Attitude Control Patents (Class 244/164)
-
Patent number: 5996942Abstract: A method in accordance with this invention for maintaining a spacecraft in a desired orbital configuration, comprises the steps of: (a) recording a history of yaw momentum stored in momentum wheels; (b) estimating an average inertial torque and momentum over a previous time interval (such as one day); (c) determining a desired change in inertial torques using PID control law; (d) commanding a change in satellite trim tab and solar array position from a desired change in solar torque; and (e) slewing the trim tab and solar array a desired amount.Type: GrantFiled: September 22, 1997Date of Patent: December 7, 1999Assignee: Space Systems/Loral, Inc.Inventors: Xen Price, Kam Chan
-
Patent number: 5992799Abstract: A spacecraft ground loop controller (GLC), located on the Earth, interfaces with a satellite ground station receiving spacecraft telemetry from the downlink baseband equipment and automatically sending spacecraft commands through the command uplink baseband equipment to control the attitude of an orbiting spacecraft and achieve partial orbit control using commanded thruster firings and magnetic torquer polarity and magnitude. A cooperative approach of using all available thrusters, of both the primary and redundant strings, provides greater fuel savings.Type: GrantFiled: February 18, 1997Date of Patent: November 30, 1999Assignee: Space Systems/Loral, Inc.Inventors: Donald W. Gamble, Mark D. McLaren, Marc Takahashi
-
Patent number: 5984238Abstract: A system for autonomous on-board determination of the position of an earth orbiting satellite consists of a biaxially measuring earth sensor which defines the z-axis (yaw axis) of the satellite and of several biaxially measuring sun sensor measuring heads which are arranged on the satellite structure such that, with the exception of earth shadow phases, they supply a direction vector measurement. Out-of-plane movement of the satellite is propagated by means of a precise model of the satellite dynamics including natural disturbance forces and thrusts during maneuvers on board. In order to compensate sensor uncertainties and the effects of thermal deformations, the satellite orbit is precisely measured at regular intervals (approximately 3 months), and by means of this information, a calibration function (time function for a day) for compensating measured values is determined.Type: GrantFiled: February 2, 1998Date of Patent: November 16, 1999Assignee: Diamler-Benz Aerospace AGInventors: Michael Surauer, Walter Fichter, Oliver Juckenhoefel
-
Patent number: 5984236Abstract: A method of simultaneously controlling East/West and North/South positioning and unloading momentum of a spacecraft while orbiting an object. The spacecraft has a thruster array and a momentum accumulator. The method entails moving said spacecraft towards a node of the orbit. At a predetermined position on the orbit, separate from the node, a thruster of the thruster array is fired so as to control the orbital position of the spacecraft. While the thruster is being fired, momentum is dumped from the momentum accumulator at the predetermined position so that any loss in control in the attitude of the spacecraft is reduced.Type: GrantFiled: December 22, 1995Date of Patent: November 16, 1999Inventors: Keith F. Keitel, Richard A. Noyola, John F. Yocum, Jr., David K. Abernethy, Bernard M. Anzel
-
Patent number: 5961077Abstract: A launcher places a first satellite practically directly on a final orbit. A second satellite carried by the same launcher is initially transferred onto a waiting orbit that is highly elliptical, having a semi-major axis situated in the initial orbital plane. The inclination and the perigee of the waiting orbit are then changed in the vicinity of the apogee of the waiting orbit, thereby placing the second satellite on an intermediate orbit. A maneuver is performed that includes at least one step making use of atmospheric braking in the vicinity of perigee of the intermediate orbit so as to lower the altitude of apogee of the intermediate orbit, and an impulse is supplied to the second satellite at apogee of the intermediate orbit so as to raise its perigee and transform the intermediate orbit into a final orbit having orbital parameters whose values are substantially different from those of the orbital parameters of the final orbit of the first satellite.Type: GrantFiled: December 30, 1997Date of Patent: October 5, 1999Assignee: Societe Nationale D'Etude et de Construction de Moteurs D'AviationInventors: Christophe Koppel, Dominique Valentian
-
Patent number: 5957411Abstract: A spacecraft includes a plurality of thrusters mounted at predetermined locations on a spacecraft structure, individual ones the plurality of thrusters being fired for generating a torque about a desired axis. The firing of a thruster is partitioned into two firings that are offset in time by an amount .DELTA.t, wherein .DELTA.t=(1/2).times.(1/F).times.SF, where F is a dominant modal frequency, in Hertz, for any particular axis (nominally the 1st mode) of the spacecraft structure, and where SF is a scale factor that is adjustable about the frequency. In a presently preferred embodiment the thruster is a low thrust thruster that is mounted on a solar array panel, and the thrusters are fired in pairs.Type: GrantFiled: December 24, 1997Date of Patent: September 28, 1999Assignee: Space Systems/Loral, Inc.Inventors: Tung Y Liu, Kam K Chan
-
