Abstract: The purpose is to develop, design a very inexpensive rocket to “clean up” the expendable debris orbiting the earth. This liter is a serious pollution of space and a threat to working satellites and the space station. Develop an inexpensive rocket engine with optimum efficiency conversion: propellants-chemicals to acquire the maximum thrust generation offering an inexpensive means to acquire low orbit to space. Parameters selected to minimize fuel consumption: by trying different hydrogen-oxygen mixture ratios. A critical design requirement for the Clean Up space engine is the ease to start and operate reliably in the cold vacuum conditions of space. The space engine will have a reliable pressure feed system, simplicity and dependability multiple starts.

Abstract: A thermal protection system and a method of manufacturing are disclosed. The thermal protection system may be configured to protect a movable joint, for example, a flexible bearing of a rocket motor nozzle. The thermal protection system includes a series of annular shims separated by a plurality of discrete spacers. Each shim of the series of annular shims may have a larger diameter than the previous shim, and the shims may nest. The shims may comprise a thermally stable material, and the discrete spacers may comprise an elastomer. Optionally, an annular bearing protector may separate the annular shims from the flexible bearing.

Abstract: A toroid pressure vessel includes a toroid body having an inner shell and an outer shell. The toroid body includes a toroid outer perimeter. The outer shell extends along the toroid outer perimeter. A planar exterior face extends along at least a portion of the outer shell and the toroid outer perimeter. A support belt circumscribes the toroid outer perimeter and is coupled along the planar exterior face. The support belt braces and supports the pressure vessel along the toroid outer perimeter against bulging force (and hoop stress) generated by pressurized fluids within the vessel. The support belt facilitates the use of thinner pressure vessel shells and thereby decreases the weight of the pressure vessel while providing a support to the outer shell that substantially prevents deformation of the planar exterior face.

Abstract: There is disclosed a vehicle and methods for maneuvering the vehicle. The vehicle may include a plurality of multiple-impulse rocket motors, each of which comprises a plurality of independently ignitable solid fuel propellant charges, and a processor that generates at least one command to ignite at least one solid fuel propellant charge of at least one of the plurality of multiple-impulse rocket motors.

Abstract: An internal combustion engine wherein a thermo potential heat flow in combustion is maximised by providing a feedback of an optimised amount of thermo potential heat flow that is modulated in the exhaust media, into the air intake, and a method of providing feedback comprises producing a shock wave of pulse of exhaust media and pulse of intake air on the opposite side of a high temperature sustainable wire screen modem thereby transferring the thermo potential heat energy flow from the exhaust media to the air intake.

Abstract: A standard 4-pole electric motor stator with capacitive plates in-line with the magnetic poles, but electrically 90° out of phase, produces two Lorentz force geometries 90° out of phase with each other (i.e., vertical, horizontal). An alternating source electrically rotates this pair of Lorentz geometries producing a propagating electromagnetic wave at the source frequency within the vacant internal cavity. Any charged particle within the cavity and along its axis will be accelerated or decelerated from an initial velocity via the Lorentz force. The rotating geometry provides for the coupling of the Lorentz force through a current loop and diamagnetism, providing acceleration and deceleration of non-charged particles. The force coupling is dependent upon the material's electromagnetic properties, the frequencies generated by the capacitive stator, and the velocity of the particles within the capacitive stator's influence.

Abstract: The invention is a universal aero, naval and space, one class integrated technology, originated from a common, fundamental unique basic absolute maximum efficiency technology, total endothermic, hyperbaric, adiabatic and stoichiometric maximum power capacity, with unlimited pressure ratio, which is the unification of the way to produce a top absolute efficiency by cryogenic electrification of all the transportation systems from ground to the entire universe.

Abstract: A torch igniter and method of cooling the torch igniter includes axially flowing a first stream of gaseous oxidizer through a torch throat and vortically flowing a second stream of the gaseous oxidizer within a combustion chamber.

Abstract: Supply of a liquid component in a combustion chamber of a rocket engine is controlled by a feed valve provided with an obturator mobile between a pen position and a closed position of at least one supply pipe, which has an inlet that communicates with a tank for containing the liquid component and an outlet that communicates with the combustion chamber; the displacement of the obturator from its closed position to its open position being triggered by a pressurized fluid supplied to the outlet of the supply pipe.

