公开(公告)号：US12123373B2

公开(公告)日：2024-10-22

申请号：US16931685

申请日：2020-07-17

Applicant: Raytheon Technologies Corporation

Inventor： Marc J. Muldoon ,  Michael E. McCune ,  Andrew J. Murphy

IPC: F02C7/36 ,  F02K3/06

CPC classification number: F02K3/06 ,  F02C7/36 ,  F05D2220/323 ,  F05D2220/36 ,  F05D2240/35 ,  F05D2240/60 ,  F05D2260/40311

Abstract: A gas turbine engine includes a fan shaft rotatable about an axis, a fan connected to the fan shaft, and an outer housing surrounding the fan to define a bypass passage. A compressor section has both a low pressure compressor that is fixed to rotate with the fan shaft and a high pressure compressor. A turbine section has a low pressure turbine driving a low speed spool and a high pressure turbine driving the high pressure compressor. The low pressure turbine includes rotating blades and a static structure aft of the rotating blades. A fan drivetrain includes a geared architecture connecting the low speed spool to the fan shaft such that the fan and low pressure compressor rotate at a lower speed than the low pressure turbine. A thrust bearing is axially aft of the geared architecture and axially forward of the high pressure compressor. A tower shaft is rotatably driven by the low speed spool.

公开(公告)号：US12116938B2

公开(公告)日：2024-10-15

申请号：US18131493

申请日：2023-04-06

Applicant: ROLLS-ROYCE plc

Inventor： Mark Spruce

IPC: F02C7/36 ,  F02K3/06

CPC classification number: F02C7/36 ,  F02K3/06 ,  F05D2200/14 ,  F05D2220/323 ,  F05D2220/36 ,  F05D2260/40311

Abstract: A gas turbine engine for an aircraft including: an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan including a plurality of fan blades; a gearbox that can receive an input from the core shaft, and can output drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft; and a fan shaft mounting structure arranged to mount the fan shaft within the engine, the fan shaft mounting structure including at least two supporting bearings connected to the fan shaft.

公开(公告)号：US20240309773A1

公开(公告)日：2024-09-19

申请号：US18600600

申请日：2024-03-08

Applicant: Rolls-Royce North American Technologies Inc.

Inventor： Robert W. Heeter ,  Daniel E. Molnar, Jr. ,  Gregory J. Hebert

IPC: F01D17/00

CPC classification number: F01D17/162 ,  F01D17/00 ,  F05D2220/36 ,  F05D2260/56 ,  F05D2260/75

Abstract: A fan assembly for a gas turbine engine includes a fan duct, a fan, and an outlet guide vane assembly located in the fan duct axially downstream of the fan. The outlet guide vane assembly includes a first variable leading edge outlet guide vane that extends radially relative to the central axis and includes a leading edge portion and a fixed aft portion, the leading edge portion including a tip segment configured to rotate about a leading edge pitch axis and a hub segment located radially inward of and separate from the tip segment, the hub segment configured to independently rotate about the leading edge pitch axis relative to the tip segment. The assembly further includes two unique actuation assemblies each including a cam rod and cams located thereon that are configured to rotate a respective tip or hub segment.

公开(公告)号：US20240295177A1

公开(公告)日：2024-09-05

申请号：US18003403

申请日：2021-06-29

Applicant: SAFRAN AIRCRAFT ENGINES

Inventor： Clement BOUROLLEAU ,  Clementine ARNOLD ,  Alexandre BRANCO ,  Serge DOMINGUES ,  Thomas DUPAYS ,  Vincent Pascal FIORE ,  Baptiste LEBLON ,  Anwer SIRAJ

IPC: F01D11/12 ,  F01D21/04

CPC classification number: F01D11/125 ,  F01D21/045 ,  F05D2220/36 ,  F05D2300/2102 ,  F05D2300/224

Abstract: A housing for an aircraft turbomachine, including an annular casing extending around an axis A and having an internal annular surface, the housing also including an annular abradable support cartridge that is fixed against the internal annular surface, the abradable support cartridge has a reinforced coating including a fibrous texture reinforcement embedded in a resin matrix, the fibrous texture reinforcement including a stack of fibrous texture plies, wherein the stack of plies includes at least one ply made of Kevlar® or of glass fibres.

公开(公告)号：US12055153B1

公开(公告)日：2024-08-06

申请号：US18528977

申请日：2023-12-05

Applicant: General Electric Company

Inventor： Suryarghya Chakrabarti ,  Nicholas M. Daggett ,  Zachary Pebley

IPC: F04D27/00 ,  F01D5/10 ,  F04D29/66

CPC classification number: F04D27/002 ,  F01D5/10 ,  F04D29/668 ,  F05D2220/36 ,  F05D2260/70 ,  F05D2260/96

Abstract: A variable pitch airfoil assembly for an engine includes a disk having an annular shape extending about an axial direction and an airfoil coupled to the disk via a platform. The airfoil extends outwardly from the disk in a radial direction and is rotatable relative to the disk about a pitch axis. The variable pitch airfoil assembly further includes a damping element positioned at least partially within the disk exterior of and adjacent to a perimeter of the platform so as to provide vibration damping by friction between the damping element, the platform, and the disk while also allowing for a pitch change of the airfoil.

