Abstract: The subject matter of this specification can be embodied in, among other things, a turbine combustor assembly includes a primary combustion chamber in fluid communication with a primary fuel injector and a primary air inlet, and an igniter carried by the primary combustion chamber and including a first igniter stage having an auxiliary combustion chamber housing comprising a mixing chamber and a tubular throat converging downstream of the mixing chamber, an ignition source projecting into the mixing chamber of the auxiliary combustion chamber, and a second igniter stage that includes an auxiliary fuel outlet manifold proximal to an outlet of the tubular throat.

Abstract: An assembly for a turbomachine nozzle having a longitudinal axis (X) includes an exhaust cone with an annular outer wall for guiding a primary air flow and an annular chamber placed radially inside said annular outer wall. The annular chamber includes an annular inner wall placed radially inside the annular outer wall of the exhaust cone. An exhaust case is placed upstream of the ejection cone. A connecting member is interposed longitudinally between the exhaust case and the exhaust cone and fastened to the exhaust case and to the annular inner wall. The assembly further includes elastic means configured to deform radially and prestressed radially between the annular outer wall and the connecting member.

Abstract: A water delivery system (18) for delivering water for injection into gas turbine engine combustor (4) includes a centrifugal pump (19) and a metering valve (23). The centrifugal pump (19) has an inlet (20) connected to a water source and a discharge (21) connected to a water supply line (22). The metering valve (23) is connected to the water supply line (22) downstream of the discharge (21) of the centrifugal pump (19). The water supply line (22) is connected to an injector nozzle (14) downstream of the metering valve (23). The metering valve (23) is operable to regulate a flow rate of water in the water supply line (22), to thereby meter an amount of water supplied to the injector nozzle (14).

Abstract: A thrust reverser for a turbojet aircraft nacelle includes a fixed structure, a movable structure and an actuator. The actuator extends along a main axis and is connected by a first connection to the fixed structure and by a second connection to the movable structure for the deployment of the movable structure between a direct jet position and a reverse jet position. The actuator is also connected to the fixed structure by a third connection arranged longitudinally between the first connection and the second connection. An axis of the third connection is radially eccentric from the main axis and the third connection allows a predetermined displacement of the actuator with respect to the fixed structure.

Abstract: Systems and methods for equalizing fluid levels within a vent line and a propellant tank in which the vent line is located are discussed herein. The vent line includes a vent valve and an equalization valve. The vent valve can be included in a vent duct of the vent line. The equalization valve is included in a bottom wall (i.e., a low point) of the vent duct of the vent line. A controller is also included in the system to instruct a vent valve and the equalization valve to open and close.

Abstract: Systems and methods for a fully reusable upper stage for a multi-stage launch vehicle are provided. The reusable upper stage uses an aerospike engine for main propulsion and for vertical landing. A heat shield can include a plurality of scarfed nozzles embedded radially around a semi-spherical surface of the heat shield, wherein inboard surfaces of the plurality of scarfed nozzles collectively define an aerospike contour. The heat shield can be actively cooled to dissipate heat encountered during reentry of the upper stage.

Abstract: The invention relates to an air/fuel injection system for a turbomachine, comprising:—an injector comprising a duct and an injection nose, arranged inside said duct, which extends from upstream to downstream along a longitudinal axis;—a mixer device comprising a bowl comprising an annular inlet, forming the inlet of the mixer device, from which there extends a conical portion flared in the downstream direction, said mixer device being arranged downstream of the injection nose; the injection system being characterized in that the injector comprises an air-injection annulus extending from the duct and from which there extends a connection ring comprising a divergent portion, said ring being arranged externally around the annular inlet of the mixer device.

Abstract: A propulsion system for an aircraft is disclosed and includes a water recovery system and a water monitoring system. The water recover system includes a condenser that is arranged along the core flow path and is configured to extract water from an exhaust gas flow. The water monitoring system includes a sensor and a controller programed to determine a condition of water and generate a prompt based on information communicated from the at least one sensor.

Abstract: A bundled tube fuel nozzle is provided. The bundled tube fuel nozzle includes a first end wall, a second end wall, an outer band that extends between the first end wall and the second end wall, and an inner band that extends between the first end wall and the second end wall. The inner band is disposed within the outer band. A bellows wall is disposed between the inner band and the outer band. The bellows wall surrounds the inner band such that a first fuel plenum is defined annularly between the inner band and the bellows wall. The inner band defines a second fuel plenum that is in fluid communication with the first fuel plenum via one or more apertures defined in the inner band. A plurality of tubes extends in an axial direction between the first end wall and the second end wall within the second fuel plenum.

Abstract: A propulsion system includes a drive shaft that rotates about an axis, a fan, a fan shaft that drives the fan about the axis, a reduction mechanism coupling the drive and fan shafts, a compression section driven by the drive shaft, and an inlet channel that extends between the fan and the compression section. The inlet channel includes inner and outer ferrules delimiting an air inlet flow path, the inner ferrule having a minimum radius. The reduction mechanism includes two reduction stages including two or more planet gears circumferentially distributed around the axis, each planet gear including a first portion meshed with the drive shaft and a second portion meshed with the fan shaft and being mounted fixed with respect to the inner ferrule and having a maximum radius greater than the minimum radius such that the air inlet flow path extends partly between the two or more planet gears.

