Abstract: Cryogenic propellant tank (1) for a spacecraft engine, comprising an external enclosure (10) defining an internal volume, characterized in that the internal volume of the tank comprises a primary volume (V1) and a secondary volume (V2) connected to the primary volume (V1) via a valve (20) configured to selectively allow a passage of fluid from the primary volume (V1) to the secondary volume (V2), or to isolate the secondary volume (V2) from the primary volume (V1), the primary volume (V1) having a primary orifice (11) adapted to be connected to a first pressurization source (41), the secondary volume (V2) having a supply orifice (4) adapted to be connected to a supply line of a spacecraft engine (30), and a secondary orifice (12) adapted to be connected to a second pressurization source (42).

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 projectile is disclosed, having: a longitudinal axis, a steering assembly, a shell body, an attitude control system, a despin module, an electromagnetic receiver and/or emitter system, and a controller. The attitude control system includes a ram air inlet in selective open fluid communication with an exhaust assembly, which includes a plurality of exhaust outlets to selectively generate each of a plurality of thrust jets from a ram air inflow provided by the ram air inlet, each thrust jet being selectively controllable via the controller. The despin module is configured for selectively de-spinning the steering assembly with respect to the shell body about the longitudinal axis. The electromagnetic receiver and/or emitter system is configured for receiving and/or emitting electromagnetic energy, and for cooperating with the controller for operating the exhaust assembly to thereby selectively provide steering control moments. Systems and methods for steering the projectile are also disclosed.

Abstract: Exemplary projectiles and methods associated therewith including embodiments formed with an internal cavity adapted to receive and retain a cryogenic material into said cavity and then generate a first internal gas upon thermal equalization with said projectile as well as a first internal gas pressure within said cavity. Exemplary embodiments include a structure adapted for maintaining structural integrity after generation of the first internal gas pressure and a second internal gas pressure that is created upon the firing of the projectile. In some embodiments, the second internal gas pressure is more than twice the first internal gas pressure. Some embodiments are adapted with a portion of the projectile formed for displacing away or laterally from an axis formed through a longitudinal center of the projectile upon an impact from striking an object after firing based in part on internal gas pressure and an impact at cavity wall section rupture zones.

Abstract: A bomb for deployment from an air vehicle includes a rocket motor for propelling the bomb. The rocket motor includes propellant which at least partially defines a void downstream of an initial burning surface. As the propellant is burned the void will be exposed, increasing the surface area of the burning surface of the propellant to increase the thrust of the rocket motor.

Abstract: A projectile launching device includes a reactive driver, a flyer housing, a flyer and a compressible buffer member. When detonated, the reactive driver will generate a detonation shock wave. The flyer housing defines a bore. The flyer is disposed in the bore and has a rear surface. The buffer member is interposed between the reactive driver and the flyer. The buffer member has a front surface in direct contact with the rear surface of the flyer. The buffer member is configured and arranged to: receive the detonation shock wave from the reactive driver; modify the detonation shock wave to generate a modified shock wave; and transmit the modified shock wave directly to the flyer to thereby propel the flyer away from the buffer member.

Abstract: A gun fired projectile includes a rocket motor housing including a pressure chamber and an exhaust nozzle. A plurality of propellant cells are positioned within the pressure chamber. The rocket motor propellant is mechanically supported during the severe gun fire event. This support may take several forms, each of which is discussed herein. The projectile further includes a support structure including one or more supports: wherein each of the one or more supports is engaged with the rocket motor housing. Each of the one or more supports is engaged with one propellant cell of the plurality of propellant cells, and each of the one or more supports suspends an individual propellant cell from the remainder of the plurality of propellant cells. All of these approaches provide the opportunity to tailor the performance of the rocket motor by combining a combination of propellant formulations and geometries to optimize the projectile performance.

Abstract: Embodiments of a vortice-amplified diffuser section for use in a buoyancy dissipater are generally described herein. The vortice-amplified diffuser section may include a plurality of diffusion ports to diffuse an expanding gas, a reduction sleeve to adjust an amount of diffusion flow, and vortex generators within at least some of the diffusion ports to generate vortices. The reduction sleeve may be configurable to block off some of the diffusion ports. The vortex generators may generate vortices of gas bubbles in the water to reduce the water's buoyancy and to inhibit movement or disrupt the operations of an errant vessel. The reduction sleeve may be used to control the size, shape, and intensity of the expanding gas bubble or bubble plume as well as to control the lethality level of the buoyancy reduction.