Patent number: 5957410Abstract: A low earth orbiting satellite control arrangement according to the invention has an earth sensor for measuring roll and pitch of the satellite relative to an earth oriented moving reference system, a spin wheel mounted along an axis perpendicular to the orbital plane of the satellite for measuring spin of the satellite, and two magnet coils for generating control. A first of the two magnet coils is aligned substantially in parallel to a roll axis of the satellite, and the second is aligned substantially parallel to the yaw axis. Alternative control algorithms are disclosed for controlling attitude, nutation and spin of the satellite based on information from the earth sensor, spin wheel and an observer which is also mounted on the satellite.Type: GrantFiled: June 5, 1996Date of Patent: September 28, 1999Assignee: Daimler-Benz Aerospace AGInventors: Ernst Bruederle, Michael Surauer, Walter Fichter
-
Patent number: 5949675Abstract: For a switching of a magnitude of system gain in the driving of a load by a control system, a method of inhibiting formation of switching transients has a step of distinguishing between an estimate of a plant state, such as the attitude of a spacecraft, and an integral control state which allows the control system to provide actuation in response to long term disturbances, such as solar pressure, in operation of the system. In the case of a feedback configuration to the system with controller in a forward branch and an estimator in a feedback branch, a loop error signal serves to drive an actuator of the load via the controller. The method includes a further step of evaluating a portion of a controller input signal which is exclusive of an integral control state and which comprises a difference between a desired state and an estimate of the plant state.Type: GrantFiled: November 1, 1996Date of Patent: September 7, 1999Assignee: Space Systems/Loral, Inc.Inventors: Thomas Joseph Holmes, David L. Cielaszyk, David J. Wirthman
-
Patent number: 5944761Abstract: A filter for use in an attitude control system which may be subjected to vibrations from disturbances which have frequency contents that vary with time. Sensors on the system produce signals that contain the disturbance frequencies and a calculator connected to the source of disturbance produces a signal which varies as a function of the disturbance. The filter receives these signals and augments the control loop gain at the disturbance frequencies thereby enabling the attitude control devices to null the disturbance effects on the attitude of the system.Type: GrantFiled: June 6, 1997Date of Patent: August 31, 1999Assignee: Honeywell Inc.Inventor: Christopher J. Heiberg
-
Patent number: 5935176Abstract: Momentum wheel speed correction in an orbital space vehicle that corrects for the oscillation resulting from spurious variations in the angular velocity of the momentum wheel. The momentum wheel's angular velocity is subject to random fluctuations due to Coulomb forces. These fluctuations are coupled into the vehicle's spin rate causing errors in the yaw rate and yaw. By adding corrective values to the determined yaw rate and yaw by selectively filtering the momentum wheel tachometer signals, pointing errors are substantially reduced or even eliminated.Type: GrantFiled: November 22, 1996Date of Patent: August 10, 1999Assignee: Lockheed Martin CorporationInventor: Marlin C. Nielson
-
Patent number: 5934619Abstract: In a process and apparatus for controlling the attitude of a spacecraft, drive signals for driving attitude control units of the spacecraft are generated based on control unit signals related to the three principle axes of a spacecraft oriented coordinate system, using the rule k=B.sup.I e+cu.sub.2. The drive matrix B.sup.I, a definite vector u.sub.2 and the constant c result from the application of a matrix decomposition according to the Singular Value Decomposition (SVD) method to a nozzle matrix B. The latter matrix contains as elements, the momentum and force vectors which can be generated by the attitude control units, and the drive matrix B.sup.I is multiplied directly with the vector e derived from the axis-related control unit signals.Type: GrantFiled: June 3, 1996Date of Patent: August 10, 1999Assignee: Daimler-Benz Aerospace AGInventors: Horst-Dieter Fischer, Joachim Chemnitz, Michael Surauer
-
Patent number: 5931419Abstract: Satellites are launched to a first orbit on a launch rocket. The satellites are connected together, forming a satellite cluster, that is propelled from the first orbit to a higher mission orbit by thrustors on each satellite. The satellite cluster is formed so that satellite cluster rotation from aerodynamic drag on the individual satellites is minimized. The thrustors are sequenced to control satellite cluster attitude during the transition to the second altitude. At the second altitude, the satellites are separated from the satellite cluster and propelled to their mission orbits using their own thrustors and attitude system.Type: GrantFiled: August 7, 1997Date of Patent: August 3, 1999Assignee: Honeywell Inc.Inventor: Gordon L. Collyer
-
