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公开(公告)号：US11946474B2
公开(公告)日：2024-04-02
申请号：US17450840
申请日：2021-10-14
CPC分类号： F04D17/025 , F01D17/26 , F01D19/00 , F02C7/264 , F02C9/18 , F05D2240/11 , F05D2270/80
摘要： A gas turbine engine includes a combustor having a combustor air inlet, an axial-centrifugal compressor, a shroud, a secondary flow duct, and a valve. The shroud surrounds at least a portion of the axial-centrifugal compressor and has a surge bleed plenum defined therein that is in fluid communication with, and receives compressed air from, the axial compressor outlet. The secondary airflow duct has a duct inlet that is in fluid communication with the surge bleed plenum, and a duct outlet that is in fluid communication with the combustor air inlet. The valve is mounted on the secondary airflow duct and is movable between a closed position, in which the secondary airflow duct does not provide fluid communication between the surge bleed plenum and the combustor air inlet, and an open position, in which the secondary airflow duct provides fluid communication between the surge bleed plenum and the combustor air inlet.
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公开(公告)号：US11015465B2
公开(公告)日：2021-05-25
申请号：US16363117
申请日：2019-03-25
摘要： A compressor section includes a shroud surface and a rotor with a blade tip that opposes the shroud surface. The rotor is configured to rotate within the shroud about an axis of rotation. Moreover, the compressor section includes a serration groove that is recessed into the shroud surface. The serration groove includes a forward portion with a forward transition and a forward surface that faces in the downstream direction. The forward transition is convexly contoured between the shroud surface and the forward surface. The serration groove includes a trailing portion with a taper surface and a trailing transition. The taper surface tapers inward as the taper surface extends from the forward surface to the trailing transition. The trailing transition is convexly contoured between the taper surface and the shroud surface.
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公开(公告)号：US20200240435A1
公开(公告)日：2020-07-30
申请号：US16681303
申请日：2019-11-12
摘要： Variable stator vane assemblies and stator vanes thereof having a local swept leading edge are provided. The variable stator vane comprises an airfoil disposed between spaced apart inner and outer buttons centered about a rotational axis. The inner and outer buttons each having a button forward edge portion and the airfoil including leading and trailing edges, pressure and suction sides, and a root and a tip. The leading edge, at least a portion of which extends forward of the buttons, includes a local aft sweep at the tip, thereby forming a locally swept tip of the leading edge thereat. The button forward edge portion of the outer button is substantially vertically aligned with the locally swept tip of the leading edge. Methods are also provided for minimizing endwall leakage in the variable stator vane assembly using the same.
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公开(公告)号：US10359051B2
公开(公告)日：2019-07-23
申请号：US15006551
申请日：2016-01-26
发明人： Raymond Gage , Bruce David Reynolds , Jeffrey D. Harrison , Michael Todd Barton , Mahmoud Mansour , Kent L. Kime
摘要： Embodiments of an impeller shroud support for disposition around an impeller are provided, as are embodiments of gas turbine engine including impeller shroud supports. In one embodiment, the impeller shroud support includes a shroud body, a support arm joined to and extending around the shroud body, and a plurality of Mid-Impeller Bleed (MIB) flow passages. Each MIB flow passage includes, in turn, an inlet formed in the shroud body and configured to receive bleed air extracted from the impeller, a throat portion, an outlet formed in the support arm and through which the bleed air is discharged, and a curved intermediate section between the inlet and the outlet. During usage of the impeller shroud support, the curved intermediate section turns the bleed air flowing through the MIB passage in a radially outward direction prior to discharge from the outlet of the MIB flow passage.
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公开(公告)号：US20180231023A1
公开(公告)日：2018-08-16
申请号：US15431890
申请日：2017-02-14
CPC分类号： F04D29/526 , F01D5/20 , F01D11/08 , F01D11/122 , F02C3/04 , F04D29/164 , F04D29/324 , F05D2220/32 , F05D2240/307 , Y02T50/673
摘要： A compressor for a turbine engine includes a shroud having a grooved section including a plurality of groove segments extending radially into a shroud surface. A rotor assembly rotatably supported in the shroud includes a rotor hub and a plurality of rotor blades. Each rotor blade extends radially from the rotor hub and terminates at a blade tip, which is spaced from the shroud surface by a tip gap and defines a non-constant clearance region between a leading edge position and a medial chord position along the blade chord at the minimum tip clearance. The rotor blades generate an aft axial fluid flow through the shroud and the grooved section is formed in the shroud surface upstream of the medial chord position within the non-constant clearance region for resisting a reverse axial fluid flow through the tip gap when the compressor section is operated at near stall conditions.
