Abstract: A hydrogen injection scheme can be employed via use of one or more injectors for injecting hydrogen to help entrain local mass while also generating local turbulence that can enhance mixing with a mixture of air and fuel to facilitate enhanced lean combustion, lower peak flame temperatures of combustion, and reduce nitrous oxide (NOx) emissions from the combustion of fuel. In some embodiments, the hydrogen can be injected to help transport heat released during combustion away from the injector to help avoid injector overheating as well. Different injectors can be utilized to provide a desired hydrogen injection scheme for a particular set of design and operational criteria for a gas turbine system or at least one combustion system that can be utilized in a gas turbine system.

Abstract: An engine can utilize a combustor to combust fuel to drive the engine. A fuel nozzle assembly can supply fuel to the combustor for combustion or ignition of the fuel. The fuel nozzle assembly can include a swirler and a fuel nozzle to supply a mixture of fuel and air for combustion, which can supply a primary fuel supply and a secondary fuel supply. Increasing efficiency and reducing emission require the use of alternative fuels, which combust at higher temperatures or burn at faster burn speeds than traditional fuels, requiring improved fuel introduction without the occurrence of flame holding or flashback.

Abstract: The present invention provides a multimode propulsion system, comprising at least one propellant ejector system, a high speed fluid ejection nozzle coupled to a propellant supply provided in the engine, and a propellant-air mixing system comprising at least one fluid intake member having an inlet end and an outlet end, the inlet end being in fluidic communication with the fluid ejection nozzle to receive the propellant ejected from the nozzle.

Abstract: A propulsion system includes a core engine generating a high energy gas flow, a condenser configured to extract water from the high energy gas flow and an evaporator transforming the extracted water into a steam flow. The steam flow is injected into a core flow path of the core engine increase mass flow.

Abstract: A resonator ring for a combustor basket, in particular for a gas turbine power plant, which has at least one ring, wherein the ring or the rings has or have apertures, wherein insert plates are connected to the ring or the rings within the aperture. A method for producing a resonator ring, in which a ring or rings is or are produced from a first metal, then insert plates made of a, preferably cast, second material are connected to the ring or the rings, and individual resonator boxes are produced.

Abstract: An engine can utilize a combustor to combust fuel to drive the engine. A fuel nozzle assembly can supply fuel to the combustor for combustion or ignition of the fuel. The fuel nozzle assembly can include a swirler and a fuel nozzle to supply a mixture of fuel and air for combustion. Increasing efficiency and meeting emission needs can be met with the use of alternative fuels, which combust at higher temperatures or higher speeds than traditional fuels, requiring improved fuel introduction without the occurrence of flame holding or flashback.

Abstract: A structure for damping at a fluid manifold assembly for an engine is generally provided. The fluid manifold assembly includes a first walled conduit defining a first fluid passage therewithin. A flow of fluid defining a first frequency is permitted through the first fluid passage. A second walled conduit includes a pair of first portions each coupled to the first walled conduit. A second portion is coupled to the pair of first portions. A second fluid passage is defined through the first portion and the second portion in fluid communication with the first fluid passage. The flow of fluid is permitted through the second fluid passage at a second frequency approximately 180 degrees out of phase from the first frequency.

Abstract: An assembly is provided for an engine. This engine assembly includes an engine structure and a fuel injector. The engine structure includes an injector aperture and a fuel supply passage. The injector aperture extends longitudinally through the engine structure along a centerline. The fuel supply passage extends laterally within the engine structure to the injector aperture. The fuel injector is mated with the injector aperture. The fuel injector includes a nozzle passage and a nozzle orifice. The nozzle passage extends longitudinally along the centerline within the fuel injector to the nozzle orifice. The nozzle passage is fluidly coupled with the fuel supply passage.

Abstract: A gas turbine engine includes a bypass duct and a rotating detonation augmentor. The bypass duct is configured to conduct air through a flow path arranged around an engine core of the gas turbine engine to provide thrust for propelling the gas turbine engine. The rotating detonation augmentor is located in the bypass duct and configured to be selectively operated to detonate fuel and a portion of the air to increase the thrust for propelling the gas turbine engine.

Abstract: A turbine engine system includes at least one hydrogen fuel tank, a core flow path heat exchanger in a core flow path; and engine systems located in the core flow path. The engine system including at least a compressor section, a combustor section having a burner, and a turbine section. The core flow path heat exchanger is arranged in the core flow path downstream of the combustor section. The hydrogen fuel is supplied from the at least one hydrogen fuel tank through a hydrogen fuel supply line, passing through the core flow path heat exchanger and then supplied into the burner for combustion.

Abstract: Dual propellant electrothermal propulsion systems with a polar propellant and a non-polar propellant are disclosed. Such a system may be used for imparting high specific impulse momentum on a spacecraft. A thruster includes a microwave source operably coupled to a heating chamber. An injector mixes polar and non-polar propellants and a nozzle at one end of the heating chamber. The thruster raises the temperature of the mixed propellants that exit the heating chamber to expand in the nozzle to generate thrust.

