Abstract: In a satellite control system, a liquid is ejected by thermal ejection from holes in a substrate structure to create a reactive force on the satellite allowing the position, such as the attitude, of the satellite to be adjusted.

Abstract: In a satellite control system, a liquid is ejected by thermal ejection from holes in a substrate structure to create a reactive force on the satellite allowing the position, such as the attitude, of the satellite to be adjusted.

Abstract: Methods and apparatus for a missile having an external system operate in conjunction with an airframe and a fluid transfer system. The airframe includes an interior surface defining a substantially enclosed internal chamber. The fluid transfer system selectively connects the internal chamber to the external system, for example to provide pressurant or coolant to the external system.

Abstract: A system for providing attitude control with respect to a spacecraft is provided. The system includes a reaction wheel control module configured to control a number of reaction wheel assemblies associated with the spacecraft in order to control attitude, and a maneuver control module configured to use a number of gimbaled Hall Current thrusters (HCTs) to control the total momentum associated with the spacecraft during an orbit transfer. The total momentum includes the momentum associated with the reaction wheel assemblies and the angular momentum of the spacecraft. Using the gimbaled HCTs to control the momentum associated with the reaction wheel assemblies during the orbit transfer results in minimal HCT gimbal stepping.

Abstract: Disclosed herein, in certain embodiments, is a method of altering the stability of unstable space debris. In some embodiments, the method further comprises changing the orbit of the unstable space debris.

Abstract: Systems and methods are disclosed employing electric propulsion stationkeeping in a cyclical manner to better match the cyclical pattern of power generated by the solar array system. For a typical orbit design, e.g. a geostationary orbit, North-South stationkeeping can be intermittently suspended, tolerating some additional drift but yielding in a very significant reduction in the required solar power system. If necessary, stationkeeping can be supplemented with a chemical thrusters during off periods for the electric propulsion. Because of this, the overall electrical power margin for the solar array system design can be reduced without compromising the mission performance.

Abstract: The present system and methods enable simultaneous momentum dumping and orbit control of a spacecraft, such as a geostationary satellite. Control equations according to the present system and methods generate accurate station-keeping commands quickly and efficiently, reducing the number of maneuvers needed to maintain station and allowing station -keeping maneuvers to be performed with a single burn. Additional benefits include increased efficiency in propellant usage, and extension of the satellite's lifespan. The present system and methods also enable tighter orbit control, reduction in transients and number of station-keeping thrusters aboard the satellite. The present methods also eliminate the need for the thrusters to point through the center of mass of the satellite, which in turn reduces the need for dedicated station-keeping thrusters. The present methods also facilitate completely autonomous orbit control and control using Attitude Control Systems (ACS).

Abstract: A tubular distributor in a missile having a right-angled hollow tubular first section, the right-angled hollow tubular first section having a first leg and a second leg, a straight hollow tubular second section, an end of the first leg of the right-angled hollow tubular first section perpendicularly connected to a first end of the hollow tubular second section, a straight hollow tubular third section, a first end of the hollow tubular third section perpendicularly connected to the first end of the straight hollow tubular second section, and a close/open valve positioned within the straight hollow tubular third section, the close/open valve positioned between a gas generator and the straight hollow tubular second section.

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: A system and a method for commanding a spacecraft to perform a three-axis maneuver purely based on “position” (i.e., attitude) measurements. Using an “inertial gimbal concept”, a set of formulae are derived that can map a set of “inertial” motion to the spacecraft body frame based on position information so that the spacecraft can perform/follow according to the desired inertial position maneuvers commands. Also, the system and method disclosed herein employ an intrusion steering law to protect the spacecraft from acquisition failure when a long sensor intrusion occurs.

Abstract: A system for providing attitude control with respect to a spacecraft is provided. The system includes a reaction wheel control module configured to control a number of reaction wheel assemblies associated with the spacecraft in order to control attitude, and a maneuver control module configured to use a number of gimbaled Hall Current thrusters (HCTs) to control the total momentum associated with the spacecraft during an orbit transfer. The total momentum includes the momentum associated with the reaction wheel assemblies and the angular momentum of the spacecraft. Using the gimbaled HCTs to control the momentum associated with the reaction wheel assemblies during the orbit transfer results in minimal HCT gimbal stepping.

