Abstract: The invention provides methods and an apparatus for more efficiently and economically producing purified water from sea water or some other salty or brackish water source. The efficiency is derived from the co-location with a power plant or other thermal generating source that will heat the feed water. Reverse osmosis membrane filtration systems work optimally when the feed water is at certain higher temperature, where that temperature is typically higher than the feed water at ambient temperatures. By using the heated sea water as the byproduct of the power plant electricity generating process and if necessary mixing it with ambient temperature sea water, if needed to lower the water temperature, and using this feed water with a higher temperature than ambient water temperature, the efficiency of the reverse osmosis system can be increased.

Abstract: The invention relates to a propulsion device comprising a gas ejection nozzle and an injection chamber for injecting at least one propellant fluid. According to the invention, the device has an induction loop which surrounds a zone of the nozzle and also an electricity generator to feed said induction loop, in particular under drive from a heat engine whose heat source is a nuclear core and whose heat sink is a cryogenic liquid which is subsequently used for propulsion.

Abstract: A thruster for providing thrust for spacecraft positioning, which has a propellant reservoir for storing propellant, a reaction chamber for discharging a vapor for providing thrust, a pump module comprising one or more micropumps for drawing propellant from the reservoir and for systematically metering propellant to the reaction chamber in a controlled manner, and a controller for actuating the pump module.

Abstract: An apparatus for generating electricity that uses at least one jet-type engine fueled with fissile material. The nuclear fueled jet engine is affixed to a connecting member that projects from a central, rotatable shaft, which is in engageable communication with an apparatus for converting the rotation of the central shaft to electricity, such as a stator and rotor combination. The engine is positioned so that the thrust produced by the jet engine causes the engine and connecting member to travel in a radial direction around the longitudinal axis of the central shaft, rotating the central shaft. As the central shaft rotates, the rotational motion of the central shaft is transmitted to conversion apparatus. An operating gas is used to cool the nuclear fueled jet engines and as the propellant for the jet engines.

Abstract: A gas, e.g. hydrogen, at relatively low pressure is directly heated by the fission fragments (FF) emitted by a thin layer of fissile material, e.g. 242mAm, deposited on the inner wall of a chamber which is kept cooled at a typical temperature of about 1,000/1,500 K. The gas is preferably emitted as capillary flow from the walls of cylindrical tubes. Its temperature progressively increases until it reaches an equilibrium value of the order of 9,500 K, at which point FF heating and radiative cooling balance. With a relatively modest surface power density at the foil of 200 W/cm2, the specific, volume-averaged power given to the H gas may be as large as 0.66 MWatt/g. Heating powers up to megawatts for each gram of gas are therefore feasible with acceptable foil surface heating. The gas heating method can be used in rocket engines for deep space propulsion.

Abstract: An air/atmosphere breathing electrically powered Hall effect thruster including a thruster duct having an inlet, an exit, and a discharge zone between the inlet and the exit for receiving air from the inlet into the discharge zone, an electrical circuit having a cathode for emitting electrons and an anode in the discharge zone for attracting the electrons from the cathode through the exit, and a magnetic circuit for establishing a magnetic field in the discharge zone radially across the duct between the anode and exit which creates an impedance to the flow of electrons toward the anode and enables ionization of the air/atmosphere moving through the discharge zone and which creates an axial electric field in the duct for accelerating ionized air/atmosphere through the exit to create thrust.

Abstract: Propellant liquid is supplied to a Colloidal Thruster for Micro Satellite vehicles in Space by capillarity induced flow through a wick element comprising a permeable porous aggregate of fibers or particles of material that is wetted by the propellant liquid. An intense electric field at the tip of the wick element dispersed the arriving liquid into a fine spray of charged droplets. Electrodes having appropriate design, location and potentials accelerate the charge droplets to high velocity, thereby providing reactive thrust to the vehicle. In this method of propellant liquid introduction the flow rate and exhaust velocity, and therefore the thrust level, are determined by the applied potential difference, thereby eliminating the need for pumps or pressurized gas and flow controllers to provide the desired flowrate for the propellant liquid.