Abstract: An electrostatic colloid thruster for implementing a method of ionizing a liquid is disclosed herein. The electrostatic colloid thruster includes an electrically conductive extractor having a plurality of holes defined therethrough; an ultrasonic atomizer having an electrically conductive atomization surface at least partially facing the extractor and being arranged relative thereto so as to define a gap; a reservoir system in fluid communication with the atomization surface; and an electrical power source in electrical communication with both the extractor and the atomization surface. The apparatus and method are generally utile in various applications including, for example, spacecraft propulsion, paint spray techniques, semiconductor fabrication, biomedical processes, and the like.

Abstract: Methods, devices, and systems relating to a sensing device are disclosed. A device may comprise a structure including a first surface and a second, opposite surface, wherein the structure comprises one or more segments. Further, the device may include a plurality of sensors disposed on the structure, wherein each segment of the one or more segments comprises a first sensor of the plurality of sensors coupled to the first surface and an associated second sensor of the plurality of sensors coupled to the second surface. Moreover, each sensor of the plurality of sensors may be configured to measure a strain exhibited on an adjacent surface of the structure at an associated segment of the one or more segments.

Abstract: A liquid propellant tank for storing a liquid propellant A and supplying vapor produced by evaporation of part of the liquid propellant A to an external location comprises a tank body for storing the liquid propellant A, and a plurality of holder plates arranged inside the tank body, around an axis L of the tank body. The holder plates cause the liquid propellant A to adhere to them by surface tension, thereby establishing, inside the tank body, a liquid propellant holding area LA in which the liquid propellant A is held and a gas accumulation area GA in which vapor produced by evaporation of part of the liquid propellant A accumulates. The tank body has a propellant inlet open into the liquid propellant holding area LA and a gas outlet open into the gas accumulation area GA.

Abstract: In some embodiments a propulsion system includes a thrust chamber having an inside wall, an expansion nozzle mounted to the thrust chamber and having an interior and having an exterior, a main propellant injector mounted to the thrust chamber to inject a fluid in the interior of the thrust chamber, the fluid comprising oxidizer, fuel and internal film coolant, the internal film coolant ranging from about 1% to about 5% of the fluid, limited coolant tubing circumscribing the exterior of the expansion nozzle to circulate an external coolant, and an injector mounted to the expansion nozzle to inject the external coolant in the interior of the expansion nozzle, the external convective coolant about 2.5% of the fluid. The system operates at lower temperatures while having conventional amounts of thrust, in which the thrust chamber can be made of thin walls of lower cost conventional metals with simple coolant tube construction.

Abstract: A propulsion system for use in a liquid or gas fluid is provided including an axially-extending funnel-shaped conduit, a flow generator, a power source, and, optionally, an airfoil-shaped wing. The funnel-shaped conduit has outer walls forming an inner fluid passageway, an upper edge defining a fluid inlet, and a lower edge defining a fluid outlet. The optional airfoil-shaped wing is connected to and circumferentially surrounds the funnel-shaped conduit upper edge. The flow generator is rotatably mounted about the axis of the funnel-shaped conduit and is configured to force the fluid from the fluid inlet rearward through the fluid outlet. A forward force is produced by the combination of both thrust from the flow generator plus the lift force created as the flow generator draws the fluid across the annular airfoil-shaped wing and inwardly through the fluid inlet forcing the fluid rearward to exit out of the fluid outlet.

Abstract: An integral composite rocket motor nozzle. The novel nozzle includes a first layer of a first reinforcement material, a second layer of a second reinforcement material, and a common matrix material surrounding the first and second reinforcement materials such that the reinforcement materials and matrix material form a single integral composite structure. In an illustrative embodiment, the first reinforcement material includes graphite fibers for providing structural support, and the second reinforcement material includes glass or quartz fibers for providing thermal insulation on a first side of the first layer. The nozzle may also include a third layer of a third reinforcement material for providing thermal insulation on a second side of the first layer. In a preferred embodiment, the first layer is shaped to form an integrated dome and nozzle structure.

Abstract: An insulation composition that comprises at least one nitrile butadiene rubber, basalt fibers, and nanoclay is disclosed. Further disclosed is an insulation composition that comprises polybenzimidazole fibers, basalt fibers, and nanoclay. The basalt fibers may be present in the insulation compositions in a range of from approximately 1% by weight to approximately 6% by weight of the total weight of the insulation composition. The nanoclay may be present in the insulation compositions in a range of from approximately 5% by weight to approximately 10% by weight of the total weight of the insulation composition. Rocket motors including the insulation compositions and methods of insulating a rocket motor are also disclosed.

