1293results about "Spacecraft guiding apparatus" patented technology

A GPS positioning system that has: at least one mobile station to measure a positional coordinate by receiving radio wave from a satellite; a plurality of fixed stations that have predetermined positional coordinates and receive radio wave from the satellite; and arithmetic processing means, which is connected with at least one mobile station and a plurality of fixed stations via communication, and that transmits correction data suitable for the mobile station to the mobile station based on positioning data transmitted from the mobile station.

A system and method which partitions a parameter estimation model, a fault detection model, and a fault classification model for a process surveillance scheme into two or more coordinated submodels together providing improved diagnostic decision making for at least one determined operating mode of an asset.

A closed-loop LRF pointing technology to measure the range of a target satellite from a chaser satellite for rendezvous is provided that includes several component technologies: LOS angle measurements of the target satellite on a visible sensor focal plane and the angles' relationships with the relative position of the target in inertial or LVLH frame, a relative navigation Kalman filter, attitude determination of the visible sensor with gyros, star trackers and a Kalman filter, pointing and rate commands for tracking the target, and an attitude controller. An analytical, steady-state, three-axis, six-state Kalman filter is provided for attitude determination. The system and its component technologies provide improved functionality and precision for relative navigation, attitude determination, pointing, and tracking for rendezvous. Kalman filters are designed specifically for the architecture of the closed-loop system to allow for pointing the laser rangefinder to a target even if a visible sensor, a laser rangefinder, gyros and a star tracker are misaligned and the LOS angle measurements from the visible sensor are interrupted.

An autonomous unmanned space flight system and planetary lander executes a discrete landing sequence including performing an initial velocity braking maneuver to remove velocity at altitude, coasting during which the planet surface is imaged and correlated to reference maps to estimate cross-track and along-track navigation errors and one or more lateral braking maneuvers are performed to reduce cross-track navigation error, and performing a terminal velocity braking maneuver(s) to reduce the along-track braking maneuver and remove the remainder of the velocity just prior to landing. A bi-propellant propulsion system provides a very high T/M ratio, at least 15:1 per nozzle. Short, high T/M divert maneuvers provide the capability to remove cross-track navigation error efficiently up to the maximum resolution of the reference maps. Short, high T/M terminal velocity braking maneuver(s) provide the capability to remove along-track navigation error to a similar resolution and remove the remaining velocity in a very short time window, approximately 3-15 seconds prior to touchdown. The propulsive efficiency frees up mass which can be allocated to a fuel to remove the unknown navigation errors, perform hazard avoidance and/or relocate the lander by flying it to another site or be allocated to additional payload.

A method and apparatus for managing separation between vehicles. A closest point of approach between a first vehicle traveling along a first path and a second vehicle traveling along a second path is predicted. A number of compensation commands for altering the first path of the first vehicle are generated using the closest point of approach and a desired level of separation between the first vehicle and the second vehicle. The number of compensation commands is integrated with a number of control commands for the first vehicle to form a final number of control commands configured to maneuver the first vehicle to substantially maintain the desired level of separation between the first vehicle and the second vehicle. A response of the first vehicle to the final number of control commands is a desired response.

An apparatus for lowering an orbit of a space object includes an envelope, an inflation system for inflating the envelope, an inflation control system for controlling the inflation system, and attachment hardware for connecting the apparatus and the space object. Inflating the envelope increases an effective drag area of the envelope for increasing atmospheric drag on the envelope.

The invention discloses an active control device of vibration simulation for bending and torsional modes of space sailboards and a method. The device symmetrically sticks multi-chip bending mode piezoelectric actuators on front and rear faces of a flexible plate, a bending mode sensor is arranged at the position horizontally closing to a fixed end by 320-25mm of the flexible plate and on the longitudinal center line of the flexible plate, torsional mode actuators composed of a plurality of piezoelectric ceramic chips are antisymmetrically stuck on both faces of the flexible plate, and torsional mode sensors composed of a plurality of the piezoelectric ceramic chips are stuck on both faces of the flexible plate. The method operates an active vibration control strategy according to information of the bending mode and trosional mode sensed by a piezoelectric sensor, and then drives a voltage amplifier through piezo-electricity to respectively drive the piezoelectric actuators for controlling multimode vibrations of bending and torsion, thereby realizing the purpose of actively inhibiting the vibrations. The invention employs the optimum distribution of piezoelectric sensor chips and driving chips, and realizes decoupling of bending modes and torsional modes of flexible cantilever plates in inspection and drive control.