Abstract: A self-powered microthermionic converter having an internal thermal power source integrated into the microthermionic converter. These converters can have high energy-conversion efficiencies over a range of operating temperatures. Microengineering techniques are used to manufacture the converter. The utilization of an internal thermal power source increases potential for mobility and incorporation into small devices. High energy efficiency is obtained by utilization of micron-scale interelectrode gap spacing. Alpha-particle emitting radioisotopes can be used for the internal thermal power source, such as curium and polonium isotopes.

Abstract: Propellant modules for Micro Pulsed Plasma Thrusters, and techniques for bundling propellant modules and for using a two-stage discharge process to increase MicroPPT propellant throughput, and decrease the output voltage required from the power-processing unit are provided.

Abstract: A propulsion system is disclosed that uses microwaves originating from a ground-based or on-board source to provide energy for propulsion. The energy produces thrust by coupling into a propellant via a microwave absorbent heat exchanger, and by subsequent conversion of the propellant thermal energy to kinetic energy through a thrust converter or nozzle.

Abstract: A thruster device and associated methods are provided. The thruster device comprises a frustoconical inner layer and a frustoconical outer layer that surrounds the inner layer such that the inner and outer layers cooperate to define a spiral flow passage therebetween. An outlet, such as a nozzle, is attached to an exit end of the thruster device such that the outlet is in fluid communication with the spiral flow passage. In one embodiment, the thruster device includes an end flow cap that is also attached to the exit end and establishes the fluid communication between the spiral flow passage and the outlet. The end flow cap includes at least one channel for directing the propellant from the spiral flow passage to the outlet. The design of the thruster device allows the frustoconical inner and outer layers to be assembled without galling or damaging the spiral flow passage.

Abstract: An apparatus and method provided for utilizing radiation pressure provided by a light wave to generate mechanical work. The apparatus includes a containment chamber constructed to contain the propagation of light waves therein along a predetermined reflected light wave path. The apparatus also an optic switch selectively operable in an open mode and a closed mode, wherein the open mode allows a light wave to enter the containment chamber and the closed mode prevents escape of the light wave from the containment chamber. Further, the apparatus has a reflective mirror positioned at one end of the containment chamber. The reflective mirror has a first reflective surface.

Abstract: Propellant liquid is supplied to a Colloidal Thruster for Micro-Satellite vehicles in Space by capillarity induced flow through a wick element comprising a permeable porous aggregate of fibers or particles of material that is wetted by the propellant liquid. An intense electric field at the tip of the wick element dispersed the arriving liquid into a fine spray of charged droplets. Electrodes having appropriate design, location and potentials accelerate the charge droplets to high velocity, thereby providing reactive thrust to the vehicle. In this method of propellant liquid introduction the flow rate and exhaust velocity, and therefore the thrust level, are determined by the applied potential difference, thereby eliminating the need for pumps or pressurized gas and flow controllers to provide the desired flow-rate for the propellant liquid.

Abstract: An electric thruster and thrust augmenter is disclosed in which intaken or compressed atmospheric gas or reaction thruster exhaust is passed through a gap space between electrodes so that the atmospheric or reaction thrust exhaust gases are subjected to an electric current of sufficient intensity to rapidly heat and expand such gasses through an exhaust nozzle to produce reaction thrust.

Abstract: An apparatus and method for positron fueled ramjet operation is provided including a convection chamber for a ramjet comprising a rate of absorption of gamma rays that yields a uniform heating of the convection chamber to a temperature sufficient to sustain ramjet operation. The convection chamber may comprise at least one open ended cylinder, a plurality of concentrically arranged cylinders, a plurality of rods, a honeycomb structure or other structures for absorbing gamma rays. A ramjet powered by gamma rays is also provided that includes a housing forming a fluid passageway. A reaction vessel is positioned within the housing, and arranged in flow communication with a source of positrons. A gamma ray absorption assembly is positioned within the fluid passageway so that as positrons interact with electrons and annihilate, gamma rays impinge upon the gamma ray absorption assembly.

