Abstract: A micro-mixer includes a plurality of fuel tubes through which air and fuel flow, a casing accommodating the plurality of fuel tubes therein, and a plurality of injection parts each formed as a pyramidal protrusion on one side of the casing and connected to a front end of the fuel tube to inject the air and the fuel, wherein each of the plurality of injection parts has an inclined injection hole formed obliquely with respect to an extension direction of the plurality of fuel tubes on at least one side of the pyramidal protrusion.

Abstract: A combustion-gas supply system and a combustion-gas supply method thereof, a device equipped with a turbine engine, and a fracturing system are provided. The combustion-gas supply system includes a main pipeline and a multi-functional pipeline; the main pipeline includes a first sub-pipeline and a second sub-pipeline; the first sub-pipeline includes a first gas intake pipe, a first gas supply valve and a first gas outlet pipe arranged in sequence; the second sub-pipeline includes a combustion-gas supply valve and a gas supply pipe, the first gas outlet pipe is connected with the combustion-gas supply valve, the gas supply pipe is configured to be connected with a turbine engine, the multi-functional pipeline includes a second gas intake pipe, a second gas supply valve and a second gas outlet pipe arranged in sequence, and the second gas outlet pipe is communicated with the first gas outlet pipe.

Abstract: A powerplant is provided that includes a pre-burner, a combustor, a power turbine, a mechanical load and a propellant system. The combustor is fluidly coupled with and downstream of the pre-burner. The power turbine is fluidly coupled with and downstream of the combustor. The mechanical load is rotatably driven by the power turbine. The propellant system is configured to direct fluid oxygen and fluid hydrogen to the pre-burner to provide an oxygen rich fuel mixture for combustion within the pre-burner. The propellant system is also configured to direct the fluid hydrogen to the combustor for combustion within the combustor with oxygen within combustion products received from the pre-burner.

Abstract: A hollow plank of a combustor liner defining a combustion chamber including an inner wall having a plurality of inner openings and one or more inner holes, an outer wall having one or more outer openings and a plurality of outer holes, a plurality of lateral walls coupled to the inner wall and the outer wall to define a cavity, and a partition wall connected to the plurality of lateral walls and dividing the cavity into a first sub-cavity and a second sub-cavity. The one or more outer openings communicate with the first sub-cavity and communicate through a plurality of tubes with the second sub-cavity. The plurality of inner openings communicate with the second sub-cavity and communicate with the first sub-cavity through one or more bypass tubes. The first sub-cavity or the second sub-cavity, or both, are frequency tuned to reduce combustion dynamic frequencies.

Abstract: An assembly is provided for an aircraft propulsion system. This aircraft propulsion system assembly includes a nacelle. The nacelle includes an inlet structure and a cowl. The inlet structure extends axially in a first direction along a centerline to an inlet structure aft end. The inlet structure includes a clip and a hoist bracket integrated with the clip. The clip is at the inlet structure aft end. The clip includes a locator receptacle. The hoist bracket is configured to provide a structural lift point for the inlet structure during assembly and/or disassembly of the nacelle. The cowl extends axially in a second direction along the centerline to a cowl forward end. The cowl includes a locator at the cowl forward end. The locator includes a protrusion received by the locator receptacle.

Abstract: Provided is a transition piece including a first flow path group formed by arraying a plurality of in-wall flow paths that extend inside a plate material forming the transition piece and a second flow path group that is positioned on a side closer to a combustor liner than the first flow path group is. Each in-wall flow path in the first flow path group and the second flow path group has an inlet that is located at one end section in a flow direction of a combustion gas and that faces a compressed air main flow path, and an outlet that is located at the other end section in the flow direction of the combustion gas and that faces a combustion gas flow path.

Abstract: An example hybrid aircraft propulsion system includes one or more power units configured to output electrical energy onto one or more electrical busses; a plurality of propulsors; and a plurality of electrical machines, each respective electrical machine configured to drive a respective propulsor of the plurality of propulsors using electrical energy received from at least one of the one or more electrical busses.

Abstract: A combustor includes an inner liner and an outer liner defining a combustion chamber. The inner liner includes an inner mesh structure, and a plurality of inner planks mounted to the inner mesh structure. The outer liner includes an outer mesh structure, and a plurality of outer planks mounted to the outer mesh structure. Each of the plurality of inner planks and outer planks includes an inner wall, an outer wall, and lateral walls defining a cavity to allow circulation of airflow within the cavity to cool down the inner wall.

Abstract: A combustor includes a liner defining a combustion chamber, and receiving an air and fuel mixing body. The mixing body includes a forward face facing into the combustion chamber. A plurality of circumferentially spaced columns about a central axis of the combustor each have at least two fuel supply passages and two air supply passages extending to the forward face. At least one of the two air supply passages and one of the fuel supply passages extending along a direction with a circumferentially extending component in a first circumferential direction. The other of the two air supply passages and the fuel supply passages extend with a circumferential component in a second opposed circumferential direction. A gas turbine engine is also disclosed.

Abstract: A turbomachine fuel injector system includes an air distributor configured to be mounted to and contained within a casing and to provide air to mix with a fuel from one or more fuel distribution systems. The system includes a fuel manifold configured to be mounted indirectly to the casing such that the fuel manifold system is contained within the casing but is independent such that fuel manifold does not touch or directly mount to either an interior of the casing or touch the air distributor within the casing to prevent or reduce thermal transfer from the air distributor to the fuel manifold.

