Abstract: An engine system is provided that includes a compressor section, a combustor section, a turbine section, a flowpath and a flow regulator. The combustor section includes a combustion chamber. The flowpath extends sequentially through the compressor section, the combustor section and the turbine section. The flow regulator is configured to open the flowpath between the compressor section and the combustion chamber during a first mode of operation. The flow regulator is configured to at least substantially close the flowpath between the compressor section and the combustion chamber during a second mode of operation.

Abstract: Provided is a burner of double tube structure including an outer tube and an inner tube. The inner tube includes a tip-side support portion fixed on the outer tube, a base-side support portion fixed on the outer tube, a main body portion supported by the tip-side support portion, bellows interposed between the main body portion and the base-side support portion, and a bulkhead member attached at a position between the tip-side support portion and the base-side support portion. Between an outer peripheral surface of the bulkhead member and an inner peripheral surface of the outer tube, a second gap smaller than a first gap between an outermost peripheral surface of the bellows and the inner peripheral surface of the outer tube is interposed.

Abstract: No known single-stage dry low emissions fuel injectors are capable of effectively operating over all ranges of hydrogen concentrations in hydrogen/natural gas fuel mixtures. Accordingly, a fuel injector is disclosed that is capable of operating in both a premix mode for fuel mixtures with lower hydrogen concentrations and a micromix mode for fuel mixtures with higher hydrogen concentrations. The fuel injector may comprise premix jets near an inlet of the fuel injector, optionally within one or more swirlers, and micromix jets near the outlet of the fuel injector. In the premix mode, fuel with lower hydrogen concentrations is provided to the premix jets, whereas in the micromix mode, fuel with higher hydrogen concentrations is provided to the micromix jets.

Abstract: A combustor casing component for a gas turbine engine which is manufactured as a single-piece. The combustor casing component has a combustor outer casing portion, an outlet guide vane outer case portion, a pre-diffuser portion, a plurality of outlet guide vanes, and an outlet guide vane inner case portion.

Abstract: A fuel system is provided for an aircraft having a fuel source. The fuel system includes a fuel oxygen reduction unit defining a liquid fuel supply path, a stripping gas supply path, a liquid fuel outlet path, and a stripping gas return path, wherein the stripping gas return path is in airflow communication with the fuel source.

Abstract: A gas turbine engine includes a first combustor having a first fuel nozzle, wherein the first fuel nozzle is configured to supply a water fuel emulsion into the first combustor. The water fuel emulsion includes a water-in-fuel (WIF) emulsion having a plurality of water droplets dispersed in a fuel, wherein the plurality of water droplets is configured to vaporize within the fuel to cause micro-explosions to atomize the fuel, and the atomized fuel is configured to combust to generate a combustion gas. The gas turbine engine further includes a turbine driven by the combustion gas from the first combustor.

Abstract: A method for tuning a gas turbine engine includes performing a sensitivity step process on a tuning parameter. An operating parameter is monitored. The gas turbine is operating in a first operational state and the operating parameter has an initial condition. The tuning parameter is selected for adjustment. The tuning parameter is adjusted by a predefined amount. The adjustment includes applying an incremental bias adjustment to a fuel flow fraction schedule. The gas turbine engine transitions to a second operational state, wherein the operating parameter has an adjusted condition. The adjusted condition and the initial condition of the operating parameter are applied to a cost function. It is then determined that the cost function results in a cost function value indicative of a decreased cost. The incremental bias adjustment and the cost function value is written to a bias look-up table and are associated with the selected tuning parameter.

Abstract: A combustion section defines an axial direction, a radial direction, and a circumferential direction. The combustion section includes a casing that defines a diffusion chamber. A combustion liner is disposed within the diffusion chamber and defines a combustion chamber. The combustion liner is spaced apart from the casing such that a passageway is defined between the combustion liner and the casing. A fuel cell assembly is disposed in the passageway. The fuel cell assembly includes a fuel cell stack having a plurality of fuel cells each extending between an inlet end and an outlet end. Each fuel cell of the plurality of fuel cells includes an air channel and a fuel channel each fluidly coupled to the combustion chamber.

Abstract: A trapped vortex reverse flow combustor for a gas turbine includes a first dome structure having a plurality of first-dome vortex driver airflow openings for providing a first vortex generating mid airflow therethrough to a trapped vortex cavity. A second dome structure is arranged downstream of the first dome structure and includes a plurality of second-dome vortex driver airflow openings providing a first vortex generating outer airflow therethrough to the trapped vortex cavity, and a plurality of primary driver airflow openings providing a primary driver airflow therethrough radially inward of the trapped vortex cavity.

Abstract: An apparatus is provided for a turbine engine. This apparatus includes a fuel-air mixer, and the fuel-air mixer includes an inner passage, a sidewall, a steam passage, a fuel nozzle and an air swirler. The inner passage extends axially along an axis within the fuel-air mixer. The sidewall extends circumferentially around and axially along the inner passage. The steam passage is embedded within the sidewall and extends along the inner passage. The fuel nozzle is configured to direct fuel into the inner passage. The air swirler is configured to direct swirled air into the inner passage for mixing with the fuel.

