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公开(公告)号：US20240060427A1
公开(公告)日：2024-02-22
申请号：US17892776
申请日：2022-08-22
发明人： Martin Drolet
IPC分类号： F01D21/00
CPC分类号： F01D21/003 , F05D2220/323 , F05D2260/80
摘要： A system for a gas turbine engine includes an engine control system. The engine control system includes a processor and a non-transitory memory storing instructions, which instructions when executed by the processor, cause the processor to: obtain a current engine installation configuration for the gas turbine engine and the aircraft, determine an expected normalized value of the engine parameter based on the current engine installation configuration and one or more of a normalized engine power (SHPN) of the gas turbine engine, an engine inlet temperature, an airspeed, or an altitude, determine a fully deteriorated engine (FDE) value of the engine parameter using the expected normalized value of the engine parameter, determine a current value of the engine parameter for the gas turbine engine, and determine the engine operating margin for the engine parameter based on the FDE value of the engine parameter and the current value of the engine parameter.
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公开(公告)号：US20210179294A1
公开(公告)日：2021-06-17
申请号：US16713273
申请日：2019-12-13
摘要： Systems and Methods are described for testing engine performance in-flight in an aircraft having a first engine and a second engine. The method comprises operating the first engine at a first power level in an output speed governing mode, operating the second engine at a second power level greater than the first power level in a core speed governing mode concurrently with the first engine operating at the first power level in the output speed governing mode, and performing an engine performance test on the second engine while the second engine is at the second power level in the core speed governing mode.
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公开(公告)号：US11619240B2
公开(公告)日：2023-04-04
申请号：US17407525
申请日：2021-08-20
发明人： Martin Drolet
摘要： There is disclosed a centrifugal compressor including an impeller rotatable about an axis and a diffuser downstream of the impeller. The diffuser has walls delimiting flow passages. Plasma actuators are positioned adjacent the walls and are operatively connectable to a source of electricity. The plasma actuators have a first electrode, a second electrode, and a dielectric layer therebetween. The first electrode is upstream of the second electrode. The first electrode is exposed to the flow passage. The second electrode is shielded from the flow passage by the dielectric layer. The plasma actuators are operable to generate an electric field through the dielectric layer. The plasma actuators are located closer to inlets of the flow passage than to outlets of the flow passages. A method of operating the compressor is disclosed.
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公开(公告)号：US11554874B2
公开(公告)日：2023-01-17
申请号：US17061657
申请日：2020-10-02
发明人： Martin Drolet , Frederic Fortin
摘要： There are described methods and systems for operating an aircraft having two or more engines. One method comprises operating the two or more engines of the aircraft in an asymmetric operating regime, wherein a first of the engines is in an active mode to provide motive power to the aircraft and a second of the engines is in a standby mode to provide substantially no motive power to the aircraft; governing the first engine in the active mode using a first governing logic; and governing the second engine in the standby mode using a second governing logic, the second governing logic based on a target compressor speed and variable geometry mechanism (VGM) settings that are adjusted using trim values dependent on at least one parameter of the second engine in the standby mode.
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公开(公告)号：US12031489B2
公开(公告)日：2024-07-09
申请号：US17382922
申请日：2021-07-22
CPC分类号： F02C9/22 , F04D27/001 , F05D2220/323 , F05D2260/81 , F05D2270/301 , F05D2270/304 , F05D2270/306 , F05D2270/71
摘要： A system and a method for configuring at least one variable geometry mechanism (VGM) of an aircraft engine are provided. Pass-off testing data for the aircraft engine is obtained, the pass-off testing data indicative of an actual value of at least one operating parameter of the aircraft engine. Based on the pass-off testing data, at least one trim value to be used to adjust a setting of the at least one VGM to bring the actual value of the at least one operating parameter towards a target value is determined, a running line of the aircraft engine being substantially constant when the actual value of the at least one operating parameter is at the target value. The setting of the at least one VGM is adjusted, during pass-off testing of the aircraft engine, using the at least one trim value.
