Abstract: A system includes a gas turbine having a turbine stage disposed in a combustion gas path, wherein the turbine stage includes turbine vanes disposed upstream from turbine blades. The system includes an isothermal expansion system coupled to the turbine stage. The isothermal expansion system includes multi-fluid injectors configured to vary axial positions of combustion within a turbine stage expansion of the turbine stage to reduce temperature variations over the turbine stage expansion, wherein at least one of the multi-fluid injectors is coupled to each of the turbine vanes. Each of the multi-fluid injectors includes a fuel port configured to inject a fuel, an oxidant port configured to inject an oxidant, and a barrier fluid port configured to inject a barrier fluid between the fuel and the oxidant, wherein the barrier fluid is configured to delay mixing between the fuel and the oxidant.
Abstract: A fuel injector for an aircraft turbine engine includes a tubular body having an axis of elongation. A first longitudinal end configured to be Supplied with fuel and a second longitudinal end configured to elect a jet of fuel. The body further includes an integrated purge-air circuit that has an internal cavity which is connected to air inlet orifices situated on the body and to at least one air outlet situated at said second end. Aire-flow disruptors are provided, projecting into said cavity.
Type:
Grant
Filed:
July 16, 2020
Date of Patent:
February 6, 2024
Assignee:
SAFRAN HELICOPTER ENGINES
Inventors:
Thomas Jean Olivier Lederlin, Denis Luc Alain Chanteloup, Simon Arthur Meilleurat
Abstract: An atomizer provides a high-quality fuel-air mixture to a gas turbine engine, by combining air input from an engine compressor and fuel input from a single low-pressure fuel supply pump. The atomizer includes an atomizer body, a main vortex chamber, a secondary vortex chamber for improving quality of the fuel-air mixture, and a fuel sleeve providing fuel to the secondary vortex chamber. The main vortex chamber includes a main outlet nozzle in fluid communication with a combustion chamber inlet of the gas turbine engine. The secondary vortex chamber includes a secondary outlet nozzle in fluid communication with the main vortex chamber. The fuel sleeve has a blind channel with a longitudinal axis and a fuel tip. The same atomizer may be used for startup mode and for all operational modes of the gas turbine engine.
Abstract: A gas turbine engine with a compressor section, a turbine section, and a combustion section located downstream from the compressor section and upstream from the turbine section, the combustion section including a dome inlet, a combustor outlet fluidly coupled to the turbine section, a liner and a dome assembly together at least partially defining a combustion chamber extending between the dome inlet and the combustor outlet, a fuel system fluidly coupled to the combustion section, the fuel system comprising a fuel supply, a primary fuel line fluidly coupling the fuel supply to the combustion section, and a reformer fluidly coupled to the fuel supply.
Type:
Grant
Filed:
January 31, 2022
Date of Patent:
January 30, 2024
Assignee:
General Electric Company
Inventors:
Narendra D. Joshi, Lawrence B. Kool, Joseph Zelina
Abstract: A system for monitoring fuel additives on board a vehicle includes a fuel line carrying fuel from a fuel source to an engine; a fuel additive sensor configured to measure concentration of additives in fuel at a point along the fuel line; a fuel additive dispenser connected in parallel to the fuel line; at least one flow control device for controlling an amount of flow from the fuel line into the fuel additive dispenser; and a controller configured to receive input from the fuel additive sensor and to control the flow control device to adjust the amount of the flow from the fuel line into the fuel additive dispenser.
Type:
Grant
Filed:
November 8, 2021
Date of Patent:
January 23, 2024
Assignee:
RTX Corporation
Inventors:
Zissis A Dardas, Haralambos Cordatos, Ying She
Abstract: A jet engine for propelling aircraft, capable of providing thrust from rest to high speeds is provided. The engine has an axial compressor (16) or several axial compressors located on the same plane and is driven by a gas generator. At the outlet of the turbine there is a gasification chamber (23) into which more fuel is injected. Combustion of the gases from the gasification chamber is performed in two combustion chambers (18) with a rectangular cross-section, separated by a central body (10). The exhaust of the gases is performed in nozzles, each with a square convergent/divergent cross-section (19) and (21). The cross-section of the throats (26) can be adjusted by means of two mobile elements (20). The final section of the central body (10) forms a wedge-shape (27), enabling the continued expansion of the exhaust gases.