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: 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: A system for providing attitude control with respect to a spacecraft is provided. The system includes a reaction wheel control module configured to control a number of reaction wheel assemblies associated with the spacecraft in order to control attitude, and a maneuver control module configured to use a number of gimbaled Hall Current thrusters (HCTs) to control the total momentum associated with the spacecraft during an orbit transfer. The total momentum includes the momentum associated with the reaction wheel assemblies and the angular momentum of the spacecraft. Using the gimbaled HCTs to control the momentum associated with the reaction wheel assemblies during the orbit transfer results in minimal HCT gimbal stepping.

Abstract: A satellite has a depletion detector arranged in a propellant line such that, when depletion is detected, the amount of propellant remaining in the propellant lines is sufficient to dispose of the satellite, and may include a margin sufficient for 6-12 months of stationkeeping. This provides a simple and reliable method of determining when decommissioning is required.

Abstract: The present system and methods enable simultaneous momentum dumping and orbit control of a spacecraft, such as a geostationary satellite. Control equations according to the present system and methods generate accurate station-keeping commands quickly and efficiently, reducing the number of maneuvers needed to maintain station and allowing station-keeping maneuvers to be performed with a single burn. Additional benefits include increased efficiency in propellant usage, and extension of the satellite's lifespan. The present system and methods also enable tighter orbit control, reduction in transients and number of station-keeping thrusters aboard the satellite. The present methods also eliminate the need for the thrusters to point through the center of mass of the satellite, which in turn reduces the need for dedicated station-keeping thrusters. The present methods also facilitate completely autonomous orbit control and angular momentum control using.

Abstract: The present system and methods enable simultaneous momentum dumping and orbit control of a spacecraft, such as a geostationary satellite. Control equations according to the present system and methods generate accurate station-keeping commands quickly and efficiently, reducing the number of maneuvers needed to maintain station and allowing station-keeping maneuvers to be performed with a single burn. Additional benefits include increased efficiency in propellant usage, and extension of the satellite's lifespan. The present system and methods also enable tighter orbit control, reduction in transients and number of station-keeping thrusters aboard the satellite. The present methods also eliminate the need for the thrusters to point through the center of mass of the satellite, which in turn reduces the need for dedicated station-keeping thrusters. The present methods also facilitate completely autonomous orbit control and control using Attitude Control Systems (ACS).

Abstract: The present invention blocks and/or attenuates the upstream travel of acoustic disturbances or sound waves from a flight vehicle or components of a flight vehicle traveling at subsonic speed using a local injection of a high molecular weight gas. Additional benefit may also be obtained by lowering the temperature of the gas. Preferably, the invention has a means of distributing the high molecular weight gas from the nose, wing, component, or other structure of the flight vehicle into the upstream or surrounding air flow. Two techniques for distribution are direct gas injection and sublimation of the high molecular weight solid material from the vehicle surface. The high molecular weight and low temperature of the gas significantly decreases the local speed of sound such that a localized region of supersonic flow and possibly shock waves are formed, preventing the upstream travel of sound waves from the flight vehicle.

Abstract: Systems and methods are disclosed employing electric propulsion stationkeeping in a cyclical manner to better match the cyclical pattern of power generated by the solar array system. For a typical orbit design, e.g. a geostationary orbit, North-South stationkeeping can be intermittently suspended, tolerating some additional drift but yielding in a very significant reduction in the required solar power system. If necessary, stationkeeping can be supplemented with a chemical thrusters during off periods for the electric propulsion. Because of this, the overall electrical power margin for the solar array system design can be reduced without compromising the mission performance.

Abstract: A method for performing east-west station keeping for a satellite in an inclined synchronous orbit is described. The method includes averaging a value of a right ascension of the ascending node for an inclination vector associated with the satellite over a period of the control cycle, and managing corrections for the satellite such that an eccentricity vector, directed at perigee, is substantially collinear with the inclination vector.

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 thrust vector actuation control system and method is configured to allow self-testing of the entire actuation system and/or its individual system components. The control system also provides real-time, continuous monitoring of actuation system status, and allows system gain and compensation parameters to be changed during vehicle operation remote from its launch site.

Abstract: A system and method for supplying thrust to a structure, such as a satellite or spacecraft, for the purposes of station keeping and attitude control of the structure in low-gravity (orbital) and zero-gravity environments. The system includes devices for emitting energy beams and targets impacted by the energy beams to cause ablation of the targets. The beam-emitting devices and targets are adapted to cooperate and cause the structure to selectively undergo translational and/or rotational motion in reaction to the motion of material ablated from the targets. The position, alignment, and/or attitude of the structure can thereby be controlled in a zero or low-gravity environment by selectively emitting the energy beams at the targets.