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公开(公告)号：US09932985B2
公开(公告)日：2018-04-03
申请号：US14612404
申请日：2015-02-03
CPC分类号： F04D27/02 , F01D25/24 , F04D19/02 , F04D27/001 , F04D27/0261 , F04D29/321 , F04D29/526 , F04D29/685 , F05D2220/32 , F05D2230/00
摘要： Multistage gas turbine engine (GTE) compressors having optimized stall enhancement feature (SEF) configurations are provided, as are methods for the production thereof. The multistage GTE compressor includes a series of axial compressor stages each containing a rotor mounted to a shaft of a gas turbine engine. In one embodiment, the method includes the steps or processes of selecting a plurality of engine speeds distributed across an operational speed range of the gas turbine engine, identifying one or more stall limiting rotors at each of the selected engine speeds, establishing an SEF configuration in which SEFs are integrated into the multistage GTE compressor at selected locations corresponding to the stall limiting rotors, and producing the multistage GTE compressor in accordance with the optimized SEF configuration.
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公开(公告)号：US20230122939A1
公开(公告)日：2023-04-20
申请号：US17450840
申请日：2021-10-14
摘要： A gas turbine engine includes a combustor having a combustor air inlet, an axial-centrifugal compressor, a shroud, a secondary flow duct, and a valve. The shroud surrounds at least a portion of the axial-centrifugal compressor and has a surge bleed plenum defined therein that is in fluid communication with, and receives compressed air from, the axial compressor outlet. The secondary airflow duct has a duct inlet that is in fluid communication with the surge bleed plenum, and a duct outlet that is in fluid communication with the combustor air inlet. The valve is mounted on the secondary airflow duct and is movable between a closed position, in which the secondary airflow duct does not provide fluid communication between the surge bleed plenum and the combustor air inlet, and an open position, in which the secondary airflow duct provides fluid communication between the surge bleed plenum and the combustor air inlet.
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公开(公告)号：US11098731B2
公开(公告)日：2021-08-24
申请号：US16732507
申请日：2020-01-02
摘要： A compressor for a turbine engine includes a shroud having a grooved section including a plurality of groove segments extending radially into a shroud surface. A rotor assembly rotatably supported in the shroud includes a rotor hub and a plurality of rotor blades. Each rotor blade extends radially from the rotor hub and terminates at a blade tip, which is spaced from the shroud surface by a tip gap and defines a non-constant clearance region between a leading edge position and a medial chord position along the blade chord at the minimum tip clearance. The rotor blades generate an aft axial fluid flow through the shroud and the grooved section is formed in the shroud surface upstream of the medial chord positon within the non-constant clearance region for resisting a reverse axial fluid flow through the tip gap when the compressor section is operated at near stall conditions.
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公开(公告)号：US20210054761A1
公开(公告)日：2021-02-25
申请号：US17090039
申请日：2020-11-05
摘要： A gas turbine engine includes a shroud with an abradable section and a non-abradable section that cooperatively define a shroud surface. The gas turbine engine also includes a rotor that is supported for rotation within the shroud to generate an aft axial fluid flow. The rotor includes a blade with a blade tip that is crowned and that opposes the abradable section and the non-abradable section of the shroud surface. A crown area of the blade tip opposes the abradable section. A casing treatment feature is provided in the non-abradable section of the shroud to oppose the blade tip of the rotor.
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10.发明申请公开(公告)号：US20200208532A1
公开(公告)日：2020-07-02
申请号：US16235876
申请日：2018-12-28
IPC分类号： F01D11/12
摘要： A gas turbine engine includes a shroud with an abradable section and a non-abradable section that cooperatively define a shroud surface. The gas turbine engine also includes a rotor that is supported for rotation within the shroud to generate an aft axial fluid flow. The rotor includes a blade with a blade tip that is crowned and that opposes the abradable section and the non-abradable section of the shroud surface. A crown area of the blade tip opposes the abradable section. A casing treatment feature is provided in the non-abradable section of the shroud to oppose the blade tip of the rotor.
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