Abstract: An igniter plug for an aircraft turbomachine includes an external electrode, an internal electrode, as well as a semiconducting body arranged between the external electrode and the internal electrode and set back from these electrodes so as to define an electrical arc forming cavity, the bottom of which is formed by an axial end surface of the semiconducting body. The free end of the external electrode is equipped with at least one scoop for discharging possible fuel residues present in the cavity.

Abstract: The present disclosure relates to a turbojet engine nacelle having an external cowling with an internal wall defining, with an internal fixed structure, an annular flow channel for a secondary airflow and an exhaust nozzle for said secondary airflow. The nozzle includes at least one opening, at least one continuous downstream end portion, downstream of said opening, and at least one door moving between a closed position and an open position allowing the passage of a part of said airflow through said opening. The nacelle is notable in that it includes a shaped device for activation of the door, during their activation from a closed position to an open position of the door, to drive the door in a combined movement that is translational in the direction running upstream from the nacelle and rotary towards the exterior of the nacelle.

Abstract: A combined cycle gas turbine plant and a corresponding method for operating such a combined cycle gas turbine plant, in which, during load operation of the combined cycle gas turbine plant, a water mass flow which is supplied according to the forced-flow principle to a waste heat steam generator is adjusted such that the evaporator heating surface of the medium pressure stage is oversupplied and thus a defined amount of excess water, which is heated in the evaporator heating surface but not vaporized, is discharged via the water-steam separator to a heat exchanger circuit for preheating fuel for the gas turbine.

Abstract: A combustor includes a plurality of bundled tube fuel nozzles which are annularly arranged around a common axial centerline. The plurality of bundled tube fuel nozzles comprise a first bundled tube fuel nozzle and a second bundled tube fuel nozzle. The first bundled tube fuel nozzle includes a first fuel plenum and a plurality of premix tubes which extend axially therethrough. The second bundled tube fuel nozzle includes a second fuel plenum and a plurality of premix tubes which extend axially therethrough. The first fuel plenum is axially offset at a predefined axial distance from the second fuel plenum so as to mitigate combustion tones within the combustor. In one embodiment, premix tube fuel ports disposed within the first and/or second fuel plenums may be axially offset with respect to each other within the corresponding fuel plenums at a predefined axial distance to mitigate combustion tones within the combustor.

Abstract: Hall thrusters with conductive coatings are disclosed. A Hall thruster comprises magnetic shielding in order to avoid collisions with the inner walls of its discharge chamber. By removing the source of erosion, the walls of the chamber can be removed reducing mass, cost and complexity of the thruster. A conductive coating, such as an aluminum coating, is deposited on inner screens between the discharge chamber and the magnetic poles of the thruster. The magnetic field within the chamber shields the conductive coating deposited on the inner and outer screens of the chamber.

Abstract: The present invention relates to an axial swirler, in particular for premixing of oxidizer and fuel in gas turbines. The axial swirler for a gas turbine burner includes a plurality of swirl vanes with a streamline cross-section being arranged around a swirler axis and extending in radial direction between an inner radius Rmin and an outer radius Rmax. Each swirl vane has a leading edge, a trailing edge, and a suction side and a pressure side extending each between the leading and trailing edges. A discharge flow angle ? between a tangent to the swirl vane camber line at its trailing edge and the swirler axis is first function of radial distance R from the swirler axis. A position of maximum camber of the swirl vane is second function of radial distance R from the swirler axis. At least one swirl vane of the first and second functions include each a respective local maximum and local minimum values along said radial distance from Rmin to Rmax. The invention also relates to a burner with such a swirler.

Abstract: Systems and methods for improving plasma propellant ablation/sublimation based systems are provided. One set of embodiments provides systems and methods for reducing carbon charring during plasma system (e.g., a plasma coating application system) propellant (e.g., a carbon-fluorine polymer) ablation and increasing heat transfer, ablation, and plasma thrust from plasma systems. In particular, one embodiment can include using a nano or micro-sized magnetic or electromagnetic field responsive material as particulates or microcapsules that are intermixed with polytetrafluoroethylene (e.g., Teflon®) nano-fibers, and using resulting fiber composites as the propellant material. Embodiments can include improved plasma system, e.g., pulsed plasma thrusters, plasma torches, plasma coating systems, etc, as well as nozzle improvements such as embodiments with magnetic structures disposed in relation to the nozzle. Alternative embodiments also include recovery and reuse systems.

Abstract: An electrothermal device (1, 100) includes a primary chamber (2) having an anode nozzle (6) provided with an inlet passage (7), a cathode tip (9) at least partially inserted into the inlet passage (7), and a primary air inlet (10) leading into the inlet passage (7), and a voltage generator (11) arranged between the anode nozzle (6) and the cathode tip (9) in such a way as to generate an electric arc (12) on the path of the primary air flow (13) injected into the primary chamber (2). It includes a secondary chamber (3) wherein a secondary air flow (15) circulates in a heat exchange relation with the heated primary air flow (14) from the primary chamber (2), the secondary air flow (15) having a lower temperature than the heated primary air flow (14) leaving the primary chamber (2).

Abstract: An ignition system for a combustor of a gas turbine engine is disclosed. The ignition system may include an igniter operatively associated with the combustor, and an electrode operatively associated with the combustor and spaced from the igniter, wherein an electrical potential is created between the igniter and the electrode to produce an electric arc therebetween.

Abstract: A combustion system includes a combustor having staging blocks and flashback detection units configured to detect flashbacks occurring in the staging blocks, and a combustion control device configured to stop a supply of a fuel for one of the staging blocks in which a flashback occurs based on a detection signal from one of the flashback detection units. The combustion control device includes a staging processing unit configured to instruct to supply the fuel or stop the supply of the fuel for each of the staging blocks based on a staging pattern in which an operating line is specified and a load adjustment unit configured to reduce an amount of load to a level at which the supply of the fuel for the one of the staging blocks is stopped when the detection signal is input from the one of the flashback detection units.