Patent number: 5927653Abstract: A two-stage wingless reusable aerospace vehicle having upper and lower stages that take off from a take-off area and separate at a separation point along a first trajectory. The separation forces are generated by air retained between the upper and lower stages, which is at a pressure higher than ambient pressure at the separation point. The lower stage is then propelled along a return trajectory to a landing area. After separation from the lower stage, the upper stage continues to an Earth orbit for deployment of a payload. After deploying the payload, the upper stage moves out of the Earth orbit, re-enters the Earth's atmosphere, and returns to the take-off and landing area. The upper and lower stages are powered by liquid oxygen and kerosene engines.Type: GrantFiled: April 17, 1996Date of Patent: July 27, 1999Assignee: Kistler Aerospace CorporationInventors: George E. Mueller, Walter P. Kistler, Thomas G. Johnson, Henry O. Pohl, Chris McLain, Allan S. Hill, Jason E. Andrews, Thomas C. Taylor, Aaron Cohen, Dale Myers, Adam P. Bruckner, Steven C. Knowles, Richard Warwick
-
Patent number: 5862495Abstract: Real time correction to ground generated satellite ephemeris implements a reference trajectory management module (334), an error estimation module (336), an parameter calculation module (338) and a position information management module (340). The reference trajectory management module (334) accesses reference trajectory data such as ground generated ephemeris data which is uplinked to a spacecraft through a communications unit (312). The position information management module (340) accesses and interprets position information such as Global Positioning System (GPS) measurement data to provide measured spacecraft positions. The parameter calculation module (338) calculates spacecraft position and velocity from the reference trajectory data.Type: GrantFiled: September 18, 1996Date of Patent: January 19, 1999Assignee: Lockheed Martin Corp.Inventors: Hunt W. Small, John E. Bergesen, Brian J. Howley
-
Patent number: 5852792Abstract: Concurrent determinations of errors for application of corrections to sensor data signals from an orbiting space vehicle to reduce or to eliminate the three components of boresight error. To locate more accurately the geographical position of targets detected by the sensors of an orbiting satellite, the pitch, roll, and yaw boresight errors caused by distortions or misalignments in the focal plane of the sensors are compensated for by comparing known position coordinates of an object (such as a star) to the observed position coordinates. The known position coordinates are converted to sensor coordinates by a series of intermediate coordinate conversions, essentially from the celestial frame of reference to an earth-centered frame of reference to a satellite attitude frame of reference to the satellite frame of reference to the sensor frame of reference.Type: GrantFiled: October 3, 1996Date of Patent: December 22, 1998Assignee: Lockheed Martin CorporationInventor: Marlin Craig Nielson
-
Patent number: 5844232Abstract: This sensor, which is connected to an attitude control system comprising a coarse attitude detector on a mobile craft, has an array of photodetectors (1) and a slit (3) interposed between the array and the sun. According to the invention, it comprises a sun sensor having a plurality of parallel slits (3, 4, 5), each slit detecting a sector (C.sub.1, A, C.sub.2) of the overall field of view for a first angular position of the sun to be measured, with a small overlap between adjacent sectors. Said sun sensor further comprises calculation means (31) which receive, on the one hand, information about the insolation of the array and, on the other hand, information originating from the attitude control system (inertial unit 32), in combination with the date, indicating in which sector of the field of view the sun is situated, and which derive from said information the angular position of the sun in said overall field of view.Type: GrantFiled: February 14, 1997Date of Patent: December 1, 1998Assignee: U.S. Philips CorporationInventor: Christian Pezant
-
Patent number: 5841018Abstract: In accordance with the disclosed method, an attitude determining device which is installed onboard a mobile craft, like an aircraft, for example, at an unknown orientation with respect to the reference coordinate system of the craft senses its installation orientation with respect to an earth frame coordinate system when the craft is at rest to obtain a static orientation measurement thereof. Thereafter, an attitude of the mobile craft with respect to the earth frame is measured with the attitude determining device and such measurement is compensated with the static orientation measurement to obtain attitude information of the craft's reference coordinate system with respect to the earth frame coordinate system. The installation orientation of the attitude determining device may be sensed while the craft is at rest in either a leveled or unleveled condition.Type: GrantFiled: December 13, 1996Date of Patent: November 24, 1998Assignee: B. F. Goodrich Avionics Systems, Inc.Inventors: Gary Stewart Watson, Krishna Devarasetty
-
Patent number: 5826828Abstract: Methods for performing attitude maneuvers of a spinning satellite without the use of thrusters, or with a minimal number of thrusters, are disclosed. The attitude maneuvers are primarily achieved through the use of gimballed momentum wheels and solar wing drives. Various maneuvers can be performed depending on whether the satellite has near-zero net momentum or significant net momentum. The maneuvers include sun acquisition, sun hold, Earth acquisition and inversion.Type: GrantFiled: February 5, 1996Date of Patent: October 27, 1998Assignee: Hughes Electronics CorporationInventors: Richard A. Fowell, John F. Yocum, Jr.
