Abstract: A gas-circulation system for conditioning a disk of an aircraft may comprise a first takeoff port configured to extract a combusted gas and a second takeoff port configured to extract an uncombusted gas. A first valve may comprise an inlet in fluid communication with the first and second takeoff ports and an outlet of the first valve in fluid communication with the disk.

Abstract: The proposed technology relates to a thrust control apparatus of a propulsion system, and more particularly, to a thrust control apparatus of a solid propulsion system equipped with an aerospike pintle nozzle. The present invention is to simultaneously control the magnitude and direction of thrust by installing a pintle and a thrust vectoring unit at the rear end of a combustion tube of a solid propulsion system.

Abstract: A system for motoring a gas turbine engine of an aircraft is provided. The system includes an air turbine starter operable to drive rotation of a starting spool of the gas turbine engine and a starter air valve operable to deliver compressed air to the air turbine starter in response to the starter air valve being open. The system also includes a metered orifice coupled in a bypass configuration around the starter air valve to deliver a reduced amount of the compressed air to the air turbine starter while the starter air valve is closed. The reduced amount of the compressed air delivered to the air turbine starter limits a motoring speed of the gas turbine engine below a resonance speed of the starting spool of the gas turbine engine.

Abstract: A flexible bearing assembly includes at least one metal end ring, a flexible bearing core having a plurality of layers of a resilient material between layers of a reinforcement material, and a phenolic composite material between and bonded to each of the at least one metal end ring and the flexible bearing core. A rocket motor assembly includes a chamber configured to contain a propellant and a movable thrust nozzle coupled to the chamber. The movable thrust nozzle includes a phenolic composite material between and bonded to each of a metal end ring and a flexible bearing core. Methods of forming a flexible bearing assembly include bonding a phenolic composite material to at least one metal end ring and bonding a flexible bearing core to the phenolic composite material. The flexible bearing core includes a plurality of layers of a resilient material between layers of a reinforcement material.

Abstract: A gas turbine engine includes a compressor, a combustor fluidly connected to the compressor via a flow path, and a turbine fluidly connected to the combustor via the flow path. At least one multi-piece structure is disposed within the flow path such that the multi-piece structure at least partially radially spans the flow path. The at least one multi-piece structure includes a fore portion defining a leading edge, a consumable aft portion defining a trailing edge, a pressure surface at least partially defined by a first surface of the fore portion and a first surface of the consumable aft portion, and a suction surface at least partially defined by a second surface of the fore portion and a second surface of the consumable aft portion.

Abstract: A cooling mechanism includes a bottom wall (22) in contact with a combustion chamber, an upper wall (30), and a cooling passage (40) arranged between the bottom wall (22) and the upper wall (30). The cooling passage (40) includes a first passage (50) extending to a first direction, a second passage (60) extending to the first direction, and a connection section (70) connected with the first passage (50) and the second passage (60). The second passage (60) is arranged to have an offset to the first passage (50) in a second direction perpendicular to the first direction and along the bottom wall (22).

Abstract: A solid rocket motor comprises a pressure vessel, a solid propellant structure within the pressure vessel, and a flight termination system overlying the pressure vessel. The flight termination system comprises a shaped charge configured and positioned to effectuate ignition of an inner portion of the solid propellant structure and a reduction in an ability of the pressure vessel to withstand a change in internal pressure. Another solid rocket motor, a multi-stage rocket motor assembly, and a method of destroying a launch vehicle in flight are also described.

Abstract: The present invention provides a thrust flow powered vehicle comprising a first thrust flow expeller for expelling a first thrust flow in a first direction, a second thrust flow expeller for expelling a second thrust flow in a second direction, the second direction being a different direction to the first direction but sharing a plane with the first direction, a thrust flow deflector surface at an angle to the plane of the first and second directions, and an outlet portion for providing an output thrust flow, such that, in use, the thrust flow deflector surface deflects at least a portion of both the first and second thrust flows to form the output thrust flow such that the output thrust flow has a component in the plane of the first and second directions, and a component out of that plane.

Abstract: A rocket throat insert including an annular body having a radially inner annular wall portion and a radially outer annular portion. The inner wall portion has a contoured radially inner surface defining a nozzle throat. The outer portion includes an annular buttressing structure supporting the inner wall portion and defining one or more insulation gaps arranged annularly around the inner wall portion. The insulation gaps restrict the radial flow of heat through the annular body.

Abstract: A thrust ring for a rocket motor for one or more of limited re-use or single use, including a rocket motor housing, the motor housing adapted to contain propellant; an aft closure with a nozzle, the aft closure connected or connectable to the housing; a forward closure connected or connectable to the housing; wherein one or both of the aft closure and the forward closure are connectable to the housing in manner adapted for one or both of limited re-use or single use.

