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公开(公告)号：US20240141839A1
公开(公告)日：2024-05-02
申请号：US17976509
申请日：2022-10-28
Applicant: Pratt & Whitney Canada Corp.
Inventor： Jeffrey Poissant , Robert Pontarelli , Cristina Crainic
CPC classification number: F02C9/28 , F02C9/44 , F05D2220/323 , F05D2270/051 , F05D2270/304 , F05D2270/311 , F05D2270/313 , F05D2270/708
Abstract: A method and system for correlating engine thrust and engine thrust lever angle (TLA) during operation of a gas turbine engine powered aircraft is provided. The method includes: a) providing current flight conditions including a Mach number of the aircraft, altitude of the aircraft, and a TLA; b) determining a corrected fan speed value based on a sensed fan speed and flight conditions; c) determining a first corrected net thrust value based on the corrected fan speed value and the Mach number; d) determining a compensation factor using the corrected fan speed value, the Mach number, and the corrected net thrust value; e) determining a second corrected net thrust value as a function of the TLA; and f) correlating an engine fan speed to the TLA using the second corrected net thrust value as a function of the TLA and the compensation factor.
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公开(公告)号：US11125168B2
公开(公告)日：2021-09-21
申请号：US16168934
申请日：2018-10-24
Applicant: United Technologies Corporation
Inventor： Christopher J. Hanlon , James F. Krenzer , Becky E. Rose , Mark F. Zelesky
Abstract: An aspect includes a dirt mitigation system for a gas turbine engine. The dirt mitigation system includes a plurality of bleeds of the gas turbine engine and a control system configured to determine a particulate ingestion estimate indicative of dirt ingested in the gas turbine engine. The control system is further configured to determine one or more operating parameters of the gas turbine engine and alter a bleed control schedule of the gas turbine engine to purge at least a portion of the dirt ingested in the gas turbine engine through one or more of the bleeds of the gas turbine engine based on the particulate ingestion estimate and the one or more operating parameters.
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公开(公告)号：US20210033031A1
公开(公告)日：2021-02-04
申请号：US16892589
申请日：2020-06-04
Applicant: ROLLS-ROYCE plc
Inventor： Caroline L. TURNER
IPC: F02C9/44
Abstract: A gas turbine engine for an aircraft, comprises a high-pressure (HP) spool comprising an HP compressor and a first electric machine driven by an HP turbine, the first electric machine having a first maximum output power; a low-pressure (LP) spool comprising an LP compressor and a second electric machine driven by an LP turbine, the second electric machine having a second maximum output power; and an engine controller configured to identify a condition to the effect that the LP turbine is operating in an unchoked regime, and, in response to an electrical power demand being between zero and the first maximum output power, only extracting electrical power from the first electric machine to meet the electrical power demand.
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公开(公告)号：US20200378315A1
公开(公告)日：2020-12-03
申请号：US16428822
申请日：2019-05-31
Applicant: Hamilton Sundstrand Corporation
Inventor： Charles E. Reuter , Ryan Susca
Abstract: A fuel system with integrated thrust control malfunction and overspeed protection includes a centrifugal pump for receiving filtered fuel and supplying fuel to a metering and pressure regulating valve, the metering and pressure regulating valve for metering discharge of filtered fuel flow to a gas turbine engine burner, a shut off valve downstream of the metering and pressure regulating valve movable from an open position to either a fully closed position configured for providing rapid shutoff or a partially closed position configured for reduction of fuel flow to the gas turbine engine burner, and a first solenoid and a second solenoid for controlling the shut off valve position.
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公开(公告)号：US10823113B2
公开(公告)日：2020-11-03
申请号：US15681531
申请日：2017-08-21
Applicant: PRATT & WHITNEY CANADA CORP.
