Abstract: An air vehicle propulsion system incorporates an engine core with a power shaft to drive an outer blade row. The power shaft extends through and is supported by a counter rotation transmission unit which drives an inner blade row in counter rotational motion to the outer blade row. The counter rotation transmission unit exchanges power from the engine core with the shaft. An actuator engages the shaft for translation from a first retracted position to a second extended position.

Abstract: An example nacelle assembly for a gas turbine engine includes a nacelle defined about an axis and defining a boundary of a fan bypass flow path. A fan variable area nozzle includes a plurality of inserts movably mounted to said nacelle. Each of the multiple of inserts is located at a circumferential position about the nacelle. The multiple of inserts are each independently moveable into the fan bypass flow path relative the nacelle to selectively vary a fan nozzle exit area.

Abstract: A fuel injector (36) for alternate fuels (26A, 26B) with different energy densities. Vanes (47B) extend radially from a fuel delivery tube structure (20B) with first and second fuel supply channels (19A, 19B). Each vane has first and second radial passages (21A, 21B) communicating with the respective fuel supply channels, and first and second sets of apertures (23A, 23B). The first fuel supply channel, first radial passage, and first apertures form a first fuel delivery pathway providing a first fuel flow rate at a given fuel delivery pathway backpressure that is essentially common to both sets of fuel delivery pathway apertures. The second fuel supply channel, second radial passage, and second apertures form a second fuel delivery pathway providing a second fuel flow rate that may be at least 1 about twice the first fuel flow rate at the given fuel delivery pathway backpressure.

Abstract: A hybrid rocket motor is manufactured by photopolymerizing the solid fuel grain in a stereolithography method, wherein fuel grains in a plastic matrix are deposited in layers for building a solid fuel rocket body in three dimensions for improved performance and for a compact design, the hybrid rocket motor including buried radial channels for defining a desired burn profile including the oxidizer to fuel burn ratio.

Abstract: A burner having two ignition electrodes and a bracket is provided. The bracket is arranged on the outer surface of the burner and the ignition electrodes are held, at in each case three points of their periphery, in a fixed position using the bracket. A gas turbine with the burner is also provided.

Abstract: A method is described for controlling load variations in a gas turbine. The method comprises reducing the flow of gaseous fuel entering the combustor to a predefined minimum value, if an increase is observed in the rotation regime of said turbine above a predefined maximum value and a total reduction in the load, activating a selective feeding sequence of the burners if the turbine is operating in normal functioning or premixed flame mode, modifying the angulation of the adjustable stator vanes, in order to reduce the speed rate of the compressor, and opening one or more anti-surge valves and one or more overboard bleeds, in order to reduce the air flow at the inlet of the combustor.

Abstract: The invention relates to a gas turbine, for energy generation, with a compressor, arranged coaxially to a rotor, mounted such as to rotate, for the compression of an inlet gaseous fluid, at least partly serving for combustion of a fuel in a subsequent annular combustion chamber, with generation of a hot working medium, with an annular diffuser arranged coaxially to the rotor, between the compressor and the annular combustion chamber, for distribution and deflection of the fluid, whereby a part of the fluid is diverted as cooling fluid for the turbine stages after the combustion chamber, by means of a dividing element, arranged in the fluid flow.

Abstract: In one embodiment, a method is provided fuel staging for a trapped vortex (TVC) combustion apparatus comprising an inlet premixer, for injecting fuel-air mixture into the inlet of the combustion apparatus and a vortex premixer, for injecting fuel-air mixture into the recirculating vortex. The combustion apparatus may be part of an engine, such as a gas turbine engine. The method comprises varying the relative proportion of mixture introduced through the inlet and vortex premixers as a function of load.

Abstract: A method and apparatus to determine a first heating value of a low BTU fuel, determine a target fuel quality level based on a state of a turbine system, control a second heating value of a high BTU fuel, and inject the high BTU fuel into the low BTU fuel to achieve the target fuel quality level.

Abstract: A gas turbine engine system comprising, a gas turbine engine including a combustion area, a laser, a fuel nozzle including an optical fiber operative to transmit light emitted from the laser to the combustion area, the optical fiber also transmitting a fuel into the combustion area, wherein the light is operative to ignite the fuel in the combustion area.

Abstract: A turbofan engine pylon structure has a fan variable area nozzle defined by a variable area flow system between a pylon intake and a pylon exhaust to selectively adjust a bypass flow through the pylon structure. The variable area flow system changes the physical area and geometry of a pylon nozzle exit area to manipulate the bypass flow by opening and closing an additional flow area of the variable area flow system.

