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公开(公告)号：US11156092B2
公开(公告)日：2021-10-26
申请号：US16269804
申请日：2019-02-07
Applicant: HONEYWELL INTERNATIONAL INC.
Inventor： Bruce David Reynolds , Richard David Conner , Michael T. Barton , Jeffrey W. Gluck , Nick Nolcheff
IPC: F01D5/06
Abstract: A method of manufacturing a multistage axial-centrifugal compressor for a gas turbine engine and a multistage axial-centrifugal compressor that includes a series of axial compressor stages each having a rotor mounted to a common shaft positioned upstream from a centrifugal compressor stage mounted to the common shaft. The method includes determining an operational rotor tip speed for each of the axial stages. The method includes comparing the operational rotor tip speed to a threshold range value and determining to machine an airfoil of the rotor for at least one of the axial stages by an arbitrary manufacturing approach based on the operational rotor tip speed as greater than the threshold range value. The method includes determining to machine an airfoil of the rotor for at least one of the axial stages by a flank manufacturing approach based on the operational rotor tip speed as less than the threshold range value.
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公开(公告)号：US20210207614A1
公开(公告)日：2021-07-08
申请号：US17060219
申请日：2020-10-01
Applicant: HONEYWELL INTERNATIONAL INC.
Inventor： David Richard Hanson , John A. Gunaraj , Jeffrey Hayes , Nick Nolcheff , John Repp
Abstract: A rotor for a compressor includes a hub and a plurality of airfoils having a root, a tip opposite the root and a span that extends from 0% at the root to 100% at the tip. Each of the airfoils is coupled to the hub at the root and is spaced apart from adjacent ones of the airfoils over the span by a throat dimension. The throat dimension has a maximum value at a spanwise location between 60% of the span and 90% of the span of the adjacent ones of the airfoils. The throat dimension between 90% of the span and the tip of the adjacent ones of the plurality of airfoils has a first value that is less than 70% of the maximum value.
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43.发明申请公开(公告)号：US20200340486A1
公开(公告)日：2020-10-29
申请号：US16924887
申请日：2020-07-09
Applicant: HONEYWELL INTERNATIONAL INC.
Inventor： Bruce David Reynolds , Nick Nolcheff
Abstract: 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：US20200182259A1
公开(公告)日：2020-06-11
申请号：US16732507
申请日：2020-01-02
Applicant: HONEYWELL INTERNATIONAL INC.
Inventor： Timothy Gentry , Bruce David Reynolds , Nick Nolcheff
Abstract: 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：US10648484B2
公开(公告)日：2020-05-12
申请号：US15431890
申请日：2017-02-14
Applicant: HONEYWELL INTERNATIONAL INC.
Inventor： Timothy Gentry , Bruce David Reynolds , Nick Nolcheff
Abstract: 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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Patent Agency Ranking
公开(公告)号：US10294965B2
公开(公告)日：2019-05-21
申请号：US15163990
申请日：2016-05-25
Applicant: HONEYWELL INTERNATIONAL INC.
Inventor： Nick Nolcheff , Michael LaMar Trego , John Repp , James Kroeger
IPC: F01D5/16 , F01D5/34 , F02C3/08 , F04D17/02 , F04D19/00 , F04D19/02 , F04D29/32 , F04D29/38 , F04D29/66
Abstract: A blisk fan is provided for a turbine engine propulsion system. The blisk fan includes a hub configured to rotate about a rotational axis at a maximum rotational speed, and a plurality of blades extending radially outward from the hub to define a fan leading edge tip diameter. Each of the blades has a first vibratory mode at a natural frequency, which is greater than a first fan order and less than a second fan order at the maximum rotational speed. The compression system preferably has a balance factor of the compression system between 1.9 and 3.2.
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公开(公告)号：US20190085765A1
公开(公告)日：2019-03-21
申请号：US15709541
申请日：2017-09-20
Applicant: HONEYWELL INTERNATIONAL INC.
