Abstract: A combustor for a gas turbine includes a combustor liner, an outer casing surrounding the combustor liner, an ignitor housing extending through an ignitor opening through the liner, and an ignitor disposed within the ignitor housing. The ignitor housing includes a housing wall including at least one airflow inlet passage on a downstream side of the housing wall and arranged within the outer flow passage, and at least one cooling passage on the downstream side within the housing wall. The at least one cooling passage extends along a length of the ignitor housing from the at least one airflow inlet passage and through the inner end of the ignitor housing, the at least one cooling passage being in fluid communication with the at least one airflow inlet passage.

Abstract: A spacecraft propulsion system including a first thruster system including a first power processing unit connected to a first thruster string via a crossover switching unit and connected to a second thruster string via the crossover switching unit. A second thruster system including a second power processing unit connected to the second thruster string via the crossover switching unit and to the first thruster string via the crossover switching unit. A controller is connected to each of the first power processing unit, the second power processing unit, and the crossover switching unit.

Abstract: A monolithic combustion liner for use in a gas turbine engine includes fuel channels integrated into the wall of the combustion liner. The integrated fuel channels can have an aerodynamic shape to reduce flow losses of cooling air flowing around the exterior of the combustion liner. The monolithic combustion liner allows more cooling air to flow around the combustion liner, increasing the cooling of the combustor region of the gas-turbine engine.

Abstract: A dome-deflector assembly for a combustor of a gas turbine includes a dome portion having a dome-side swirler opening therethrough, and a deflector portion having a deflector-side swirler opening therethrough. The dome portion and the deflector portion are connected together to form a dome-deflector cavity therebetween. The deflector portion includes a first-side surface and a second-side surface, the second-side surface being arranged within the dome-deflector cavity, and the second-side surface including a plurality of stress-relief contours arranged in a pattern about the deflector-side swirler opening.

Abstract: A combustor wall includes a panel, a boss, a wall aperture and multiple cooling apertures. The panel extends axially along and circumferentially about an axial centerline. The panel extends radially between a first panel surface and a second panel surface. The first panel surface forms a peripheral boundary of a combustion chamber. The boss projects out from the second panel surface. The boss extends circumferentially around and forms an outer peripheral boundary of the wall aperture. The wall aperture extends along a wall aperture centerline through the combustor wall to the first panel surface. The cooling apertures are arranged circumferentially about the wall aperture. Each of the cooling apertures extends along a cooling aperture centerline through the panel and/or the boss to the wall aperture. The cooling aperture centerline of a first of the cooling apertures is angularly offset from the wall aperture centerline by a first acute angle.

Abstract: Disclosed are a jet nozzle, and a combustor and a gas turbine including the jet nozzle. The jet nozzle is formed on an inner surface of a combustor to supply fluid into an interior of a liner of the combustor in a direction intersecting with a progress direction of combustion gases primarily combusted in a combustion chamber in the interior of the liner. The jet nozzle includes a body section provided with a fluid inlet on one end, through which fluid is introduced, and a fluid outlet on the other end, through which the fluid introduced through the fluid inlet is discharged, and a swirler section provided inside the body section to allow a portion of the fluid introduced into one side of the body section to rotate in a clockwise direction and to allow the rest of the fluid introduced into the other side of the body section to rotate in a counterclockwise direction.

Abstract: Provided are systems and methods of controlling load share in a multi-engine propulsion system. A load share participation value indicative of a load share participation regime for a first engine may be received and a load sharing requirement control command for the first engine may be output. The load share participation value may depend or be based at least in part on one or more operating conditions of a second engine. The load sharing requirement control command for the first engine may depend on whether the load share participation value indicates a full-participation load share regime for the first engine, a partial-participation load share regime for the first engine, or a non-participation load share regime for the first engine.

Abstract: The present disclosure relates to a system for supplying fuel to a turbomachine. In some embodiments, a fuel circuit includes a pressurisation valve at an outlet of the system and a pump. The circuit may include a flow rate sensor arranged between the outlet of the pump and the pressurisation valve. In some embodiments, the flow rate sensor may include a sliding drawer, a restoring spring, and a sensor for detecting the position of said drawer in order to indicate the flow passing through the flow rate sensor. The system may include a device arranged to drive the pump with a controllable rotational speed and a control configured to control the device on the basis of a measurement supplied by the flow rate sensor, in such a way as to adapt the rotational speed of the pump shaft.

Abstract: A method of manufacturing a nested combustor liner includes manufacturing a nested combustor liner into a green state including a plurality of annular interior walls radially adjacent to one another and circumferentially surrounding an exhaust duct aperture and a plurality of annular exterior walls radially adjacent to one another and radially spaced apart from and circumferentially surrounding the plurality of annular interior walls and an ignitor wall attached to a first annular interior wall at a first interior end, extending radially toward and attached to a first annular exterior wall at a first exterior end. The method includes assembling the plurality of annular interior walls and the plurality of annular exterior walls, forming an assembled combustor liner. The method includes densifying the assembled combustor liner.

Abstract: A liner assembly for a combustor. The liner assembly includes a liner defining a combustion chamber of the combustor. The liner includes a looped section at a forward end of the liner. The looped section includes one or more bends such that the looped section forms a compliant joint and vibrations are dampened through the liner downstream of the looped section.

