3568 results about "Spacecraft" patented technology

This invention is directed to a 3-dimensional imaging lidar, which utilizes modest power kHz rate lasers, array detectors, photon-counting multi-channel timing receivers, and dual wedge optical scanners with transmitter point-ahead correction to provide contiguous high spatial resolution mapping of surface features including ground, water, man-made objects, vegetation and submerged surfaces from an aircraft or a spacecraft.

The invention relates to a temperature measurement and calibration platform in space vacuum environment. The temperature measurement and calibration platform is favorable for realizing the simultaneous calibration of contact type temperature measurement and non-contact type temperature measurement, so the temperature measurement and calibration platform is served for heat vacuum and heat balance experiments of spacecrafts such as satellites, spaceship and the like. The temperature measurement and calibration platform comprises a constant temperature bath, wherein a double-sub-cavity vacuum cavity, the double-sub-cavity vacuum cavity comprises a first vacuum cavity body and a second vacuum cavity body, the first vacuum cavity body and the second vacuum cavity body are connected with a vacuum pumping device through a three-way valve, standard temperature indicator sensors are respectively arranged on the outer wall of the first vacuum cavity body and on the outer wall of the second vacuum cavity body, the standard temperature indicator sensors are connected with a temperature secondary meter, a laser light path reflecting device is arranged in the vacuum cavity of the first vacuum cavity body for calibrating a non-contact type temperature measuring system based on the tunable diode laser absorption spectrum technology, and the vacuum cavity of the second vacuum cavity body is used for accommodating a temperature sensor for calibrating a contact type temperature measuring system adopting the temperature sensor.

A system for capturing and docking an active craft to a passive craft has a first docking assembly on the active craft with a first contact member and a spike projecting outwardly, a second docking assembly on the passive craft having a second contact member and a flexible net deployed over a target area with an open mesh for capturing the end of the spike of the active craft, and a motorized net drive for reeling in the net and active craft to mate with the passive craft's docking assembly. The spike has extendable tabs to allow it to become engaged with the net. The net's center is coupled to a net spool for reeling in. An alignment funnel has inclined walls to guide the net and captured spike towards the net spool. The passive craft's docking assembly includes circumferentially spaced preload wedges which are driven to lock the wedges against the contact member of the active craft. The active craft's docking assembly includes a rotary table and drive for rotating it to a predetermined angular alignment position, and mating connectors are then engaged with each other. The system may be used for docking spacecraft in zero or low-gravity environments, as well as for docking underwater vehicles, docking of ancillary craft to a mother craft in subsonic flight, in-flight refueling systems, etc.

Methods and apparatus are disclosed for treating motion sickness. In a preferred embodiment a method of the invention comprises operating eyewear having shutter lenses to open said shutter lenses at a selected operating frequency ranging from within about 3 Hz to about 50 Hz. The shutter lenses are opened for a short duration at the selected operating frequency wherein the duration is selected to prevent retinal slip. The shutter lenses may be operated at a relatively slow frequency of about 4 Hz when the user is in passive activity such as riding in a boat or car or in limited motion situations in a spacecraft. The shutter lenses may be operated at faster frequencies related to motion of the user's head when the user is active.

A communications system, apparatus, method, and computer program product for inter-satellite and inter-spacecraft crosslinks (ISL) with non-ISL optimized antennas on spacecraft. The system includes a mobile communications platform that includes an ISL antenna configured to transmit information to a target satellite through a non-ISL antenna of the target satellite. The mobile communications platform is configured to relay transmissions through the non-ISL antenna of the target satellite to another communications platform. The mobile communications platform includes a controller configured to determine a location of the mobile platform; determine whether the target satellite is within communications range; and prepare a signal for relayed transmissions through a non-ISL antenna of the target satellite to another communications platform in a signal format that is decipherable by this other communications platform.