Inventor： Nick Nolcheff , John Meier , James Laffan , Alan D. Hemmingson , Cristian Anghel , Mike Koerner
Abstract: A propulsion and electric power generation system includes a gas turbine propulsion engine, an electrical generator, an aircraft power distribution system, a plurality of auxiliary fans, and a controller. The gas turbine propulsion engine includes at least a low-pressure turbine coupled to a fan via a low-pressure spool, and the low-pressure turbine is configured to generate mechanical power. The electrical generator is directly connected to the low-pressure spool and generates a total amount of electrical power (Pe). The aircraft power distribution system receives a first fraction (Pa) of the total amount of electrical power. The auxiliary fans receive a second fraction (Pf) of the total amount of electrical power. The controller is configured to control a ratio of Pf to Pa (Pf/Pa) such that the ratio spans a range from less than 0.6 to at least 0.9.
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公开(公告)号：US20170284226A1
公开(公告)日：2017-10-05
申请号：US15085625
申请日：2016-03-30
Applicant: HONEYWELL INTERNATIONAL INC.
Inventor： Shakeel Nasir , John Schugardt , Nick Nolcheff , John Meier , Yogendra Yogi Sheoran , Shezan Kanjiyani , Daniel Aukland
IPC: F01D25/32
CPC classification number: F01D25/32 , F02C7/36 , F04D29/441 , F04D29/701 , F05D2220/32 , F05D2250/52 , F05D2260/40311 , F05D2260/607 , Y02T50/675
Abstract: A turbine engine incorporating a fine particle separation means includes a radial compressor that rotates about a longitudinal axis, a radially-oriented diffuser located downstream and radially outward, with respect to the longitudinal axis, from the radial compressor, and a flow path positioned downstream and radially outward, with respect to the longitudinal axis, from the diffuser, wherein the flow path comprises an outer annular wall and an inner annular wall between which the compressed air flows, and wherein the flow path comprises an arc the redirects the compressed air from flowing in a substantially radial flow direction to a substantially axial flow direction. The turbine engine further includes an extraction slot in the outer annular wall that fluidly connects with a scavenge plenum, the scavenge plenum being positioned adjacent to and radially outward from the outer annular wall at a position downstream axially along the flow path from the arc.
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公开(公告)号：US20150198090A1
公开(公告)日：2015-07-16
申请号：US14154428
申请日：2014-01-14
Applicant: Honeywell International Inc.
Inventor： Jeff Howe , Harry Lester Kington , Nick Nolcheff
CPC classification number: F02C7/052 , B03C3/08 , B03C3/12 , B03C3/361 , B03C3/41 , B03C2201/10 , B64D2033/0246 , F05D2260/607 , F23R3/04 , Y02T50/675
Abstract: An inlet particle separator system for an engine includes an inner flowpath section, an outer flowpath section, a splitter, a first electrostatic discharge device, and a second electrostatic discharge device. The outer flowpath section surrounds at least a portion of the inner flowpath section and is spaced apart therefrom to define a passageway having an air inlet. The splitter is disposed downstream of the air inlet and extends into the passageway to divide the passageway into a scavenge flow path and an engine flow path. The first electrostatic charge device is disposed between the air inlet and the splitter and is electrostatically charged to a first polarity. The second electrostatic charge device is disposed downstream of the first electrostatic charge device and is electrostatically charged to a second polarity that is opposite to the first polarity.
Abstract translation: 用于发动机的入口颗粒分离器系统包括内部流路部分，外部流路部分，分流器，第一静电放电装置和第二静电放电装置。 外部流路部分围绕内部流路部分的至少一部分并且与其间隔开以限定具有空气入口的通道。 分流器设置在空气入口的下游并延伸到通道中，以将通道分成清流流动路径和发动机流动路径。 第一静电充电装置设置在空气入口和分流器之间，并被静电充电至第一极性。 第二静电充电装置设置在第一静电充电装置的下游，并被静电充电至与第一极性相反的第二极性。
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公开(公告)号：US11466694B2
公开(公告)日：2022-10-11
申请号：US16924887
申请日：2020-07-09
Applicant: HONEYWELL INTERNATIONAL INC.
Inventor： Bruce David Reynolds , Nick Nolcheff
Abstract: 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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