Abstract: A gun fired projectile includes a rocket motor housing including a pressure chamber and an exhaust nozzle. A plurality of propellant cells are positioned within the pressure chamber. The rocket motor propellant is mechanically supported during the severe gun fire event. This support may take several forms, each of which is discussed herein. The projectile further includes a support structure including one or more supports: wherein each of the one or more supports is engaged with the rocket motor housing. Each of the one or more supports is engaged with one propellant cell of the plurality of propellant cells, and each of the one or more supports suspends an individual propellant cell from the remainder of the plurality of propellant cells. All of these approaches provide the opportunity to tailor the performance of the rocket motor by combining a combination of propellant formulations and geometries to optimize the projectile performance.

Abstract: Embodiments of a buoyancy dissipater and method for deterring an errant vessel are generally described herein. In some embodiments, a volume of gas is generated from a propellant and diffused below a waterline of a vessel. The resulting gas bubble dissipates the buoyancy of the vessel providing a non-lethal deterring effect. The buoyancy dissipater includes a diffuser having radially-positioned diffusion ports to radially diffuse the gas generated by a gas generator below the waterline. In some embodiments, the radially-positioned diffusion ports comprise holes positioned radially around the diffuser to diffuse the gas around the buoyancy dissipater.

Abstract: A method for operating a thrust-generating supercavitating projectile involves launching the projectile in water from rest at the maximum available thrust, maintaining that thrust until supercavitating movement begins, and then reducing thrust to a near-minimum amount that is required to maintain supercavitating movement of the projectile.

Abstract: A method for operating a thrust-generating supercavitating projectile involves launching the projectile at a velocity above the minimum required to maintain supercavitating movement, delaying initiation of thrust until the projectile slows to a velocity that is near that minimum velocity, and then applying thrust to maintain the near-minimum velocity until a target is reached.

Abstract: Embodiments of a buoyancy dissipater and method for deterring a vessel are generally described herein. In some embodiments, a volume of gas is generated from a propellant and diffused below a waterline of a vessel. The resulting gas bubble dissipates the buoyancy of the vessel providing a non-lethal deterring effect.

Abstract: A projectile may be used in a boat tail or base bleed configuration. The projectile includes a warhead casing having a nose end and a base end; a payload disposed in the warhead casing; a motor body attached to the base end of the warhead casing and having an igniter disposed therein; a base closure attached to the motor body, an exterior surface of the base closure being tapered to form a boat tail, the base closure and the motor body defining an interior volume; a propellant disposed in the interior volume, the propellant being used for base bleed operation; and a converter attached to the base closure and extending into the interior volume.

Abstract: A supercavitating projectile is disclosed and includes a combustion chamber, a forward-directed jet nozzle and a comparatively larger gas duct/rear-directed jet nozzle. The combustion chamber is filled with a propellant having a hollowed core. The core serves as a pathway to fluidly allow combustion gases to the jet nozzles. In operation, the propellant combusts to form gasses forced forward through the forward-directed nozzle to generate a virtual cavitator in the form of a ventilation gas bubble. Combusted gasses are also forced out the rear-directed nozzle forming a propulsion jet. The projectile therefore uses the rear-directed jet to maintain a cruise velocity approximate to the launch velocity and employs a source of ventilation gas using the forward-directed jet for supercavitating of the projectile.

Abstract: A scramjet has a cowl, a center structure, and a plurality of wide pylons connecting the cowl to the center structure, with scramjet engines positioned between adjacent pylons. Leading surfaces of adjacent pylons converge to one another to provide side wall compression to air entering the engines. The center structure includes a fore body, a center body and an aft body that, with the pylons, define a basic structure either formed entirely from one piece or several securely connected pieces. A method of testing the scramjet projectile comprises using a gun to accelerate the scramjet projectile to the takeover velocity of the engines.

Abstract: An embodiment of the rocket motor of this invention employs an insensitive munitions approach that, when subjected to elevated external temperatures, is activated by thermal expansion of the main propellant and gas generation from a secondary insensitive munitions charge. In a preferred embodiment, the rocket motor also includes a pressure equalizing system that accommodates changing temperature conditions during storage as well as varying gas pressure inherent in gun-launched systems in a manner that allows for thinner case cylinder design and increased propellant volume.

Abstract: The present invention comprises a device which is inert and “thermally-equivalent” to actual ordnance. The device can travel with live ordnance and track the propellant temperatures in order to get a more precise propellant temperature for the ordnance. The device comprises a thermally equivalent inert grain instrumented with thermocouples, connected to a data recorder, and packaged in scaled-down ordnance hardware. The hardware is scaled down to enable the device to more easily travel with live munitions.

