Abstract: A low orbit optical imaging satellite has a long thin satellite body housing an optical telescope arrangement. A major part of the telescope arrangement has its optical axis roughly parallel to the direction of elongation and includes a mirror arrangement deployed to direct a line of sight of the optical telescope out sideways from the direction of elongation. The transverse dimensions of the satellite body are preferably minimized to be close to the optical aperture dimension of the optical telescope, thereby providing a high ballistic coefficient and high orbit lifetime for orbits in the low thermosphere.

Abstract: a solar powered spacecraft power system including a solar photovoltaic array, an electric propulsion system connected directly to the solar photovoltaic array in parallel with the spacecraft power system; the electric propulsion system including a Hall effect thruster, a thruster power supply for driving the thruster; a sensor for sensing the power output of the solar array and a controller responsive to the power output of the solar array and configured to periodically adjust an operating parameter of the thruster to operate the thruster at the maximum available output power of the solar array including comparing a previous solar array output power level with a later solar array output power level, and incrementing the operating parameter with a positive value if the later is greater and with a negative value if the later is lesser; a solar powered spacecraft power system including a solar photovoltaic array, an electric propulsion system connected directly to the solar photovoltaic array, and a power manage

Abstract: The invention relates to a drive arrangement in a spacecraft, comprising several drive units (TW1,TW2,TW3), several individually controllable drive units that can be continuously applied to a common, constant voltage potential (HV), and a control of the axial thrust in the respective drive units is achieved due to the fact that the production of plasma in the respective drive units is individually controlled. In particular, the time-variable control of the production of plasma occurs by the time-variable control of the flow of neutral working gas (AG) in the ionisation chamber (IK).

Abstract: A device for reducing the aerodynamic drag of a vehicle includes at least one masking element for at least one part of a nozzle of the vehicle's engine, the masking element having a resorbable material designed to be eliminated in the nozzle's flow once the engine is ignited. A space craft includes the nozzle attached to the fuselage of the space craft.

Abstract: An afterbody device for a spacecraft fitted with at least one rocket engine at the rear of the craft includes at least one movable cover element designed to take a first position, masking and reducing the vehicle's rear drag, where it prolongs the vehicle's fuselage around at least one part of a rocket engine nozzle of the vehicle and extends beyond the rear of the vehicle's fuselage, and to take a second position fully deployed, increasing the vehicle's aerodynamic drag.

Abstract: A reliable and inexpensive attitude control system uses a plurality of pitch-over thrusters to perform rapid and precise attitude maneuvers for a flight vehicle. The pitch-over thrusters create rotational moments that directly pitch and yaw the flight vehicle. The use of very simple thrusters and control techniques provides for a reliable and cost effective solution. The ability to perform overlapping pitch and yaw maneuvers with single-shot fixed-impulse thrusters provides for high-speed maneuverability.

Abstract: The structure provided multi-Levels of the center fuselage, two front wings, Extended mid-two wings and two rear wings with two vertical winglets. The fuselage has a cross section of half circle shaped fuselage. The first level comprises multi-fuel tanks, water tanks, Helium gas tanks, hydraulic, Pneumatic systems and plurality of the multi-landing gear bays and multi-cargo Compartments. The first level comprises a top cockpit and the plurality of the multi-passenger Cabins. The second level is comprises multi-central posts, multi-guy wires on the top center Beams, mid level, lower level center beams, leak proof sealed floor gates for the Multi-gas envelopes structure, under belly multi-pontoon air bags. The first level further comprises multi-jet power plants with multi-chute flaps. The second level may comprises, carbon fiber honey comb sandwich composite Multi-envelopes and multi-level soft gas envelopes.

Abstract: A fixed sized bell rocket nozzle is lined with a layer of combustible material that is ignited during launch ignition and burns to outgas into the rocket exhaust for spatially variably confining the exhaust and perfecting an effective variably sized altitude compensating exhaust nozzle that maximizes lift during the launch of a spacecraft into orbit.

