公开(公告)号：US20250012236A1

公开(公告)日：2025-01-09

申请号：US18430907

申请日：2024-02-02

Applicant: General Electric Company

Inventor： Scott Alan Schimmels ,  Jeffrey Douglas Rambo ,  Timothy Richard DePuy ,  Steven B. Morris

IPC: F02K3/115 ,  F02K3/04

Abstract: A heat exchanger positioned within an annular duct of a gas turbine engine is provided. The heat exchanger extends substantially continuously along the circumferential direction and defining a heat exchanger height equal to at least 10% of a duct height. An effective transmission loss (ETL) for the heat exchanger positioned within the annular duct is between 5 decibels and 1 decibels for an operating condition of the gas turbine engine. The heat exchanger includes a heat transfer section defining an acoustic length (Li), and wherein an Operational Acoustic Reduction Ratio (OARR) is greater than or equal to 0.75 to achieve the ETL at the operating condition.

公开(公告)号：US12123432B2

公开(公告)日：2024-10-22

申请号：US18201875

申请日：2023-05-25

Applicant: RAYTHEON TECHNOLOGIES CORPORATION ,  MTU Aero Engines AG

Inventor： Bruce L. Morin ,  Detlef Korte

IPC: F02C7/36 ,  F01D5/06 ,  F01D5/14 ,  F01D15/12 ,  F01D25/24 ,  F02C3/04 ,  F02C3/107 ,  F02K3/04 ,  F02K3/06 ,  F04D25/04 ,  F04D29/053 ,  F04D29/32 ,  F04D29/66 ,  G06F30/17

CPC classification number: F04D29/663 ,  F01D5/06 ,  F01D5/14 ,  F01D15/12 ,  F01D25/24 ,  F02C3/04 ,  F02C3/107 ,  F02C7/36 ,  F02K3/04 ,  F02K3/06 ,  F04D25/045 ,  F04D29/053 ,  F04D29/321 ,  F04D29/325 ,  G06F30/17 ,  F05D2220/323 ,  F05D2230/50 ,  F05D2260/40311 ,  F05D2260/96 ,  F05D2270/304 ,  F05D2270/333

Abstract: An aircraft system includes, among other things, an aircraft and a gas turbine engine coupled to the aircraft. The gas turbine engine includes a propulsor section including a propulsor, a compressor section, a turbine section including a first turbine and a second turbine, and a gear reduction between the propulsor and the second turbine. The second turbine includes a number of turbine blades in each of a plurality of rows of the second turbine. The second turbine blades operating at least some of the time at a rotational speed. The number of blades and the rotational speed being such that the following formula holds true for a majority of the blade rows of the second turbine: 5500 Hz≤(number of blades×speed)/60 sec≤10000 Hz. The gas turbine engine is rated to produce 15,000 pounds of thrust or more.

公开(公告)号：US12116940B2

公开(公告)日：2024-10-15

申请号：US17215575

申请日：2021-03-29

Applicant: Pratt & Whitney Canada Corp.

Inventor： Philippe Beauchesne-Martel ,  Poi Loon Tang ,  Andrew Thompson ,  Ghislain Plante

IPC: F02C9/20 ,  F02K3/04 ,  F04D27/02

CPC classification number: F02C9/20 ,  F02K3/04 ,  F04D27/0246 ,  F05D2270/023 ,  F05D2270/052 ,  F05D2270/053 ,  F05D2270/3011 ,  F05D2270/303 ,  F05D2270/304 ,  F05D2270/335 ,  F05D2270/71

Abstract: Herein provided are methods and systems for controlling an engine having a variable geometry mechanism. A pressure ratio between a first pressure at an inlet of the engine and a predetermined reference pressure is determined. An output power for the engine is determined. The output power is adjusted based at least in part on the pressure ratio to obtain a corrected output power. A position control signal for a variable geometry mechanism of the engine is generated based on the corrected output power and the pressure ratio. The position control signal is output to a controller of the engine to control the variable geometry mechanism.