Abstract: A combustor component of a combustor for combusting fuel to produce a combustion gas includes a combustion cylinder forming a passage for the combustion gas, a first acoustic device which internally includes a first cavity communicating with the passage via a first through hole formed in the combustion cylinder, a second acoustic device which is located on a radially outer side of the first acoustic device so as to cover the first acoustic device, and internally includes a second cavity communicating with the passage via a second through hole formed in the combustion cylinder, and a first communication passage causing the first cavity and an outer space of the combustion cylinder to communicate with each other without via the first through hole and the second cavity.

Abstract: An aircraft including a fuselage, a wing structure, at least one turbomachine running on hydrogen and generating thrust at a propulsion unit distant from the fuselage, at least one fuel tank positioned in the fuselage and configured to store hydrogen in the cryogenic state, at least one hydrogen supply device connecting the fuel tank and the turbomachine and including at least one pump positioned in the fuselage in the vicinity of the fuel tank, at least one hydrogen heating system positioned in the fuselage in the vicinity of the pump. This solution makes it possible to reduce a length of the complex double-walled pipes configured for carrying the hydrogen in the cryogenic state between the fuel tank and the hydrogen heating system.

Abstract: A particle separator includes a separator body in a primary fluid passageway of a machine. The primary fluid passageway includes one or more bleed holes through which a diverted portion of the fluid flowing in the primary fluid passageway toward a volume of the machine is diverted into an auxiliary flow passageway that bypasses the volume and directs the diverted portion of the fluid toward one or more other components of the machine. The separator body is coupled with the inner wall and/or outer wall of the primary fluid passageway. The separator body includes an upstream edge positioned to separate at least some particles carried by the fluid from the fluid as the diverted portion of the fluid bends around and flows over the at least one upstream edge of the separator body and into the auxiliary flow passageway.

Abstract: A system for bleeding air from a core flow path of a gas turbine engine is disclosed. In various embodiments, the system includes a bleed valve having a bleed valve inlet configured to receive a bleed air from a first access point to the core flow path and a bleed valve outlet; and an air motor having a first air motor inlet configured to receive the bleed air from the bleed valve outlet and a first air motor outlet configured to exhaust the bleed air, the air motor configured to pump the bleed air from the core flow path of the gas turbine engine.

Abstract: A liner for a combustor in a gas turbine engine and a related method. The liner includes a liner body having a cold side and a hot side. The liner includes a dilution passage having a concatenated geometry extending through the liner body. The concatenated geometry has a plurality of discrete dilution holes, an annular slot, and a plurality of dilution inserts. The dilution passage is configured (i) to integrate a first dilution air flow flowing through the dilution passage from the cold side to the hot side and a second dilution air flow flowing through the dilution passage from the cold side to the hot side into an integrated dilution air flow, and (ii) to inject the integrated dilution air flow into a core primary combustion zone of the combustor to attain a predetermined combustion state of the combustor.

Abstract: The present disclosure is directed to a gas turbine engine assembly having a compressor configured to increase pressure of incoming air, a combustion chamber, at least one turbine coupled to a generator, a torsional damper, and a controller. The combustion chamber is configured to receive a pressurized air stream from the compressor. Further, fuel is injected into the pressurized air in the combustion chamber and ignited so as to raise a temperature and energy level of the pressurized air. The turbine is operatively coupled to the combustion chamber so as to receive combustion products that flow from the combustion chamber. The generator is coupled to the turbine via a shaft. Thus, the torsional damper is configured to dampen torsional oscillations of the generator. Moreover, the controller is configured to provide additional damping control to the generator.

Abstract: A gas turbine engine includes a core engine that includes a core flow path where air is compressed in a compressor section, communicated to a combustor section, mixed with an ammonia based fuel and ignited to generate a high energy combusted gas flow that is expanded through a turbine section. The turbine section is mechanically coupled to drive the compressor section. An ammonia flow path communicates an ammonia flow to the combustor section. A cracking device is disposed in the ammonia flow path. The cracking device is configured to decompose the ammonia flow into a fuel flow containing hydrogen (H2). At least one heat exchanger is upstream of the cracking device that provides thermal communication between the ammonia flow and a working fluid flow such that the ammonia fluid flow accepts thermal energy from the working fluid flow.

Abstract: Electric aircraft power plants and associated methods are provided. One power plant includes an emergency power unit (EPU) for providing electric power in the event of a malfunction of a battery pack of an electric aircraft to permit the electric aircraft to make an emergency maneuver. The EPU includes a rocket engine for generating a stream of exhaust fluid using a rocket propellant, a turbine operatively connected to extract energy from the stream of exhaust fluid generated by the rocket engine, and an electric generator operatively connected to be driven by the turbine and to supply electric power to an electric motor propelling the electric aircraft.

Abstract: An aircraft propulsion system includes a gas turbine engine; a generator; a storage battery; a motor which drives a rotor, using at least one of the electric power which is output from the generator and the electric power which is output from the storage battery; a detection unit which detects the number of revolutions of the engine shaft; an engine control unit which controls at least a fuel flow rate of the gas turbine engine; and a generator control unit which controls the operation of the generator. When the number of revolutions satisfies a predetermined condition, at least the generator control unit executes a control for reducing a sudden change in the number of revolutions.

Abstract: Systems and methods for preventing fuel leakage in a gas turbine engine are provided. A fuel accumulation system includes a control valve section fluidly coupled to a fuel manifold passage and an accumulator valve section fluidly coupled at a first side to the control valve section. The control valve section is configured to control expansion of a fluid flowing in the fuel manifold passage. The accumulator valve section is configured to receive fluid expanded in the fuel manifold passage via the control valve section.