Abstract: An apparatus and method including a deployable spacecraft mount for electric propulsion is disclosed. A typical apparatus includes a spacecraft body, a deployable element having at least two basic positions including a compact stowed position and a deployed position of the element and an electric thruster disposed on the deployable element where the electric thruster is disposed to mitigate negative plume effects in the deployed position that would be present in stowed position. The deployable element can be a radiator and can optionally be disposed on a second deployable payload module. Further, deployable elements can be selective deployed such that some electric thrusters can be used to assist in transfer orbit while undeployed elements help retain heat.

Abstract: A method and an apparatus for controlling the attitude and momentum of a spacecraft while deploying an appendage from the spacecraft. The method comprises the steps of predicting an environmental torque the spacecraft will be subjected to during deployment of the appendage, computing a magnitude and a direction of momentum to add to the spacecraft to at least partially oppose the predicted environmental torque, and storing the computed magnitude and direction of momentum in at least one of the momentum wheels before deploying the appendage.

Abstract: A satellite life extension spacecraft, comprising a mechanical implement adapted for connection to a parent spacecraft, a first and second boom, a first thruster pod and a second thruster pod attached to the first and second boom, the first pod positioned relative to the second pod such that the center of mass of a parent/child spacecraft can be determined relative to the first and second pod.

Abstract: A system for counteracting a disturbance in a spacecraft includes a biasing apparatus that is coupled to the spacecraft and a spacecraft controller within the spacecraft. The disturbance has a known sign, magnitude and time. The biasing apparatus controls the biasing apparatus to place the spacecraft in a first dynamic state or position as a function of the sign, magnitude, and time of the disturbance. The controller also controls the spacecraft to a second dynamic state as a function of the known sign, magnitude, and time so that the spacecraft is oriented in a position other than the desired orientation so that after the disturbance the spacecraft is oriented in the desired orientation in response to the disturbance. The biasing apparatus may comprise a momentum wheel and the disturbance may comprise thrusting firing used for controlling momentum dumping.

Abstract: A velocity change (?V) thruster is operated on a spacecraft, which unavoidably causes attitude error. A reaction wheel (RWA) corrects the attitude. At the beginning of the thruster maneuver, the total attitude control momentum required to at least correct for the ?V thruster attitude errors over the duration of the entire maneuver is determined, and the RWA momentum may also be determined. Attitude control thrusters (REAs) are operated. The REAs are operated to correct at least the net ?V thruster induced attitude error, and preferably also to reset the RWA to its nominal momentum. The maneuver may be stationkeeping.

Abstract: A method for increasing the roll offset operating range for a spacecraft using an earth sensor operating in single scan mode includes the steps of moving the spacecraft to a first roll position, which has a roll angle that will cause the earth sensor to have a desired standard chord, switching the earth sensor to single scan mode by deselecting one of the earth sensor scans, whereby switching the earth sensor to a single scan mode, the earth sensor standard chord is locked at or near the desired standard chord. After the desired standard chord is set, the spacecraft is moved to a second roll position, which is a desired roll offset operating position for the spacecraft, and the earth sensor roll output (generated by using the single scan mode) is used to calculate the spacecraft roll at the roll offset operating position.

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 providing propulsive power that utilizes a novel swirl generator for rapidly and efficiently atomizing, vaporizing, as necessary, and mixing a fuel into an oxidant. The swirl generator converts an oxidant flow into a turbulent, three-dimensional flowfield into which the fuel is introduced. The swirl generator effects a toroidal outer recirculation zone and a central recirculation zone, which is positioned within the outer recirculation zone. These recirculation zones are configured in a backward-flowing manner that carries heat and combustion byproducts upstream where they are employed to continuously ignite a combustible fuel/oxidizer mixture in adjacent shear layers. The swirl generator is compatible with the throttle range of conventional gas turbine engines, provides smooth combustion with no instabilities and minimum total pressure losses, enables significant reductions the in L/D ratio of the combustor and is readily packaged into various applications.

Abstract: A velocity change (&Dgr;V) thruster is operated on a spacecraft, which unavoidably causes attitude error. A reaction wheel (RWA) corrects the attitude. At the beginning of the thruster maneuver, the total attitude control momentum required to at least correct for the &Dgr;V thruster attitude errors over the duration of the entire maneuver is determined, and the RWA momentum may also be determined. Attitude control thrusters (REAs) are operated. The REAs are operated to correct at least the net &Dgr;V thruster induced attitude error, and preferably also to reset the RWA to its nominal momentum. The maneuver may be stationkeeping.