Abstract: Various embodiments of a vortex hybrid motor system are described herein. In some embodiments, the vortex hybrid motor system can include a control system, a vortex hybrid motor, and an oxidizer injector. The oxidizer injector can be in fluid communication with a combustion zone defined by a fuel core and/or housing of the vortex hybrid motor. In some embodiments, at least one material regression sensor can be positioned along the fuel core and sensed data from the material regression sensors can be provided to the control system for determining one or more characteristics associated with the fuel core. The control system can control, based on the analyzed sensed data, the oxidizer injector for modulating an oxidizer flow rate delivered to the combustion zone to achieve a desired oxidizer-to-fuel ratio.
Type:
Grant
Filed:
April 12, 2022
Date of Patent:
January 23, 2024
Assignee:
Sierra Space Corporation
Inventors:
Brian Richard Pomeroy, Martin Chiaverini, Jesse Morgan Warrick
Abstract: A burner assembly includes a plurality of burners for mixing fuel and air. Each of the plurality of burners includes: a fuel nozzle for injecting the fuel; a mixing passage supplied with the fuel and the air; and a support portion connecting a passage wall of the mixing passage and the fuel nozzle to support the fuel nozzle.
Abstract: A combustor includes an outer can into which fuel is introduced, an outer head disposed on a front side of the outer can, an inner can disposed inside of the outer can and having a combustion chamber in which a fuel-air mixture is combusted, and an inner head disposed to mix the fuel and the compressed air and supply the mixture into the inner can. The inner head includes a head plate covering a front side of the inner can, and nozzle assemblies disposed to mix the fuel and the compressed air and supply the mixture rearwards. The nozzle assembly includes a nozzle head into which fuel is introduced and nozzles. The nozzles each is coupled between the nozzle head and the head plate to mix the fuel and the compressed air and supply the mixture rearwards. The nozzles each has a shape with a diameter decreasing and increasing toward the rear side thereof.
Abstract: An installation for supplying a cryogenic fuel to a combustion chamber of a turbine engine of an aircraft including a tank, a mixing chamber, and one or more heat exchangers. The tank stores cryogenic fuel in a and connects to a combustion chamber for supplying the combustion chamber with cryogenic fuel in the supercritical or gaseous state. The one or more heat exchangers are provided between the cryogenic fuel and air of an air-conditioning circuit of the aircraft, mounted in a line connecting the tank for cryogenic fuel to the mixing chamber and in a line to be connected to the air-conditioning circuit of the aircraft, the heat exchange taking place therein so as to cool the air of the air-conditioning circuit of the aircraft and to increase the temperature of the cryogenic fuel coming from the tank.
Abstract: A gas turbine engine including a compressor section, a combustor for combusting a fuel, and a turbine. Compressed air flows through a combustion liner of the combustor in a bulk airflow direction. The combustor includes a primary fuel nozzle and a secondary fuel nozzle. The secondary fuel nozzle is downstream of the primary fuel nozzle in the bulk airflow direction. The primary fuel nozzle is configured to inject a primary portion of the fuel into a primary combustion zone, and the secondary fuel nozzle is configured to inject a secondary portion of the fuel into a secondary combustion zone. The secondary combustion zone is located downstream of the primary combustion zone in the bulk airflow direction. The fuel may be one of diatomic hydrogen fuel and a hydrogen enriched fuel.
Abstract: A pre-vaporisation tube for a turbine engine combustion chamber includes a main body ROOM defining a first inner duct configured to have an injector mounted therein. The tube includes a first end attached to a wall of the chamber, and at least two end pieces are arranged at a second end of the body and define second inner ducts. The end pieces include first portions and second portions, respectively. The second portions each include two coaxial cylindrical walls which are inner and outer coaxial cylindrical walls, respectively, and which define an annular cavity therebetween. The inner wall defines an inner passage and has first openings for fluid communication between the passage and the annular cavity.
Type:
Grant
Filed:
October 5, 2020
Date of Patent:
December 12, 2023
Assignee:
SAFRAN HELICOPTER ENGINES
Inventors:
Thomas Jean Olivier Lederlin, Christophe Laurent, Guillaume Gerard Joel Mauries
Abstract: In one aspect, the present disclosure is directed a gas turbine engine including a compressor, a combustor, and a turbine in a serial flow arrangement. The compressor includes a casing having an inner wall and an outer wall. The inner wall defines a passageway for airflow through the compressor. The casing defines a bleed cavity between the inner wall and the outer wall. The inner wall has an opening. The casing defining a channel between the opening and the bleed cavity to direct airflow into the bleed cavity. The channel is defined by a first wall and a second wall of the casing. The second wall being downstream of the first wall. A surface of the first wall has a pattern to reduce flow separation in the channel.