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 invention relates to a self-contained device (4) for blanking out the light rays from at least one star (1), which device (4) comprises means (43) for controlling propulsion means (44) which are in turn capable of moving or halting the device (4) in space and/or on a pseudo-orbit in space around an observation telescope (2) comprising an observation opening (21), characterized in that it comprises a blanking-out screen (40), wherein the means (43) for controlling the propulsion means (44) are further capable of positioning the screen (40) on the sighting axis (3) between the telescope (2) and the star (1) during an observation period such that the light rays from the star (1) are at least partially blanked out during said observation period as far as said observation opening (21) is concerned. The invention further relates to an assembly comprising at least one such device and a method for using said device or said assembly.

Abstract: The invention relates to a craft comprising an energy storage system of the RFCS (regenerative fuel cell system) type and a system for pressurizing the propellants used for propulsion. The craft also comprises a plurality of means (31 to 38) for coupling the tanks of the fuel cell to the propellant-pressurizing tanks of the propulsion system of the craft when the tanks of the propellant-pressurizing system are no longer used. The invention also relates to the method of coupling the tanks of the propellant-pressurizing means to the energy storage system. The invention applies particularly to telecommunication satellites and more generally to spacecraft.

Abstract: A high-efficiency spacecraft propulsion system, including includes electric pumps inserted in the oxidizer and fuel lines that increase liquid apogee engine (“LAE”) operating pressure and reduce tank-operating pressure. An on-board computer generates pump drive signals, in response to measured oxidizer and fuel line pressures, that are input to the pump controller electronics. The controller electronics provides current drives to the pump motors. The system uses an LAE that can operate at higher thrust chamber pressures (e.g., 500 psia) than standard LAEs, where pump-fed rocket motors have integrated turbopumps that are fuel operated. The turbopump increases LAE complexity and cost, reduces fuel efficiency, and is not compatible with active thrust and mixture ratio control.

Abstract: An extraterrestrial vehicle propulsion system uses a frusto-conical body member that is spun at high speeds by a drive member. A battery, which is rechargeable by a solar cell array, provides power to the drive member.

Abstract: A method of controlling attitude of a spacecraft during a transfer orbit operation is provided. The method includes providing a slow spin rate, determining the attitude of the spacecraft using a unified sensor set, and controlling the attitude of the spacecraft using a unified control law. The use of a unified set of sensors and a unified control law reduces spacecraft complexity, cost, and weight.

Abstract: Disclosed is a space propulsion device capable of continuous operation using a solid propellant having high density and excellent handleability.

Abstract: A preferred In Orbit Transportation & Recovery System (IOSTAR™) (10) includes a space tug powered by a nuclear reactor (19). The IOSTAR™ includes a collapsible boom (11) connected at one end to a propellant tank (13) which stores fuel for an electric propulsion system (12). This end of the boom (11) is equipped with docking hardware (14) that is able to grasp and hold a satellite (15) and as a means to refill the tank (13). Radiator panels (16) mounted on the boom (11) dissipate heat from the reactor (19). A radiation shield (20) is situated next to the reactor (19) to protect the satellite payload (15) at the far end of the boom (11). The IOSTAR™ (10) will be capable of accomplishing rendezvous and docking maneuvers which will enable it to move spacecraft between a low Earth parking orbit and positions in higher orbits or to other locations in our Solar System.

Abstract: A preferred In Orbit Transportation & Recovery System (IOSTAR™) (10) includes a space tug powered by a nuclear reactor (19). The IOSTAR™ includes a collapsible boom (11) connected at one end to a propellant tank (13) which stores fuel for an electric propulsion system (12). This end of the boom (11) is equipped with docking hardware (14) that is able to grasp and hold a satellite (15) and as a means to refill the tank (13). Radiator panels (16) mounted on the boom (11) dissipate heat from the reactor (19). A radiation shield (20) is situated next to the reactor (19) to protect the satellite payload (15) at the far end of the boom (11). The IOSTAR™ (10) will be capable of accomplishing rendezvous and docking maneuvers which will enable it to move spacecraft between a low Earth parking orbit and positions in higher orbits or to other locations in our Solar System.