Abstract: A system for controlling attitude about pitch, yaw, and roll axes and axial thrust of a body. The system comprises a main thrust generator located on an aft portion of the body and at least one reaction control system (RCS) located on a forward portion of the body. A thrust vector controller (TVC) is connectable to the main thrust generator and at least one RCS controller is connectable to the RCS. The RCS controller and the TVC are synchronized to adjust the direction of the principal line of thrust through the body center of gravity.

Abstract: An attitude control system for a spacecraft that includes a north side, a south side, an east side, a west side, an earth facing side, and an anti-earth facing side. The attitude control system includes a plurality of gimbaled thrusters interconnected with the spacecraft and arranged to produce thrust in a direction parallel to a roll-pitch plane of the spacecraft.

Abstract: A versatile cavity actuator. The versatile cavity actuator includes a cavity having one or more polymer-based sidewalls. An energetic material is disposed therein. A heater is disposed on or within the cavity. In a specific embodiment, the cavity includes a thermally insulating base positioned beneath the heater, which is positioned near the bottom of the cavity. The polymer-based sidewalls are constructed from a photo-curable epoxy, which is disposed on a substrate via microelectromechanical processes. The sidewalls are angled or parabolic and are constructed via a low-temperature lithographic spin process compatible with post integrated circuit processing.

Abstract: An integrated detonation element or firing element including a base member, e.g., a silicon member, and a reaction region associated therewith, is provided. The reaction region includes porous silicon and an oxidizing agent for silicon. An arrangement is provided with which a chemical reaction is initiated between the oxidizing agent and the porous silicon. The detonation or firing element is suitable principally for use in a microreactor; in a microbooster, e.g., for course correction of satellites; as a firing element in a gas generator for a belt tensioner or an airbag, e.g., in motor vehicles; or as a primer for the ignition of explosive charges.

Abstract: The present invention provides a maneuver device for an artificial satellite, which causes small attitude error during maneuver and which requires a shorter period of setting time for obtaining a target attitude. The maneuver device is provided with: a feed forward torque instruction signal generator 8 which outputs a feed forward torque instruction signal 11 based on a maneuver plan; a thruster 10 which outputs control torque based on the feed forward torque instruction signal 11; and an attitude control signal calculator 6 to which an attitude angle and an angular velocity of the artificial satellite as well as a target attitude angle and a target angular velocity are input and which outputs an attitude control signal 13. The maneuver device is further provided with a disturbance compensating signal calculator to which the feed forward torque instruction signal 11 and a detected angular velocity signal 16 are input, and which generates and outputs a disturbance compensating signal 12.

Abstract: Methods and structures are provided that enhance the accuracy of the service attitude of an inclined-orbit spacecraft and, thereby, facilitate reduction of service error between a communication service area and the spacecraft's payload beam. The enhancement is realized by configuring a beacon-receiving antenna to have a beacon-receiving field-of-view that substantially matches a beacon-station window. Preferably, the beacon-receiving field-of-view is elongated and tilted to enhance its match with the beacon-station window in both size and orientation. The goals are also realized by configuring the beacon-receiving antenna to have a beacon-receiving field-of-view that is substantially smaller than the beacon-station window and successively steering a beacon-receiving boresight to successive beacon-receiving attitudes that maintain the beacon station within the beacon-receiving field-of-view over each solar day.

Abstract: A multi-stage pilot valve is disclosed for selectively supplying a working fluid to and venting from a reaction jet main stage actuation chamber. In a preferred embodiment of the invention, the pilot valve comprises a solenoid actuated ball and socket flapper valve having a pressure inlet, an exhaust outlet and a service port. The first stage service port is in fluid communication with a second stage actuation chamber. A piston disposed in the second stage actuation chamber operatively engages a ball member of a ball-and-seat type valve comprising the second stage valve. The second stage valve also comprises a pressure inlet, an exhaust outlet and a service port. The piston and ball are sized relative to each other such that when the second stage actuation chamber is pressurized by the first stage, the force is sufficient to seat the ball against the second stage pressure inlet.

Abstract: Various ways of integrating pulsed plasma thrusters with a spacecraft. In one embodiment, pulsed plasma thrusters are mounted on inflatable struts/booms. In another embodiment, pulsed plasma thrusters are mounted within a recess formed on an exterior surface of the spacecraft body.

Abstract: An electric propulsion (EP) device is used to enable, a stationkeeping satellite to track a prescribed stationkeeping Earth orbit. EP propellant and electric power are throttled to vary the thrust and specific impulse of the EP device. A solar array provides electrical power during each Earth day cycle with excess power above that needed by the spacecraft stored by a battery. Software control manages the voltage, current and burn time to minimize propellant usage and impact to the system.

