-
公开(公告)号：US12292196B2
公开(公告)日：2025-05-06
申请号：US18345941
申请日：2023-06-30
Applicant: RAYTHEON TECHNOLOGIES CORPORATION
Inventor： Timothy S. Snyder
Abstract: A fuel nozzle guide assembly for a gas turbine engine is provided. The fuel nozzle guide assembly includes a housing; a tube arranged in the housing and defining a portion of a first fluid passage therein, the first fluid passage configured to include a first fluid, wherein a second fluid passage is defined, in part, between an exterior surface of the tube and an interior surface of the housing, the second fluid passage configured to include a second fluid; an air inflow tube, the air inflow tube defining a central air passage and configured to include a third fluid; an air inflow assembly defining a third fluid passage and configured to include a fourth fluid; and a nozzle outlet configured to receive each of the first fluid, the second fluid, the third fluid, and the fourth fluid to cause mixing thereof.
-
公开(公告)号：US12291337B2
公开(公告)日：2025-05-06
申请号：US17675138
申请日：2022-02-18
Applicant: Raytheon Technologies Corporation
Inventor： Alan F. Retersdorf
Abstract: An environmental control system (ECS) for use with a gas turbine engine has an air cycle system (ACS) and a vapor cycle system (VCS). The VCS has along a vapor compression flowpath: a VCS compressor; a heat donor leg of a VCS condenser; an expansion device; and a heat receiving leg of a VCS evaporator. The ACS has along a bleed flowpath: a bleed air inlet; a primary heat exchanger; an ACS compressor; a secondary heat exchanger; a turbine coupled to the ACS compressor to drive the ACS compressor; a heat donor leg of the VCS evaporator; a water collector; and a heat receiving leg of the VCS condenser.
-
公开(公告)号：US12270543B2
公开(公告)日：2025-04-08
申请号：US17408049
申请日：2021-08-20
Applicant: Raytheon Technologies Corporation
Inventor： Lawrence A. Binek , Sean R. Jackson
Abstract: A monolithic combustion liner for use in a gas turbine engine includes fuel channels integrated into the wall of the combustion liner. The integrated fuel channels can have an aerodynamic shape to reduce flow losses of cooling air flowing around the exterior of the combustion liner. The monolithic combustion liner allows more cooling air to flow around the combustion liner, increasing the cooling of the combustor region of the gas-turbine engine.
-
公开(公告)号：US12216184B2
公开(公告)日：2025-02-04
申请号：US18310937
申请日：2023-05-02
Applicant: Raytheon Technologies Corporation
Inventor： Andrew DeBiccari , Zhong Ouyang
Abstract: A system for magnetically inspecting a metallic component uses a manipulator configured to manipulate a relative position between a part fixture that holds the metallic component and a probe fixture that holds a magnetic probe, thereby causing the probe tip to trace an inspection route along the surface of the metallic component so that the probe tip contacts the metallic component such that an angular difference between the probe axis and a vector normal to the surface is less than a predetermined angle delta. The magnetic probe has a probe tip that measures magnetic permeability of the metallic component along the inspection route, which the controller receives. A method of performing the magnetic inspection is also disclosed.
-
公开(公告)号：US12215865B2
公开(公告)日：2025-02-04
申请号：US17714064
申请日：2022-04-05
Applicant: Raytheon Technologies Corporation
Inventor： Tianli Zhu , Richard Wesley Jackson, III , John A. Sharon , James T. Beals , Brian T. Hazel
IPC: F23R3/00
Abstract: A gas turbine engine component having a substrate; a thermal barrier coating on the substrate having a porous microstructure; and a reflective layer conforming to the porous microstructure of the thermal barrier coating, wherein the reflective layer comprises a conforming nanolaminate defined by alternating layers of platinum group metal materials selected from the group consisting of platinum group metal-based alloys, platinum group metal intermetallic compounds, mixtures of platinum group metal with metal oxides and combinations thereof. A capping layer can be added over the reflective layer. A supporting layer can be added between the reflective layer and the thermal barrier coating. A process is also disclosed.