Abstract: An embodiment of the rocket motor of this invention employs an insensitive munitions approach that, when subjected to elevated external temperatures, is activated by thermal expansion of the main propellant and gas generation from a secondary insensitive munitions charge. In a preferred embodiment, the rocket motor also includes a pressure equalizing system that accommodates changing temperature conditions during storage as well as varying gas pressure inherent in gun-launched systems in a manner that allows for thinner case cylinder design and increased propellant volume.

Abstract: An embodiment of the rocket motor of this invention employs an insensitive munitions approach that, when subjected to elevated external temperatures, is activated by thermal expansion of the main propellant and gas generation from a secondary insensitive munitions charge. In a preferred embodiment, the rocket motor also includes a pressure equalizing system that accommodates changing temperature conditions during storage as well as varying gas pressure inherent in gun-launched systems in a manner that allows for thinner case cylinder design and increased propellant volume.

Abstract: A propulsion-assisted projectile has a body, a cowl forming a combustion section and a nozzle section. The body has a fuel reservoir within a central portion of the body, and a fuel activation system located along the central axis of the body and having a portion of the fuel activation system within the fuel reservoir. The fuel activation system has a fuel release piston with a forward sealing member where the fuel release piston is adapted to be moved when the forward sealing member is impacted with an air flow, and an air-flow channel adapted to conduct ambient air during flight to the fuel release piston.

Abstract: A rocket according to the invention includes a motor having a fuel chamber and an oxidizer tank, a nose portion provided forward of the motor, and a nozzle surrounded by fins provided aft of the motor. The fuel chamber is provided with a relatively central solid propellant, and a hybrid fuel surrounds the solid propellant. The oxidizer tank is filled with a reactant and coupled to the forward end of the fuel chamber. A pathway is provided between the tank and the fuel chamber for the passage of reactant therethrough. At least one of a valve and a barrier is coupled in the pathway to prevent passage of the reactant into the fuel chamber until after the solid propellant is at least partially consumed. After the solid propellant is at least partially consumed, the valve is opened and/or the barrier is removed to permit the passage of reactant into the fuel chamber.

Abstract: One embodiment of this novel rocket system employs at least one rocket motor closure that includes low and high load capability retainers that, respectively, provide an insensitive munitions system during storage and positive locking of a nozzle assembly at lunch. Another embodiment of this novel rocket involves an insensitive munitions approach that does not require use of such retainers by developing different normal and abnormal burn paths. The rocket motor also includes a pressure equalizing system that accommodates changing temperature conditions during storage as well varying gas pressure inherent in gun-launched systems in a manner that allows for thinner case cylinder design and increased propellant volume.

Abstract: Gas generator apparatus usable with a ducted rocket propellant having a high exponent burn rate, a burn rate that is burn chamber pressure sensitive to a greater degree than are normally used propellants. The gas generator includes elementary pressure regulating elements and apparatus for damping movement of the pressure regulating elements to remove oscillatory components of motion therefrom. Burn rate operating point determination and the stability characteristics thereof are described along with the effects of variations in propellant burn rate exponent and regulating element parameters.

Abstract: A rocket system employing at least one rocket motor closure that includes low and high load capability retainers that, respectively, provide an insensitive munitions system during storage and positive locking of a nozzle assembly at launch. The rocket motor also includes a pressure equalizing system that accommodates changing temperature conditions during storage as well as varying gas pressure inherent in gun-launched rocket systems in a manner that allows for thinner case cylinder design and increased propellant volume.

Abstract: The disclosure pertains to a projectile guided by successive lateral gas jets by way of pairs of impellers. The impellers of one pair are fired so as to exert opposite moments of rotation. According to the disclosure, the impellers are brought together in two groups. Of the impellers of one pair, one is in one group and the other is in the other group. The impellers of each of the groups are separated by diode bulkheads that withstand the pressure of the impeller that is located on one side of the bulkhead and is fired first, but does not withstand the firing of the impeller that is located on the other side of the bulkhead and is fired second. The impellers of one group lead into one and the same nozzle. Thus, the making of the guidance impellers is simplified at the same time as the control of the guidance is improved.

Abstract: A launching device with a work-controlled gas generator employs a dampening piston in combination with an accumulator volume to provide substantially uniform acceleration of a launching piston.

Abstract: A pyrotechnically driven device for restricting the area of the nozzle throat in a solid propellant rocket motor is disclosed. In one embodiment, the device includes a piston support mounted to the rocket motor in the exit cone. A restrictor is attached to a piston which is supported by the piston support. The restrictor is movable from a retracted position outside the nozzle throat to a restricting position within the throat upon actuation of the device. The piston and attached restrictor are driven by a pyrotechnic driving means that is actuatable separately from the solid propellant of the rocket motor. A locking means locks the restrictor in its restricting position.