Abstract: Methods and structures are provided that enhance the accuracy of the service attitude of an inclined-orbit spacecraft and, thereby, facilitate reduction of service error between a communication service area and the spacecraft's payload beam. The enhancement is realized by configuring a beacon-receiving antenna to have a beacon-receiving field-of-view that substantially matches a beacon-station window. Preferably, the beacon-receiving field-of-view is elongated and tilted to enhance its match with the beacon-station window in both size and orientation. The goals are also realized by configuring the beacon-receiving antenna to have a beacon-receiving field-of-view that is substantially smaller than the beacon-station window and successively steering a beacon-receiving boresight to successive beacon-receiving attitudes that maintain the beacon station within the beacon-receiving field-of-view over each solar day.

Abstract: The present invention provides methods of and apparatus for determining the inertial attitude of an aerospace vehicle. In one embodiment, the invention provides a rotational astronomical object-sighting concept to determine the inertial attitude of an axis of the aerospace vehicle without the star identification or dragback. In another embodiment, the invention provides an attitude measurement apparatus comprising a high sensitivity optical sensor and a low power inertial sensor.

Abstract: A thruster system for spacecraft includes a thruster, a thruster mounting boom, a gimbals mechanism connecting the thruster to the thruster mounting boom, and a pivot mechanism connected to the thruster mounting boom, where the pivot mechanism attaches the thruster mounting boom to a spacecraft. A method includes steps of providing a first thruster mounted on a pivoting thruster mounting boom attached to a spacecraft; using the first thruster at a first position, for example, a north pointing position, to provide orbit control; repositioning the first thruster to a second position, for example, a south pointing position; and using the first thruster at the second position to provide orbit control so as to provide redundancy for a failed second thruster by repositioning the first thruster.

Abstract: An electric propulsion (EP) device is used to enable a stationkeeping satellite to track a prescribed stationkeeping Earth orbit. EP propellant and electric power are throttled to vary the thrust and specific impulse of the EP device. A solar array provides electrical power during each Earth day cycle with excess power above that needed by the spacecraft stored by a battery. Software control manages the voltage, current and burn time to minimize propellant usage and impact to the system.

Abstract: A satellite attitude control system includes a programmed processor system which includes a gyroscopic actuator first control stage for changing the attitude of the satellite and a reaction wheel second control stage for assuring that pointing of the satellite is accurate and stable. The method is intended to be used for a satellite including the two control stages indicated above, which it uses selectively for the operations indicated above.

Abstract: A microthruster having an inverted exhaust system traps burst diaphragm fragments providing a clean exhaust while an exhaust port provides increased back pressure for efficient combustion of a propellant charge in a fuel cell. A converging diverging micronozzle provides a predictable exhaust vector for improved microthrusting well suited for propulsion system on small spacecraft.

Abstract: A method for controlling spacecraft velocity. A thruster torque gimbal angle command is calculated to provide a desired thruster torque demand command including a desired momentum adjust torque and a desired attitude control torque. A velocity change error between a commanded velocity change and an actual velocity change, a commanded force vector based on the error to cause the actual velocity change to approach a commanded velocity change, and force gimbal angle commands and force thrust commands to achieve the commanded force vector are calculated; and/or an error between gimbal angles and thrust levels and reference gimbal angles and thrust levels, reference gimbal angle commands and reference thrust commands to drive actual gimbal angles and thrust levels toward reference gimbal angles and reference thrust levels are calculated. A total gimbal angle command is calculated by adding the torque gimbal command with at least one of the force gimbal angle command and the reference gimbal angle command.

Abstract: The present invention provides a maneuver device for an artificial satellite, which causes small attitude error during maneuver and which requires a shorter period of setting time for obtaining a target attitude.

Abstract: Spacecraft thruster systems are provided that can realize station changing, station keeping and momentum dumping in spacecraft while minimizing loss of spacecraft service time. The systems are formed with pairs of E and W thrusters, NE and SE thrusters and NW and SW thrusters whose thrust directions are each defined by respective polar angles &PHgr; and slew angles &agr;. The E and W thrusters are especially suited for rapid station changing while the remaining thrusters generate normal, tangential and radial thrust components required for station keeping. Because all thrusters are directed through the spacecraft's center of mass during station changing and station keeping firings, the spacecraft's attitude is not disturbed and loss of service time is avoided. The thrusters are preferably gimbaled so that they can track spatial changes of the spacecraft's center of mass or, for momentum dumping, be offset from that center of mass.