3022results about "Cosmonautic parts" patented technology

A solar cell regulator in a nanosatellite includes a pulse width modulated DC-DC boost converter and a peak power tracking controller for converting solar cell power to bus power for charging of system batteries and powering loads while the controller controls the pulse width modulation operation of the converter for sensing solar cell currents and voltages along a power characteristic curve of the solar cell for peak power tracking, for determining any power data point, including a peak power point, an open circuit voltage point, and a short circuit current point along the power characteristic curve of the solar cell, and for communicating the power data to a satellite processor for monitoring the performance of the solar cell during operational use of the satellite.

The present invention relates to light-weight thin-film photovoltaic cells, methods for making cells, modules made from cells, and methods for making modules from cells. The invention teaches a manner in which individual cells may be bonded to one another, eliminating the need for an additional support substrate and interconnecting leads, thus reducing the overall weight and thickness of individual cells and modules of the cells.

A foldable member in the shape of a tube with at least one predetermined hinge area along the length of the tube and opposing sets of elongated slots in the tube at the hinge area thereof forming separated longitudinal strips of tube material between the slots which fold when subjected to localized buckling forces, each slot of each set of elongated slots separated longitudinally along the length of the tube from each adjacent slot by a bridge element of tube material. A collapsible structure constructed of such foldable members.

This invention is a method and supporting apparatus for autonomously capturing, servicing and de-orbiting a free-flying spacecraft, such as a satellite, using robotics. The capture of the spacecraft includes the steps of optically seeking and ranging the satellite using LIDAR; and matching tumble rates, rendezvousing and berthing with the satellite. Servicing of the spacecraft may be done using supervised autonomy, which is allowing a robot to execute a sequence of instructions without intervention from a remote human-occupied location. These instructions may be packaged at the remote station in a script and uplinked to the robot for execution upon remote command giving authority to proceed. Alternately, the instructions may be generated by Artificial Intelligence (AI) logic onboard the robot. In either case, the remote operator maintains the ability to abort an instruction or script at any time, as well as the ability to intervene using manual override to teleoperate the robot. In one embodiment, a vehicle used for carrying out the method of this invention comprises an ejection module, which includes the robot, and a de-orbit module. Once servicing is completed by the robot, the ejection module separates from the de-orbit module, leaving the de-orbit module attached to the satellite for de-orbiting the same at a future time. Upon separation, the ejection module can either de-orbit itself or rendezvous with another satellite for servicing. The ability to de-orbit a spacecraft further allows the opportunity to direct the landing of the spent satellite in a safe location away from population centers, such as the ocean.

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.

Provided is a satellite simulation system based on component-based satellite modeling. The system includes: a user interface unit for receiving simulation control commands and data and parameter required for simulation from a user; a satellite model unit for individually storing information dependant on the satellite, characteristics of the simulation object model and parameter information based on the characteristics, and performing simulation upon receipt of simulation control commands; and a simulation kernel unit for creating a schedule control command for simulation control of the satellite model unit, the onboard simulation unit and the external interface unit by the control command receiving/transmitting from/to the control command/telemetry to satellite control system, performing control and collecting and managing simulation results.

A rocket-powered vehicle used for entertainment, namely races, exhibitions, competitions, and revenue-generating events.

A method includes: a) producing initial scheduling plans based on requests related to tasks to be performed within a time period by a remote sensing satellite; wherein each initial scheduling plan schedules respective tasks, which do not conflict with each other in time and in using satellite resources; and each task is scheduled in at least one of the initial scheduling plans; b) applying a genetic-algorithm-based processing to the initial scheduling plans to produce a genetic-algorithm-based scheduling plan which is for mission objectives, and complies with constraints related to the satellite resources, to the tasks, and to the time period; and c) applying a simulated-annealing-based processing to the genetic-algorithm-based scheduling plan to produce a simulated-annealing-based scheduling plan that fits the mission objectives, complies with the constraints, and in which a larger number of tasks is scheduled than in the genetic-algorithm-based scheduling plan.

The present invention provides an emulation system having a control system that allows the testing of a satellite control system with all of its hardware in place, i.e. fully integrated. The emulation methodology is applicable to the case of either a rigid spacecraft or a flexible spacecraft, provided that the spacecraft's sensors and actuators are stowed to the rigid part of spacecraft in the case of a flexible spacecraft. Practically, the latter condition is not restrictive, as the actuators and sensors are usually placed rigidly in the satellite bus, while the satellite solar panels constitute the flexible elements. The control system is used to tune the mass properties and dynamic behaviour of a rigid ground-spacecraft in a 1-G environment to those of a flight-spacecraft in 0-G. A six-axis force/moment sensor is placed at an interface of the ground-spacecraft and a manipulator. Signals received from the force/moment sensor, and in some cases signals relating to the position and velocity of manipulator joints, are received into the control system.