Abstract: A gas turbine engine includes a compressor section with a compressor rotor shaft assembly including a plurality of compressor rotors longitudinally spaced apart from each other via respective ones of a plurality of shaft sections of a rotor shaft, each compressor rotor of the plurality of compressor rotors having a plurality of rotor vanes extending radially outward therefrom and being circumferentially spaced about the compressor rotor. A stator shroud assembly has a stator shroud casing surrounding the compressor rotor shaft assembly, a compressor flow passage being defined between the compressor rotor shaft assembly and the stator shroud casing. A pressurized air source generates a flow of pressurized air to be provided to the compressor section, and a plurality of pressurized airflow nozzles are connected with the pressurized air source and provide the flow of pressurized air into the compressor flow passage to cause the compressor rotor shaft assembly to rotate.

Abstract: An aircraft engine, comprising: a first wall and a second wall defining a gaspath between the first wall and the second wall, the gaspath extending around a central axis, each of the first wall and the second wall having a gaspath side facing the gaspath and an opposed side facing away from the gaspath; struts circumferentially distributed around the central axis, the struts extending across the gaspath, a strut of the struts having an airfoil including a leading edge, a trailing edge, a first end secured to the first wall, and a second end secured to the second wall; and protrusions extending from a baseline surface of the opposed side of the first wall, the protrusions increasing a thickness of the first wall, a protrusion of the protrusions overlapping a location where the leading edge or the trailing edge meets the gaspath side of the first wall.

Abstract: A fuel system for a gas turbine engine is provided. The gas turbine engine including an exhaust assembly. The exhaust assembly including a tail cone and defining a radial direction, a circumferential direction, an axial direction, a working gas flow path, and an outer bypass flow path, the fuel system including: a hydrogen fuel tank for holding a hydrogen fuel; a heat exchanger configured to be coupled to or integrated into the tail cone of the exhaust assembly, the heat exchanger including: a first axial flow path in flow communication with the hydrogen fuel tank; a second axial flow path in flow communication with a combustion section of the gas turbine engine; and a radial flow path.

Abstract: An aircraft heat exchanger system according to an exemplary embodiment of this disclosure, among other possible things includes a first heat exchanger assembly that is disposed in an inlet duct assembly, and a skin heat exchanger assembly is in thermal communication with an outer surface of an aircraft structure. The skin heat exchanger is in fluid communication with the first heat exchanger such that a working fluid is communicated therebetween.

Abstract: A combustor structure of an embodiment is disposed to penetrate, from a direction perpendicular to an axial direction of a turbine rotor in a supercritical CO2 gas turbine which uses supercritical CO2 for a working fluid, a casing of the supercritical CO2 gas turbine. The combustor structure includes a plurality of combustors. Each of the combustors includes: a combustor liner in a cylindrical shape, which combusts fuel and an oxidant; a fuel supply part which is provided at an upstream end of the combustor liner and supplies the fuel into the combustor liner; and an oxidant supply part which is provided at the upstream end of the combustor liner and supplies the oxidant into the combustor liner.

Abstract: A combustor includes a combustion tube which has a tubular shape around an axis; a secondary fuel nozzle which is able to inject a secondary fuel in a radially inner direction with respect to the axis, inside the combustion tube; a fuel manifold which is disposed external to the combustion tube, and defines a fuel space which is able to temporarily store the secondary fuel; and a support which supports the fuel manifold with a gap extending radially from the combustion tube. The support has a support plate which is separated from the combustion tube by the gap and spreads in a circumferential direction with respect to the axis, and a support leg which is attached to the combustion tube and supports the support plate to define the gap between the support plate and the combustion tube.

Abstract: A microthruster for controlling small spacecraft and methods for manufacturing the same are disclosed. Embodiments of the microthruster include one or more nozzle throats with cross sectional areas of at most 20 ?m2, and some with 6 ?m2. Some embodiments include heaters that heat water in one or more reservoirs to increase pressure in the reservoirs and eject the water from the one or more nozzle throats. Some embodiments are manufactured by etching channels into one or more layers of material, and still further embodiments are manufactured by forming the nozzle throats and/or the reservoirs between two layers of material. Some microthruster embodiments are flat in shape with the nozzle throats ejecting water out the thin sides of the microthruster. Still further embodiments are formed by etching channels into one layer of material, printing a heater onto another layer of material, and bonding the two layers together.

Abstract: A gas turbine engine system includes a gas turbine engine, an accessory gearbox, and a flywheel powered barring engine system. The gas turbine engine includes a rotor and a turbine. The accessory gearbox is connected with the rotor. The flywheel powered barring engine system is connected with the accessory gearbox and configured to rotate the rotor after shut down of the gas turbine engine.

Abstract: A system and method for ram air intake for pulse combustion systems is disclosed that improves the ability of pulse combustions to ingest air into the inlet pipe when the pulse combustion system is moving in a direction opposite the direction the open end of the inlet pipe is facing and the system and method includes the ability to increase the thrust output from the pulse combustion system.

Abstract: A fluid injection system can include a main flow line, a primary flow line connected to the main flow line, and a primary flow valve disposed between the primary flow line and the main flow line and configured to allow injectant flow to the primary flow line from the main flow line in an open primary flow valve state, and to prevent injectant flow to the primary flow line from the main flow line in a closed primary flow valve state. The system can include a secondary flow line connected to the main flow line and a secondary flow valve disposed between the secondary flow line and the main flow line and configured to selectively allow injectant flow to the secondary flow line from the main flow line in an open secondary flow valve state, and to prevent injectant flow to the secondary flow line from the main flow line in a closed secondary flow valve state.