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: Methods for controlling actuating components of turbine engines using an adaptive damping lag filter are provided. The adaptive filter includes features that filter out insignificant changes in actuator demand, respond to fast transient conditions to follow demanded positon of the actuating component more closely, and adapts the gain of the output position to avoid stall conditions.

Abstract: A method includes delivering fuel and a first portion of oxidant as a rich mixture to the first zone of the combustion chamber. The fuel includes ammonia (NH3). The method further includes burning the rich mixture in the first zone. Combustion gases containing nitrogen oxides (NOx) are produced. The method further includes delivering a second portion of oxidant into the second zone to break down unburned ammonia into ammonia intermediates in the second zone. The nitrogen oxides (NOx) are consumed by reacting with the ammonia intermediates in the second zone. Byproduct hydrogen is produced as a result of breaking down the unburned ammonia into the ammonia intermediates. The method further includes delivering a third portion of oxidant into the third zone. The byproduct hydrogen is burned in the third zone. The third portion of oxidant is greater than the first portion of oxidant and the second portion of oxidant.

Abstract: A turbine engine can include a compressor section, a combustion section, and a turbine section in serial flow arrangement. A combustor in the combustion section can include a combustor liner at least partially defining a combustion chamber. The combustor liner can include at least one aperture fluidly coupled to the combustion chamber.

Abstract: The present technology is directed to plasma systems and associated methods, including propulsion systems for flight vehicles. A representative system includes a plurality of coils. The coils include a first coil positioned along a force axis, a second coil positioned along the force axis and spaced apart from the first coil, and a third coil that is magnetically shielded. A controller is operatively coupled to the coils and is configured to (a) increase energy to the first coil to generate a magnetic field in a portion of the plasma adjacent to the first coil, (b) decrease energy to the first coil and increase energy to the second coil to translate the resulting superposed magnetic field through the plasma to a position adjacent the second coil, and (c) transfer energy from the second coil to the third coil and decrease energy to the second coil to reduce the magnetic field in the plasma.

Abstract: Transducer assemblies that selectively measure a fuel pressure within a fuel supply system of an aircraft, aircraft that include the transducer assemblies, and methods of utilizing a transducer assembly. The transducer assemblies include an electronic pressure transducer, which is configured to measure the fuel pressure and to generate a fuel pressure signal indicative of the fuel pressure, and an electronic pressure display, which is configured to receive the fuel pressure signal and to display the fuel pressure responsive to receipt of the fuel pressure signal. The transducer assemblies also include a power supply structure and an attachment structure, which is configured to operatively interconnect the electronic pressure transducer to the aircraft. The attachment structure includes a fitting, which is configured to interface with a pressure switch receptacle of the aircraft, and a valve actuation structure, which is configured to selectively actuate a receptacle valve of the pressure switch receptacle.

Abstract: An electrothermal plasma production device is presented. The plasma production device includes: a plasma production chamber; an RF antenna external to the plasma production chamber; a propellant tank and flow regulator external to the plasma production chamber and in communication with the plasma production chamber; and a plenum disposed between the propellant tank and the plasma production chamber. The RF antenna, in combination with an AC power source, is configured to provide an RF energy to an interior region of the plasma production chamber and to an interior region of the plenum with sufficient power to ionize at least some of the propellant in the plenum. The plasma production chamber is configured to include a propellant injector for receiving propellant at a first closed end of the plasma production chamber.

Abstract: A method for controlling an overdetermined system with multiple power-restricted actuators that perform a primary task and non-primary tasks, including: a) determining a pseudo-control command based on a physical model of the system, which pseudo-control command represents the torques and a total thrust force acting on the system, b) determining a control matrix, c) dissociating the control matrix into sub control matrices, wherein the sub control matrices and the corresponding sub pseudo-control commands correspond to the primary task for i=1 and for i>1 correspond to the non-primary task(s) and a priority of the non-primary tasks decreases with increasing index i, d) determining actuator control commands for solving the primary task, e) projecting the non-primary tasks into the null space of the primary task, and into respective null spaces of all of the non-primary tasks of higher priority, if present, and f) providing the actuator control commands from d) and e) at the actuators.

Abstract: A mass propelled device is described. The mass propelled device includes an evaporation chamber. The evaporation chamber has at least one transparent substrate configured to receive light on a first surface and a plurality of nanostructures disposed on a second surface opposite the first surface. The plurality of nanostructures excite electrons in response to light being provided to the first surface. The mass propelled device also includes propellent storage to store propellent and a propellent delivery system to provide propellent from the propellent storage component to the evaporation chamber. The evaporation chamber is configured to heat the propellent using electrons. The mass propelled device also includes at least one nozzle configured to exhaust heated propellent from the evaporation chamber in order to produce thrust.

Abstract: Provided herein are various leptonic power sources, leptonic control systems, and leptonic-powered engines. An apparatus includes a leptonic source configured to emit beam electrons to ionize a material into a plasma according to a selectable ionization degree and deposit charge onto a plurality of cathodes in a progressively more negatively charged arrangement to establish an electric field of a selectable intensity in the plasma.