Abstract: A preferred In Orbit Transportation & Recovery System (IOSTAR™) (10) includes a space tug powered by a nuclear reactor (19). The IOSTAR™ includes a collapsible boom (11) connected at one end to a propellant tank (13) which stores fuel for an electric propulsion system (12). This end of the boom (11) is equipped with docking hardware (14) that is able to grasp and hold a satellite (15) and as a means to refill the tank (13). Radiator panels (16) mounted on the boom (11) dissipate heat from the reactor (19). A radiation shield (20) is situated next to the reactor (19) to protect the satellite payload (15) at the far end of the boom (11). The IOSTAR™ (10) will be capable of accomplishing rendezvous and docking maneuvers which will enable it to move spacecraft between a low Earth parking orbit and positions in higher orbits or to other locations in our Solar System.

Abstract: A low orbit optical imaging satellite has a long thin satellite body housing an optical telescope arrangement. A major part of the telescope arrangement has its optical axis roughly parallel to the direction of elongation and includes a mirror arrangement deployed to direct a line of sight of the optical telescope out sideways from the direction of elongation. The transverse dimensions of the satellite body are preferably minimized to be close to the optical aperture dimension of the optical telescope, thereby providing a high ballistic coefficient and high orbit lifetime for orbits in the low thermosphere.

Abstract: A method is disclosed for placing a satellite in an operational orbit The satellite is equipped with its own satellite propulsion system as well as a detachable separate propulsion device The satellite and separate propulsion device are launched into a transfer orbit by means of a space launcher The separate propulsion device is controlled by a satellite. The satellite is transferred from the transfer orbit to an intermediate orbit by means of the separate propulsion device. The separate propulsion device is separated from the satellite in the intermediate orbit.

Abstract: A preferred In Orbit Transportation & Recovery System (IOSTAR™) (10) includes a space tug powered by a nuclear reactor (19). The IOSTAR™ includes a collapsible boom (11) connected at one end to a propellant tank (13) which stores fuel for an electric propulsion system (12). This end of the boom (11) is equipped with docking hardware (14) that is able to grasp and hold a satellite (15) and as a means to refill the tank (13). Radiator panels (16) mounted on the boom (11) dissipate heat from the reactor (19). A radiation shield (20) is situated next to the reactor (19) to protect the satellite payload (15) at the far end of the boom (11). The IOSTAR™ (10) will be capable of accomplishing rendezvous and docking maneuvers which will enable it to move spacecraft between a low Earth parking orbit and positions in higher orbits or to other locations in our Solar System.

Abstract: A rocket includes a front warhead, a rear propellant actuator, and a propulsion enhancement arrangement which comprises at least two propulsion impulse generators longitudinally cascadedly mounted between the warhead and the rear propellant actuator, wherein each propulsion impulse generator comprises a storage barrel for storing a predetermined amount of explosive materials to provide propelling impulse upon controlled explosion thereof, and a time-internal control arrangement comprising a time-interval control device mounted on a bottom of the front warhead, a time-interval trigger wire operatively connected with the time-interval control device, and at least two time-interval triggers mounted along the time-interval trigger wire and inside of each of the storage chambers respectively, wherein the time-interval triggers trigger the explosion of the explosive material stored in the storage chamber from the bottom propelling impulse generator to the top propelling impulse generator one by one at a predetermin

Abstract: A digital programmable driver for a solenoid valve comprises at least one valve driver circuit for opening and closing the valve and a controller capable of sending an output signal to the valve driver circuit. The valve driver circuit comprises a power switch for providing current to the valve in order to open the valve. The output signal sent by the controller directs the power switch of the valve driver circuit to provide current to the valve, thereby causing the valve to open. The controller receives a plurality of parameter values, and responsive to the plurality of parameter values, calculates a rise time, a pull-in duty cycle, a fall time, and a hold duty cycle so as to modify the output signal sent to the valve driver circuit. The plurality of parameter values comprises valve resistance, valve inductance, supply voltage, valve pull-in current, valve hold current, and valve pull-in time.

