Abstract: The dipole drive is a new propulsion system which uses ambient space plasma as propellant, thereby avoiding the need to carry any of its own. The dipole drive is constructed from two parallel screens, one charged positive, the other negative, creating an electric field between them with no significant field outside. Ambient solar wind protons entering the dipole drive field from the negative screen side are reflected out, with the angle of incidence equaling the angle of reflection, thereby providing lift if the screen is placed at an angle to the plasma wind. Protons entering from the positive side are accelerated out the negative screen, producing thrust. The dipole drive can achieve more than 3 mN/kWe in interplanetary space and better than 10 mN/kWe in Earth, Venus, Mars, or Jupiter orbit and offers potential as a means of achieving ultra-high velocities necessary for interstellar flight.

Abstract: A gas turbine engine includes: a compressor, a combustor, and a turbine arranged in sequential flow relationship along a primary flowpath, the turbine being connected in mechanical driving relationship to the compressor, so as to define at least one engine rotor that is rotatable about a centerline axis of the engine; a secondary flowpath connected in flow communication with the primary flowpath; and an air cycle machine including an air cycle rotor carrying an at least one air cycle compressor and at least one air cycle expander, wherein: the air cycle rotor is coupled in mechanical driving relationship with the at least one engine rotor; the air cycle rotor is coupled in fluid flow communication with the secondary flowpath; and the air cycle rotor is coupled in fluid flow communication with at least one heat exchanger.

Abstract: A turbine engine system includes a gutter and a gear train with an axial centerline. The gutter is disposed radially outside of the axial centerline. The gutter includes an inner surface and a channel that receives fluid directed out of the gear train. The inner surface at least partially defines a bore in which the gear train is arranged. The channel extends radially into the gutter from the inner surface, and circumferentially to a channel outlet. The bore has a cross-sectional bore area, and the channel has a cross-sectional channel area that is substantially equal to or less than about two percent of the bore area.

Abstract: A portable fuel preservation system is disclosed. The portable fuel preservation system may comprise a switch box configured to be coupled to an integrated fuel pump and control of a gas turbine engine. The switch box may comprise a circuit configured to cause a metering valve and a solenoid valve of the integrated fuel pump and control to open. A driver may be configured to inject preservation fluid into the integrated fuel pump and control.

Abstract: A fuel nozzle apparatus for a gas turbine engine includes: a first pilot fuel injector disposed on a centerline axis of the fuel nozzle which defines a direction of air flow through the fuel nozzle, the first pilot fuel injector being of a pressure atomizing type; an annular second pilot fuel injector at least partially surrounding the first pilot fuel injector, the second pilot fuel injector being of an air blast type and having a fuel outlet disposed axially downstream and radially outboard of the first pilot fuel injector; an annular venturi surrounding the first and second pilot fuel injectors, the venturi including a throat of minimum diameter; an array of inner swirl vanes extending between the first pilot fuel injector and the second pilot fuel injector; and an array of outer swirl vanes extending between the second pilot fuel injector and the venturi.

Abstract: Disclosed is a system including a turbine having a plurality of spokes. The first spoke end is coupled to a shaft and the second spoke end is coupled to a blade of a plurality of blades. Each blade is a cup with an open surface. A dispenser includes a combustion chamber. An air injector is configured to inject air into the combustion chamber. A fuel injector is configured to inject fuel into the combustion chamber. An ignitor is configured to supply a spark for combustion in the combustion chamber. A nozzle directs discharged fluid toward the open surface of the blade to drive the turbine. A housing encloses the second nozzle end and the plurality of blades, and has an exhaust pipe. A controller is configured to control the air injector, the fuel injector and the ignitor.