-
Patent number: 5826831Abstract: A geostationary orbital box is divided by radially-oriented planes to form orbital sub-boxes. A thruster system facilitates station keeping satellites in each of these orbital sub-boxes. The thruster system includes thruster pairs that are positioned on an anti-nadir face of each satellite and directed through the satellite's center of mass. In addition, multiple satellites can be stationed in each of the sub-boxes by causing them to rotate in a chosen one of a clockwise and a counterclockwise direction and to be spaced from each other in at least one of a radial direction and a normal direction. To enhance the fine-grain control of each of the satellites, the thruster system preferably uses ion-propulsion thrusters.Type: GrantFiled: August 22, 1996Date of Patent: October 27, 1998Assignee: Hughes ElectronicsInventor: Bernard M. Anzel
-
Patent number: 5826829Abstract: An active attitude control system for a spacecraft having first, second, and third mutually perpendicular axes utilizes four flywheels, at a minimum, which can selectively provide options of fully redundant momentum bias, 3/4 redundant momentum bias, or fully redundant zero momentum bias. A plurality of reaction wheels are mounted on the spacecraft and rotatable on spin axes in a fixed configuration for together maintaining full three-axis control of the spacecraft to a predetermined attitude. A momentum wheel is also rotatable on the spacecraft about the first axis for maintaining gyroscopic stiffness. In the event of a failure of the momentum wheel, the reaction wheels may have a combined angular momentum sufficient to maintain the gyroscopic stiffness lost by the failure of the momentum wheel while maintaining full three-axis control of the spacecraft to a predetermined attitude.Type: GrantFiled: July 15, 1996Date of Patent: October 27, 1998Assignee: Space Systems/Loral Inc.Inventor: Thomas Joseph Holmes
-
Patent number: 5822515Abstract: In a spacecraft having devices which are operative in plural modes, particularly an earth sensor operative in a north hemispherical inhibit mode, a south hemispherical inhibit mode, and a normal mode, a watchdog circuit* is employed to detect the presence of a mode command issued to the device. The command may be issued by a station on the earth, and communicated via a telemetry link to the spacecraft. The watchdog circuit may be part of a central processing unit (CPU) with a computer on board the spacecraft, the computer serving to perform various data processing operations for the mission of the spacecraft. The watchdog circuit stores the mode command and then repeats the command during a succession of repetitions.Type: GrantFiled: February 10, 1997Date of Patent: October 13, 1998Assignee: Space Systems/Loral, Inc.Inventor: Michel B. Baylocq
-
Patent number: 5816540Abstract: A spacecraft traveling in a volume of space receiving radiation from the sun and encountering an undesired force. The spacecraft has a solar panel movably attached to said spacecraft. The spacecraft further includes a solar panel controller which controls the movement of the solar panel, wherein the controller moves the solar panel in a time modulated manner so that a torque is generated which compensates for the undesired force.Type: GrantFiled: December 22, 1995Date of Patent: October 6, 1998Assignee: Hughes ElectronicsInventors: John R. Murphy, Ross Crowley
-
Patent number: 5806801Abstract: The present invention provides a method and system for formationkeeping between two or more orbiting spacecraft by controlling the surface area of the spacecraft facing the direction of perturbative forces acting on the spacecraft. The actual position of each spacecraft is sensed and the distance between the spacecraft is computed. If this separating distance exceeds acceptable tolerances, the affected spacecraft are commanded to change their orientation so that the total surface area facing the direction of motion is altered. Alternatively, the total surface area of the spacecraft facing the direction of solar pressure can be altered. As a result, the forces acting on the spacecraft are altered which alters spacecraft position and velocity. In accordance with another aspect of the invention, surface area variations of individual spacecraft are used to maintain the planar separation between adjacent constellations of orbiting spacecraft at desired distances.Type: GrantFiled: August 5, 1996Date of Patent: September 15, 1998Assignee: Orbital Sciences CorporationInventors: David A. Steffy, Gregg E. Burgess, Maria J. Evans
-
Patent number: 5803407Abstract: The life of a target satellite is modified, i.e., extended or terminated by docking an extension spacecraft with the target satellite to form a docked satellite-spacecraft combination. The extension spacecraft is docked with and mechanically connected to the target satellite and includes guidance, navigation, and control systems for performing the rendezvous and docking maneuvers and for controlling the position of the docked spacecraft-satellite combination. The extension spacecraft also includes an onboard propellant supply for accomplishing the rendezvous and docking of the spacecraft with the satellite and for controlling the position of the docked spacecraft-satellite combination. A remote cockpit system is provided to permit human control of the extension spacecraft during proximity operations.Type: GrantFiled: April 21, 1995Date of Patent: September 8, 1998Inventor: David R. Scott
-