Abstract: The invention provides a novel, low-cost, spin-stabilized lander architecture capable (with appropriate system scaling tailored to the attributes of the target) of performing a soft-landing on a solar-system body such as Earth's Moon, Mars, Venus, the moons of Mars, Jupiter, Saturn, Uranus and Neptune, selected near-Earth and main-belt asteroids, comets and Kuiper belt objects and even large human-made objects, and also moving about on the surface of the target solar-system body after the initial landing in movement akin to hopping.

Abstract: In an aerospace vehicle in which a satellite is detachably adapted to the first step rocket and the second step rocket, an electronic device 65 is installed in the satellite 60, wherein the electronic device 65 controls the first step rocket 10 and the second step rocket 20 before the satellite is detached from the first step rocket 10 and the second step rocket 20. Thereby, it is unnecessary to provide sensors, radio devices and electronic devices for exclusive use with respect to the first step rocket 10 and the second step rocket so that a manufacturing cost of the rockets 10 and 20 can be reduced and a total weight of the rockets 10 and 20 can become lighter by omitting these equipments. Thus, an aerospace vehicle system within the rockets 10 and 20 can be simplified and a launch of the aerospace vehicle can be prepared within a short period.

Abstract: A preferred In Orbit Transportation & Recovery System (IOSTAR™) (10) includes a space tug powered by a nuclear reactor (19). The IOSTAR™ includes a collapsible boom (11) connected at one end to a propellant tank (13) which stores fuel for an electric propulsion system (12). This end of the boom (11) is equipped with docking hardware (14) that is able to grasp and hold a satellite (15) and as a means to refill the tank (13). Radiator panels (16) mounted on the boom (11) dissipate heat from the reactor (19). A radiation shield (20) is situated next to the reactor (19) to protect the satellite payload (15) at the far end of the boom (11). The IOSTAR™ (10) will be capable of accomplishing rendezvous and docking maneuvers which will enable it to move spacecraft between a low Earth parking orbit and positions in higher orbits or to other locations in our Solar System.

Abstract: A two-stage space vehicle is provided for achieving low earth orbit that includes a carrier and an orbiter. The carrier has a large airfoil area relative to its dry weight, and the thrust of the rocket engine is controlled such that the carrier achieves a launch altitude and speed for the orbiter without exceeding a wingloading pressure beyond 3,500 Pa., which allows the carrier to be inexpensively constructed. Liquid propellant for the rocket engine is advantageously stored within the relatively large airfoil of the carrier, which is preferably a delta wing. The ratio of airfoil area to dry weight is about 33 m2 per ton, which allows the carrier to descend and land after launching the orbiter with low wingloading on the order of 300 Pa.

Abstract: A vehicle and method for enabling propulsive flight from suborbital altitudes and velocities directly to far space, or beyond Low Earth Orbit (LEO), preferably without requiring injection into or refueling in LEO. The vehicle is preferably reusable and can withstand re-entry speeds and temperatures higher than those that are typical for LEO vehicles. The vehicle preferably lands horizontally, for example on a runway. The vehicle forms the upper stage of a booster vehicle system which can be launched either from the ground or from a subsonic air platform. The vehicle may optionally be used for LEO missions.

Abstract: The hybrid rocket system of this invention is characterized by use of an oxidizer tank having a cylindrical midsection surrounded by a skirt and bonded thereto by a layer of elastomeric adhesive. The skirt outer surface is in turn adhesively secured to a spacecraft inner surface. An elongated solid-fuel motor case is mechanically rigidly secured to a central rear surface of the tank, and the case terminates in a throat and nozzle. The elastomeric-adhesive bonding of tank to skirt, and rigid adhesion of skirt to spacecraft forms the sole support for the rocket system, and separate support for the motor case is not required.

Abstract: An experiment system with six different re-entry experiment locations for testing high temperature re-entry materials, creating new thermal protection systems, proving innovative new concepts for spacecraft exterior surfaces and the incremental development of next generation aerospace materials. A commercial transportation system to and from orbit provides a 24-hour return cycle for the experiments on a surface actually re-entering the earth's atmosphere. Now using existing doors, hatches and other points on the reusable launch vehicle's exterior, the actual re-entry environment is experienced by test specimens with quick turn around for a wide variety of different re-entry temperatures ranges for broad testing and development purposes. The reusable launch vehicle launches, remains in orbit for 24 hours and returns to provide an actual test environment for the exterior experiment system.