Abstract: An electric thruster has at least a portion of a surface of at least one of the component elements of its housing surface treated to increase the thermal transmission at the surface. The surface treatment is selected to increase the thermal absorption and thence the absorption of heat at the interiorly facing surfaces of the treated component, and/or to increase the thermal emissivity and thence radiation heat loss at the exteriorly facing surfaces of the treated component, and/or to increase the effective surface area through which heat is absorbed or emitted. The surface-treated housing components are assembled together with a cathode assembly, an ionization chamber, a propellant gas source, and a magnetic structure to form an electric thruster.

Abstract: The invention provides an improved spacecraft thruster, either Hall effect or ion effect, using gaseous propellant converted from solid iodine. A heated tank contains iodine crystals, which tank connects to a thrust chamber by a feed tube. A filter is mounted at the input end of the feed tube, proximate the tank, which filter is warmed by a heat control. A mass flow controller is mounted in the feed tube between the tank and the chamber and is heated by a temperature controller, such controller having a shut-off valve and means to control the flow rate of gaseous propellant to the thruster chamber.

Abstract: An electric thruster has at least a portion of a surface of at least one of the component elements of its housing surface treated to increase the thermal transmission at the surface. The surface treatment is selected to increase the thermal absorption and thence the absorption of heat at the interiorly facing surfaces of the treated component, and/or to increase the thermal emissivity and thence radiation heat loss at the exteriorly facing surfaces of the treated component, and/or to increase the effective surface area through which heat is absorbed or emitted. The surface-treated housing components are assembled together with a cathode assembly, an ionization chamber, a propellant gas source, and a magnetic structure to form an electric thruster.

Abstract: A combined generator device for use in a satellite capable of generating electrical power and thrust from an optical solar image. The combined generator includes a first receiver positioned to receive and store the outer, low-intensity region of the optical solar image. An electrical power generator thermally coupled to the first receiver uses the stored thermal energy to produce electricity. The combined generator also includes a thrust generator having a second receiver. The second receiver is positioned to receive and absorb the inner, high-intensity region of the optical solar image. A heat exchanger of the combined generator transfers the thermal energy from the second receiver into a supply of propellant which is vented to a nozzle to produce thrust.

Abstract: A combined generator device for use in a satellite capable of generating electrical power and thrust from an optical solar image. The combined generator includes a first receiver positioned to receive and store the outer, low-intensity region of the optical solar image. An electrical power generator thermally coupled to the first receiver uses the stored thermal energy to produce electricity. The combined generator also includes a thrust generator having a second receiver. The second receiver is positioned to receive and absorb the inner, high-intensity region of the optical solar image. A heat exchanger of the combined generator transfers the thermal energy from the second receiver into a supply of propellant which is vented to a nozzle to produce thrust.

Abstract: A modular solar thermal rocket that receives and absorbs solar energy and then acts as a heat exchanger to provide propulsive thrust. Thermal energy storage modules receive and store solar energy via thermal energy storage elements provided in each module. The solar energy is focused into a cavity defined by the modules by a solar energy secondary concentrator. A preheater is positioned adjacent the solar energy secondary concentrator and is in fluid communication with a propellant supply and one end of the thermal energy storage modules. A propulsion nozzle is in fluid communication with the opposite end of the thermal energy storage modules. Stored propellant is directed through the preheater and the thermal energy storage modules where it is heated to a high temperature. The propellant is then directed to the propulsion nozzle where it is exhausted into space to provide propulsive thrust.

Abstract: A spacecraft component (20) includes a plurality of parts, e.g., a first part (22) and a second part (23) connected to one another, at least the first part (22) being mounted for movement with respect to another of the parts, e.g., the second part (23). At least one microthruster (27) is provided on the first part for actuation thereof and includes a chamber (2, 14), a fluid provided in the chamber and a diaphragm (3, 13) covering an opening in the chamber (2, 14).

Abstract: A spacecraft thruster and propulsion method comprising a laser directed at an ablation target, wherein when the laser is operating, material ablates from the ablation target, thereby generating a thrust vector on the spacecraft.