The present application discloses an open system architecture and software system for plug and play modular mission payloads in aerial vehicles. The improved software moves the control function of mission payloads away from the ground station and into the aerial vehicle. The plug & play web-based payload interface software resides in a payload interface controller in the vehicle, and this is networked via a uniform resource locator (URL) address to a ground control station. Consequently, when new payloads are added to the system, integration issues and costs are minimized.

A method and apparatus for deploying a group of panels. An apparatus comprises a group of panels in a folded configuration against a side of a spacecraft, a group of flexible members connected to the group of panels, and an interface system associated with the group of panels and the group of flexible members. The interface system is configured to move the group of panels from the folded configuration to a deployed configuration when the group of flexible members is extended from the spacecraft.

The invention discloses a satellite posture all-round controlling method based on a magnetic moment device and a flywheel, relating to an all-round posture controlling method for completing a satellite orbit-injection phase by using the magnetic moment device and the flywheel. The invention solves the problems of low reliability and short service life of the traditional satellite posture all-round controlling technology. The satellite posture all-round controlling method comprises the following steps of: 1, setting controller parameters according to the requirement of a control system; 2, measuring a geomagnetic field intensity vector Bb, a satellite angular velocity vector Wb and a solar azimuth, and sending the measured data to a satellite controller; 3, calculating an expected control moment vector Tm and a control magnetic moment vector Mm, and sending the control magnetic moment vector Mm to the magnetic moment device; 4, acquiring an effective solar azimuth vector Alfa; 5, calculating a control input moment vector Tw and sending to the flywheel; and 6, jointly completing the satellite posture all-round control by the magnetic moment device according to the control magnetic moment vector Mm and the flywheel according to the control input moment vector Tw. The invention is suitable for the field of satellite posture control.

A routing tool is disclosed. In one embodiment, the method and system include receiving flight data and geographic data in the aircraft, and generating route data based on the flight data and the geographic data. The route data provides information about attainable landing areas for the aircraft.

An omni-directional aircraft with flight capabilities surpassing those of a regular VTOL or helicopter, being able to take full advantage of the simultaneous six degrees of freedom of motion possible in the atmosphere, undergoing any desired combination of translational and rotational movement, and keeping station in the air in any arbitrary attitude. In the preferred embodiment, the flying object comprises a cubic body on the six faces of which are mounted six propulsion units, such that the propellers on each pair of opposite faces are coplanar with each other and with the main axis passing through the centers of these opposite faces, and their thrusts act along the direction of another main axis, each of the three pairs of propellers acting along a different one of the cube's three main axes. The thrust from each motion-inducing assembly being continuously variable and reversible, the resultant translational and rotational thrust vectors can be positioned arbitrarily within their respective solid envelopes. A control element with equal freedom of motion allows intuitive piloting of the vehicle. In an alternate embodiment, the propulsion units are disposed along the sides of a tetrahedron.

Various embodiments of the present invention include an attitude control system for use with small satellites. According to various embodiments, the system allows rapid retargeting (e.g., high slew rates) and full three-axis attitude control of small satellites using a compact actuation system. In certain embodiments, the compact actuation system includes a plurality of single-gimbaled control moment gyroscopes (SGCMG) arranged in a pyramidal configuration that are disposed within a small satellite.