Abstract: A propulsive element usable for producing a jet of fluid using a pressurized fluid. An inlet receives the pressurized fluid; a propulsive element passageway extends from the inlet; two main outlets are in fluid communication with the propulsive element passageway and located substantially opposed to the inlet relative to the propulsive element passageway. The two main outlets are configured and sized for releasing each a respective main jet portion when the pressurized fluid is injected in the inlet, the two main jet portions being each substantially divergent, the two main jet portions creating a low pressure zone therebetween. An auxiliary outlet is located between the two main outlets, the auxiliary outlet being in fluid communication with the propulsive element passageway and located substantially opposed to the inlet relative to the propulsive element passageway.

Abstract: A pressure vessel for propellants, an explosion preventing method of the same, and a manufacturing method of the same are disclosed. The pressure vessel for propellants comprises a body provided with a cylindrical portion between a forward dome and an aft dome, and has an inner space where propellants are arranged. The pressure vessel also includes an insulation layer disposed on an inner wall of the body and configured to insulate the body from the inner space when the propellants are ignited; and a nozzle mounted at the aft dome and through which combustion material from the propellants is exhausted out. The body comprises a first hybrid fiber layer which forms the cylindrical portion, the forward dome, and the aft dome, and has a mechanical property lowered at a temperature more than a specific temperature such that the forward dome and the aft dome collapse by an inner pressure when the propellants are abnormally combusted.

Abstract: The sun imparts 174 petawatt per second on the earth, and a large portion of this energy is absorbed by the earth's atmosphere in the form of translational energy for the gaseous molecules, i.e. continuous random motion in the average speed range of 500 meters per second on earth's surface. This invention utilizes a partition with large number of through-holes which all have the characteristic of providing greater cross section for gas molecules to transit from one side to the other than the reverse, thus creating a higher statistical probability for the molecules to move from one side of the partition to the other side.

Abstract: A propulsion system for use in a liquid or gas fluid is provided including an axially-extending funnel-shaped conduit, a flow generator, a power source, and, optionally, an airfoil-shaped wing. The funnel-shaped conduit has outer walls forming an inner fluid passageway, an upper edge defining a fluid inlet, and a lower edge defining a fluid outlet. The optional airfoil-shaped wing is connected to and circumferentially surrounds the funnel-shaped conduit upper edge. The flow generator is rotatably mounted about the axis of the funnel-shaped conduit and is configured to force the fluid from the fluid inlet rearward through the fluid outlet. A forward force is produced by the combination of both thrust from the flow generator plus the lift force created as the flow generator draws the fluid across the annular airfoil-shaped wing and inwardly through the fluid inlet forcing the fluid rearward to exit out of the fluid outlet.

Abstract: A propulsion system and method are described configured to exert a force upon a vehicle. The system includes a concentrated mass which may be discharged from the vehicle by a propellant typically via an ejection tube. The system is optimized so that the discharged concentrated-mass imparts a large impulse upon the vehicle. The system and method may be used to alter the momentum of vehicles for propulsion, attitude correction, vehicle separation and the like.

Abstract: The present invention provides a heretofore-unknown use for zirconium nitride as a hydrogen peroxide compatible protective coating that was discovered to be useful to protect components that catalyze the decomposition of hydrogen peroxide or corrode when exposed to hydrogen peroxide. A zirconium nitride coating of the invention may be applied to a variety of substrates (e.g., metals) using art-recognized techniques, such as plasma vapor deposition. The present invention further provides components and articles of manufacture having hydrogen peroxide compatibility, particularly components for use in aerospace and industrial manufacturing applications. The zirconium nitride barrier coating of the invention provides protection from corrosion by reaction with hydrogen peroxide, as well as prevention of hydrogen peroxide decomposition.

Abstract: An insulation composition that comprises at least one nitrile butadiene rubber, basalt fibers, and nanoclay is disclosed. Further disclosed is an insulation composition that comprises polybenzimidazole fibers, basalt fibers, and nanoclay. The basalt fibers may be present in the insulation compositions in a range of from approximately 1% by weight to approximately 6% by weight of the total weight of the insulation composition. The nanoclay may be present in the insulation compositions in a range of from approximately 5% by weight to approximately 10% by weight of the total weight of the insulation composition. Rocket motors including the insulation compositions and methods of insulating a rocket motor are also disclosed.