Abstract: Pulsed detonation engines (PDEs) are adapted for use in reaction control systems (RCS), such as thrusters for orbital correction and control (e.g., for earth-orbiting satellites), divert thrust generation and control for space-based interceptor devices, and for missile trajectory correction and motion control. According to one aspect of the invention, PDEs are adapted for motion control of so-called “kill vehicles,” which are small devices, typically launched from satellites, for strategic missile defense.

Abstract: A combination for performing a variety of functions includes (a) apparatus for moving a projectile or other mass along an arcuate path and moving the path substantially radially along a local radius of curvature and (b) a tool, vehicle or other article for receiving such projectile and being moved thereby.

Abstract: In one aspect, a space vehicle includes a structure configured to expand from a first configuration to a second configuration and at least two equipment compartments attached to the periphery of the structure. The structure includes at least one of an antenna mesh, a light-shielding mesh, an optical reflector mesh and a net. In another aspect, a space vehicle includes an antenna structure configured to expand from a first configuration to a second configuration and at least two equipment compartments attached to the periphery of the antenna structure. At least one of the at least two equipment components include a solar panel, a propulsion system and an antenna feed. In a further aspect, a space vehicle includes a payload element and at least two spacecraft-support structures attached to the periphery of the payload element. At least one of the at least two spacecraft-support structures includes a propulsion system and a tracking, telemetry and control system.

Abstract: A preferred In Orbit Transportation & Recovery System (IOSTAR™)(10) includes a space tug powered by a nuclear reactor (19). The IOSTAR™ includes a collapsible boom (11) connected at one end to a propellant tank (13) which stores fuel for an electric propulsion system (12). This end of the boom (11) is equipped with docking hardware (14) that is able to grasp and hold a satellite (15) and as a means to refill the tank (13). Radiator panels (16) mounted on the boom (11) dissipate heat from the reactor (19). A radiation shield (20) is situated next to the reactor (19) to protect the satellite payload (15) at the far end of the boom (11). The IOSTAR™ (10) will be capable of accomplishing rendezvous and docking maneuvers which will enable it to move spacecraft between a low Earth parking orbit and positions in higher orbits or to other locations in our Solar System.

Abstract: An orbital stage system has an orbital stage and one or more launch stages. The orbital stage incorporates an orbital maneuvering system (OMS) and an abort propulsion system which both utilize the same propellants, propellant tankage, and propellant pressurization system, but which employ radically different engines. The OMS engines are comprised of at least two engines which have a combined thrust in the neighborhood of 1/10 the weight of the orbital stage, an area ratio of 50 or more and an operating life of many hundred seconds, preferably many thousands of seconds or more. The abort engine may be a single engine and typically has a thrust of three, four, or more times the weight of the vehicle and an area ratio in the neighborhood of two and an operating life of at most a few tens of seconds.

Abstract: A preferred In Orbit Transportation & Recovery System (IOSTAR™)(10) includes a space tug powered by a nuclear reactor (19). The IOSTAR™ includes a collapsible boom (11) connected at one end to a propellant tank (13) which stores fuel for an electric propulsion system (12). This end of the boom (11) is equipped with docking hardware (14) that is able to grasp and hold a satellite (15) and as a means to refill the tank (13). Radiator panels (16) mounted on the boom (11) dissipate heat from the reactor (19). A radiation shield (20) is situated next to the reactor (19) to protect the satellite payload (15) at the far end of the boom (11). The IOSTAR™ (10) will be capable of accomplishing rendezvous and docking maneuvers which will enable it to move fully deployed spacecraft between a low Earth parking orbit and positions in higher orbits or to other locations in our Solar System.

