331results about "Cosmonautic ground equipments" patented technology

An optical communication system using a high altitude tethered balloon 10 that operates above most clouds and atmospheric turbulence. An optical communication system includes a balloon 10 with an optical communication payload 30, a fiber optic cable attached to the tether 12, an automated winch system 14, and a ground station 28. The balloon 10 is designed for sustained flight at, and recovery from, high altitude using the automated winch system 14. An acquisition, tracking and pointing (ATP) system 22 enables a balloon-based optical transceiver 24 to maintain line-of-sight optical communications with an overhead satellite 34. The optical link between the satellite 34 and the balloon 10 is an open channel. Data is transmitted to the ground station 28 from the balloon 10 through the secure fiber optic closed channel attached to the tether 12.

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.

The present invention provides a method, system and apparatus for robotic refueling of satellites. The system may include a dedicated refueling satellite launched directly from either earth, or alternatively it could be launched from another larger mother spacecraft or space station in which the refueling satellite is ferried into space in the case of the larger space craft or it may be stored on the space station until needed from which it can be launched. The system includes a robotic arm, suitable tools which can be affixed to the end effector of the robotic arm required for accessing, opening and closing the fuel fill valve on the satellite being serviced, storage and retrieval stations on a tool caddy on which the tools and various fuel fill valve caps are stored. The system is under teleoperation by a remotely located operator, for example located on earth, in the mother station or in the space station. Cameras are included focussed on the robotic arm and end effector and images are transmitted to the operator to allow the operator to direct and control the refueling procedure. The system may also be configured to be operated autonomously under computer control.

Power supply satellites may be launched to LEO and boosted to GEO using power generated on board from solar insolation. A cluster of power production satellites may be operated as a phased antenna array to deliver power to one or more ground-based facilities, which may be located in different time zones.

A flying vehicle-launching apparatus and method for launching a winged flying vehicle by accelerating the vehicle in a predetermined direction includes a flying vehicle-supporting platform, a guide device, and a driving device. The flying vehicle-supporting platform supports the winged flying vehicle so that it can lift off therefrom. The guide device supports the flying vehicle-supporting platform across two or more separated and parallel guide ways and guides the flying vehicle-supporting platform along the guide ways. In the flying vehicle-launching method, a propulsive engine of the winged flying vehicle is started after initiating acceleration with a flying vehicle-accelerating device. The winged flying vehicle is released from the flying vehicle-accelerating device in a predetermined angle of attack that is positive to a direction of acceleration after reaching a predetermined velocity.

This invention is a launch-to-space, refuel and return system which comprises: a unique circular and straight maglev track and maglev sled spacecraft launcher; a single stage-to-LEO spacecraft (Mother Spacecraft); an orbiting space platform (OSP); a currently utilized Space Shuttle is altered to improve its safety, range, longevity and versatility which serves as a prototype for the Mother Spacecraft design. The Mother Spacecraft is releasably attached to the maglev sled and propulsion for the maglev sled is supplied by magnetic propulsion and on-board rocket propulsion. The collapsible OSP can fit into the Space Shuttle storage bay and is remotely deployed. The OSP is a refueling depot for the Mother Spacecraft, the Space Shuttle, as well as a rendezvous point for various space missions. A single, liquid, non-cryogenic rocket fuel is standardized throughout the entire system which eliminates: all vertical launches; all expendables; all solid rocket propellants.

A flying vehicle-launching apparatus and method for launching a winged flying vehicle by accelerating the vehicle in a predetermined direction includes a flying vehicle-supporting platform, a guide device, and a driving device. The flying vehicle-supporting platform supports the winged flying vehicle so that it can lift off therefrom. The guide device supports the flying vehicle-supporting platform across two or more separated and parallel guide ways and guides the flying vehicle-supporting platform along the guide ways.In the flying vehicle-launching method, a propulsive engine of the winged flying vehicle is started after initiating acceleration with a flying vehicle-accelerating device. The winged flying vehicle is released from the flying vehicle-accelerating device in a predetermined angle of attack that is positive to a direction of acceleration after reaching a predetermined velocity.

A rocket launch system (1) includes a tubular rocket launcher carriage (2) with electromotive cableway traction drives (26) conveyed beneath a two axis pivot (63) anchored to the earth, elevated into a co-axial transfer tube (124, 143) leading to three primary tether cables (27) whose weight is offset by balloons (164). The carriage is conveyed to a docking station (166) supported into the stratosphere by a pair of secondary cables (184) suspended under an attachment frame (162) for tensioning balloons. The carriage is engaged by a carriage end gripper (196) guided by two secondary and two tertiary cables (186) and lifted by a lower hoist (198) guided by the secondary cables. This lower hoist is supported by an upper hoist (168) suspended from the tensioning balloons attachment frame. The carriage, which engages a lift ring (183) guided by two secondary cables, is elevated further, rotated as desired, with rocket release and nearly recoilless ejection during freefall of the carriage, with engine ignition occurring at a safe distance.