A ceramic composite tile armor which is reinforced at the more vulnerable joint and free edge areas, using glass or ceramic strips or overlays bonded with an adhesive to the outer surface of the tile joints and free edges. This reinforcement provides improved ballistic threat protection for ground vehicle, aircraft, watercraft, spacecraft, and body (personnel) ceramic tile armor applications. Glass or ceramic overlay strips assist in fracturing impacting projectiles that strike the tile joints or free edges. The substrate laminate backing can then capture fragments of the projectile and broken ceramic and prevent penetration. The invention provides improved protection over conventional joint and edge enhancements with higher reliability of accurate positioning over joint and free-edge areas, with less added weight, and at lower associated production costs.

The invention relates to a spacecraft assembly simulation technique-based virtual assembly system. The virtual assembly system comprises a CAD modeling module, a virtual assembly planning module and an assembly process design module, wherein the CAD modeling module is used for designing components and tooling tools, and assembling the components together by defining a series of matching constraint relations so as to obtain an assembling model of products; the virtual assembly planning module is used for establishing a geometrical constraint-based virtual environment, planning an optimized assembly sequence according to disassembly directions, disassembly tools and precedence constraint information recorded in the virtual environment for execution, verifying the optimized assembly sequence, and performing simulation evaluation in the virtual environment; and the assembly process design module is used for designing a product structure tree and a process step catalog, editing process steps and assembling a technological process. The CAD modeling module is connected with the virtual assembly planning module through a CAD modeling interface so that models generated in the CAD modeling module are guided into the virtual assembly planning module; and the process planning result in the virtual assembly planning module is guided into the assembly process design module through a Mockup interface.

The invention relates to an autonomously identifying and capturing method of a non-cooperative target of a space robot, comprising the main steps of (1) pose measurement based on stereoscopic vision, (2) autonomous path planning of the target capture of the space robot and (3) coordinative control of a space robot system, and the like. The pose measurement based on the stereoscopic vision is realized by processing images of a left camera and a right camera in real time, and computing the pose of a non-cooperative target star relative to a base and a tail end, wherein the processing comprises smoothing filtering, edge detection, linear extraction, and the like. The autonomous path planning of the target capture of the space robot comprises is realized by planning the motion tracks of joints in real time according to the pose measurement results. The coordinative control of the space robot system is realized by coordinately controlling mechanical arms and the base to realize the optimal control property of the whole system. In the autonomously identifying and capturing method, a self part of a spacecraft is directly used as an identifying and capturing object without installing a marker or a comer reflector on the target star or knowing the geometric dimension of the object, and the planned path can effectively avoid the singular point of dynamics and kinematics.

The invention provides a fault predicting and diagnosing method suitable for a dynamic complex system. The method can be applied in the field of fault prediction and diagnosis of dynamic complex systems of spacecrafts and the like. The method comprises the following steps of: performing failure mode and effect analysis (FMEA) on the dynamic complex system to obtain a main fault mode and corresponding performance detection parameters, dividing the performance detection parameters into slowly variable data and fast variable data, pre-processing the performance detection parameters, establishingan autoregressive moving average model (ARMA) aiming at the slowly variable data to perform time sequence prediction, establishing a multi-resolution wavelet neural network aiming at the fast variable data to perform time sequence prediction, performing fault early warning on the time sequence prediction results by establishing a prediction interval model, and performing fault diagnosis by establishing a D-S (Dempster-Shafer) evidence theory-based multi-signal fusion model. The method can be used for predicting and diagnosing the faults of the dynamic complex system with high precision, and has strong universality.

A method for implementing a modular platform for the construction of satellites and other spacecraft based on modular platform architecture, the method comprising: (a) identifying a plurality of functional elements and their associated functional routines that may be operable within at least one satellite; (b) associating the functional routines with one another in a strategic manner; (c) dividing the functional routines to define a plurality of subsystems; and (d) deriving a plurality of modules from the plurality of subsystems, each of the modules being configured to operably interface with at least one other module to construct a working satellite capable of carrying out a pre-determined number of the functional routines.

