公开(公告)号：US11939924B2

公开(公告)日：2024-03-26

申请号：US17958682

申请日：2022-10-03

Applicant: Raytheon Technologies Corporation

Inventor： Nicholas W. Kantany ,  Edward Boucher ,  Kristine Marie Carnavos ,  Robert Russell Mayer ,  Donald W. Peters

IPC: F02C7/28 ,  F02C9/20 ,  F02K3/075 ,  F02K3/115

CPC classification number: F02C7/28 ,  F02C9/20 ,  F02K3/075 ,  F02K3/115

Abstract: A system for modulating air flow in a gas turbine engine is provided. The system may include a seal wall comprising an opening, a seal door configured to slideably engage the seal wall, and an actuator configured to move the seal door over the opening. In various embodiments, the system may include a surface forward of the seal door. The seal door may be configured to seal a passage through the surface and the opening of the seal wall. A track may be disposed under the seal door. The track may comprise cobalt. Rollers may be coupled to the seal door with the rollers on the track. The seal door may comprise a nickel-chromium alloy. A sync ring may be coupled to the seal door. The actuator may be coupled through the sync ring to the seal door.

公开(公告)号：US11846250B2

公开(公告)日：2023-12-19

申请号：US17907587

申请日：2021-03-19

Applicant: SAFRAN AIRCRAFT ENGINES

Inventor： Alexandre Gérard François Couilleaux ,  Mathieu Lasne

IPC: F02K3/115 ,  F28F27/00 ,  F02C7/18

CPC classification number: F02K3/115 ,  F05D2220/323 ,  F05D2260/213

Abstract: A turbofan aircraft propulsion assembly includes a retractable heat exchanger. The heat exchanger can be deployed in a secondary flow path of the propulsion assembly so as to cool a fluid circulating in a circuit of a transfer module of the heat exchanger using a secondary flow circulating in the secondary flow path.

公开(公告)号：US11828228B2

公开(公告)日：2023-11-28

申请号：US18110664

申请日：2023-02-16

Applicant: ROLLS-ROYCE PLC

Inventor： Natalie C Wong ,  Thomas S Binnington ,  David A Jones ,  Daniel Blacker

IPC: F02C7/18 ,  F02C7/14 ,  F02C7/04 ,  B64D27/10 ,  B64D33/02 ,  B64D33/04 ,  B64D33/08 ,  B64D27/20 ,  B64D33/10 ,  F02K3/06 ,  F02K3/075 ,  F02C7/08 ,  F02C7/12 ,  F01D9/06 ,  F02K3/115

CPC classification number: F02C7/04 ,  B64D27/10 ,  B64D27/20 ,  B64D33/02 ,  B64D33/04 ,  B64D33/08 ,  B64D33/10 ,  F01D9/065 ,  F02C7/08 ,  F02C7/12 ,  F02C7/18 ,  F02K3/06 ,  F02K3/075 ,  B64D2033/024 ,  B64D2033/0286 ,  F02C7/14 ,  F02K3/115 ,  F05D2220/323 ,  F05D2220/36 ,  F05D2260/213

Abstract: An aircraft includes a machine body that encloses a turbofan gas turbine engine and a plurality of ancillary systems. The turbofan gas turbine engine includes, in axial flow sequence, a first heat exchanger module, a fan assembly, a compressor module, a combustor module, a turbine module, and an exhaust module. The aircraft includes a second heat exchanger module. The machine body comprises a single fluid inlet aperture, with the fluid inlet aperture being configured to allow a fluid cooling flow to enter the machine body and to pass through the first heat exchanger module. When a temperature of the fluid cooling flow is less than a temperature of a fluid to be cooled, the fluid to be cooled is directed to the first heat exchanger module, and when a temperature of the fluid cooling flow is greater than a temperature of the fluid to be cooled, the fluid to be cooled is directed to the second heat exchanger module and cooled using a fuel supply for the gas turbine engine.

公开(公告)号：US20230193822A1

公开(公告)日：2023-06-22

申请号：US18110664

申请日：2023-02-16

Applicant: ROLLS-ROYCE PLC

Inventor： Natalie C WONG ,  Thomas S BINNINGTON ,  David A JONES ,  Daniel BLACKER

IPC: F02C7/04 ,  F02C7/18 ,  B64D27/10 ,  B64D33/02 ,  B64D33/04 ,  B64D33/08 ,  B64D27/20 ,  B64D33/10 ,  F02K3/06 ,  F02K3/075 ,  F02C7/08 ,  F02C7/12 ,  F02C7/14 ,  F01D9/06

