Abstract: This application describes creating, modifying, and bending electromagnetic solitons at large scales for the various applications. An electromagnetic soliton generator system controls the magnetic soliton such that the orientation, rotation rate, pitch angle, and magnetic field strength of the solitons are modified to provide the described standing waves and generate a magnetic flux differential.

Abstract: A method of controlling a gas turbine engine includes the steps of: detecting a shaft break event in a shaft connecting a compressor of the gas turbine engine to a turbine of the gas turbine engine; and in response to this detection, activating a shaft break mitigation system which introduces a fluid into a gas flow of the gas turbine engine downstream of the turbine, or increases an amount of a fluid being provided into the gas flow of the gas turbine engine downstream of the turbine, whereby the fluid reduces an effective area of a nozzle for the gas flow so as to reduce the mass flow rate of the gas flow through the turbine.

Abstract: Combustors and gas turbines are provided. A combustor includes an axial centerline and an end cover. The combustor further includes at least one fuel nozzle that extends from the end cover and at is least partially surrounded by a combustion liner. The combustion liner extends between the at least one fuel nozzle and an aft frame and that defines a combustion chamber. An outer sleeve is spaced apart from and surrounds the combustion liner such that an annulus is defined therebetween. The outer sleeve defines at least one aperture. A wake energizer is mounted on the outer sleeve. The wake energizer defines at least one passage that is angled with respect to the axial centerline of the combustor. The at least one passage aligns and is in fluid communication with the at least one aperture of the outer sleeve.

Abstract: A system and method for vectoring the thrust of a supersonic, air-breathing engine. A thrust vectoring mechanism uses two asymmetrically staggered ramps; one placed at the throat, the other positioned at the exit lip of the nozzle of the engine to re-direct exhaust flow off-axis with the nozzle.

Abstract: An embodiment of a torch igniter for a combustor of a gas turbine engine includes a combustion chamber oriented about an axis, a cap defining an axially upstream end of the combustion chamber, a tip defining the axially downstream end of the combustion chamber, an igniter wall extending from the cap to the tip and defining a radial extent of the combustion chamber, a structural wall coaxial with and surrounding the igniter wall, an outlet passage defined by the igniter wall within the tip, a glow plug housing configured to receive a glow plug and allow an innermost end of the glow plug to extend into the combustion chamber, and a cooling system. The cooling system includes an air inlet formed within an exterior of the structural wall, a cooling channel forming a flow path through the structural wall at the glow plug housing, and an air passage.

Abstract: A combustor nozzle and a combustor for a gas turbine including the same are provided. The combustor nozzle may include a plurality of vanes disposed radially on an outer circumferential surface thereof, each vane including an internal cavity to which fuel is supplied, wherein the vane may include an airfoil in cross-section, wherein at least one of a pressure surface and/or a suction surface of the airfoil is disposed in a height direction thereof with a plurality of nozzle holes communicating with the cavity, wherein the vane may include a V-shaped nozzle hole including a pair of outlets on both sides of the airfoil based on a leading edge of the airfoil, wherein the V-shaped nozzle hole may form an angle (?) diverging outward with respect to a horizontal plane across the leading edge, and an acute angle (?) radially outward with respect to a vertical plane perpendicular to the horizontal plane.

Abstract: A method of reducing noise from a combustor of a gas turbine engine includes the steps of establishing a maximum noise limit that may be for a particular frequency range. A primary fuel flow percentage, which may be emitted from a fuel nozzle arrangement having various groupings of simplex and duplex nozzles, is then established. An immersion depth measured between an aft rim of a swirler and a distal tip of the fuel nozzles may then be reduced thereby reducing the noise amplitude.

Abstract: A Hall-effect thruster assembly includes a plurality of magnetic sources for creating a magnetic circuit. The plurality of magnetic sources are positioned between a first end and a second, opposite end of the Hall-effect thruster. The plurality of magnetic sources define a longitudinal axis extending through the first end and the second end. The first end is configured as a discharge end. A mount assembly is coupled to the second end. The mount assembly is configured to secure the plurality of magnetic sources to a spacecraft. A magnetic element is supported by the mount assembly. The magnetic element is positioned relative to the plurality of magnetic sources by the mount assembly.

Abstract: A hybrid propulsion system for a small satellite package consisting of a main rocket motor containing a solid propellant with multiple oxidizer tanks positioned to direct oxidizer into the rocker motor, thereby producing a desired thrust necessary for orbit insertion and/or orbit correction. Additionally, oxidizers can serve a dual function in controlling cold fuel thrusters for attitude adjustment.

Abstract: A cartridge for attachment to the inner surface of a single-walled combustion liner of an annular combustor is provided. The cartridge includes at least one chute projecting into the combustion chamber for a counter swirl effect for improved fuel/air mixing in the combustion chamber and at least two studs projecting through associated stud holes in the liner. A method for attaching a cartridge to the inner surface of a single-walled combustion liner is also provided.

Abstract: An exemplary gas turbine engine includes a fan section including a fan rotor and at least one fan blade. A fan pressure ratio across the at least one fan blade is less than 1.45, noninclusive of the pressure across any fan exit guide vane system. The engine further includes a low-pressure compressor having a low-pressure compressor rotor that rotates together with the fan rotor at a common speed in operation, and a geared architecture that drives the low-pressure compressor rotor and the fan rotor. The geared architecture has a gear reduction ratio of greater than 2.5. The engine further includes a high-pressure compressor having a pressure ratio greater than 20, a low-pressure turbine having a pressure ratio greater than 5, and a bypass ratio greater than 10.