Patent number: 5799904Abstract: A spacecraft (10) has an attitude sensor (20) mounted on its body (12). The sensor must be maintained near a temperature setpoint. Each sensor (20) produces its own temperature-indicative signal. Each attitude sensor is coupled to a thermally conductive instrument platform (18). A standoff (21) supports the platform (18) away from a baseplate (16) and the spacecraft body (12). The standoff (21) includes a thermally conductive portion (22) adjacent the platform (18), and a nonconductive portion (24) remote from the platform. An electric heater (26) is connected to the thermally conductive portion (22) of the standoff (21). A temperature sensor (28) thermally coupled to the platform (18) generates platform temperature signals. A filter (212) high-pass filters either the platform temperature signals or the attitude sensor temperature signals, to form filtered signals. A combining circuit (218) combines the filtered signals with the other temperature signals, to make composite temperature signals.Type: GrantFiled: April 25, 1996Date of Patent: September 1, 1998Assignee: Lockheed Martin Corp.Inventors: Neil Evan Goodzeit, Arthur Jon Throckmorton
-
Patent number: 5794891Abstract: The attitude of a satellite placed on a non-heliosynchronous earth orbit in a plane that is inclined relative to the equatorial plane of the earth is controlled for efficient use of solar panels and radiators. The satellite has a structure, solar panels apt to be rotated with respect to the structure rotation and two opposed radiators each fixed on one of two opposed faces of the satellite structure which are orthogonal to the rotation axis. One of the radiators has greater emissivity than the other. A yaw axis bound to the structure satellite and orthogonal to rotation axis is aimed towards the earth. The solar panels of the satellite are maintained in an optimum orientation relative to the sun by rotating them. At least during periods of each year when an angle .beta.Type: GrantFiled: January 3, 1996Date of Patent: August 18, 1998Assignee: Matra Marconi Space FranceInventors: Bernard Polle, Marcel Billand, Benoit Hanin
-
Patent number: 5788188Abstract: For controlling the attitude of the body of an earth satellite placed on a low orbit, values of components of the geomagnetic field of the earth are measured along three axes of a frame of reference bound to the body. The values are derivated with respect to time, and multiplied by a gain. Currents responsive to the multiplicated derivatives are passed through magnetic torquers located along the three axes of the body to create magnetic torques that bias the body to a fixed angular position relative to the field lines of the geomagnetic field. Such steps are continuously carried out during eclipse periods. Out of eclipse periods, the pitch of the body is controlled by modifying an internal momentum in response to a signal provided by a solar sensor, so as to maintain solar generators carried by the body of the satellite oriented towards the Sun.Type: GrantFiled: December 5, 1996Date of Patent: August 4, 1998Assignee: Matra Marconi Space FranceInventor: Patrice Damilano
-
Patent number: 5787368Abstract: An attitude or orientation control system for spacecraft that provides for yaw axis error reduction and momentum dumping using a simple calculation involving only the momentum wheel speeds as input and which outputs two axis (roll and yaw) torque commands. The system utilizes the spacecraft 1) momentum wheels that store spacecraft momentum, 2) processing means for the calculation of control torques, and 3) torque actuators, e.g., magnetic torquers, that can provide a torque axis anywhere in the spacecraft roll/yaw plane, and involves three fundamental steps, 1) the input of the momentum wheel speed mearsurements, and 2) the calculation of the required torques, using the wheel speed measurements, and and 3) the outputting of the required torques in a form appropriate for controlling the available actuators.Type: GrantFiled: November 3, 1995Date of Patent: July 28, 1998Assignee: Space Systems/Loral, Inc.Inventors: Donald W. Gamble, Xenophon H. Price, Kam Kin Chan
-
Patent number: 5765780Abstract: A method of simultaneously performing a translational maneuver of a spacecraft by a thruster and dumping momentum from the spacecraft during a time period P. The method entails aligning the thruster along a thrust vector which is fixed during the time period P, wherein the thrust vector is aligned with the center of mass of the spacecraft at a time P/2, and firing the thruster throughout the time period P.Type: GrantFiled: December 22, 1995Date of Patent: June 16, 1998Assignee: Hughes Electronics CorporationInventors: Michael F. Barskey, John F. Yocum, Jr.
-
Patent number: 5758846Abstract: A method and system are disclosed for inverting a satellite spinning about a first desired spin axis to spin about a second desired spin axis substantially antiparallel to the first desired spin axis. A tumbling motion is induced in the satellite so that a spin axis of the satellite oscillates between the first desired spin axis and the second desired spin axis. The tumbling motion is induced by sensing at least one component of the angular rate vector and controlling a single degree of freedom momentum storage device based upon the at least one component of the angular rate vector. The single degree of freedom momentum storage device has an orientation of variation substantially perpendicular to the desired spin axis. The single degree of freedom momentum storage device is controlled so that the first desired spin axis is made an intermediate inertia axis of an effective inertia matrix.Type: GrantFiled: March 13, 1997Date of Patent: June 2, 1998Assignee: Hughes Electronics CorporationInventor: Richard A. Fowell
-
Patent number: 5758260Abstract: A satellite communication system has at least one satellite (1) with an antenna that generates a moving beam pattern on the surface of the earth. The beam pattern (3) is comprised of a plurality of sub-beams (4). A method of this invention determines an attitude correction signal for the satellite by the steps of: (a) providing at least one reference transmitter (10) at a known location on the surface of the earth; (b) transmitting at least one signal from the at least reference transmitter into at least one of the sub-beams; (c) receiving the at least one signal with the satellite antenna and transponding the received at least one signal to a ground station (8).Type: GrantFiled: August 23, 1995Date of Patent: May 26, 1998Assignee: Globalstar L.P.Inventor: Robert A. Wiedeman