Various embodiments relate to systems and methods for simultaneously switching input image streams to output devices, while providing optional image processing functions on the image streams. Certain embodiments may provide vision systems and methods suitable for use in vehicles, particularly windowless vehicles, such as armored ground vehicles, submerged watercraft, and spacecraft. Some embodiments may enable sharing of image streams (e.g., with one or more other vehicles), generation of panoramic views (e.g., from various camera feeds), intelligent encoding of image streams, and implementation of security features based on image streams.

An optical scanner system for contiguous three-dimensional topographic or volumetric imaging of a surface from an aircraft or spacecraft is disclosed. A servo controller synchronizes the rotation rates of a pair of wedge scanners with high precision to the multi-kilohertz laser fire rate producing an infinite variety of well-controlled scan patterns. This causes the beam pattern to be laid down in precisely the same way on each scan cycle, eliminating the need to record the orientations of the wedges accurately on every laser fire, thereby reducing ancillary data storage or transmission requirements by two to three orders of magnitude and greatly simplifying data preprocessing and analysis. The described system also uses a holographic element to split the laser beam into an array that is then scanned in an arbitrary pattern. This provides more uniform signal strength to the various imaging detector channels and reduces the level of optical crosstalk between channels, resulting in a higher fidelity three-dimensional image.

Automatic homing systems are disclosed, operable between pairs of objects. One or both of the objects in the pair may be moving and/or unmanned. Examples of a pair of objects between which the inventive automatic homing systems may be operated are ground vehicles in leader-follower configuration, as well as aircraft, spacecraft, watercraft. The automatic homing systems are based on a relatively simple, elegant concept of rightness/leftness and a line-of-sight link between a respective leader vehicle and follower vehicle. Complex EO/IR cameras, LIDAR, RADAR, and/or SONAR-based sensor systems can be avoided. Convoys of vehicles in leader-follower configuration advantageously may be operated without needing human operators in the follower vehicles.

The invention provides a joint measurement method based on a laser radar and a binocular visible light camera to obtain accurate and dense three-dimensional information simply and efficiently. The joint measurement method comprises assuming that the laser radar is a camera device with a fixed internal reference; directly projecting three-dimensional point cloud data into a two-dimensional image, calculating rotation and translation relationships between the laser radar device and the binocular camera by using an image processing method and the matching between the two-dimensional images; creatively introducing an idea for obtaining a matrix norm and a matrix trace to solve a rotation matrix; finally fusing laser radar and binocular stereo vision point cloud data. The method not only can obtain an accurate position and attitude information, but also can reconstruct the special texture and feature information of a target surface, which has a high application value for spacecraft dockingand hostile satellite capture in the military field and workpiece measurement and unmanned driving in the civilian field.

The invention relates to SINS/GPS magnetic compass integrated navigation system data fusion method. It includes the following steps: processing position speed error combination for the SINS and GPS; estimating carrier position, speed, and altitude error by the first Kalman filtering to feed back and rectify SINS; calculating direction cosine matrix from carrier coordinate system to level coordinate system by outputted roll angle and pitch angle to assist magnetic compass to gain carrier magnetic course angle, and course angle; processing course angle error combination for the SINS and magnetic compass; estimating carrier altitude error by the second Kalman filtering to feed back and rectify SINS; outputting carrier position, speed, and altitude information from SINS to user. The invention has the advantages of simple computing, high precision, strong fault tolerant ability and wide application. Thus it can be used to increase magnetic compass altitude fixing precision, navigation precision of the SINS/GPS magnetic compass integrated navigation system used in space craft, aircraft, shipping, or surface car.

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.

A spaced based local area network system for providing dynamically allocated downlink services for spacecraft in Geosynchronous Earth Orbit. The system includes at least two client satellites which host an inter-satellite communications payload; a hub satellite which hosts an inter-satellite communications payload, a downlink communications payload, and an aggregator payload; and a ground station which transmits data to and receives data from the hub satellite. The hub satellite, using the aggregator payload, aggregates at least two data flows received from the at least two client satellites over an inter-satellite link, by the inter-satellite communications payload, into a data stream that is downlinked to the ground station, using the downlink communications payload, over a space-to-ground link.