Abstract: A spin-stabilized projectile with a correctable trajectory, wherein the projectile possesses lateral or transverse thrust-propulsion devices which are distributed about the circumference thereof, each incorporating an impulse-generating charge inclusive an electrically-activatable detonator arranged beneath a cover which is radially ejectable relative to the longitudinal axis of the projectile. The detonator, with the presence of a connecting cable which initially extends coaxially with the cover and then thereafter extends at an angle relative thereto, is arranged in the transitional region intermediate the impulse-generating charge and the cover, whereby the cover, with conformance of its external contour to the contour of the encompassing casing surface of the projectile, is fixed to the structure of the projectile across a rupture location, which is hermetically sealed radially inwardly of the impulse-generating charge with respect to the center of the projectile.

Abstract: This invention relates to a course correction unit, specifically but not exclusively for use on a spin stabilized guided projectile which can produce high speed instantaneous course correction.

Abstract: A device for the limitation of gas pressure in a projectile having an outer shell by the production of a control stream comprises a combustion chamber disposed inside the projectile; the combustion chamber having a combustion chamber lid; control lines connected to be changed by gas from the combustion chamber for transferring the gas pressure from the combustion chamber; monostable fluidic elements being arranged in the control lines in a region of the combustion chamber lid for assisting in confining the control stream; branching lines disposed in the lid and connected with the monostable fluidic elements, the branching lines including vent openings in the outer shell for releasing gas to the atmosphere and an overpressure element for permitting gas to be expelled from the combustion chamber when an overpressure exists therein, whereby gas expelled via the overpressure element assists in the production of the control stream.

Abstract: An improved rocket staging system for missiles and the like wherein a carriage borne rocket engine assembly is sequentially employed within separate, generally aligned oxidizer stages which are generally coaxially disposed about the central rocket engine and its associated carriage. A central fuel tank is surrounded by several separate, cooperating, generally ring-shaped oxidizer tanks generally coaxially disposed about the rocket periphery. A plurality of oxidizer delivery lines run through each of the outer tanks and up to the top of the fuel tank, where a flexible hose brings oxidizer down to the engine carriage. As fuel is consumed, the rocket motor carriage slides upwardly inside the fuel tank in response to thrust. When the carriage is firmly seated inside the next higher oxidizer tank and all of the propellant has been removed from the lowest tank, the lowest tank is jettisoned to discard unnecessary mass.

Abstract: A tandem arrangement of a flare and a booster rocket motor is held together by a hollow coupler having hinged spring-loaded wall segments resembling barrel staves that are released on launch to function as a braking device, similar to a drogue parachute. The coupler also contains the main parachute which is released when the rocket motor is separated from the flare. The spring loading on the coupler wall segments controls the parachute drag area as a function of air speed enabling a relatively constant deceleration rate to be realized during the transition from high speed to the final descent velocity.

Abstract: A gun-launched ramjet projectile follows a classical vacuum ballistic trajectory by maintaining a thrust-drag balance. The thrust of the ramjet is varied by an air door which bleeds air from the duct that leads from the supersonic diffuser at the front of the projectile to the combustion chamber. The control system for the air door includes an actuator and a pair of accelerometers, the output of which is integrated to provide an indication of velocity. The first accelerometer operates during the launching of the projectile in the gun to measure the launch velocity. A second, and more sensitive, accelerometer measures in-flight velocity changes due to atmospheric conditions. Both of the measured velocities are compared with prerecorded standards and error signals are derived representing any velocity deviations. These error signals are combined and then used to drive the air doors which varies the thrust of the ramjet so that it follows the classical vacuum ballistic trajectory.

Abstract: A gun-launched ramjet projectile follows a classical vacuum ballistic trajectory by maintaining a thrust-drag balance. The thrust of the ramjet is varied by an air door which bleeds air from the duct that leads from the supersonic diffuser at the front of the projectile to the combustion chamber. The control system for the air door includes an actuator and a pair of accelerometers, the output of which is integrated to provide an indication of velocity. The first accelerometer operates during the launching of the projectile in the gun to measure the launch velocity. A second, and more sensitive, accelerometer measures in-flight velocity changes due to atmospheric conditions. Both of the measured velocities are compared with prerecorded standards and error signals are derived representing any velocity deviations. These error signals are combined and then used to drive the air doors which varies the thrust of the ramjet so that it follows the classical vacuum ballistic trajectory.