Abstract: For a plasma accelerator arrangement in particular for use as an ion thruster in a spacecraft, a structure is proposed in connection with which an accelerated electron beam is admitted into an ionization chamber with fuel gas, and is guided through the ionization chamber in the form of a focused beam against an electric deceleration field, said electric deceleration field acting at the same time as an acceleration field for the fuel ions produced by ionization. The arrangement generates a focused beam of a largely neutral plasma with a high degree of efficiency. Configurations for electric and magnetic fields for guiding and focusing the beams are given by way of example.

Abstract: Propellant modules for Micro Pulsed Plasma Thrusters, and techniques for bundling propellant modules and for using a two-stage discharge process to increase MicroPPT propellant throughput, and decrease the output voltage required from the power-processing unit are provided.

Abstract: A electric thruster and thrust augmenter is disclosed in which intaken or compressed atmospheric gas or reaction thruster exhaust is passed through a gap space between electrodes so that the atmospheric or reaction thrust exhaust gases are subjected to an electric current of sufficient intensity to rapidly heat and expand such gasses through an exhaust nozzle to produce reaction thrust.

Abstract: The propulsion system of the present invention includes charged particles held in a containment structure at high temperatures. Propellant is injected into the containment system where it collides elastically with the charged particles. The propellant is heated until it achieves a high velocity. It can then be ejected to produce thrust. The propulsion system can include thrusters to propel, for example, space launch vehicles and jet aircraft.

Abstract: The invention relates to a propulsion device comprising a gas ejection nozzle and an injection chamber for injecting at least one propellant fluid. According to the invention, the device has an induction loop which surrounds a zone of the nozzle and also an electricity generator to feed said induction loop, in particular under drive from a heat engine whose heat source is a nuclear core and whose heat sink is a cryogenic liquid which is subsequently used for propulsion.

Abstract: A modular solar thermal rocket that receives and absorbs solar energy and then acts as a heat exchanger to provide propulsive thrust. Thermal energy storage modules receive and store solar energy via thermal energy storage elements provided in each module. The solar energy is focused into a cavity defined by the modules by a solar energy secondary concentrator. A preheater is positioned adjacent the solar energy secondary concentrator and is in fluid communication with a propellant supply and one end of the thermal energy storage modules. A propulsion nozzle is in fluid communication with the opposite end of the thermal energy storage modules. Stored propellant is directed through the preheater and the thermal energy storage modules where it is heated to a high temperature. The propellant is then directed to the propulsion nozzle where it is exhausted into space to provide propulsive thrust.

Abstract: A thruster device and associated methods are provided. The thruster device comprises a frustoconical inner layer and a frustoconical outer layer that surrounds the inner layer such that the inner and outer layers cooperate to define a spiral flow passage therebetween. An outlet, such as a nozzle, is attached to an exit end of the thruster device such that the outlet is in fluid communication with the spiral flow passage. In one embodiment, the thruster device includes an end flow cap that is also attached to the exit end and establishes the fluid communication between the spiral flow passage and the outlet. The end flow cap includes at least one channel for directing the propellant from the spiral flow passage to the outlet. The design of the thruster device allows the frustoconical inner and outer layers to be assembled without galling or damaging the spiral flow passage.

Abstract: A modular solar thermal rocket that receives and absorbs solar energy and then acts as a heat exchanger to provide propulsive thrust. Thermal energy storage modules receive and store solar energy via thermal energy storage elements provided in each module. The solar energy is focused into a cavity defined by the modules by a solar energy secondary concentrator. A preheater is positioned adjacent the solar energy secondary concentrator and is in fluid communication with a propellant supply and one end of the thermal energy storage modules. A propulsion nozzle is in fluid communication with the opposite end of the thermal energy storage modules. Stored propellant is directed through the preheater and the thermal energy storage modules where it is heated to a high temperature. The propellant is then directed to the propulsion nozzle where it is exhausted into space to provide propulsive thrust.