The invention relates to a spacecraft attitude integral sliding mode fault tolerance control method taking consideration of a performer fault and provides a robustness attitude active fault tolerance control method based on an integral sliding mode surface for problems of the performer fault, external disturbance and control moment amplitude limits in a spacecraft attitude control process. The method comprises steps that firstly, a spacecraft attitude dynamics model taking consideration of the performer fault and containing external disturbance is established; secondly, on the condition that a performer is not in fault, a designed nominal controller can guarantee system stability, and input saturation amplitude limits can be easily satisfied through adjusting controller parameters; lastly, the fault information is introduced to design an integral sliding mode controller, robustness of external disturbance and the performer fault can be effectively improved, system stability is analyzed on the basis of an Lyapunov method. The method is advantaged in that stability of the attitude control system is guaranteed when a spacecraft operating on orbit generates the performer fault, and relatively strong fault tolerance capability and external disturbance robustness are realized.

A satellite or spacecraft in low Earth orbit, when in eclipse and not illuminated by sunlight, represents a low-bandwidth datastream through modulation of a source of uncollimated light, such as a set of light emitting diodes. Transmission modes include generating patterns in the color, intensity or polarization of a light source, or precise timing control of a strobe signal. A ground station tracks the satellite in a telescope, stewing the telescope to follow the satellite as it moves across the sky, recording light generated by the satellite with a light sensor such as a video camera. The datastream is regenerated through analysis of recorded video. Satellite downlink is initiated by detection of eclipse, by command via radio uplink, or in response to a periodic automatic timer. In an alternative embodiment, the satellite represents the datastream through modulation of its effective albedo or reflective flux during periods of solar illumination.

The invention discloses a small-size spacecraft butt-joint mechanism, relates to a spacecraft butt-joint mechanism and aims at solving the problems that an existing small-size spacecraft butt-joint mechanism is fierce in collision, low in positioning precision, and large in control difficulty in the butt-joint process and unstable enough after butt-joint. A connecting sleeve and three arc-shaped enclosing plates are arranged between an upper driven disc and a lower driven disc; a first guide plate is arranged at one end of each arc-shaped enclosing plate; a second guide plate is arranged at the other end of each arc-shaped enclosing plate; the adjacent first guide plates and three second guide plates form a V-shaped positioning slot; three brackets are uniformly distributed between the upper drive disc and the lower drive disc along the same circumference; a stepped motor is fixed in the center of the lower drive disc; the stepped motor is connected with a lead screw; a threaded lifting disc is screwed on the lead screw; the lower end of each locking claw penetrates through a slide path and is hinged with the threaded lifting disc, and each locking claw is contacted with the surface of a roller; the upper end of each locking claw is arranged in the corresponding V-shaped positioning slot. The small-size spacecraft butt-joint mechanism disclosed by the invention is mainly used for capturing and butt-joint of a small-size spacecraft.

A system and method are disclosed for computing a vehicle's motion in four dimensions (e.g., three spatial dimensions and time) and reliably predicting the vehicle's arrival time at a predetermined location, by providing a graphical display to an operator of the vehicle's progress that enables the operator to adjust the vehicle's movement and achieve the desired arrival time. Specifically, a system and method are disclosed for computing the movement of an aircraft in four dimensions, predicting its arrival time at a predetermined waypoint, and displaying (in a highly intuitive format) the aircraft's progress in achieving that desired arrival time. The pilot can then adjust the movement (e.g., speed) of the aircraft in accordance with the parameter(s) displayed, in order to achieve the desired arrival time. Thus, for example, numerous aircraft could be scheduled to arrive at a specific final approach waypoint at a predetermined rate (e.g., one aircraft per minute), which would enable the traffic controllers to optimize runway traffic without having to stack the aircraft in holding patterns and thereby waste fuel. Notably, although an example of an aircraft navigation and control system and method is disclosed, the system and method can be implemented for any type of vehicle (e.g., aircraft, spacecraft, ship, submarine, bus, train, automobile, etc.) whose operator desires to reach a particular location at a specified time.