-
Patent Agency Ranking
公开(公告)号：US12209997B2
公开(公告)日：2025-01-28
申请号：US17730114
申请日：2022-04-26
Applicant: RAYTHEON TECHNOLOGIES CORPORATION
Inventor： Derek J. Michaels , Elizabeth F. Vinson , Zhong Ouyang
Abstract: An inspection device is disclosed herein. The inspection device may comprise: a support structure; a motor; a shaft operably coupled to the motor, the shaft extending from a first side of the support structure to a second side of the support structure, the shaft configured to couple to a bladed rotor; an optical device moveably coupled to the support structure; and a broad-band energy source configured to generate acoustic energy through the bladed rotor during inspection.
-
公开(公告)号：US20250027422A1
公开(公告)日：2025-01-23
申请号：US18224908
申请日：2023-07-21
Applicant: Raytheon Technologies Corporation
Inventor： Andrew E. Breault , William K. Ackermann
Abstract: A buffer air assembly for an aircraft engine includes a low-pressure header, a high-pressure header, a low-pressure bleed air source, a high-pressure bleed air source, and an electric buffer compressor. The low-pressure header is connected to at least one low-pressure bearing compartment. The high-pressure header is connected to at least one high-pressure bearing compartment. The low-pressure bleed air source is connected to the low-pressure header. The low-pressure bleed air source is configured to direct a low-pressure buffering air to the at least one low-pressure bearing compartment through the low-pressure header. The high-pressure bleed air source is configured to direct a high-pressure buffering air to the at least one high-pressure bearing compartment through the high-pressure header. The electric buffer compressor is connected to the low-pressure header and the high-pressure header. The electric buffer compressor is configured to direct pressurized buffering air to the low-pressure header and the high-pressure header.
-
公开(公告)号：US20250026482A1
公开(公告)日：2025-01-23
申请号：US18224430
申请日：2023-07-20
Applicant: Raytheon Technologies Corporation
Inventor： Cameron B. Sahirad , Alek Gavrilovski , Lichu Zhao
Abstract: A method of producing an operational data for an aircraft turbine engine is provided that including: sensing an inlet airflow to an aircraft turbine engine for CMAS particulate matter, the sensing performed during one or more ground portions of a flight operational cycle of the turbine engine, the sensing producing sensor signals indicative of the presence or absence of the CMAS particulate matter; determining a presence or absence of an exposure by the turbine engine to a CMAS environment based on the sensor signals; and producing an operational data indicative of the presence or absence of said exposure by the turbine engine to said CMAS environment.
-
9.发明授权公开(公告)号：US12203392B2
公开(公告)日：2025-01-21
申请号：US18117919
申请日：2023-03-06
Applicant: Raytheon Technologies Corporation
Inventor： Jon E. Sobanski
Abstract: A turbine engine with an axis is provided. This turbine engine includes a fan section, a turbine engine core, a bypass flowpath, an engine housing, an evaporator, a condenser and a core flowpath. The turbine engine core is configured to power the fan section. The turbine engine core includes a core compressor section, a core combustor section and a core turbine section. The bypass flowpath is fluidly coupled with and downstream of the fan section. The engine housing includes a cavity radially outboard of and axially overlapping the fan section and/or the bypass flowpath. The evaporator module is within the cavity. The condenser module is within the cavity. The core flowpath extends sequentially through the core compressor section, the core combustor section, the core turbine section, the evaporator module and the condenser module.
-
公开(公告)号：US20250020558A1
公开(公告)日：2025-01-16
申请号：US18352945
申请日：2023-07-14
Applicant: Raytheon Technologies Corporation
Inventor： Garrett KERNOZICKY , Jesse R. BOYER , Ashwin RAGHAVAN , Peter John BREITZMANN , Praba Kharan BAPTIST , Michael L. VUKOVINSKY , Lawrence A BINEK
IPC: G01N3/02
Abstract: A method for repairing a stator stage is disclosed herein. The method includes receiving a stator stage including a plurality of stator vanes disposed between an outer diameter and an inner diameter, analyzing the stator stage for defects, determining there is a first defect in a first segment of a the stator stage, the first segment including at least one of a first portion of the outer diameter, a first portion of the inner diameter, or a first stator vane, removing the first segment from the stator stage forming a void in the stator stage, manufacturing a second segment that is the same size as the first segment, attaching the second segment to the stator stage to fill the void, performing a blending process on the stator stage including the second segment to smooth the plurality of stator vanes including the second stator vane.
-
-
-
-
-
-
-
-
-
Search Results
2520
records
(took 0.023 seconds)