Inventor： Sylvain Lamarre , Simon Lopez , Nicolas Des Roches-Dionne , Michael Conciatori
Abstract: Systems and methods for limiting power of a gas turbine engine for an aircraft are described herein. A blade angle of a propeller blade of the engine and a commanded power for the engine are obtained. A thrust transition direction is determined. The commanded power is compared to a selected threshold based on the blade angle and the thrust transition direction. Power to the engine is limited when the commanded power exceeds the selected threshold.
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Patent Agency Ranking
公开(公告)号：US10731570B2
公开(公告)日：2020-08-04
申请号：US15609875
申请日：2017-05-31
Applicant: PRATT & WHITNEY CANADA CORP.
Inventor： Ninad Joshi , Sid-Ali Meslioui , Tony Yee
Abstract: Herein provided are methods and systems for reducing an acoustic signature of a gas turbine engine. An acceleration command for the engine is received. In response to receiving the acceleration command: a fuel flow to the engine is increased for a first predetermined time period; subsequent to the first predetermined time period, the fuel flow to the engine is reduced for a second predetermined time period; and subsequent to the second predetermined time period, the fuel flow to the engine is increased for a third time period.
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公开(公告)号：US20200232394A1
公开(公告)日：2020-07-23
申请号：US16745454
申请日：2020-01-17
Applicant: United Technologies Corporation
Inventor： Neal R. Herring
Abstract: A gas turbine engine includes a main compressor section, a combustor, and a main turbine section. A fuel pump delivers fuel to the combustor. A tap taps air compressed by the main compressor section, and is connected for delivering the tapped air through a first heat exchanger and to a boost compressor. Air downstream of the boost compressor is connected to cool a component. Driving compressed air is connected to be delivered to a power turbine. The power turbine is connected to drive both the boost compressor and the fuel pump.
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公开(公告)号：US10647414B2
公开(公告)日：2020-05-12
申请号：US15443826
申请日：2017-02-27
Applicant: Textron Innovations Inc.
Abstract: A fly-by-wire system for a rotorcraft includes a computing device having control laws. The control laws are operable to engage a roll command or a yaw command in response to deflection of a beep switch of a pilot control assembly, wherein a roll angle for the roll command or a yaw rate for the yaw command is determined based on forward airspeed of the rotorcraft. The beep switch may be disposed on a collective control of the pilot control assembly. The control laws are further operable to disengage the roll command or the yaw command in response to the beep switch being returned from a deflected position to a neutral position. In representative aspects, the roll angle or the yaw rate may correspond to a standard rate turn (e.g., 3° per second).
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公开(公告)号：US10227933B2
公开(公告)日：2019-03-12
申请号：US14620271
申请日：2015-02-12
Applicant: UNITED TECHNOLOGIES CORPORATION
Inventor： Kurt J. Sobanski , Richard P. Meisner , Robert J. Bengtson
Abstract: A method of controlling thrust for a gas turbine engine of an aircraft is provided. The method includes determining a fan speed required for minimum thrust to achieve an aircraft operation. The method also includes determining an excess amount of thrust generated by the gas turbine engine. The method also includes reducing the amount of thrust generated by the gas turbine engine.
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公开(公告)号：US20170198644A1
公开(公告)日：2017-07-13
申请号：US15314048
申请日：2015-05-13
Applicant: SAFRAN AIRCRAFT ENGINES
Inventor： Amaury OLIVIER , Christophe JAVELOT , Darragh MCGRATH
CPC classification number: F02C9/44 , F02C9/26 , F02C9/28 , F05D2270/051 , F05D2270/101
Abstract: A method for controlling thrust from a turbojet that is fuel flow rate regulated by a high limit value for providing protection against surging of a compressor of the turbojet is provided. The method includes: obtaining a first thrust value corresponding to a first operating point of the compressor on the high limit value, the high limit value taking account of an underestimate of the fuel flow rate; controlling the turbojet to reach the first thrust value; monitoring the turbojet to detect underspeed of the compressor; and where applicable: obtaining a second thrust value corresponding to a second operating point that guarantees a predetermined margin relative to the high limit value so as to obtain protection against underspeed of the turbojet; and controlling the turbojet to reach the second value.
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