Abstract: Provided is a thrust control apparatus having a plug area variable spike pintle nozzle, in which an aerospike-shaped pintle can be used in a pintle nozzle for precise thrust control. The thrust control apparatus includes a combustion chamber, a motor housing, and an outer pintle nozzle. The motor housing has a driving motor within the combustion chamber. The outer pintle nozzle is moved in a lateral direction by the driving motor.

Abstract: A solid rocket motor includes a first solid propellant and a second solid propellant at least partially surrounding the first solid propellant. The second solid propellant is resistant to fragment impact and the first solid propellant has a higher impulse than the second solid propellant.

Abstract: A fuel nozzle is provided for a combustor of a gas turbine engine. The fuel nozzle includes an outer air swirler along an axis. The outer air swirler defines an outer annular air passage. An inner air swirler along the axis defines an annular fuel gas passage around the axis between the outer air swirler and the inner air swirler. An annular liquid passage is defined between the inner air swirler and an air inflow tube. A tube is within a housing to define an annular gas passage around the tube. The tube is operable to communicate a liquid into the annular liquid passage and the annular gas passage is operable to communicate a gas into the annular fuel gas passage.

Abstract: A monopropellant rocket engine in which nitrous oxide is the propellant includes a catalyst bed on an end of a diverging combustion chamber to inject a first flow of monopropellant to ignite the monopropellant, followed by a series of additional monopropellant injections along the diverging combustion chamber where the previously injected monopropellant decomposes to be used to ignite the downstream injected monopropellant. When all monopropellant has been decomposed, the monopropellant is passed through a throat and nozzle to produce thrust. A control valve is used to regulate an amount of monopropellant injected at each point in order to regulate an amount of thrust.

Abstract: The invention relates to the field of feeding devices, in particular for feeding rockets on the ground. A feeding device (4) of the invention comprises at least two mutually complementary feeding connectors (5, 6), a breakable connection member (15) connecting these two feeding connectors (5, 6) together and presenting a breakable section (15c) between these two feeding connectors (5, 6), and a force transmission member (16) connected to said breakable section (15c) in such a manner as to transmit a breaking load thereto in order to unlock the connection between the two feeding connectors (5, 6). During unlocking, a breaking load transmitted by the force transmission member (16) thus serves to break the breakable section (15c).

Abstract: A combustor assembly for use in a gas turbine engine is provided. The combustor assembly includes a dome assembly, a fuel nozzle coupled to the dome assembly, and a cowl assembly. The cowl assembly includes a radially inner cowl coupled to the dome assembly, and a radially outer cowl including a first cowl member and a second cowl member each positioned about the fuel nozzle. The first and second cowl members substantially align to define a substantially continuous flow path configured to direct a flow of air about the dome assembly.

Abstract: A rocket motor includes a case and first and second nozzles in the case. The first nozzle is disposed in the second nozzle. The first nozzle includes a forward leg, a rear leg, and an intermediate leg. The intermediate leg has a convex conical geometry, and the forward leg has a forward lip that is spaced from the case. The rear leg has a rear lip that is spaced from the case. The forward leg and the rear leg at least partially define a flow passage through the first nozzle. The first nozzle is exclusively secured by the intermediate leg to at least one of the case or the second nozzle. At least a portion of a fragmentation system is disposed between the first and second nozzles.

Abstract: A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a platform having an outer surface and an inner surface that axially extend between a leading edge portion and a trailing edge portion. At least one augmentation feature is disposed on at least the leading edge portion or the trailing edge portion of the outer surface of the platform.

Abstract: A thruster has an additively-manufactured housing that includes an integrally-formed nozzle with a burst disk in it. The housing is part of a casing that surrounds and encloses a propellant that is burned to produce pressurized gases that burst the burst disk and produce thrust. The thruster may be placed in a receptacle that defines a recess for receiving the thruster. The receptacle also may be additively manufactured. The thruster and the recess both may be cylindrical, with the housing being closely fit with the cylindrical walls of the receptacle. This may allow some of the structural loads on the housing, such as loads produced by the combustion of the propellant, to be transferred to the adjoining walls of the receptacle. This enables the housing to have less structural strength than if it were to have to contain the pressure from the propellant all on its own.

Abstract: An aircraft comprising a fuselage and propelled by a turbine engine having two coaxial and contrarotating fans, the turbine engine comprising a power turbine having two contrarotating rotors, one of which drives a fan upstream from the turbine, the other a fan downstream from the turbine, each fan comprising a ring of blades, and the assembly of the fans and the power turbine being incorporated at the rear of the fuselage, in the extension of same. The aircraft comprises, for at least one of the fans, a device for blocking the rotation of the fan and a device configured to modify the pitch of the blades of the fan in such a way as to make it operate as a flow straightener with respect to the other fan.