Abstract: Nozzles that offer shape variability to maintain or purposely change the pressure drop across the throat are obtained by constructing the nozzles with components that change their shape, angle, or curvature in response to temperature changes that occur during the flow of combustion products through the nozzle. The temperature change may be the gradual heating of the nozzle wall from hot combustion gases, and the shape change may result in a decrease in the throat diameter or an expansion of the throat diameter. A decrease in throat diameter will be useful when the depletion of propellant as burning proceeds causes a drop in the pressure or flow rate of the combustion gas and there is a need to compensate for this drop to maintain the pressure drop across the throat. An increase in throat diameter will be useful when an initial high thrust is no longer needed and depletion of the fuel by itself is insufficient to lower the thrust to its desired reduced level.

Abstract: Control and/or drive device for a flying body for ejecting hot gas streams of a combusted fuel combination of at least a first and second component. Device includes a first hollow chamber body structured and arranged to contain first component, a second hollow chamber body structured and arranged to contain second component, a controllable fuel valve arranged between first hollow chamber body and second hollow chamber body to control feed of first component to second hollow chamber body, and a plurality of outlets structured and arranged to eject respective hot gas streams for influencing a flight path of flying body. Second hollow chamber body is formed as a combustion chamber for combusting the at least first and second components within second hollow chamber body to generate respective hot gas streams, and plurality of outlets are connected to the second hollow chamber body.

Abstract: The invention concerns an electrothermal thruster (1) comprising a heating section (20) through which a propellant fluid (4) is intended to pass so that it is heated before it is exhausted, the heating section (20) being supplied by an electric power source (22). According to the invention, the electric power source (22) comprises photovoltaic cells (24) mounted on the heat exchanger (10) through which the propellant fluid (4) is intended to pass before it reaches the heating section (20). Application to the area of spacecraft propulsion.

Abstract: Coaxial micronozzle designs are presented. Initial coaxial micronozzle designs utilizing center-body geometries to better exploit pressure thrust show a potential threefold increase in specific impulse under vacuum and near vacuum conditions.

Abstract: A rocket engine has spaced apart inner and outer skins each unitarily formed in one piece from carbon fiber fabric. The inner skin is formed on a two-part mold that is separated and removed from the inner skin after it is cured. An oxidizer ring encircles the bottom of the engine and is in flow communication with flow channels between the skins. Oxidizer tubes are connected at one end to the ring and at their other end to support brackets on the engine. The oxidizer ring is formed as an integral part of the engine by extending the outer skin over an inflatable mold, which is deflated and removed from the ring after the ring is cured. The oxidizer tube is formed on a mold with a rigid spine that holds the shape of the mold until the oxidizer tube is cured. The spine and the mold are then removed.

Abstract: Methods and systems are provided for propelling a vehicle. In an embodiment, by way of example only, a method includes flowing a decomposed hydroxylammonium nitrite (HAN)-based propellant into a chamber, introducing an aspirated non-polar fuel into the chamber, and combusting the decomposed HAN-based propellant and the aspirated non-polar fuel to produce an exhaust gas.

Abstract: An integrated composite structure with a graded coefficient of thermal expansion (CTE) is formed by selecting a plurality of layers of materials with a graded CTE and using build-up (bottom-up) fabrication approaches such as metal deposition or powder metallurgy to produce a CTE-graded layered composite preform, which is then consolidated and heat treated to create the CTE graded integrated composite billet or near net shape. The integrated composite billet or near net shape is then processed to produce a first surface for attachment of a first structural member having a first CTE and to produce a second surface of for attachment of a second structural member having a second CTE.

Abstract: A rocket engine having a combustion chamber, a throat, and an exhaust bell is made with spaced apart inner and outer skins each unitarily formed in one piece of carbon fiber fabric. Longitudinal ribs in the space between the skins reinforce the engine and divide the space into a plurality of flow channels. An oxidizer ring at the bottom of the exhaust bell is in fluid flow communication with the flow channels, and one or more oxidizer tubes are connected tangentially at one end to the ring to supply oxidizer to the ring and thence to the flow channels. The oxidizer tubes are connected at their other end to the engine above the throat, further reinforcing the engine. An igniter is in the combustion chamber, and ignition fuel ports are directed toward the igniter to provide a soft start ignition.

Abstract: A cured ablative composite assembly comprises a housing enclosing a pair of ablative composite sub-assemblies joined by a film adhesive. The cured ablative composite assembly is made by surface treating both ablative composite sub-assemblies in preparation for joining; coupling one ablative composite sub-assembly to another ablative composite sub-assembly with a film adhesive and enclosing the uncured ablative composite assembly within a housing; and depositing the combination of the housing and uncured ablative composite assembly in a ventilated oven with a load applied to the combination housing and uncured ablative composite assembly. The film adhesive is cured providing a portion of a hot gas valve suitable for use in tactical missiles. The film adhesive does not erode at the high temperatures (5000° F.) encountered in hot gas rocket exhausts, thereby providing a seal that offers high strength, pressure-tight joints.