Abstract: A fuel supply circuit including a supply duct in which a first flow (Fl) is able to circulate, at least one supply pump for circulating the fuel from a tank to a metering unit, and a recirculation duct for a second flow upstream of the pump, wherein the ducts (20, 60) discharge in different directions upstream of the pump; the circuit includes a dispensing device including an internal flow duct for the first flow (Fl), an external flow channel for the second flow and at least one passage orifice for the second flow, which communicates fluidically with the channel; and the at least one orifice (810) and the duct are coaxial and respectively discharge upstream of the pump.

Abstract: An engine can utilize a combustor to combust fuel to drive the engine. A fuel nozzle assembly can supply fuel to the combustor for combustion or ignition of the fuel. The fuel nozzle assembly can include a swirler and a fuel nozzle to supply a mixture of fuel and air for combustion. The fuel nozzle can include both a primary and secondary fuel passage, and an additional air passage to provide for greater flame control, fuel provision, or local fuel and air mixing prior to combustion.

Abstract: A heating system for a component within a compartment of a spacecraft includes a fuel source, a gas generator, and a heat sink. The fuel source includes hydrazine. The gas generator is in fluid communication with the fuel source. The gas generator includes a catalyst. The catalyst is configured to decompose the hydrazine and generate an exhaust gas. The heat sink is thermally coupled to the gas generator and configured to receive heat from the exhaust gas of the gas generator. The heat sink is thermally coupled to the component within the compartment of the spacecraft to transfer heat from the exhaust gas to the component.

Abstract: A method for controlling a gas turbine, having a measurement step, a prediction step which is carried out after the measurement step, and a control step which is carried out after the prediction step. In the measurement step, a state variable of a combustion within a gas turbine is measured. In the prediction step, a future combustion dynamic is predicted using the measured state variable. In the control step, a control signal is output using the prediction of the future combustion dynamic.

Abstract: Disclosed is a fuel mixed injection method for a gas turbine. The method includes the following steps: arranging a secondary fuel injection nozzle and a secondary air injection nozzle on a secondary combustion section, wherein the secondary fuel injection nozzles is closer to a main combustion section than the secondary air injection nozzle; and injecting secondary fuel and secondary primary air in sequence through the secondary fuel injection nozzle and the secondary air injection nozzle, respectively, thus enabling the secondary fuel to spontaneously combust in a mainstream high-temperature flue gas atmosphere to form a transverse jet flame and increase the flame lift-off height.

Abstract: A combustor for a gas turbine has a combustor liner including an outer liner and an inner liner, a combustion chamber being defined between the outer liner and the inner liner. The combustion chamber includes a primary combustion zone at an upstream end of the combustion chamber. The combustor also includes an outer liner expanded primary volume portion, and a secondary outer liner portion. One of the outer liner expanded primary volume portion and the secondary outer liner portion is movable to adjust a volume of the primary combustion zone by opening and closing access to the outer liner expanded primary volume portion so as to increase and to decrease the volume of the primary combustion zone.

Abstract: An atomizer provides a high-quality fuel-air mixture to a gas turbine engine, by combining air input from an engine compressor and fuel input from a single low-pressure fuel supply pump. The atomizer includes an atomizer body, a main vortex chamber, a secondary vortex chamber for improving quality of the fuel-air mixture, and a fuel sleeve providing fuel to the secondary vortex chamber. The main vortex chamber includes a main outlet nozzle in fluid communication with a combustion chamber inlet of the gas turbine engine. The secondary vortex chamber includes a secondary outlet nozzle in fluid communication with the main vortex chamber. The fuel sleeve has a blind channel with a longitudinal axis and a fuel tip. The same atomizer may be used for startup mode and for all operational modes of the gas turbine engine.

Abstract: A system includes a gas turbine having a turbine stage disposed in a combustion gas path, wherein the turbine stage includes turbine vanes disposed upstream from turbine blades. The system includes an isothermal expansion system coupled to the turbine stage. The isothermal expansion system includes multi-fluid injectors configured to vary axial positions of combustion within a turbine stage expansion of the turbine stage to reduce temperature variations over the turbine stage expansion, wherein at least one of the multi-fluid injectors is coupled to each of the turbine vanes. Each of the multi-fluid injectors includes a fuel port configured to inject a fuel, an oxidant port configured to inject an oxidant, and a barrier fluid port configured to inject a barrier fluid between the fuel and the oxidant, wherein the barrier fluid is configured to delay mixing between the fuel and the oxidant.

Abstract: A fuel injector for an aircraft turbine engine includes a tubular body having an axis of elongation. A first longitudinal end configured to be Supplied with fuel and a second longitudinal end configured to elect a jet of fuel. The body further includes an integrated purge-air circuit that has an internal cavity which is connected to air inlet orifices situated on the body and to at least one air outlet situated at said second end. Aire-flow disruptors are provided, projecting into said cavity.

Abstract: A gas turbine engine with a compressor section, a turbine section, and a combustion section located downstream from the compressor section and upstream from the turbine section, the combustion section including a dome inlet, a combustor outlet fluidly coupled to the turbine section, a liner and a dome assembly together at least partially defining a combustion chamber extending between the dome inlet and the combustor outlet, a fuel system fluidly coupled to the combustion section, the fuel system comprising a fuel supply, a primary fuel line fluidly coupling the fuel supply to the combustion section, and a reformer fluidly coupled to the fuel supply.