-
Patent number: 5749545Abstract: An autonomous on-board satellite control system is to achieve autonomous orientation control and autonomous determination of the satellite's altitude and location in relation to the Earth's longitude and latitude grid. This is done with the aid of the following elements: an Earth sensor (1), a Pole-star sensor (2), a computer (4), a timing device (6) and actuator units (7). The system also includes a navigational star sensor (3) and a storage device (5), while the computer is designed so as to facilitate supplementary determinations. The orientation of the satellite is controlled by superimposing the general sensory plane (16) of the sensors (1 and 2) with the plane of the angle "center of Earth--satellite--Pole star" which defines latitude. The geovertical (11) rotates about a line to the Pole star (12) as the satellite (8) moves in its orbit (9) and this is equivalent to the revolution of the stars in the field of vision of the sensors (2 and 3).Type: GrantFiled: August 23, 1995Date of Patent: May 12, 1998Inventor: Sevastian Dmitrievich Gnatjuk
-
Patent number: 5738309Abstract: A method and system of orienting a payload of an orbiting spacecraft (14) to maintain a desired pointing profile in the presence of orbit inclination. A cone (12) is determined which is traced in inertial space by a pitch axis of the payload to maintain the desired pointing profile throughout an orbit. A bias momentum vector of the spacecraft (14) is oriented at an attitude which lies along the cone (12). The attitude has a nonzero angle with respect to a plane spanned by an orbit normal vector and an equatorial normal vector. The payload is rotated about a single body-fixed axis perpendicular to the pitch axis to align the pitch axis along the cone (12). As a result, the desired pointing profile is maintained throughout the inclined orbit.Type: GrantFiled: February 28, 1996Date of Patent: April 14, 1998Assignee: Hughes ElectronicsInventor: Richard A. Fowell
-
Patent number: 5716031Abstract: An artificial satellite equipped with roll, yaw and pitch rudders (4, 4', 5) to impose on it rotational forces around its three axes due to the resistance of the rarefied air which surrounds the satellite (compensation for the dynamic inertia of rotating objects onboard the satellite can also be ensured). This invention applies to low-orbit satellites.Type: GrantFiled: January 16, 1996Date of Patent: February 10, 1998Assignee: Centre National d'Etudes SpatialesInventor: Paul Duchon
-
Patent number: 5687084Abstract: A technique for maintaining a satellite in an assigned orbit without control or intervention from the ground. Autonomously obtained navigational data provide a measurement of the actual orbit in which the satellite is traveling. So long as the measured orbit conforms to a desired orbit to within a preselected tolerance, periodic corrections of equal magnitude are made to the satellite's velocity, based on a prediction of the effect of atmospheric drag on the orbit. Measurement of the orbit is made by observation of the time that the satellite passes a reference point in the orbit, such as by crossing the ascending node. If the measured orbit departs from the desired orbit by more than the preselected tolerance, a velocity correction of a magnitude different from the one based on prediction is applied to the satellite. For a decaying orbit, the magnitude of the velocity correction is increased above the correction value based on prediction.Type: GrantFiled: April 16, 1996Date of Patent: November 11, 1997Assignee: Microcosm, Inc.Inventor: James R. Wertz
-
Patent number: 5681011Abstract: The present invention comprises a method for injecting a heavier payload into orbit than is possible using a traditional method and the same rocket booster launch vehicle. The method of the present invention does not utilize a parking orbit and does not perform orbital injection at perigee or apogee of the desired orbit. Rather, in the present method the flight path angle is positive at the final lower stage booster burn so as to boost the perigee kick motor and payload into a suborbital trajectory having a low perigee, which may be below the surface of the Earth. In the preferred embodiment, the launch vehicle maintains a negative flight path angle during the PKM burn. The PKM burn does not occur at perigee, but perigee is at the desired location when the PKM burn is complete. Use of this method increases the payload capacity of some launch vehicles by up to fifty percent (50%).Type: GrantFiled: August 24, 1993Date of Patent: October 28, 1997Assignee: Orbital Sciences CorporationInventor: Scott R. Frazier
-
Patent number: 5667171Abstract: Methods and systems for stabilizing satellite spin about an intermediate inertia axis (Z) are disclosed. A set of gyros (22) sense the X component and the Y component of the angular velocity of the satellite body. A single degree of freedom momentum wheel (26) has a fixed transverse orientation with respect to the intermediate axis in order to store momentum. In one embodiment, the momentum wheel (26) is oriented to store momentum parallel to the Y axis. A tachometer (30) senses the rotation rate of the momentum wheel (26). A processor (24) forms a control signal representative of a control torque to be applied to the momentum wheel (26). The control torque is based upon the X component and the Y component of angular velocity of the satellite, and the angular velocity of the momentum wheel (26).Type: GrantFiled: April 28, 1995Date of Patent: September 16, 1997Assignee: Hughes Aircraft CompanyInventors: Richard A. Fowell, John F. Yocum
-