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.

The invention discloses a novel aircraft main bearing structure, comprising a pushing cabin structure, a star-shaped truss structure (6) and an electronic cabin main framework structure; the upper part of the propelling cabin structure is connected with the electronic cabin main framework and the lower part thereof is taken as a whole star precision detection benchmark surface and an installation surface used for the propelling cabin instruments; the star-shaped truss structure (6) is a load installation frame supported by a supporting rod and arranged in the electronic cabin main framework structure and used for installing effective loads; the lower end of the supporting rod is fixedly connected with the upper part of the propelling cabin structure; as the main bearing structure adopts an external framework side plate structure and an internal start-shaped truss structure to form a mixed main structure with dual bearing paths (internal bearing path and external bearing path) so as to transmit the whole star load, the force transmission path is short and continuous, the structure is compact, the center of mass is low, and the main bearing structure meets the requirement of the whole satellite of high rigidity and small inertia.

The invention relates to a flexible satellite locus linearization attitude control method based on a disturbance observer. The invention aims at solving problems that a single locus linearization control method is poor in capability of inhibiting interference, is poorer in robustness, and does not consider external interference and the impact from flexible accessories. The method comprises the steps: employing Euler angles for describing attitudes of a spacecraft, employing an idea of equivalent disturbance, and building a flexible spacecraft dynamics and kinetics equation; solving the pseudo-inverse of a controlled object under the condition of neglecting equivalent disturbance, designing a quasi-differentiator of a specific type, and obtaining the nominal control of an expected locus; and designing a linear time varying adjuster through proportion-integration control. The method gives consideration to the influence of equivalent disturbance, designs the disturbance observer, and guarantees the asymptotic convergence of a tracking error of a flexible spacecraft. The method improves the anti-interference capability of a system, and improves the robustness of the system. The method is used in the attitude control field of flexible satellites.

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.

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.

Multifunctional structures and methods of manufacturing multifunctional structures which function as both electronic devices and load-bearing elements are disclosed. The load-bearing elements are designed to have electronic functionality using electronics designed to be load-bearing. The method of manufacturing the multifunctional structure comprises forming an electronic element directly on at least one ply of arbitrarily shaped load-bearing material using conventional lithographic techniques and/or direct write fabrication techniques, and assembling at least two plies of arbitrarily shaped load-bearing material into a multifunctional structure. The multifunctional structure may be part of an aerospace structure, part of a land vehicle, pan of a watercraft or part of a spacecraft.

The invention relates to an autonomous integrated navigation system which belongs to the technical field of navigation systems. The SINS (Strapdown Inertial Navigation System)/SAR (Synthetic Aperture Radar)/CNS (Celestial Navigation System) integrated navigation system takes SINS as a main navigation system and SAR and CNS as aided navigation systems and is established by the following steps: firstly, designing SINS/SAR and SINS/CNS navigation sub-filters, calculating to obtain two groups of local optimal estimation values and local optimal error covariance matrixes of the integrated navigation system state, then transmitting the two groups of local optimal estimation values into a main filter by a federal filter technology for fusion to obtain an overall optimal estimation value and an overall optimal error covariance matrix, and finally, performing real-time correction on the error according to the overall optimal estimation value so as to obtain an optimal estimation fusion algorithm of the SINS/SAR/CNS integrated navigation system. The autonomous integrated navigation system, disclosed by the invention, is less in calculation amount and high in reliability, is applicable to aircrafts in near space, aircrafts flying back and forth in the aerospace, aircrafts for carrying ballistic missiles, orbit spacecrafts and the like, and has wide application prospect.

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.