A person can safely descend from a burning high-rise building or the like using an inflatable flying body that has a hollow conical form in an inflated deployed condition, but is deflated and folded into a backpack form in a stowed condition. The flying body includes an upper stabilizing ring, a lower nose structure with a pneumatic damping body, spoke struts extending conically therebetween, a cover skin covering the abovementioned inflatable components to form the conical outer surface and provide aerodynamic braking drag, and gas generators to inflate the inflatable components. A person straps on the apparatus in the stowed backpack form and pulls a handle to actuate the gas generators for inflating the apparatus, whereby the expanding apparatus ejects the person from the building and then orients itself in a nose-down attitude during the descent. The pneumatic damping body damps and dissipates the landing impact energy.

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 network architecture and protocol provides “plug and play” spacecraft capabilities. A distributed-control architecture is used, wherein each component operates semi-autonomously, and interacts with other components on a task/resource level. Each component announces its requirement for system resources as a request to the network, and components that can provide some or all of the requested resources respond to the request. An arbitration device centralizes and coordinates requests for critical and/or singular resources, such as requests for a specific orientation of the spacecraft. To facilitate such distributed control, requests are made in advance of the requirement for the resource, and include a time interval during which the resource is required. A configuration and test system is provided to process the mission requirements and provide a set of components that can be configured to satisfy the requirements.

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.

A double-sided photovoltaic package with an incident photovoltaic cell and a reflective photovoltaic cell. Both photovoltaic cells have a corresponding absorbing surface for absorbing solar radiation. The incident photovoltaic cell and the reflective photovoltaic cell are arranged so that when the absorbing surface of the incident photovoltaic cell is located to receive incident non-reflected solar radiation the absorbing surface of the reflective photovoltaic cell is located to receive reflected solar radiation. The structure of the incident photovoltaic cell is adapted to convert non-reflected light to electrical energy and the structure of the reflective photovoltaic cell is adapted to convert reflected light to electrical energy. Additionally, in the preferred embodiment, the incident photovoltaic cell and the reflective photovoltaic cell both comprise inverted metamorphic multijunction photovoltaic cells. Furthermore, a plurality of double-sided photovoltaic packages according to the present invention may be interconnected in a string formation and mounted on a transparent panel to form an array.

The invention relates to a method for automatic fault diagnosis of a satellite and a diagnostic system thereof. The method comprises the following steps: dividing the satellite fault into a component level fault, a subsystem level fault and a system level fault; adopting a fault detection method based on a fuzzy set to fuzzify a precise sensor sampling value; defining different alarm weight values for fuzzy subset based on the membership grade belonging to each fuzzy subset; and determining whether to alarm or not by calculating the relation between the summation of the alarm weight values in a time window and an alarm threshold value. The method can effectively overcome the defects of a common fault detection method and effectively reduce the false alarm rate and the rate of missing report; and when certain fault is caused by a plurality of sub-faults, the method performs heuristic search according to the reliability of a system corresponding to each sub-fault and searches a sub-fault node corresponding to a system with low reliability preferentially. Besides, in order to adapt to the synergistic diagnosis between distributed satellites in the future, a synergistic interface is designed to provide diagnosis services for other satellites, thus the aim of the invention is achieved.

Thrust vectoring enhances aircraft and spacecraft maneuverability. The propulsion systems of aircraft, using gas turbine engines, and spacecraft, using solid or liquid rockets, generate compressed gases that are directed through a nozzle to generate forward thrust. This patent describes opening ports to release gases from aircraft and spacecraft in a direction other than the direction of forward propulsion. By altering the direction of gas flows leaving the craft pitch, roll, yaw and attitude control is obtained. To control gas flow from ports a variable throat thrust vectoring nozzle is described. Independent operation of ports, rotational nozzles, laterally moving flaps and rotational vanes assists the thrust vectoring of gases leaving aircraft or spacecraft ports. The ability to direct gases enhances maneuverability during vertical takeoff and landing, forward flight and space flight.

The invention provides a seven-revolute pair extensible unit and a space extensible mechanism using same, and relates to a single-closed loop and single-freedom degree space extensible mechanism aiming at solving the problem that the existing five-revolute pair mechanism is rigorous in geometric parameter requirement, poor in extensible performance, low in reliability, and not suitable for a large-size space extensible mechanism. A plurality of first extensible units are single-closed loop and single-freedom degree extensible units, a plurality of first extensible unit ring arrays are arranged on a common base, and every two adjoining first extensible units are formed into a first space extensible mechanism by a common first connecting rod or a common sixth connecting rod; and a plurality of second extensible units are single-closed loop and single-freedom degree space extensible units, a plurality of second extensible unit ring arrays are arranged on the common base, and every two adjoining second extensible units are formed into a second space extensible mechanism by the common first connecting rod or the common second connecting rod. The invention is suitable for a support frame of a satellite parabolic antenna and a solar battery array.