CPC classification number: F02C7/04 ,  F02C7/18 ,  B64D27/10 ,  B64D33/02 ,  B64D33/04 ,  B64D33/08 ,  B64D27/20 ,  B64D33/10 ,  F02K3/06 ,  F02K3/075 ,  F02C7/08 ,  F02C7/12 ,  F02C7/14 ,  F01D9/065 ,  F05D2260/213 ,  B64D2033/0286 ,  F05D2220/323 ,  B64D2033/024 ,  F05D2220/36 ,  F02K3/115

Abstract: An aircraft includes a machine body that encloses a turbofan gas turbine engine and a plurality of ancillary systems. The turbofan gas turbine engine includes, in axial flow sequence, a first heat exchanger module, a fan assembly, a compressor module, a combustor module, a turbine module, and an exhaust module. The aircraft includes a second heat exchanger module. The machine body comprises a single fluid inlet aperture, with the fluid inlet aperture being configured to allow a fluid cooling flow to enter the machine body and to pass through the first heat exchanger module. When a temperature of the fluid cooling flow is less than a temperature of a fluid to be cooled, the fluid to be cooled is directed to the first heat exchanger module, and when a temperature of the fluid cooling flow is greater than a temperature of the fluid to be cooled, the fluid to be cooled is directed to the second heat exchanger module and cooled using a fuel supply for the gas turbine engine.

公开(公告)号：US20190128189A1

公开(公告)日：2019-05-02

申请号：US15796991

申请日：2017-10-30

Applicant: General Electric Company

Inventor： Jeffrey Douglas Rambo

IPC: F02C7/18 ,  F02K3/06 ,  F01D17/16 ,  F02C7/14

CPC classification number: F02C7/185 ,  F01D17/16 ,  F01D17/162 ,  F02C7/14 ,  F02C7/18 ,  F02K3/06 ,  F02K3/075 ,  F02K3/077 ,  F02K3/115 ,  F05D2220/323 ,  F05D2240/12 ,  F05D2260/213

Abstract: A gas turbine engine includes an outer nacelle; a fan at least partially surrounded by the outer nacelle; and a turbomachine drivingly coupled to the fan and at least partially surrounded by the outer nacelle. The outer nacelle defines a bypass airflow passage with the turbomachine. The turbomachine includes a compressor section defining in part a core air flowpath. The turbomachine also includes a heat sink heat exchanger; and a thermal management duct assembly defining a thermal management duct flowpath extending between an inlet and an outlet and positioned between the core air flowpath and the bypass airflow passage along the radial direction, the outlet selectively in airflow communication with a core compartment of the turbomachine, and the heat sink heat exchanger positioned in thermal communication with the thermal management duct flowpath for transferring heat to an airflow through the thermal management duct flowpath during operation.

公开(公告)号：US10054054B2

公开(公告)日：2018-08-21

申请号：US14847463

申请日：2015-09-08

Applicant: Rolls-Royce Deutschland Ltd & Co KG

Inventor： Matthijs Van Der Woude

IPC: F02C7/18 ,  F01D11/24 ,  F02K3/115 ,  F01D25/24

CPC classification number: F02C7/18 ,  F01D11/24 ,  F01D25/24 ,  F02C7/185 ,  F02K3/115 ,  F05D2260/232 ,  F05D2260/60 ,  Y02T50/671 ,  Y02T50/675

Abstract: An air guiding device for an aircraft engine, characterized in that it has at least one valve element which is moveable between a first position, in particular an open position, and a second position, in particular a closed position, wherein in the first position of the at least one valve element, through a combined effect with at least one first air guiding element that is arranged inside the air flow, a cooling air flow can be guided from an air flow into a hollow space between the housing and the core engine of the aircraft engine, and the at least one first air guiding element is aligned at least approximately in parallel to the air flow for the purpose of reducing flow resistance. The invention also relates to an aircraft engine comprising at least one air guiding device.

公开(公告)号：US09945247B2

公开(公告)日：2018-04-17

申请号：US14813755

申请日：2015-07-30

Applicant: ROLLS-ROYCE PLC

Inventor： Ajith Appukuttan ,  Benjamin Keeler ,  Paolo Vanacore

IPC: F01D25/02 ,  F01D9/02 ,  F01D15/08 ,  F01D17/10 ,  F01D25/10 ,  F02C7/141 ,  F02K3/115 ,  F02C7/047 ,  B64D13/06

CPC classification number: F01D25/02 ,  B64D2013/0607 ,  F01D9/02 ,  F01D15/08 ,  F01D17/105 ,  F01D25/10 ,  F02C7/047 ,  F02C7/141 ,  F02K3/115 ,  F05D2220/32 ,  F05D2240/10 ,  F05D2260/208 ,  Y02T50/675

Abstract: An anti-icing system for an engine section stator of a gas turbine engine. The system includes an environmental control system pre-cooler heat exchange system and a conduit. The environmental control system pre-cooler heat exchange system is configured to exchange heat between air bled from a compressor of the engine and bypass duct air. The conduit is configured to exchange heat from the pre-cooler heat exchange system to a heat transfer medium. The conduit is also configured to transfer the heat from the heat transfer medium to the engine section stator.