Abstract: A fluid cooling system for use in a gas turbine engine including a fan casing circumscribing a core gas turbine engine includes a heat source configured to transfer heat to a heat transfer fluid and a primary heat exchanger coupled in flow communication with the heat source. The primary heat exchanger is configured to channel the heat transfer fluid therethrough and is coupled to the fan casing. The fluid cooling system also includes a secondary heat exchanger coupled in flow communication with the primary heat exchanger. The secondary heat exchanger is configured to channel the heat transfer fluid therethrough and is coupled to the core gas turbine engine. The fluid cooling system also includes a bypass mechanism coupled in flow communication with the secondary heat exchanger. The bypass mechanism is selectively moveable based on a temperature of a fluid medium to control a cooling airflow through the secondary heat exchanger.

Abstract: A nacelle for an engine includes an exterior housing a least partially surrounding the engine. The nacelle further includes a front portion proximate an engine face of the engine. The front portion of the nacelle defines an opening into an interior of the nacelle. The nacelle further includes a seal disposed proximate the opening. The seal is configured to selectively allow air into the interior of the nacelle. The nacelle further includes a heat exchanger disposed within the interior of the nacelle. The heat exchanger is configured to exchange heat between a fluid flowing within the heat exchanger and air at the interior of the nacelle.

Abstract: A rich-quench-lean combustor assembly for a gas turbine engine includes a liner extending between a forward end and an aft end. The liner includes a plurality of quench air jets positioned between the forward end and the aft end. The combustor assembly additionally includes a dome attached to or formed integrally with the liner, the dome and the liner together defining at least in part a combustion chamber. A fuel nozzle is attached to the dome, the fuel nozzle configured as a premix fuel nozzle for providing a substantially homogenous mixture of fuel and air to the combustion chamber, the mixture of fuel and air having an equivalence ratio of at least 1.5.

Abstract: A fuel nozzle with turning guide is included in a gas turbine. The turning guide is disposed in an air inlet of the fuel nozzle to distribute a flow of compressed air and includes at least one of a turning separator, an inner separator, and an outer separator. The fuel nozzle includes a central body having an outer wall; a shroud concentrically disposed with respect to the central body and configured to surround the central body while maintaining a space for an air passage between an inner wall of the shroud and the outer wall of the central body; a rim formed on one end of the shroud and forming an air inlet communicating with the air passage; and a turning guide including a turning separator disposed in the air inlet, to make the air flow uniform, thereby suppressing the creation of an air pocket.

Abstract: A turbofan gas turbine engine includes low pressure (LP) motor-generator operable in either a generation mode or a drive mode, an high pressure (HP) motor-generator operable in a either a generation mode or a drive mode, and a power control module for selectively operating the LP and HP motor-generators according to operational conditions of the engine.

Abstract: A bearing assembly may comprise a housing, a slinger disk, and a bearing. The housing may be configured to receive a shaft and having at least one scavenge passage between an outer housing portion and an inner housing portion. The slinger disk may have a plurality of flow guides configured to direct the lubricant flow toward the at least one scavenge channel. The scavenge passage may be configured to provide a lubricant flow through the bearing. Lubricant flow may be provided in a forward direction, through the bearing and opposite a windage direction.

Abstract: A combustor nozzle assembly and a gas turbine having the combustor nozzle assembly includes a nozzle casing for receiving compressed air from a compressor, a curved surface at a corner continuing from a surface of the inner surface thereof that is perpendicular to the direction of the compressed air, a fuel spray hole formed at the curved surface, and a fuel chamber having a predetermined volume and formed inside the curved surface to supply fuel into the nozzle casing through the fuel spray hole.

Abstract: To improve air/fuel mixing at low speed, the invention concerns an aerodynamic injection system (30) for a combustion chamber (8) of an aircraft turbine engine, the system comprising a central body (70) comprising a first air circulation space (72) and a swirler (74) allowing air to penetrate into the first air circulation space (72), the hollow body (70) also comprising a fuel circuit (92, 93, 82) arranged at least partially around the first space (72), the circuit comprising a second annular space (82) for a fuel film to circulate and fuel supply structures (92, 93) communicating with the second space (82), the spaces (72, 82) opening into a third air and fuel mixing space (86). According to the invention, the system comprises fuel injection openings (90) that bring the fuel circuit into communication with the first space (72).