An automatic celestial navigation system for navigating both night and day by observation of K-band or H-band infrared light from multiple stars. In a first set of preferred embodiments three relatively large aperture telescopes are rigidly mounted on a movable platform such as a ship or airplane with each telescope being directed at a substantially different portion of sky. Embodiments in this first set tend to be relatively large and heavy, such as about one cubic meter and about 60 pounds. In a second set of preferred embodiments one or more smaller aperture telescopes are pivotably mounted on a movable platform such as a ship, airplane or missile so that the telescope or telescopes can be pivoted to point toward specific regions of the sky. Embodiments of this second set are mechanically more complicated than those of the first set, but are much smaller and lighter and are especially useful for guidance of aircraft and missiles. Telescope optics focus (on to a pixel array of a sensor) H-band or K-band light from one or more stars in the field of view of each telescope. Each system also includes an inclinometer, an accurate timing device and a computer processor having access to catalogued infrared star charts. The processor for each system is programmed with special algorithms to use image data from the infrared sensors, inclination information from the inclinometer, time information from the timing device and the catalogued star charts information to determine positions of the platform. Direction information from two stars is needed for locating the platform with respect to the celestial sphere. The computer is also preferably programmed to use this celestial position information to calculate latitude and longitude which may be displayed on a display device such as a monitor or used by a guidance control system. These embodiments are jam proof and insensitive to radio frequency interference. These systems provide efficient alternatives to GPS when GPS is unavailable and can be used for periodic augmentation of inertial navigation systems.

A sub-orbital, high altitude communications system comprising at least two ground stations and at least one high altitude relay station. Each of the ground stations including means for sending and receiving telecommunications signals. The relay stations include means for receiving and sending telecommunications signals from and to said ground stations and from and to other relay stations. Means are provided for controlling the lateral and vertical movement of the relay stations so that a predetermined altitude and location of each of said relay stations can be achieved and maintained Means are provided for receiving the relay stations so that they can be serviced for reuse.

Apparatus and methods for performing satellite proximity operations such as inspection, recovery and life extension of a target satellite through operation of a “Satellite Inspection Recovery and Extension” (“SIRE”) spacecraft which can be operated in the following modes (teleoperated, automatic, and autonomous). The SIRE concept further consists of those methods and techniques used to perform certain (on-orbit) operations including, but not limited to, the inspection, servicing, recovery, and lifetime extension of satellites, spacecraft, space systems, space platforms, and other vehicles and objects in space, collectively defined as “target satellites”. The three basic types of SIRE proximity missions are defined as “Lifetime Extension”, “Recovery”, and “Utility”. A remote cockpit system is provided to permit human control of the SIRE spacecraft during proximity operations.

A microvalve and microthruster are provided. Both include a housing with an interior in which an armature can travel and abut a valve seat to provide an electromagnetic valve. The armature acts as a valve body that is maintained in a or normally closed position by a permanent magnet, and is opened by interaction with a selectively activatable electromagnet. A microthruster is provided by fashioning the valve discharge as a nozzle, preferably of a material that is permanently and selectively magnetizable, so that the nozzle functions as the permanent magnet and which magnetism can be adjusted to assure proper opening and closing of the valve portion. The valve body is guided without any sliding fit mechanism and is suitable for controlling gas or liquid fluid flows.

The invention relates to a different-surface crossover quick-change track fixed time stable posture pointing direction tracking control method. The different-surface crossover quick-change track fixed time stable posture pointing direction tracking control method aims at solving the problems that the uncertainty of inertia of a spacecraft is not considered in the prior art, the convergence time can not be freely adjusted depending on the state initial value, and compensating moment generated in the singular direction of a flywheel needs to be designed artificially. The method comprises the particular steps that 1, a tracking satellite and a target satellite are supposed to be located on a different-surface crossover track, and the expected posture needs to be determined; 2, an expected posture tracking control rule is designed; 3, buffeting of the expected posture tracking control rule is eliminated; 4, the expected posture of the crossed points of the tracking satellite and the target satellite changes along with distance between the crossed points of the tracking satellite and the target satellite, a configuration scheme of an execution mechanism is determined according to the expected posture tracking control rule, and the expected posture control torque is solved. The different-surface crossover quick-change track fixed time stable posture pointing direction tracking control method is applied to the field of satellite control.