Abstract: An In Orbit Transportation & Recovery System (IOSTAR™) (10) is disclosed. One preferred embodiment of the present invention comprises a space tug powered by a nuclear reactor (19). The IOSTAR™ includes a collapsible boom(11) connected at one end to a propellant tank (13) which stores fuel for an electric propulsion system (12). This end of the boom (11) is equipped with docking hardware (14) that is able to grasp and hold a satellite (15) and as a means to refill the tank (13). Radiator panels (16) mounted on the boom (11) dissipate heat from the reactor (19). A radiation shield (20) is situated next to the reactor (19) to protect the satellite payload (15) at the far end of the boom (11). The IOSTAR™ (10) will be capable of accomplishing rendezvous and docking maneuvers which will enable it to move spacecraft between a low Earth parking orbit and positions in higher orbits or to other locations in our Solar System.

Abstract: An annular fairing having aerodynamic, thermal, structural and acoustic attributes couples a launch abort motor to a space vehicle having a payload of concern mounted on top of a rocket propulsion system. A first end of the annular fairing is fixedly attached to the launch abort motor while a second end of the annular fairing is attached in a releasable fashion to an aft region of the payload. The annular fairing increases in diameter between its first and second ends.

Abstract: A spacecraft system that includes a primary space vehicle and a secondary space vehicle, both of which are designed to optimize payload capacity and launch weight of the primary space vehicle. The primary and secondary space vehicles combine to form an on-orbit space vehicle capable of performing functions and maneuvers that exceed the physical capabilities of the primary space vehicle at the time of its launch. The spacecraft system is designed to minimize propellant containment-related disturbances while maintaining a standard level of functionality. The primary space vehicle is designed to be incapable of independently performing a propellant-intensive orbit change maneuver. Instead the primary space vehicle is designed to couple to a secondary space vehicle having propellant and thrust capability sufficient to perform an orbit change maneuver when the primary and secondary space vehicles are coupled.

Abstract: An environmentally hardened architecture comprises a hybrid processor, a high speed bus having environmentally-sensitive interfaces, an environmentally hardened bus having environmentally-hardened interfaces, and an environmentally-hardened processor communicatively coupled to an environmentally-sensitive interface of the high speed bus and communicatively coupled to an environmentally-hardened interface of the environmentally hardened bus. The hybrid processor includes an environmentally-hardened processing section and an environmentally-sensitive processing section. At least one environmentally-sensitive interface is configured to pass data to and from the environmentally-sensitive processing section and another environmentally-sensitive interface is configured to pass data to and from the environmentally-hardened processing section of the hybrid processor. An environmentally-hardened interface is configured to pass data to and from the environmentally-hardened processing section of the hybrid processor.

Abstract: A reversible aerospace plane includes an air intake at a first end of the aerospace plane, at least one heat exchanger disposed in the aerospace plane, an engine at a second end of the aerospace plane, wherein the aerospace plane is configured to accelerate in a first direction and configured to glide and land in a second direction, wherein the second direction is substantially in a reverse direction from the first direction.

Abstract: Fuel less ACS for solar sails, using furl- and unfurl able ballast-sail-foil-segments for simultaneously displacement of center of mass and center of solar radiation pressure into opposite directions to each other. Solar-Sail-Launch-System for direct launch of the System-Sail including already docked in daughter units and payload in the sailcraft's central docking station. Unlike todays launch able solar sail designs the System-Sail features ample solar cell arrays and additional SEP-thruster-units for steering and propulsion, while the solar sail rather serves for longtime fuel less attitude controls and station keeping. The SEP-Sailcraft may also serve as a carrier-ship for daughter-units in asteroid exploration missions and is able to deliver prospector landers back to LEO with furled in foils.

Abstract: The invention is a system and method for propelling and slowing down spacecraft and other space systems and objects using the thrust generated from the direct laser photon momentum transfer between two platforms to and from unprecedented high speeds approaching the speed of light. The thrust from the direct laser photon momentum is amplified in an intracavity arrangement, in which laser photons bounce between two high reflectance mirrors separately located in two platforms. The laser gain medium is typically located between two mirrors, and amplifies the intracavity photon power, thus creating amplified thrust. This intracavity medium location arrangement offers two critical advantages: 1) the ability to maintain the intracavity photon power constant when the distance between the mirrors rapidly changes; and 2) the ability to overcome the power loss mechanisms, such as scattering and absorption.