公开(公告)号：US20180038243A1

公开(公告)日：2018-02-08

申请号：US15229466

申请日：2016-08-05

Applicant: GENERAL ELECTRIC COMPANY

Inventor： Jeffrey Douglas Rambo ,  Andrew Martin ,  Curtis Walton Stover ,  Jared Matthew Wolfe ,  Mohammed El Hacin Sennoun

IPC: F01D25/14 ,  F01D17/10 ,  F01D17/14 ,  F04D29/54 ,  F04D19/00 ,  F02K3/06 ,  F01D25/12 ,  F01D17/16

CPC classification number: F01D25/14 ,  F01D17/105 ,  F01D17/145 ,  F01D17/162 ,  F01D25/12 ,  F02C6/08 ,  F02C7/14 ,  F02C7/18 ,  F02K3/06 ,  F02K3/115 ,  F04D19/002 ,  F04D29/542 ,  F05D2220/32 ,  F05D2260/213 ,  F05D2260/601 ,  F05D2260/98 ,  Y02T50/676

Abstract: A heat exchanger assembly for a gas turbine engine that includes an outer engine case. The heat exchanger assembly includes at least one cooling channel, the at least one cooling channel is configured to receive a flow of fluid to be cooled. At least one first coolant flow duct that is configured to receive a flow of a first coolant, wherein the at least one cooling channel is disposed between a first inlet and a first outlet. The heat exchanger assembly further include at least one second coolant flow duct that is configured to receive a flow of a second coolant, wherein the at least one cooling channel is disposed between a second inlet and a second outlet.

公开(公告)号：US09885284B2

公开(公告)日：2018-02-06

申请号：US14623037

申请日：2015-02-16

Applicant: Airbus Operations (SAS)

Inventor： Olivier Pelagatti ,  Thomas Deguin ,  Ronan Bodet

IPC: F02C6/08 ,  F02C3/10 ,  F02C6/06 ,  B64D13/04 ,  F02C6/04 ,  B64D13/02 ,  B64D13/06 ,  F02C9/18

CPC classification number: F02C6/08 ,  B64D13/02 ,  B64D13/04 ,  B64D13/06 ,  B64D2013/0644 ,  B64D2013/0648 ,  F02C3/10 ,  F02C6/04 ,  F02C6/06 ,  F02C9/18 ,  F02K3/115

Abstract: A turbojet comprising a high-pressure compressor with several compression stages in which the pressure rises in a direction of air flow, from a low pressure to a high pressure, passing through intermediate pressures, and a system to bleed air in the turbojet and deliver the air to an air system. The bleeding system comprises a first air intake to bleed air at low pressure, a second air intake to bleed air at an intermediate pressure, a first valve and a second valve having inlets connected to the first air intake, a compressor having an inlet connected to the second valve outlet, a high-pressure valve connected to the second air intake, and a control valve connected to the air system, and a controller to control the opening and closing of the valves depending on the pressures at the first and second air intakes.

公开(公告)号：US09856793B2

公开(公告)日：2018-01-02

申请号：US14745564

申请日：2015-06-22

Applicant: United Technologies Corporation

Inventor： Mark F. Zelesky ,  Gabriel L. Suciu ,  Brian D. Merry

IPC: F02C6/04 ,  F02C6/08 ,  F02K3/115 ,  F02C7/143

CPC classification number: F02C6/08 ,  F02C7/143 ,  F02K3/115 ,  F05D2260/211 ,  Y02T50/676

Abstract: A gas turbine engine comprises a main compressor section having a high pressure compressor with a downstream discharge, and more upstream locations. A turbine section has a high pressure turbine. A tap taps air from at least one of the more upstream locations in the compressor section, passes the tapped air through a heat exchanger and then to a cooling compressor. The cooling compressor compresses air downstream of the heat exchanger, and delivers air into the high pressure turbine. The cooling compressor includes a downstream connection that delivers discharge pressure air to an upstream location in the high pressure turbine and a second tap from an intermediate pressure location within the cooling compressor. The second tap is connected to a downstream location within the high pressure turbine. An intercooling system for a gas turbine engine is also disclosed.