Abstract: A combustor liner for a gas turbine engine, the combustor liner including a panel configured to at least partially define a combustion chamber. The combustor liner further includes a shell configured to mount to the panel and form a gap between the panel and the shell. The panel includes a stud and a plurality of a stand-off pins proximate to the stud defining a cavity therebetween. The shell includes a plurality of angled impingement holes located away from the cavity but extending through the shell at an orientation such that cooling air passing through the angled impingement holes is directed towards the cavity between adjacent stand-off pins and at an acute angle relative to the stud.

Abstract: A method to attenuate noise associated with operation of a gas turbine engine during a fracturing operation may include connecting an inlet of a first elongated plenum to an exhaust duct of the gas turbine engine. The method further may include connecting a noise attenuator to a distal end of the first elongated plenum. The noise attenuator may include first and second silencer hoods. The method also may include pivoting the first and second silencer hoods from stowed positions to operative positions so that the second silencer hood in combination with the first silencer hood define a second elongated plenum positioned to receive exhaust gas via the first elongated plenum and the exhaust duct. The first elongated plenum and the second elongated plenum may increase an effective length of the exhaust duct and may reduce sound emission associated with operation of the gas turbine engine.

Abstract: A system for propelling a nanosatellite, including a pair of separated electrodes defining an ignition space therebetween a power source operationally connected to the pair of separated electrodes. Also included is a liquid propellant reservoir a pump reconnected in fluidic communication with reservoir and the ignition space and an electronic controller operationally corrected to the power source and to the pump.

Abstract: A combustion staging system for fuel injectors of a multi-stage combustor of a gas turbine engine. The system includes plural fuel injectors, each having respective pilot and mains injection stages. It further includes a splitting unit which, to perform staging control of the combustor, receives a metered fuel flow and, for pilot and mains operation, controllably splits the received fuel flow into a pilot flow for injecting at the pilot stages of the injectors and a mains flow for injecting at the mains stages of the injectors, and for pilot-only operation, controllably splits the received fuel flow into a first part of the pilot flow for injecting at the pilot stages of a first portion of the injectors and a second part of the pilot flow for injecting at the pilot stages of a second portion of the injectors.

Abstract: Some embodiments of the invention include a thruster system comprising a thruster and a pulsing power supply. The thruster may include a gas inlet port; a plasma jet outlet; and a first electrode. In some embodiments, the pulsing power supply may provide an electrical potential to the first electrode with a pulse repetition frequency greater than 10 kHz, a voltage greater than 5 kilovolts. In some embodiments, the pressure downstream from the thruster can be less than 10 Torr. In some embodiments, when a plasma is produced within the thruster by energizing a gas flowing into the thruster through the gas inlet port, the plasma is expelled from the thruster through the plasma jet outlet.

Abstract: A matrix thruster that may be used to reposition and/or stabilize a CubeSAT satellite. The matrix thruster includes a conductive plate with an opening, a plurality of wires within the opening, a power supply electrically connected to the conductive plate or each of the plurality of wires via an inductor, and an electrical switch. The electrical switch creates a current change that creates an electric potential spike across the inductor. The electric potential spike across the inductor initiates an arc discharge between one of the wires and the conductive plate, which forms plasma that ejects cathode particles from the matrix thruster. Using multiple wires (e.g., four titanium wires) extends the lifetime of the thruster, as each wire restores an inter-electrode film needed for the other wires to continue generating plasma.

Abstract: An insert for an impingement plate and an impingement plate including the insert are disclosed. The impingement plate includes a plurality of cooling holes therein. The insert includes a body including an opening extending longitudinally therethrough. The body also includes a discharge end configured for positioning in an impingement air plenum between the impingement plate and an upstream surface of a combustor cap plate; a flexible insertion end configured for insertion into a respective cooling hole of the plurality of cooling holes; and a fixation element between the discharge end and the flexible insertion end. The fixation element has an outer dimension configured to fixedly couple the body in the respective cooling hole.

Abstract: A gas turbine engine including: an exhaust diffuser including an inner tube and an outer tube that form therebetween an annular exhaust passage; a bearing chamber formed radially inside the inner tube for accommodating a bearing that supports a rotor of a turbine; a plurality of hollow struts extending across the exhaust passage; an oil introduction passage extending through one of the struts for introducing oil to be supplied to the bearing chamber; an oil drain passage extending through one of the struts for draining the oil from an exhaust oil inlet opened on a bottom surface of the bearing chamber; and an oil discharge passage for discharging a portion of the oil having passed through the oil introduction passage toward the oil drain inlet.

Abstract: A narrow channel Hall thruster comprising a thruster body with a magnetic circuit, an annular thruster channel having a channel width of less than 3 mm formed within the magnetic circuit, an annular anode, a cathode positioned externally to the thruster, and configured for electron emission, a power supply applying a positive potential to the anode, such that a plasma discharge can be generated in the annular thruster channel, and another power supply applying a negative potential to the cathode, relative to the thruster body and the anode. The second power supply reduces its negative voltage output to the cathode when the current supplied by the anode power supply exceeds a predetermined level, indicating that the discharge has reached a stable initiated condition. The reduction of the voltage output of the second power supply can be achieved either by self-regulation, or by use of a current limit circuit.