Abstract: Pulsed detonation engines (PDEs) are adapted for use in reaction control systems (RCS), such as thrusters for orbital correction and control (e.g., for earth-orbiting satellites), divert thrust generation and control for space-based interceptor devices, and for missile trajectory correction and motion control. According to one aspect of the invention, PDEs are adapted for motion control of so-called “kill vehicles,” which are small devices, typically launched from satellites, for strategic missile defense.

Abstract: A digital programmable driver for a solenoid valve comprises at least one valve driver circuit for opening and closing the valve and a controller capable of sending an output signal to the valve driver circuit. The valve driver circuit comprises a power switch for providing current to the valve in order to open the valve. The output signal sent by the controller directs the power switch of the valve driver circuit to provide current to the valve, thereby causing the valve to open. The controller receives a plurality of parameter values, and responsive to the plurality of parameter values, calculates a rise time, a pull-in duty cycle, a fall time, and a hold duty cycle so as to modify the output signal sent to the valve driver circuit. The plurality of parameter values comprises valve resistance, valve inductance, supply voltage, valve pull-in current, valve hold current, and valve pull-in time.

Abstract: A toroidally shaped magnetic device with distortion-free exit and access ports provides toroidal magnetic ionic drive systems for vehicles that achieve a more operationally efficient, uniform and stronger radial magnetic field. This is accomplished by magnetizing a group of magic cylinder sections in a cylindrical direction and affixing them to a uniformly magnetized cylindrical shell with no magnetic field in its central cavity to produce composite cylindrical magnetic segments magnetized in a cylindrical direction with a central cavity and a uniform interior magnetic field. The composite segments are then bent into a toroidal tube and configured in such a way that longitudinal slots can be removed from the outer surface for exit and access ports. Other embodiments include a magnetic propulsion system for space vehicles and methods for magnetizing the toroidal ionic drive structure for a vehicle.

Abstract: The present invention relates to an aircraft and to a method of getting the aircraft onto station. According to the invention, said aircraft (1) includes propulsion means (2) capable only of enabling said aircraft (1) to move and to orient itself at high altitude, and said aircraft (1) is taken to its station in the high atmosphere, in particular in the stratosphere, by means of an independent transporter (3).

Abstract: A system and a method for an ambient atmosphere ion thruster having a pair of permeable electrical members for accelerating ambient atmosphere ions which enter the thruster due to a craft's relative velocity. Neutral ambient atmosphere molecules in the intake mass flux may also be ionized and subsequently accelerated by the pair of permeable electrical members. Acceleration of any entering mass comprises an exhaust mass flux which then produces a net thrust. Such net thrust is then vectored by configuring and orienting the thrusters for imparting axial and lateral accelerations as well as pitch, yaw, and roll controls. The present invention is operational in proximity to any celestial body having a sensible atmosphere during a portion of the free trajectory or during orbiting at altitude. Useful net thrust is thereby produced without need for carrying reaction mass onboard a spacecraft.

Abstract: Pulsed detonation engines (PDEs) are adapted for use in reaction control systems (RCS), such as thrusters for orbital correction and control (e.g., for earth-orbiting satellites), divert thrust generation and control for space-based interceptor devices, and for missile trajectory correction and motion control. According to one aspect of the invention, PDEs are adapted for motion control of so-called “kill vehicles,” which are small devices, typically launched from satellites, for strategic missile defense.

Abstract: A system and a method for an ambient atmosphere ion thruster having a pair of permeable electrical members for accelerating ambient atmosphere ions which enter the thruster due to a craft's relative velocity. Neutral ambient atmosphere molecules in the intake mass flux may also be ionized and subsequently accelerated by the pair of permeable electrical members. Acceleration of any entering mass comprises an exhaust mass flux which then produces a net thrust. Such net thrust is then vectored by configuring and orienting the thrusters for imparting axial and lateral accelerations as well as pitch, yaw, and roll controls. The present invention is operational in proximity to any celestial body having a sensible atmosphere during a portion of the free trajectory or during orbiting at altitude. Useful net thrust is thereby produced without need for carrying reaction mass onboard a spacecraft.