Abstract: An ignition system for a rocket engine utilizes the pressure energy in a propellant flow. The propellant flow generates an oscillating pressure force in a resonance system which is then transmitted to a piezoelectric system. The electrical pulses are utilized to generate a spark in an igniter system spark gap, resulting in ignition. Since the spark energy production is driven by the resonance of the propellant flow, a fully passive auto-ignition system is provided. Once ignition occurs, the resultant backpressure in the combustion chamber “detunes” the resonance phenomena and spark production stops. Furthermore, should the engine flame out, spark production would automatically resume as the propellant valves remain open thereby providing relight capability.

Abstract: The thrust of a rocket motor can be varied while maintaining efficiency over a range of pressure ratios using a design that allows for changing the relative position of a plug and a combustion chamber exit. The plug or the chamber exit may be attached to an adaptive control system for position modification. The plug may be positioned in a plug nozzle configuration or in an expansion-deflection (ED) configuration. In either configuration, the elongated downstream portion of the plug allows for efficiency over a wide range of pressure ratios, while ability to change plug position with respect to the chamber exit allows adjustment of rocket thrust.

Abstract: Gas turbine engine systems involving hydrostatic face seals are provided. In this regard, representative compressor assembly for a gas turbine engine includes a compressor having a hydrostatic seal formed by a seal face and a seal runner.

Abstract: A multi function rural fuel platform has a biofuel production unit and a diesel engine incorporated onto a transportable platform, wherein said biofuel production unit produces fuel for said diesel engine.

Abstract: Disclosed is a propulsion system having a structural configuration that provides easy and convenient access to the interior regions of a liquid fuel tank and a hybrid rocket motor case. In one embodiment, the propulsion system comprises: a hybrid rocket motor case and a fuel tank coupled to the hybrid motor case. The motor case is configured to hold solid rocket fuel and the fuel tank defines an internal volume configured to hold a fluid oxidizer. A bulkhead is interposed between the motor case and the fuel tank, wherein at least one access passageway extends through the bulkhead. The access passageway provides exterior access to the interior volume of the motor case or the internal volume of the storage tank while the hybrid rocket motor is coupled to the fuel tank.

Abstract: In an electronic control system, three-dimensional data sets (data cubes L11-L14) for different power-limiting engine parameters are stored at different engine ratings for setting a respective maximum engine power in relation to flight altitude (ALT), ambient temperature (DTAMB) and flight Mach number (MN) to separately calculate a maximum power corresponding to each parameter. For power reduction due to air bleed, corresponding separate data cubes (L21-L24) are stored. In a comparator, the respective smallest power reduction value is determined and, subsequently, the fuel supply is set according to that power reduction value. Under conforming boundary conditions, the same data cubes (L24) can be stored for different engine ratings. The system requires low storage capacity and low calculating effort.

Abstract: A propulsion device contains a working mass within an enclosed plenum, where the working mass is recirculated through the plenum to produce a net force vector. The working mass is accelerated around a first bend or change in direction within the plenum by a working mass driver. The temperature of the working mass is then reduced and the working mass is slowed. The working mass is then accelerated around a second bend or change in direction within the plenum resulting in two opposing force vectors. When combined, the two opposing force vectors produce a net force, and thus, acceleration of the propulsion device. The propulsion device may be used in connection with a space vehicle or satellite.

Abstract: A thrust chamber assembly cooling system includes a twisted ribbon/wire of any one of many variety of cross-sectional shapes located within a nozzle assembly cooling passage to direct the coolant to flow in a swirling manner which induces mixing and breaks up the boundary layer in the coolant passage to enhance the convective heat transfer.

Abstract: Disclosed is a propulsion system having a structural configuration that provides easy and convenient access to the interior regions of a liquid fuel tank and a hybrid rocket motor case. The system operates with a high oxidizer-to-fuel ratio and a high bulk density propellant combination that has a near uniform specific impulse over a large oxidizer-to-fuel ratio range. The system has an increased propellant mass fraction and reduced propellant residuals. This improves the performance of the hybrid propulsion system.