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公开(公告)号：US11667392B2
公开(公告)日：2023-06-06
申请号：US16447187
申请日：2019-06-20
发明人： Martin Drolet , Sylvain Desnoyers
CPC分类号： B64D31/00 , G01K13/02 , G01L5/133 , B64D33/02 , B64D2033/0253 , G01K2205/02
摘要： Systems and methods for operating a rotorcraft engine are described herein. Measurements indicative of at least one of current temperature and current pressure at an inlet of the engine are obtained from at least one sensor while the rotorcraft is in flight. At least one current inlet loss is determined from the measurements. Current available engine power of the rotorcraft engine is determined based on the at least one current inlet losses. A visual indication of the current available engine power is produced via a flight display.
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公开(公告)号：US11549446B1
公开(公告)日：2023-01-10
申请号：US17665184
申请日：2022-02-04
摘要： A gas turbine engine includes a compressor, an annular casing surrounding the compressor, and a bleed flow adapter mounted to an exterior side of the annular casing. The annular casing includes the exterior side and an interior side. The interior side surrounds a compressor bleed cavity located downstream of at least a portion of the compressor. The bleed flow adapter is in fluid communication with the bleed cavity. The bleed flow adapter includes an inlet end, an outlet end, and an inner diameter surface extending between the inlet end and the outlet end. The inner diameter surface defines a bleed passage. The bleed flow adapter further includes a fluid port formed through the inner diameter surface. The gas turbine engine further includes a bleed flow measurement system including a first pressure sensor in fluid communication with the bleed passage of the bleed flow adapter via the fluid port.
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公开(公告)号：US12025060B2
公开(公告)日：2024-07-02
申请号：US17324404
申请日：2021-05-19
发明人： Martin Drolet , Yves Cloutier
CPC分类号： F02C9/28 , F02C7/36 , F05D2220/323 , F05D2270/021 , F05D2270/101 , F05D2270/3011 , F05D2270/303 , F05D2270/3061 , F05D2270/71
摘要： Methods and systems for operating an aircraft engine having a compressor are described. The method comprises determining, based on actual operating parameters of the aircraft engine, a compressor mass flow limit for an aerodynamic stability of the aircraft engine; determining an actual compressor mass flow of the compressor of the aircraft engine, wherein the actual compressor mass flow is based on measured values of the aircraft engine; comparing the actual compressor mass flow to the compressor mass flow limit; and governing operation of the aircraft engine to cause an alternative compressor mass flow when the actual compressor mass flow reaches or is anticipated to reach the compressor mass flow limit.
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公开(公告)号：US11739699B2
公开(公告)日：2023-08-29
申请号：US17521352
申请日：2021-11-08
发明人： Martin Drolet , Yves Cloutier
CPC分类号： F02C9/22 , F02C9/18 , F02C9/20 , F05D2200/13 , F05D2200/14 , F05D2200/211 , F05D2220/323 , F05D2270/303 , F05D2270/3015 , F05D2270/3061 , F05D2270/335 , F05D2270/71
摘要： The method can include determining a mass flow rate W of working fluid circulating through the compressor stage, determining a control parameter value associated to the geometrical configuration of the variable geometry element based on the determined value of mass flow rate W; and changing the geometrical configuration of the variable geometry element in accordance with the determined control parameter value.
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公开(公告)号：US11427353B2
公开(公告)日：2022-08-30
申请号：US16713273
申请日：2019-12-13
摘要： Systems and Methods are described for testing engine performance in-flight in an aircraft having a first engine and a second engine. The method comprises operating the first engine at a first power level in an output speed governing mode, operating the second engine at a second power level greater than the first power level in a core speed governing mode concurrently with the first engine operating at the first power level in the output speed governing mode, and performing an engine performance test on the second engine while the second engine is at the second power level in the core speed governing mode.
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