A practical orbit raising method and system wherein a satellite quickly escapes the Van Allen radiation belts and payload mass and mission life are maximized. A satellite is launched that contains high thrust chemical propulsion thrusters, high specific impulse electric propulsion thrusters and a solar array. The satellite quickly escapes the Van Allen radiation belts by firing the high thrust chemical propulsion thrusters at apogees of intermediate orbits, starting from the transfer orbit initiated by a launch vehicle, to successively raise the perigees until the perigee clears the Van Allen radiation belts. The payload mass and mission life are maximized by firing high specific impulse electric propulsion thrusters to raise the satellite to near synchronous orbit, while steering the thrust vector and solar array to maintain the sun's illumination on the solar array. The chemical and/or electric propulsion thrusters are then fired to achieve geosynchronous orbit.

An automated docking system for space vehicles that includes a plurality of antennas on each of a target vehicle and a chase vehicle. A pseudo random code is transmitted via one of the antennas located on one of the vehicles and received by one of the antennas located on the other vehicle. The pseudo random code is then sent back to the original vehicle via transmissions from the plurality of antennas on the second vehicle and received by the plurality of antennas on the original vehicle. The distance from each of the antennas on the other vehicle to each of the antennas on the originating vehicle can be measured in this fashion. The antennas on each of the vehicles are located in a spaced-apart arrangement so that the angular orientation or attitude of each of the vehicles to each other can also be determined. A plurality of video cameras is provided on the exterior of one of the vehicles and video information from these cameras is transmitted to the other vehicle for display to operators in that vehicle. Commands between the vehicles can also be communicated.

Methods and apparatus are provided for a traffic warning system (TWS) for light aircraft. The TWS comprises a processor coupled to a transceiver, adapted to measure signal strength and send/receive messages containing station ID and preferably altitude and position data. Memory, display and various flight data instruments, such as GPS, altimeter, etc., are also coupled to the processor. The transceiver-processor automatically identifies TWS equipped aircraft within range using an ad-hoc network and exchanges ID and position information. The processor determines range from signal strength and/or received position information and, given enough data, determines direction, altitude, speed, etc., of the other aircraft, which it presents to the pilot. These values and their rate of change are compared to stored alarm thresholds, and the pilot is warned when another aircraft triggers the threshold. Evasive action is recommended where possible.

A system for, and method of recovering a solar-powered spacecraft from an anomaly that renders the attitude of the spacecraft unknown includes maintaining a power-safe attitude by switching between two orthogonal axes using solar panel current sensors. The system and method may also include simultaneously determining spacecraft attitude using a star sensor. The system is applicable to spacecraft operating in a solar wing-stowed configuration.

The invention relates to a method for controlling a rigid spacecraft for target attitude tracking, and belongs to the technical field of the high-precision and high-stability attitude tracking control of spacecrafts. The method solves the problem that when the attitude tracking spacecraft runs in a low orbit in outer space, the conventional control method cannot eliminate the inherent flutter of a sliding mode variable structure. The method comprises the following steps: 1, establishing a kinetic model and a kinematic model of the rigid spacecraft; 2, setting an attitude tracking error and anexpected attitude parameter of the rigid spacecraft, and combining the attitude tracking error and the expectation attitude parameter with the kinetic model and the kinematic model to establish a mathematical model for the attitude tracking; 3, adopting a control algorithm of a sliding mode variable structure controller to adjust a control law of the mathematical model which is established in thestep 2 and is used for the attitude tracking, and simultaneously combining an observation result of a disturbance observer to modify the control law; and 4, controlling the rigid spacecraft by using the modified control law obtained in the step 3 to realize the attitude tracking. The method is suitable for the attitude tracking of targets running in the outer space.