公开(公告)号：US12116929B2

公开(公告)日：2024-10-15

申请号：US17578795

申请日：2022-01-19

Applicant: General Electric Company

Inventor： Brandon Wayne Miller ,  Scott David Hunter ,  Patrick Michael Marrinan ,  Scott Alan Schimmels

IPC: F02C6/08 ,  B64D13/06 ,  F02C9/00 ,  F02C9/18 ,  F02K3/04

CPC classification number: F02C6/08 ,  B64D13/06 ,  F02C9/00 ,  F02C9/18 ,  F02K3/04 ,  B64D2013/0603 ,  B64D2013/0607 ,  B64D2013/0618 ,  B64D2013/0648 ,  F05D2220/32 ,  F05D2220/76

Abstract: A gas turbine engine includes a turbomachine, the turbomachine defining a core flow therethrough during operation. A first heat exchange assembly is in fluid communication with the turbomachine for receiving a first bleed flow from the turbomachine. A second heat exchange assembly is in fluid communication with the turbomachine for receiving a second bleed flow from the turbomachine. A first flow outlet is provided for receiving the first bleed flow from the first heat exchange assembly and providing the first bleed flow to a first aircraft flow assembly. A second flow outlet is provided for receiving the second bleed flow and providing the second bleed flow from the second heat exchange assembly to a second aircraft flow assembly.

公开(公告)号：US20240228048A9

公开(公告)日：2024-07-11

申请号：US18493186

申请日：2023-10-24

Applicant: Boom Technology, Inc.

Inventor： Nathaniel Blake Scholl

IPC: B64D27/16 ,  B64C15/14 ,  B64C30/00 ,  F02C7/32 ,  F02K3/04

CPC classification number: B64D27/16 ,  B64C15/14 ,  B64C30/00 ,  F02C7/32 ,  F02K3/04

Abstract: An aircraft propulsion system having a variable airflow capture area is provided. The propulsion system includes a main propulsion source and an auxiliary propulsion source. In a first mode, the auxiliary propulsion source is stowed within an aerodynamic profile of the aircraft, and the main propulsion source provides all of the propulsion force for powering flight of the aircraft. In a second mode, the auxiliary propulsion source is deployed to augment the airflow capture area of the main propulsion source and increase an overall airflow capture area of the propulsion system. In the second mode, the auxiliary power source is operated by power extracted from the main propulsion source. The main propulsion source may include one or more low bypass ratio engines. The auxiliary power source may include one or more auxiliary thrust fans coupled at a plurality of locations on the aircraft.

公开(公告)号：US20240092494A1

公开(公告)日：2024-03-21

申请号：US18513916

申请日：2023-11-20

Applicant: General Electric Company

Inventor： Randy M. Vondrell ,  Anthony Austin Bouvette ,  Glenn David Crabtree

IPC: B64D27/24 ,  B60L50/16 ,  B64C21/06 ,  B64D27/12 ,  B64D27/18 ,  F02C6/20 ,  F02K3/04 ,  H02G5/10 ,  H02K7/18

CPC classification number: B64D27/24 ,  B60L50/16 ,  B64C21/06 ,  B64D27/12 ,  B64D27/18 ,  F02C6/206 ,  F02K3/04 ,  H02G5/10 ,  H02K7/1823 ,  B64D2027/026

Abstract: A propulsion system for an aircraft can include an electric power source and an electric propulsion assembly having an electric motor and a propulsor. The propulsor can be powered by the electric motor. An electric power bus can electrically connect the electric power source to the electric propulsion assembly. The electric power source can be configured to provide electrical power to the electric power bus. An inverter converter controller can be positioned along the electric power bus and can be electrically connected to the electric power source at a location downstream of the electric power source and upstream of the electric propulsion assembly.