Abstract: A space vehicle engine (10) comprising a chemical thruster having a nozzle (30) for ejecting combustion gas, together with a Hall effect thruster. The engine is arranged in such a manner that the nozzle serves as the ejection channel for particles ejected by the Hall effect thruster when it is in operation. The engine can deliver high thrust with low specific impulse or relatively low thrust with large specific impulse.

Abstract: A system and method provides a fault-tolerant multi-channel pulsed plasma thruster system utilizing a control unit and an embedded real time application manipulating low-level timing events with programming, with clear examples of completely flexible control techniques of a scalable micropropulsion system having many pulsed plasma thruster channels, taking into account system aging behavior and specific mission utilization requirements that may change in the mission lifetime. The system and method also covers an architecture lending itself suitable for design of a dedicated FGPA or ASIC that would tightly integrate many channels of thruster components to build a robust, resilient and versatile micropropulsion subsystem for space applications, and indirectly for advanced multi-channel spacecraft instrumentation.

Abstract: The present invention is a spacecraft electromagnetic propulsion system which provides thrust without expelling propellant. The system includes a multi-element capacitor with rotating segmented cathodes, and electromagnetic coils for generating magnetic fields. Interactions between the electromagnetic fields and the rotating charged capacitor elements produce unidirectional forces, while reaction momentum is carried away by the Poynting Vector fields in conformity with the currently understood principles of modern physics.

Abstract: An assembly for a turbojet engine includes a pylon, a nacelle supported by the pylon, and a locking device interposed between the pylon and a cowl of the nacelle. The cowl slides between operating and maintenance positions. The locking device is movable between active and inhibited positions, and the device allows an operation mode in the active position and a maintenance mode in the inhibited position. In particular, the locking device comprises a rocker to maintain the cowl to the pylon, and the rocker includes an inhibiting lever and an activation lever. The inhibiting and activation levers are mounted on a switching shaft of the locking device. The inhibiting lever is set in rotation through an actuator, providing direct access from a bypass flow duct in the active position while the activation lever is set in rotation from the bypass duct in the inhibited position.

Abstract: Provided is a vapor jetting device including: a box-shaped casing; a propellant holding unit that is placed in the casing and forms a space for holding a propellant; a gas storing unit that is placed in the casing, forms a space that is divided from the propellant holding unit by a partition wall including a communication hole, and stores gas; a machine housing unit that is placed in the casing, forms a space that is divided from the propellant holding unit and the gas storing unit by partition walls, and houses machines; a nozzle that is connected to the casing and ejects the gas to an outside; a gas flow path that is formed in the casing and supplies the gas stored in the gas storing unit to the nozzle; and a heater that is placed in the casing and heats at least the gas storing unit.

Abstract: Thruster grid clear circuits and methods to clear thruster grids are disclosed. An example apparatus includes a low voltage grid clear circuit to apply first energy to a grid at a first voltage, and a high voltage grid clear circuit to detect a failure of the applied energy to clear a short circuit condition of the grid and to apply second energy to the grid at a second voltage higher than the first voltage.

Abstract: A variable bleed valve includes a variable bleed valve door disposed in a bleed inlet in transition duct and rotatable about an axis at or near a forward end of door. Valve is operable to open and close a bleed slot open to a booster bleed flowpath located outwardly of transition duct and having an airflow restrictor. Airflow restrictor may be variable such as a flapper valve. Booster bleed flowpath may pass through a bifurcated bleed duct including an inner passage radially bound by spaced apart inner and middle bleed walls and an outer passage radially bound by middle bleed wall and an outer bleed wall. Airflow restrictor may include an outlet area smaller than an inlet area of inner passage. Door seals against middle bleed wall to open inner passage and against outer bleed wall to open both passages.

Abstract: Plenums for bifurcated ducts and bifurcated ducts are provided for stabilizing flow therethrough. The plenum comprises an outer cylindrical body intersected by a pair of exhaust duct stubs that are configured to be coupled to a corresponding pair of exhaust ducts and an inner body. The outer cylindrical body includes an axial rear end portion. The inner body is disposed in the axial rear end portion and increases in diameter in the aft direction. The inner body comprises one of a generally axi-symmetrical inner body or a non-axi-symmetrical inner body. The bifurcated duct comprises the plenum and the pair of exhaust ducts. Exhaust systems are also provided.

Abstract: A nuclear thermal rocket with a superconducting electric motor driven boost pump submerged within a tank of liquid hydrogen, where the boost pump is driven by both an electric motor and a turbine. The boost pump can be submerged in liquid hydrogen so that the electric motor operates as a superconducting motor. Also, a turbopump for a rocket engine can include both a turbine and an electric motor to drive the liquid oxidizer and liquid fuel pumps of the turbopump.

Abstract: A combustor cap assembly includes an impingement plate coupled to an annular shroud and a cap plate which is coupled to the impingement plate to form an impingement air plenum therebetween. The combustor cap assembly further includes a flow conditioning plate coupled to a forward end portion of the shroud. The flow conditioning plate includes an inner band portion, an outer band portion and an annular portion. The annular portion defines a plurality of flow conditioning passages. The inner band portion at least partially defines a cooling air plenum within the combustor cap assembly. The inner band portion defines an exhaust channel which is in fluid communication with the impingement air plenum and an exhaust outlet. The flow conditioning plate further defines a cooling air passage which provides for cooling air flow into the cooling air plenum.