Abstract: A combustion assembly for a gas turbine includes a flame tube and a fuel supply including a flow rate limiter supplying an injector. The flow rate limiter and the fuel injector are formed in one piece.

Abstract: A control device includes a suction amount controller and a bleed flow controller. When an actual output is in a second output range from a second output to a third output less than the second output, the suction amount controller controls a suction amount so that the suction amount becomes constant even if the actual output decreases. The bleed flow controller controls a bleed flow so that the bleed flow gradually increases as the actual output decreases.

Abstract: A combined heat and power (CHP) system and a method of operating is provided. The CHP system includes an oxygen-air separation device. A combustion chamber has an inlet coupled to the oxygen-air separation device. The combustion chamber receives an expanded heated CO2 working gas for combustion of solid fuel and generates a CO2 exhaust gas. The expanded heated CO2 working gas comprising at least a portion of the CO2 exhaust gas. A heat exchanger receives combustion gases and outputs a cooled CO2 working gas. A turbine is provided having a compression section and an expansion section, the compression section configured to receive and compress the cooled CO2 working gas and to direct it to the heat exchanger. The turbine receives a heated CO2 working gas at the expansion section and expands the heated CO2 working gas. In an embodiment, CO2 products may be generated from the system.

Abstract: Aircraft propulsion systems and aircraft are described. The aircraft propulsion systems include aircraft systems having at least one hydrogen tank and an aircraft-systems heat exchanger and engine systems having at least a main engine core, a high pressure pump, a hydrogen-air heat exchanger, and an expander, wherein the main engine core comprises a compressor section, a combustor section having a burner, and a turbine section. Hydrogen is supplied from the at least one hydrogen tank through a hydrogen flow path, passing through the aircraft-systems heat exchanger, the high pressure pump, the hydrogen-air heat exchanger, and the expander, prior to being injected into the burner for combustion.

Abstract: In one aspect of the present disclosure, there is provided a gas turbine engine. The gas turbine engine includes a primary gas path having, in fluid series communication: an air inlet, a compressor fluidly connected to the air inlet, a combustor fluidly connected to an outlet of the compressor, and a turbine section fluidly connected to an outlet of the combustor section. In embodiments, a hydrogen expansion turbine is in fluid communication to receive hydrogen from the gaseous hydrogen outlet of the heat exchanger. In certain embodiments, the gas turbine engine includes a heat exchanger having a gas conduit fluidly connected to the primary gas path, and a fluid conduit in fluid isolation from the gas conduit and in thermal communication with the gas conduit.

Abstract: An assembly is provided for a turbine engine. This assembly includes a fuel injector assembly, and the fuel injector assembly includes a fuel passage, an air passage and a steam passage. The fuel injector assembly is configured to direct fuel out of the fuel passage along an axis into a volume as a fuel flow. The fuel injector assembly is configured to direct air out of the air passage along the axis into the volume as an air flow, where the air flow circumscribes the fuel flow. The fuel injector assembly is configured to direct steam out of the steam passage along the axis into the volume as a steam flow, where the steam flow circumscribes the fuel flow and provides a radial buffer between the fuel flow and the air flow.

Abstract: A composite airfoil assembly for a turbine engine. The composite airfoil assembly has an airfoil outer surface defining opposing pressure and suction sides, which extend between a leading edge and a trailing edge. The composite airfoil assembly includes a composite airfoil having a woven core and a skin applied to at least a portion of a core exterior. The composite airfoil assembly also includes a cladding coupled to a skin outer surface where at least one edge or portion of the cladding is located adjacent at least one of the trailing edge, a root, or a tip, wherein the cladding includes a tapered portion.

Abstract: A device for distributing at least one fuel flow for a circuit for supplying fuel to a combustion chamber of a turbo machine, in particular of an aircraft, the device including a rotating main body extending about a longitudinal axis, the main body including an internal flow duct for a first flow of fuel, extending through the main body along the axis, a first annular portion configured to form an outlet passage for a second flow of fuel, and including at least one orifice for the second flow of fuel to pass through, and an external annular groove in fluidic communication with the at least one orifice in the first portion.

Abstract: An aircraft propulsion system includes a core engine that includes a core flow path where a core airflow is compressed in a compressor section, communicated to a combustor section, mixed with fuel and ignited to generate a gas flow that is expanded through a turbine section for powering a primary propulsor. The aircraft propulsion system further includes a tap that is at a location upstream of the combustor section where a bleed airflow is drawn, a heat exchanger where the bleed airflow is heated by the gas flow, a power turbine through which heated bleed airflow is expanded to generate a work output, and a secondary propulsor that is driven by the work output that is generated by the power turbine.