Abstract: A nacelle assembly includes an inlet section having a plurality of discrete sections. Each of the plurality of discrete sections includes an adaptive structure. A thickness of each of the plurality of discrete sections is selectively adjustable between a first position and a second position to influence the adaptive structure of each of the plurality of discrete sections.

Abstract: A method of controlling mixing of a flow exiting a downstream end of a primary nozzle associated with a jet engine. The method may involve coupling a shape memory alloy (SMA) element to a mixing structure disposed at the downstream edge of the primary nozzle. An electrical signal may be applied to the SMA element to heat the SMA element and induce a phase change in the SMA element. The phase change may cause an axial length of the SMA element to constrict, to cause movement of the mixing structure into a path of the flow exiting the primary nozzle.

Abstract: The present invention describes a hybrid rocket motor that includes a first solid reactant and at least one thrust nozzle and at least one moveable combustion control member within the hybrid rocket motor that restricts the contact of a first fluid reactant in the combustion chamber. In this way, it is then possible to regulate the exposure of the solid reactant to the fluid reactant and thus control thrust.

Abstract: A vectoring nozzle with external actuation generates thrust vectoring by applying mechanical or fluidic actuation, or both, on the nozzle deck, external sidewalls, and/or air vehicle aft body to produce changes in the aft body flowfield and/or exhaust plume. An external mechanical sidewall may be integrated into a nozzle deck or side walls without the need for engine bleed to supply fluid injectors. An external fluidic vectoring system uses injectors or plasma devices located aft of the nozzle exit to vector the exhaust plume with no external moving parts. Elements of both mechanical and fluidic systems may be combined for a given application.

Abstract: A nozzle is provided and includes an outer annulus defined by an exterior wall and an interior wall and including air inlets through which air flows to a fuel mixing zone and a combustion zone, an inner annulus disposed within the interior wall and including a fuel volume into which fuel is fed to a distal end thereof, which is adjacent to and isolated from the combustion zone and an airflow line, disposed between the fuel volume and the interior wall, through which air flows to the combustion zone with the airflow line and the combustion zone isolated from the fuel volume, and a fuse configured to melt during a flameholding incident and to form a breach through which fuel flows from the fuel volume, bypassing the fuel mixing zone, to a fuel burning zone downstream from the fuel mixing zone.

Abstract: An air-blast fuel nozzle assembly includes a housing having an inner surface defining an interior chamber around a central axis. The inner surface terminates in an exit aperture. An air swirler pneumatically communicates with the interior chamber and has vanes operative to impart a swirling motion to air passing across the vanes and into the interior chamber. A fuel injection assembly includes a nozzle portion extending along the central axis and having a plurality of outlets circumferentially-arranged around the central axis that are directed into the interior chamber toward the inner surface of the housing. A shield extends radially outwardly from the nozzle portion upstream of the plurality of outlets. The shield extends between a base at the nozzle section and a free tip end such that the shield extends partially across the interior chamber toward the inner surface.

Abstract: A nozzle arrangement for use in a gas thruster is presented. At least one heater micro structure (20) is arranged in a stagnation chamber (12) of the gas thruster. The heater microstructure (20) comprises a core of silicon or a silicon compound coated by a surface metal or metal compound coating. The heater microstructure (20) is manufactured in silicon or a silicon compound and covered by a surface metal coating. The heater microstructure (20) is mounted in the stagnation chamber (12) before or after the coverage of the surface metal or metal compound coating. The coverage is performed by heating the heater microstructure and flowing a gas comprising low quantities of a metal compound. The compound decomposes at the heated heater microstructure (20), forming the surface metal or metal compound coating. The same principles of coating can be used for repairing the heater microstructure (20) in situ.

Abstract: A system that includes a turbine fuel nozzle comprising an air-fuel premixer. The air-fuel premixed includes a swirl vane configured to swirl fuel and air in a downstream direction, wherein the swirl vane comprises an internal coolant path from a downstream end portion in an upstream direction through a substantial length of the swirl vane.

Abstract: A combined cycle power generating plant that can realize a reduction in the start-up time includes a gas turbine, a steam turbine, a steam supplying section that supplies steam to the gas turbine, a first steam pipe that directs steam from the steam supplying section to the gas turbine, a second steam pipe that directs steam from the gas turbine to the steam turbine, a first release section that carries out control such that the supply destination of steam that is directed to the first steam pipe is one of either the outside of the gas turbine or the outside of the first steam pipe, a second release section that carries out control such that the supply destination of steam that is directed to the second steam pipe is one of either the outside of the steam turbine or the outside of the second steam pipe, and a bypass pipe that directs at least a portion of the steam inside the first steam pipe to the second steam pipe between the gas turbine and the second steam pipe.