Abstract: A combustion staging system is provided for fuel injectors of a multi-stage combustor of a gas turbine engine. The system has a splitting unit which receives a metered total fuel flow and controllably splits the metered total fuel flow into out-going pilot and mains fuel flows to perform pilot-only and pilot-and-mains staging control of the combustor. The splitting unit includes a metering valve and a spill valve, a first portion of the total metered fuel flow received by the splitting unit being an inflow to the metering valve and a second portion of the total metered fuel flow received by the splitting unit being an inflow to the spill valve. The metering valve is configured to controllably determine a fuel flow rate of a metered outflow formed from the first portion of the total metered fuel flow.

Abstract: A gas turbine engine of an aircraft includes: an engine core having a turbine including a lowest pressure rotor stage, a turbine diameter, a fan including a plurality of fan blades extending from a hub, an annular fan face at a leading edge of the fan defining a fan tip radius at the fan face; wherein a downstream blockage ratio is defined as: the ? ? turbine ? ? diameter ? ? at ? ? an ? ? axial ? ? location of ? ? the ? ? lowest ? ? pressure ? ? rotor ? ? stage a ? ? distance ? ? f ? rom ? ? a ? ? ground ? ? plane ? ? to ? ? the ? ? wing ? and wherein an engine blockage ratio of: ( 2 × the fan tip radius/the engine length ) the ? ? downstream ? ? blockage ? ? ratio is in the range from 2.5 to 4.

Abstract: The present invention provides methods and system for power generation using a high efficiency combustor in combination with a CO2 circulating fluid. The methods and systems advantageously can make use of a low pressure ratio power turbine and an economizer heat exchanger in specific embodiments. Additional low grade heat from an external source can be used to provide part of an amount of heat needed for heating the recycle CO2 circulating fluid. Fuel derived CO2 can be captured and delivered at pipeline pressure. Other impurities can be captured.

Abstract: A combustion staging system is provided for fuel injectors of a multi-stage combustor of a gas turbine engine. The system has a splitting unit which receives a metered total fuel flow and controllably splits the metered total fuel flow into out-going pilot and mains fuel flows to perform pilot-only and pilot-and-mains staging control of the combustor. The system further has pilot and mains fuel manifolds which respectively receive the pilot and mains fuel flows. The system further has a plurality of mains flow scheduling valves which distribute the mains fuel flow from the mains fuel manifold to mains discharge orifices of respective injectors of the combustor. The system further has servo line which extends to the mains flow scheduling valves, the servo line being controllably fillable with pressurised fuel to change the pressure in the servo line relative to the pressure in the mains fuel manifold.

Abstract: A mobile fracking system and methods may include a gas turbine housed at least partially inside a trailer and an exhaust attenuation system configured to receive exhaust gas from the gas turbine. The exhaust attenuation system may include a lower elongated plenum configured to receive exhaust gas from the gas turbine and an upper noise attenuation system that is movably connected relative to a distal end of the lower elongated plenum.

Abstract: The present disclosure is directed to a fuel nozzle including a center body having a tube shape and a ring manifold disposed at an aft end of the center body. The fuel nozzle also includes an inner tube extending axially through the ring manifold and disposed within the center body. The inner tube is in fluid communication with a diluent supply. The fuel nozzle further includes a fuel tube extending helically around a portion of the inner tube. The fuel tube fluidly couples a fuel plenum of the ring manifold to a liquid fuel supply. Furthermore, the fuel nozzle includes a plurality of fuel injectors circumferentially spaced within an outer band of the ring manifold and in fluid communication with the fuel plenum. Each fuel injector is oriented to direct atomized liquid fuel radially outward from the center body. The ring manifold and the inner tube are thermally decoupled.