公开(公告)号：US11840958B2

公开(公告)日：2023-12-12

申请号：US17393508

申请日：2021-08-04

Applicant: RTX Corporation

Inventor： Daniel Bernard Kupratis ,  Paul R. Hanrahan ,  Christopher J. Hanlon

IPC: F02C6/02 ,  F02C3/10 ,  F02C7/36 ,  F02C3/14 ,  F02C7/04 ,  F02K3/04

CPC classification number: F02C3/10 ,  F02C6/02 ,  F02C7/36 ,  F02C3/14 ,  F02C7/04 ,  F02K3/04

Abstract: A gas turbine engine includes a fan positioned at an engine central longitudinal axis, and a fan drive turbine located at the engine central longitudinal axis and configured to drive rotation of the fan. A gas generator is non-coaxial with the fan drive turbine and operably connected to the fan drive turbine such that exhaust from the gas generator drives rotation of the fan drive turbine. An auxiliary power core is located at the engine central longitudinal axis, and one or more bleed passages connect the gas generator and the auxiliary power core. The one or more bleed passages are configured to selectably combine a bleed airflow from the gas generator and an auxiliary core airflow at the auxiliary power core to direct the combined airflow to the fan drive turbine to increase output of the fan drive turbine.

公开(公告)号：US11629666B2

公开(公告)日：2023-04-18

申请号：US17184903

申请日：2021-02-25

Applicant: ISRAEL AEROSPACE INDUSTRIES LTD.

Inventor： Aviad Brandstein ,  Avi Ponchek

IPC: F02K3/04 ,  F01D5/02 ,  F01D9/02 ,  F01D15/10 ,  F02C3/04 ,  F02K3/06

Abstract: A method for converting a turbofan engine including providing a turbofan engine and converting the turbofan engine. The turbofan engine includes a core engine (including at least one high pressure spool assembly and a combustion chamber), and an unmodified fan configured for providing at least a bypass flow bypassing the core engine, the fan being mechanically coupled to a low pressure turbine that is in turn driven by the core engine. The conversion includes modifying or replacing the unmodified fan to provide a modified fan, the modified fan configured for generating a reduced bypass flow with respect to said fan bypass flow during operation of the converted turbofan engine corresponding to at least one set of engine conditions, enabling said low pressure turbine to generate an excess shaft power above a baseline shaft power required for driving the modified fan during operation of the converted turbofan engine.

公开(公告)号：US11572834B2

公开(公告)日：2023-02-07

申请号：US17152139

申请日：2021-01-19

Applicant: Rolls-Royce plc

Inventor： Manuel Diaz Brito

IPC: F02C7/18 ,  F02K3/04

Abstract: Gas turbine engine including a nacelle and an engine core within the nacelle. The engine core defines a principal rotational axis along its length. The engine core and nacelle define a bypass passage therebetween. The gas turbine engine further includes a cooling system including a cooling duct, which duct defines an inlet for receiving bypass air from the bypass passage at an upstream location and an outlet for discharging the bypass air at a downstream location. The cooling duct extends, relative to the principal axis, axially and circumferentially around a section of the engine core. The cooling duct comprises: first portion that extends at least axially relative to the principal rotational axis; second portion downstream of the first portion that extends circumferentially around the engine core relative to the principal rotational axis; and third portion downstream of second portion that extends at least axially relative to the principal rotational axis.

公开(公告)号：US11525406B2

公开(公告)日：2022-12-13

申请号：US17023860

申请日：2020-09-17

Applicant: Raytheon Technologies Corporation

Inventor： Michael E. McCune

IPC: F02C7/36 ,  F16H1/28 ,  F01D15/12 ,  F02K3/072 ,  F16H57/00 ,  F16H1/46 ,  F02K3/04

Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section, a compressor section, and a turbine section including a fan drive turbine that drives the fan through a gear reduction. The gear reduction includes at least two double helical gears in meshed engagement, each of the at least two double helical gears having a first plurality of gear teeth separated from a second plurality of gear teeth such that a first end of the first plurality of gear teeth and a first end of the second plurality of gear teeth are spaced apart by an axial distance. Each of the first plurality of gear teeth is offset a first circumferential offset distance in relation to the next gear tooth of the second plurality of gear teeth when moving in a circumferential direction relative to respective axes.