Abstract: The optical system includes an optical source capable of producing light having physical characteristics sufficient for optically driven chemical disassociation of a fuel; and, an optical delivery system for providing optical delivery of light from the optical source to a chemical fuel source. The optical system is used for a propulsion producing engine having a chemical fuel source. The endothermic fuel generation is produced by the non-linear, optical interaction of the produced light with the fuel, thereby leading to molecular dissociation of the fuel.

Abstract: A method and system are provided for propelling an aerodynamic vehicle into space. The aerodynamic vehicle uses a nuclear-based thermal rocket (NTR) propulsion system capable of producing a hydrogen exhaust. A flow of air is introduced into the hydrogen exhaust to augment the thrust force at speeds of the vehicle up to approximately Mach 6. When the speed of the vehicle is approximately Mach 6 and the altitude of the vehicle is approximately 40 kilometers, the flow of air is stopped and the vehicle is propelled into space using only the NTR.

Abstract: A micro-colloid thruster system may be fabricated using micro electromechanical system (MEMS) fabrication techniques. A beam of charged droplets may be extracted from an emitter tip in an emitter array by an extractor electrode and accelerated by an accelerator electrode to produce thrust. The micro-colloid thruster system may be used as the main propulsion system for microspacecraft and for precision maneuvers in larger spacecraft.

Abstract: A thermal solar rocket that includes a solar energy receiver having two sections (a thermal energy storage section and a direct gain section), a solar concentrator, and a propulsion nozzle. In one embodiment, the focus of the solar energy between the storage section and the direct gain section is controlled by mechanical means such as movable insulation. In another embodiment, the focus of the solar energy between the storage section and the direct gain section is controlled by an optical switch in the form of relative motion between the solar concentrator and the solar energy receiver.

Abstract: Both chemical and nuclear rocket engines waste huge amounts of energy. This waste can be decreased by application of light-conducting structures.

Abstract: An engine is disclosed, including a controllable output plasma generator, a controllable heater for selectably raising a temperature of the plasma connected to an outlet of the plasma generator, and a nozzle connected to an outlet of the heater, through which heated plasma is discharged to provide thrust. In one embodiment, the source of plasma is a helicon generator. In one embodiment, the heater is an ion cyclotron resonator. In one embodiment, the nozzle is a radially diverging magnetic field disposed on a discharge side of the heater so that helically travelling particles in the heater exit the heater at high axial velocity. A particular embodiment includes control circuits for selectably directing a portion of radio frequency power from an RF generator to the helicon generator and to the cyclotron resonator so that the thrust output and the specific impulse of the engine can be selectively controlled. A method of propelling a vehicle is also disclosed.

Abstract: A solar thermal engine for propelling and powering a craft. The solar thermal engine includes a housing, a propellant annulus, a plurality of static power converters and an electrical energy storage device. The housing has an optical cavity for receiving a beam of concentrated sunlight and converting the beam into ambient thermal energy. The propellant annulus is coupled to the housing and is selectively operable in a heating mode wherein the propellant annulus transmits at least a first portion of the ambient thermal energy to heat a flow of propellant. The plurality of static power converters are coupled to the housing and receive the first portion of the ambient thermal energy when the propellant annulus is not operated in the heating mode. The plurality of static power converters employ the first portion of the ambient thermal energy to generate electrical energy.

Abstract: A thruster system includes a power supply and a pulse forming circuit coupled to the power supply. The pulse forming circuit includes a capacitor and first and second diodes. The positively-charged plate of the capacitor is coupled to the anode of the first diode, the negatively-charged of the capacitor is coupled to the anode of the second diode, and the cathode of the first diode is coupled to the cathode of the second diode. A low-impedance thruster is coupled in parallel to the second diode.

Abstract: A radio frequency (RF) plasma thruster for use in electric propulsion for spacecraft. The thruster operates by heating plasma in a magnetic field, which then flows out along magnetic field lines, producing axial thrust. The present invention greatly increases the efficiency of the RF plasma thruster compared to previous thrusters of this type, while retaining the advantages of RF plasma thrusters over other types of electric and chemical propulsion systems. The present invention utilizes a lower hybrid wave for heating of the electrons, rather than electron cyclotron resonance (ECR) heating. The lower hybrid wave is used because it creates high-density plasmas and the antennas used to couple RF energy to the plasma are relatively simple to construct. This allows much better efficiency because no hot electron population is created to siphon off much of the RF power applied to the plasma.