公开(公告)号：US12044167B2

公开(公告)日：2024-07-23

申请号：US18171073

申请日：2023-02-17

Applicant: RTX Corporation

Inventor： Colin G. Amadon ,  Michael C. Firnhaber

IPC: F02C7/04 ,  F04D19/00 ,  F04D29/54

CPC classification number: F02C7/04 ,  F04D29/545 ,  F04D19/002 ,  F05D2220/36 ,  F05D2240/12 ,  F05D2260/606

Abstract: A stator configuration for a gas turbine engine including: a splitter segment, the splitter segment extending from a forward end to a rearward end, the splitter segment being configured to split an incoming fluid flow into a first, radially outer stream and a second, radially inner stream; a forward most first stator extending radially outwardly from the splitter segment, the forward most first stator being completely located downstream from the forward end of the splitter segment; a forward most second stator extending radially inwardly from the splitter segment, the splitter segment, the forward most first stator and the forward most second stator being formed as a single, integrally formed structure, the forward most first stator positioned closer to the forward end relative to a distance between the forward most second stator and the forward end and the forward most second stator positioned closer to the rearward end relative to a distance between the forward most first stator and the rearward end, wherein the splitter segment defines a radially outward boundary of the second, radially inner stream such that an outer tip of a rotor passes in close proximity to the splitter segment at the radially outward boundary of the second, radially inner stream defined by the splitter segment; and a sacrificial rub layer located within a pocket formed in a radially inner surface of the splitter segment that faces the outer tip of the rotor at the radially outward boundary of the second, radially inner stream defined by the splitter segment.

公开(公告)号：US12037921B2

公开(公告)日：2024-07-16

申请号：US17817511

申请日：2022-08-04

Applicant: General Electric Company

Inventor： Nicholas J. Kray ,  David W. Crall ,  Anthony L. DiPietro

IPC: F01D5/14 ,  F01D25/32 ,  F02K3/06

CPC classification number: F01D5/146 ,  F02K3/06 ,  F01D25/32 ,  F05D2220/32 ,  F05D2220/36 ,  F05D2240/30

Abstract: A fan for a turbine engine having a core air flowpath and a bypass air flowpath. The fan includes a disk, a plurality of fan blades, and a plurality of splitter blades. The disk rotates about a centerline axis. The plurality of fan blades are coupled to the disk. The plurality of splitter blades are coupled to the disk and positioned between the plurality of fan blades. The plurality of splitter blades are axially aligned on the disk at a splitter blade axial distance from an annular inlet of the core air flowpath such that the plurality of splitter blades direct debris that passes through the fan into the bypass air flowpath.

公开(公告)号：US20240229655A9

公开(公告)日：2024-07-11

申请号：US17769617

申请日：2020-10-16

Applicant: SAFRAN AIRCRAFT ENGINES

Inventor： Vivien Mickaël COURTIER ,  Christophe Paul JACQUEMARD ,  Vincent JOUDON

IPC: F01D5/30 ,  B64C11/06 ,  F01D5/32 ,  F01D7/00

CPC classification number: F01D5/3007 ,  B64C11/06 ,  F01D5/326 ,  F01D7/00 ,  F05D2220/36

Abstract: Assembly for a turbomachine blade, comprising: a fastener (9) defining a pocket (10) for receiving a blade root (11), and comprising two flanks defining between them a passage leading into the pocket (10) and forming stops preventing the root (11) from exiting from the pocket (10) via the passage, a shim (26a, 26b) adapted to be received in the pocket (10) with the root (11), the shim (26a, 26b) having a first support surface (28a, 28b) to bear on the root (11) and a second support surface (30a, 30b) opposite the first support surface (28a, 28b) to bear against the fastener (9), the support surfaces being fixed relative to each other and oriented such that movement of the shim (26a, 26b) in the pocket (10) relative to the fastener (9) varies the value of force exerted by the shim (26a, 26b) on the root (11).

公开(公告)号：US12031503B2

公开(公告)日：2024-07-09

申请号：US17545310

申请日：2021-12-08

Applicant: Rolls-Royce Deutschland Ltd & Co KG

Inventor： Tomasz Burza ,  Moritz Wirth

IPC: F02K3/06 ,  F01D5/02 ,  F01D25/32

CPC classification number: F02K3/06 ,  F01D5/02 ,  F01D25/32 ,  F05D2220/36 ,  F05D2260/31 ,  F05D2260/40311

Abstract: A fan of a gas turbine engine, which has a fan disk with a multiplicity of fastening elements which are spaced apart in a circumferential direction and which project axially forwardly from the fan disk, and a nose cone which is arranged upstream of the fan disk and which is connected by means of the fastening elements to the fan disk. On an axially front side of the fan disk, there is formed a periphery which runs in encircling fashion in the circumferential direction and which runs radially at the inside in relation to the axially rear end region of the nose cone, wherein the periphery which runs in encircling fashion in the circumferential direction forms a concave indentation, in such a way that water which ingresses into a gap between the axially rear edge of the nose cone and the fan disk passes into the concave indentation.

公开(公告)号：US20240209777A1

公开(公告)日：2024-06-27

申请号：US18455137

申请日：2023-08-24

Applicant: General Electric Company

Inventor： Arthur William SIBBACH ,  Brandon Wayne MILLER

IPC: F02C7/05 ,  F02C6/20 ,  F02C7/04 ,  F02K3/00

CPC classification number: F02C7/05 ,  F02C6/20 ,  F02C7/04 ,  F02K3/00 ,  F05D2220/36

Abstract: A turbofan engine is provided. The turbofan engine includes a fan comprising a plurality of fan blades; a turbomachine operably coupled to the fan for driving the fan, the turbomachine comprising a compressor section, a combustion section, and a turbine section in serial flow order and together defining a core air flowpath; a nacelle surrounding and at least partially enclosing the fan; an inlet pre-swirl feature located upstream of the plurality of fan blades, the inlet pre-swirl feature attached to or integrated into the nacelle; and a means for directing incoming objects towards an outer portion of the turbofan engine in communication with the inlet pre-swirl feature.