Type:
Grant
Filed:
May 27, 2022
Date of Patent:
November 28, 2023
Assignee:
General Electric Company
Inventors:
Sesha Subramanian, Ravikanth Avancha, Atanu Saha
Abstract: An example engine for producing thrust includes: a fuel supply to supply a fuel; a chamber fluidly coupled to the fuel supply to receive the fuel; an induction heating assembly operatively coupled to the chamber to inductively energize the fuel in the chamber; and an exhaust nozzle coupled to the chamber to receive energized fuel from the chamber to produce thrust.
Abstract: An intake device for a gas turbine engine includes a snorkel and a particle separator. The snorkel is configured to be mounted to a panel defining at least a portion of a gas flow path within the gas turbine engine. The snorkel includes a tubular body extending between a closed end and an open end opposite the closed end. The snorkel further includes an inlet aperture formed through the tubular body adjacent the closed end. At least a portion of the snorkel is configured to be disposed within the gas flow path. The particle separator is mounted to the snorkel downstream of the inlet aperture. The particle separator includes at least one gas flow passage extending between a flow inlet and a flow outlet. The at least one gas flow passage is configured to remove particulate matter from the at least one gas flow passage upstream of the flow outlet.
Type:
Grant
Filed:
May 6, 2022
Date of Patent:
November 21, 2023
Assignee:
PRATT & WHITNEY CANADA CORP.
Inventors:
Julien Girard, Sylvain Lamarre, Xiaoliu Liu, David Koo, Kevin Nguyen, Liam McPherson, AnnMarie Unnippillil
Abstract: A turbine engine that includes an engine core having at least a compressor section and a combustion section. The combustion section includes a combustor. The combustor section or combustor includes a fuel-air mixing assembly fluidly coupled to the compressor section. The fuel-air mixing assembly includes an outer wall, a center body at least partially circumscribed by the outer wall, and an annular flow passage between the outer wall and center body. At least one fuel orifice includes a fuel outlet fluidly coupled to the annular flow passage.
Type:
Grant
Filed:
January 27, 2022
Date of Patent:
November 14, 2023
Assignee:
General Electric Company
Inventors:
Manampathy G. Giridharan, Ajoy Patra, Pradeep Naik, R Narasimha Chiranthan, Perumallu Vukanti, Michael T. Bucaro
Abstract: A fuel supply circuit for an aeronautical cryogenic turbomachine including: at least one cryogenic reservoir containing a liquid fuel topped with a boil-off gas and including a high-pressure liquid pump to supply at least one main propulsion device of the turbomachine with liquid fuel, an auxiliary turbomachine including an electric generator, a gas compressor to supply the auxiliary turbomachine with gaseous fuel, and a buffer gas reservoir connecting the gas compressor to the cryogenic reservoir to reinject gas into the cryogenic reservoir in order to maintain the pressure in the cryogenic reservoir above a predefined value, the electric generator supplying the high-pressure liquid pump and the gas compressor.
Type:
Grant
Filed:
July 21, 2021
Date of Patent:
November 7, 2023
Assignees:
SAFRAN, ARIANEGROUP SAS
Inventors:
Pierre-Alain Lambert, Hugo Pierre Mohamed Jouan, Samer Maalouf, Louis-Vianney Mabille De La Paumeliere, Carlos Alberto Cruz, Davide Duri
Abstract: An assembly is provided for a gas turbine engine. This assembly includes a combustor and a fuel conduit. The combustor includes a combustor wall that forms a peripheral boundary of a combustion chamber within the combustor. The fuel conduit extends along and is formed integral with the combustor wall. The fuel conduit is disposed outside of the combustion chamber.
Type:
Grant
Filed:
November 24, 2021
Date of Patent:
November 7, 2023
Assignee:
RTX CORPORATION
Inventors:
Lawrence A. Binek, Timothy S. Snyder, Sean R. Jackson
Abstract: A method and system for enhancing power generated by a gas turbine system. The system may include a turbine inlet cooling system and a wet compression air fogging system. Air entering the gas turbine system is cooled by the turbine inlet cooling system and the wet compression air fogging system. The wet compression air fogging system may increase the mass flow rate of the air entering the gas turbine system at the compressor.