Abstract: A solar thermal rocket that includes a thermal energy storage section, a radiant inter-heater, a primary solar concentrator, and a propulsion nozzle. The primary solar concentrator is selectively movable to direct solar energy to either the thermal energy storage section or to the radiant inter-heater. The thermal energy storage section, along with insulation, is arranged to define a cavity such that a focused beam of solar rays can enter the cavity through an aperture in the insulation. The thermal energy storage section typically absorbs and stores solar energy during the non-propulsion portion of the orbital period. The solar rays are captured and absorbed and thereby heat the thermal energy storage section to very high temperatures. A radiant inter-heater directly receives concentrated solar rays and transfers the heat to the propellant during the propulsion phase.

Abstract: A single energy storage capacitor is used to apply electrical energy to selected propellant modules, each module generating thrust in a desired direction by the high voltage induced vaporization of the copolymer spacer material in a conventional coaxial cable within the selected module. A lightweight, low cost, and EMI free thruster is thus provided.

Abstract: The invention provides a microthruster which includes a housing having a propellant container and a discharge section, means to heat the propellant to drive vapors toward the discharge section, a heating element to heat the vapors and one or more ports to discharge the vapors to provide thrust. Further provided is a method for powering a microthruster by flowing gas in free molecular flow therein so that a plurality of molecules of the gas contact a heating element before discharge from the microthruster.

Abstract: A pulse plasma thruster (50) utilizes a vapor producing solid (54) and a micro-sized heater (52) to produce a high pressure vapor that is directed into an ignition chamber (58) and to a thrust discharge chamber (70). The thrust discharge chamber (70) comprises two oppositely disposed electrode plates (72, 74) and oppositely disposed fuel propellants sources (60, 62). The passageway (56) leading from vapor producing solid (54) to the thrust discharge chamber (70) is configured to permit uniform feeding of the vapors to the thrust discharge chamber (70). A pair of electrode terminals (82, 84) extend from the electrode plates (72, 74) and through a housing (88). A power source (100) is coupled to the terminals (82, 84) and provides the ignition signals necessary to cause a spark and a breakdown to a useful plasma arc by controlling the voltage-current shape of the ignition signal.

Abstract: Microwave-Electro-Thermal thrusters produce a high temperature rocket exhaust by sending microwaves into a resonant cavity where an excited mode then creates an electrodeless discharge that heats gaseous fuel. Heretofore, the microwave power coupling between the microwave generator and the resonant cavity and plasma has consisted of rigid waveguide with impedance matching equipment. This waveguide and impedance matching hardware greatly adds to the weight and size of the system making it impractical for spaceflight. The foregoing problems are overcome by greatly reducing or eliminating the waveguide and impedance matching equipment.

Abstract: In drive circuitry for powering an arc jet or ion engine aboard a spacecraft carrying solar panels, a start-up circuit interconnects an array of series-parallel connected solar panels to establish and maintain an arc and plasma in the engine. In the start-up circuit, an electric switch is placed in parallel with an electrode assembly of the engine. One terminal of the engine connects directly with the solar panel array, and another terminal of the engine connects via an inductor to the solar panel array. A capacitor connects across terminals of the solar panel array. A series connection of solar panels increases the voltage from that which is available from a single panel. Conduction of current from the solar panel via the inductor is initiated by the switch to store magnetic energy within the conductor. Thereupon, upon an opening of the switch, the inductor produces a relatively high voltage which strikes an arc in the engine.

Abstract: A spacecraft attitude and altitude control system utilizes sets of three pulsed plasma thrusters connected to a single controller. The single controller controls the operation of each thruster in the set. The control of a set of three thrusters in the set makes it possible to provide a component of thrust along any one of three desired axes. This configuration reduces the total weight of a spacecraft since only one controller and its associated electronics is required for each set of thrusters rather than a controller for each thruster. The thrusters are positioned about the spacecraft such that the effect of the thrusters is balanced.