Patent number: 5655735Abstract: A low cost, fuel efficient solution to the problem of avoiding high yaw errors in communications satellites after station keeping maneuvers is achieved by adding a selectable function to the LTMM controller which allows it to calculate roll momentum based on measured yaw attitude error data from the yaw DIRA. The so calculated roll momentum is used to trigger roll unloads, which will reduce the yaw attitude error. This solution has the advantages that: it maintains yaw attitude error within the pointing budget; the fuel penalty is negligible; it can be made fully automatic; it can be disabled and not used; and, it requires only a very small addition to the firmware.Type: GrantFiled: July 3, 1995Date of Patent: August 12, 1997Assignee: Space Systems Loral, Inc.Inventors: David J. Wirthman, John S. Higham, Michel B. Baylocq, Peter Y. Chu
-
Patent number: 5651515Abstract: At the end of the operative life of the spacecraft, when all the propellant has been consumed for the normal attitude and orbit control operations, the arcjet thrusters are connected to a feed line for being fed with the helium still remaining in the pressurant tank and the arcjet thrusters are then fired to perform the complete re-orbiting maneuver until the spacecraft is placed in a graveyard orbit.Type: GrantFiled: January 30, 1995Date of Patent: July 29, 1997Assignee: Agence Spatiale EuropeenneInventors: Giorgio Saccoccia, Fabrizio Paganucci, Fabrizio Scortecci
-
Patent number: 5646847Abstract: A three-axis stabilized spacecraft includes a plurality of primary attitude control thrusters, the torque vectors of which lie in, or parallel to a primary plane. It also includes at least two more secondary attitude control thrusters, the torque vectors of which lie in a secondary plane which is not parallel to the primary plane. The control system produces attitude error signals, which are processed with a PID characteristic to produce impulse demand signals, all in known fashion. The impulse demand signals are transformed into an auxiliary coordinate system, in which two of the three auxiliary axes lie in the primary plane, and the third is orthogonal thereto. One of the secondary thrusters is selected, which has, along the third auxiliary axis, the largest torque magnitude and the same sign as the transformed impulse demand.Type: GrantFiled: August 25, 1995Date of Patent: July 8, 1997Assignee: Martin Marietta Corp.Inventors: Santosh Ratan, Neil Evan Goodzeit
-
Patent number: 5610820Abstract: A zero-momentum spacecraft's attitude is controlled by determining the torque required about a control axis to maintain the desired attitude, and, during each of recurrent control cycles, enabling a magnetic torquer if the torque demand exceeds a threshold. During each of the control cycles, thruster(s) are enabled to make up the difference between the torque demand and the estimated torque produced by the magnetic torquer. In determining the torque demand, the attitude rate signal is low-pass filtered to reduce noise, and the control loop bandwidth is maintained by totalling the estimated torque applied by the magnetic torquer and thrusters, integrating and high-pass filtering the estimated torque signals, and adding the filtered estimated torque with the filtered attitude rate signals to generate low-noise attitude rate signals. A three-axis system is described.Type: GrantFiled: March 23, 1995Date of Patent: March 11, 1997Assignee: Martin Marietta Corp.Inventors: Uday J. Shankar, Neil E. Goodzeit, George E. Schmidt, Jr.
-
Patent number: 5608634Abstract: A spacecraft is controlled by a composite attitude signal including two components with different passbands. In one embodiment, a spacecraft attitude control system includes an attitude sensor and a controller which provides a time derivative function. High frequency noise components of the sensed attitude signal are enhanced by the derivative, and tend to cause attitude jitter or excess power consumption. The jitter is reduced by low pass filtration of the sensor signal, but this undesirably reduces the high frequency response of the attitude sensor. The high frequency response is restored by high pass filter coupled to a reaction or momentum wheel tachometer, which produces a signal representative of the high frequency components of the body rate. A summing circuit couples together the filtered attitude sensor signals with the high frequency components of the wheel speed signal to produce a relatively noise free broadband body rate signal.Type: GrantFiled: June 12, 1992Date of Patent: March 4, 1997Assignee: Martin Marietta Corp.Inventors: Neil E. Goodzeit, Michael A. Paluszek
-
Patent number: 5597141Abstract: A dynamic balance mechanism for balance control of a gyro stabilized (spinning) satellite is disclosed. An elongated gear rack is attached to the spacecraft. A movable mass is mounted by guide rollers on the gear rack and translates along the gear rack according to requisite electronic commands. The movable mass includes a housing, a stepper motor, a rotary potentiometer, a cable reel, a pair of gear heads and a pinion gear. The pinion gear meshes with the rack and is driven by the stepper motor. The potentiometer measures the position of the movable mass on the gear rack. The cable reel saves space and minimizes harness jamming conditions. The invention secures better weight efficiency.Type: GrantFiled: July 25, 1994Date of Patent: January 28, 1997Assignee: Hughes Aircraft CompanyInventor: Allen G. Storaasli
-
Patent number: 5597142Abstract: A spacecraft orientation procedure, in accordance with a first embodiment of the invention, is practiced with a sun sensor to bring the x (roll) axis of the spacecraft parallel to a ray of the sun, and with a gyro sensor and an earth sensor of the spacecraft in conjunction with one instruction provided either autonomously or by a ground tracking station regarding an orientation of a spacecraft reference plane to enable locating the earth by the earth sensor. Furthermore, in accordance with a second embodiment of the invention, the orientation is established without aid from the ground tracking station by use of at least one telemetry and command antenna having a continuous field of view, as measured in one plane, which is greater than a semicircle. In the second embodiment, the orientation procedure provides for rotation of the spacecraft about the x axis for a scanning of the antenna to intercept command signals broadcast from the earth, thereby to locate the earth in a first reference plane.Type: GrantFiled: March 6, 1995Date of Patent: January 28, 1997Assignee: Space Systems/Loral, Inc.Inventors: Yat F. Leung, Scott W. Tilley