Abstract: A variable area flow duct and method. In one embodiment the flow duct includes a plurality of helical vanes arranged around an interior surface wall of the flow duct. The vanes cause secondary flow vortices to be developed in the vicinity of each of the vanes that effectively reduce the interior cross-sectional area of the duct that a primary flow sees as it flows through the duct. In various embodiments fluidic injection is employed to suppress the formation of the secondary flow vortices during certain phases of operation, for example, during an afterburn phase of operation of a jet aircraft engine which the flow duct is being used with. In another embodiment, an ablative coating is used over the vanes to suppress the formation of the secondary flow vortices. The ablative material is removed by the hot fluid flow during an afterburn phase of operation, thus exposing the vanes and enabling the subsequent formation of secondary flow vortices to narrow the cross-sectional area of the throat.

Abstract: An insulation composition that comprises at least one nitrile butadiene rubber (“NBR”) having an acrylonitrile content that ranges from approximately 26% by weight to approximately 35% by weight and polybenzoxazole (“PBO”) fibers. The NBR may be a copolymer of acrylonitrile and butadiene and may be present in the insulation composition in a range of from approximately 45% by weight to approximately 56% by weight of a total weight of the insulation composition. The PBO fibers may be present in a range of from approximately 3% by weight to approximately 10% by weight of a total weight of the insulation composition. A rocket motor including the insulation composition and a method of insulating a rocket motor are also disclosed.

Abstract: An armor module and an explosive material therefore, which explosive material is normally retained at a less sensitive position, and upon demand it is modified into a more sensitive position, where its initiation ability is upgraded and wherein the modification is carried out by heating the explosive material.

Abstract: An exhaust nozzle bell for a rocket engine including a first part arranged on a motor of the rocket engine, and a second part coupled to the first part. The second part having a stowed position in which the second part surrounds the first part and is positioned closer to the motor, and an operating position in which the first part and the second part form a continuous shape and the second part is arranged farther from the motor. Moreover, the nozzle can include an extension mechanism structured and arranged to extend the second part from the stowed position to the operating position, the extension mechanism including a plurality of swiveling extension arms, wherein the extension arms have first and second ends. Furthermore, each of the extension arms comprise one of a sliding and rolling element on an end facing the second part of the exhaust nozzle bell.

Abstract: A supersonic combustion apparatus and method for using the same including a fixed geometric nozzle having a converging area, throat, and a diverging area, at least one fuel injection means and at least one flame stabilization means located in the divergent area, and an exit plane adjacent and downstream to the diverging area, where an initial first injection/flame stabilization means is located in the diverging area and the exit plane Mach is varied by heat addition in the diverging area by at least one more fuel injection means.

Abstract: A side thruster module, comprises: a cavity-type body skin extending in a longitudinal direction; a first thruster arranged in the body skin and extending in a longitudinal direction; and a conversion nozzle arranged in the body skin and extending in a radial direction perpendicular to the longitudinal direction, for converting a direction of a thrust generated from the first thruster in the longitudinal direction into the radial direction. A large number of thrusters can be mounted at the side thruster module thus to generate a high thrust, and the side thruster module can be slim in the radial direction perpendicular to the longitudinal direction.

Abstract: A propulsion device contains a working mass within an enclosed plenum, where the working mass is circulated through the plenum to produce a net force vector. The working mass is accelerated around a first bend or change in direction within the plenum by a working mass driver. The temperature of the working mass is then reduced and the working mass is slowed. The working mass is then accelerated around a second bend or change in direction within the plenum resulting in two opposing force vectors. When combined, the two opposing force vectors produce a net force, and thus, acceleration of the propulsion device.

Abstract: This invention describes a miniaturized hybrid diesel-electric engine formed by a closed-loop system powered by plasma-aided combustion of JP-8 fuel (or other hydrocarbon fuels) working in tandem with a vapor cycle utilizing miniaturized expanders and condensers. The output of this engine is electric power and mechanical work. Water, or organic fluids, heated by the combustion product developed inside a special burner, undergoes an explosive, quasi-supersonic conversion to steam. This steam drives a high-speed turbine connected together with a gas turbine outputting shaft work. This work output is utilized to power internal subsystems, cool down the miniaturized condensers, and to produce torque and electric power. The dimensions of this miniaturized hybrid-engine are so compact that it can fit inside the battery compartment of most applications requiring high-density miniaturized power sources.

Abstract: A device and method for guiding or steering projectiles (self-propelled or non-self-propelled), or missiles, and for steering a supersonic projectile, or a missile, having a nose, generally in the shape of a cone, having a more or less pointed end, and capable of creating a plasma discharge near the end of the projectile over a limited sector of the outer surface of nose.