A method for controlling an actuator of a vehicle comprises providing a dynamic condition sensor generating a vehicle movement signal and a position sensor for generating a reported position. A processor is coupled to the inertial sensor and the position sensor and comprises an estimator, a position measurement predictor having a filter, a comparator and a control shaping block, said estimator generating a vehicle position based upon the dynamic condition sensor, said position measurement predictor generating an estimated position measurement in response to the reported vehicle position and a matched frequency response to the movement signal, said control shaping block generating an actuator control signal in response to a comparison of the estimated position measurement and the reported vehicle position.

An iterative method enabling a request plan to be established for an observation satellite. A plan consists in a succession of requests which are associated with pluralities of opportunities for satisfying said requests. The plan must also comply with a plurality of constraints. Each iteration k of the iterative method is made up of the following steps:new opportunity is selected;a provisional plan is derived from the preceding plan k-1 as calculated in-the preceding iteration, and from the new opportunity;provisional plan is verified for compliance with said plurality of constraints;the quality of said provisional plan is evaluated; andit is determined whether the provisional plan should be confirmed as plan k as a function of the quality of said provisional plan and of the quality of said preceding plan k-1.In the method, it is decided whether or not to confirm the provisional plan by a probabilistic metaheuristic of the simulated annealing type, and by the rule for constructing the provisional plan as a function of the newly selected opportunity.

The invention relates to a tracking control method and especially relates to a distributed limited time tracking control method for formation-flying satellites. The invention aims at solving problems that the communication among formation-flying satellites is limited in a conventional formation-flying satellite control method, the control input upper limit information of a pilotage satellite is difficult to obtain, the inevitable perturbation of all satellites in a system cannot be considered, and the formation is longer in time. The method comprises the following steps: 1, building a double-satellite relative movement dynamic model; 2, building a relative dynamic model of the formation-flying satellites relative to a reference point; 3, designing a distributed limited time tracking control rule. The method solves a problem that the communication among formation-flying satellites is limited in the conventional satellite formation control method, considers the inevitable perturbation of all satellites in the system, and can complete the formation in a shorter time. The method can be used in the field of tracking control.

The invention discloses flywheel based attitude maneuvering control device and method for successive approaching of a satellite rounding an instantaneous Euler shaft and relates to control device and method for satellite attitude adjustment. The invention is provided for solving the problems of great fuel consumption, short service life of a satellite, complex configuration of an air injecting control system and difficult reduction of size and weight of the satellite existing in the realization of satellite wide-angle attitude maneuvering by adopting air injecting control. The method comprises the following steps of: setting a parameter of the control device according to the system requirement of the control device and obtaining attitude deviation angular velocity according to a motion equation; and expressing a relationship of the instantaneous Euler shaft and a deviation angle of the current attitude and a target attitude of the satellite by an attitude error quaternion to obtain a control signal, calculating to obtain a flywheel control input moment vector calculated by a satellite controller, to be used as a data control command as a basis for generating moment by a back action flywheel. The invention does not consume other resources on the satellite or consume fuel, prolongs the service life of the satellite and can be widely suitable for various satellites needing attitude maneuvering.

Provided are an automatic operation system for automating satellite control operation, a method thereof, and an automatic satellite ground control system using the same. The object of the present research is to provide an automatic operation system that can reduce operation manpower and cost by automatically operating an entire satellite ground control system, which requires a plurality of operators to be properly operated, all the times, a method thereof, and an automatic satellite ground control system using the same. The satellite ground control system of the present research includes: a mission timeline receiving unit, an operation procedure editing unit, an operation procedure analyzing and code transforming unit, an operation procedure executing unit, a subsystem process state monitoring unit, and an access managing unit for providing interface with each of the subsystems.

Boeing achieves satellite diversity by having a large discrimination angle for a MEO constellation of communications satellites to limit interference in the Ku-band with GSO communications systems. Each satellite entering an exclusion zone over a GSO ground station terminates all transmissions to provide EPFD within acceptable limits. The Boeing MEO constellation preferably comprises 20 satellites, 5 in each orbit. The four orbits are inclined at about 57° with respect to the equator. Services include an Integrated Digital Service (IDS) and Backhaul Data Service (BDS) to accommodate the needs of different users.

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.