-
Patent number: 5597143Abstract: The invention provides an improved process and apparatus for controlling the attitude of a three-axis stabilized spacecraft. Separate dead zone elements are used to define threshold values for limiting nutation amplitude and deviation of the spin direction of the spacecraft from a desired orientation, and a continuous check is performed to determine whether the nutation amplitude and deviation of the spacecraft exceed the threshold values. A control intervention is initiated to reduce the nutation, even when the spin direction is within the range limited by the applicable threshold value.Type: GrantFiled: June 10, 1994Date of Patent: January 28, 1997Assignee: Deutsche Aerospace AGInventors: Michael Surauer, Helmut Bittner
-
Patent number: 5568904Abstract: A satellite perigee velocity augmentation method wherein a satellite is steered throughout the perigee velocity augmentation maneuver such that its thrust vector is always oriented approximately opposite to the direction of motion of the satellite. The satellite is pitched around a predetermined axis (e.g. the Y-axis) such that the thrust vector is always pointing approximately opposite to the direction of motion of the satellite. The satellite is steered during the perigee velocity augmentation maneuver in order to achieve the desired attitude wherein the thrust vector is always oriented approximately opposite to the direction of motion of the satellite is accomplished by pitching the satellite around the predetermined axis such that the thrust vector is always pointing approximately opposite to the direction of motion of the satellite. By steering the satellite in accordance with the present perigee velocity augmentation method, overall efficiency achieved is increased.Type: GrantFiled: August 28, 1992Date of Patent: October 29, 1996Assignee: Space Systems/Loral, Inc.Inventors: J. Kurt Brock, Darren R. Stratemeier, Eugene L. Williams
-
Patent number: 5563794Abstract: A control system (40) and procedure (50) that uses repetitive control to control repetitive error sources such as result from the phenomenon of thermal shock experienced by a spacecraft (10). Repetitive control is a "learning algorithm" that substantially eliminates errors in a stable control system (43) that performs highly repeatable tasks. The repetitive control system (40) and procedure (50) employs a repetitive signal generator (41) (G.sub.r), to store a sensed short-term error signal during each cycle and to process it to generate a signal that compensates for the error. The repetitive control system and procedure integrates (52) the error signal over several cycles, multiplies (53) the integrated value by a predetermined gain factor, and sums (54) it with the sensed error signal before the sensed error reaches the stable controller (43). An output filter (42), G.sub.Type: GrantFiled: November 22, 1993Date of Patent: October 8, 1996Assignee: Hughes Aircraft CompanyInventors: Christopher M. Cosner, Stuart F. Bockman
-
Patent number: 5557988Abstract: A centripetal impeller to impart a force and having an attachment piece defining an axis of revolution. The impeller has a movable arm pivotable relative to the attachment piece so that the movable arm is a distance r from the axis of revolution. The impeller includes a rotation device coupled to the attachment piece causing the attachment piece to rotate along the axis of revolution. A variable force device is attached to the movable arm to produce a force on the movable arm toward the axis of revolution so as to cause a net force along the axis of revolution.Type: GrantFiled: November 29, 1994Date of Patent: September 24, 1996Inventor: John C. Claxton
-
Patent number: 5556058Abstract: A method and system for determining the general three-axis attitude measurement for a spacecraft is disclosed. The system provides a cost-effective attitude measurement system for geosynchronous momentum bias spacecraft which need a continuously updated, moderately accurate, yaw measurement in addition to the usual roll and pitch measurements. The continuous yaw measurement enables control of the spacecraft yaw axis in the presence of disturbance torques. A sun sensor generates a signal which may be used to calibrate a continuously updated yaw measurement derived from the combination of signals generated by an Earth sensor and a space-to-ground link.Type: GrantFiled: May 16, 1994Date of Patent: September 17, 1996Assignee: Hughes ElectronicsInventor: Douglas J. Bender
-
Patent number: 5546309Abstract: An attitude sensing system utilizing simplified techniques and apparatus includes a Kalman filter which receives signals from an inertial measurement unit, a GPS receiver, and an integrated optical assembly. The output vector of the filter includes estimates of attitude misalignments and estimates of gyro drifts corresponding to the axes of the inertial measurement unit. The optical assembly includes a sensor array providing signals to the filter representing detection of the Earth's horizon or the center of the Sun. More particularly, a local vertical vector, computed from fore and aft detections of the Earth's horizon, is used in combination with GPS received signals to initially determine attitude by means of gyrocompassing. This attitude information is thereafter maintained by the inertial measurement unit and azimuth error resulting from drift of the inertial measurement unit and the initial gyrocompassing error is corrected by detections of the Earth's horizon and the Sun.Type: GrantFiled: October 20, 1993Date of Patent: August 13, 1996Assignee: The Charles Stark Draper Laboratory, Inc.Inventors: William M. Johnson, Howard Musoff