifty and two hundred fifty feet per second. 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. A high-pressure shaft is coupled to the high-pressure compressor and the high-pressure turbine and is characterized by a high-speed shaft rating from 1.5 to 6.2, and a ratio of LCORE/DCORE is from 2.1 to 4.3.
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