Advanced spaceport information management methods and systems are disclosed. In one embodiment, a method includes coupling a test system to the payload and transmitting one or more test signals that emulate an anticipated condition from the test system to the payload. One or more responsive signals are received from the payload into the test system and are analyzed to determine whether one or more of the responsive signals comprises an anomalous signal. At least one of the steps of transmitting, receiving, analyzing and determining includes transmitting at least one of the test signals and the responsive signals via a communications link from a payload processing facility to a remotely located facility. In one particular embodiment, the communications link is an Internet link from a payload processing facility to a remotely located facility (e.g. a launch facility, university, etc.).

This invention will allow a reusable thrust-powered sled mounted on an inclined track to launch spacecraft or airborne vehicles from earth at supersonic speeds using existing technology properly integrated into an inclined track system. If launched up a tunnel track, a rear blast shield can trap the rocket exhaust to provide a pneumatic boost upon launch. The sled can also launch ramjet or scramjet powered vehicles from earth by achieving the Mach 2+ speed necessary to ignite their engines. This system is much safer than the traditional method of launching rockets since weather is less a factor and the launch can be aborted if problems develop. Moreover, it is far less costly since the engines on the sled can be reused hours after a launch and the track can accommodate a variety of sleds to launch objects of many different sizes.

A rocket landing stabilization system can include one or more upright support structures such as posts, columns, or walls, from which one or more stabilizing elements can be supported. The stabilizing elements can be used to stabilize a rocket as it lands at a landing site. The rocket landing stabilization system can also include a cradle, funnel, or cone to catch or otherwise support a rocket as it lands at the landing site. The rocket landing stabilization system can be located on land or at sea.

The invention discloses a quick fault detection system for spacecrafts. The quick fault detection system comprises a measurement module, a fault detection and diagnosis module, an interface module and an output module. State signals of sensors can be received by the measurement module and can be inputted into signal conditioning circuits of a LABVIEW platform in the measurement module so as to obtain measurement data of electric analog signals; the measurement data can be received by the interface module and can be integrated according to the time so as to obtain matrix forms; information fusion can be carried out on the integrated measurement data by the fault detection and diagnosis module so that the measurement data can be converted into CLIPS fact structures, eligible measurement data can be deduced by inference machines according to rules of an expert knowledge base, positioning and detection can be carried out by the aid of fault tree rules, and maintenance suggestion can be obtained; positioning and detection results and the maintenance suggestion can be displayed by the output module. The quick fault detection system has the advantages that knowledge is easy to acquire and analyze, and the quick fault detection system is high in rule inference efficiency; effects of predicting faults can be realized according to judgment on current working states of the spacecrafts, accordingly, possible consequence can be eliminated in advance, and the quick fault detection system has low hardware requirements.

A tool for fuelling an unprepared or partially prepared client satellite required for accessing, opening and closing a fuel fill valve on the satellite being serviced, storage and retrieval stations on a tool caddy on which the tools and various fuel fill valve caps are stored. The tool includes interchangeable socket modules, a support frame, a socket module holder mechanism, a socket drive mechanism, a clamping mechanism to secure the tool to a reaction area on the fuel fill valve and oxidizer fill valve, and a valve actuation mechanism. The interchangeable socket modules include a first module for removing and replacing the access valve cap, a second module for coupling the fuel fill line to the fuel tank, a third module for engaging the access valve cap on the oxidizer fill valve, and a fourth module for coupling the oxidizer fill line to the oxidizer tank.

The invention discloses a self-adaptive propulsion mounting system for an erecting frame of a launch vehicle. The self-adaptive propulsion mounting system comprises a reversal docking device and a transport vehicle, wherein the erecting frame is arranged on the transport vehicle; a plurality of lateral fine-tuning mechanisms are respectively mounted on the two sides of the erecting frame; a plurality of universal swing mechanisms are mounted at the top surface of the transport vehicle, and the universal swing mechanisms are abutted against the bottom end surface of the erecting frame; a propulsion mounting support is fixedly connected to the erecting frame; the propulsion mounting support is hinged to the output end of a propulsion mounting oil cylinder; through a hinge seat A, the tail end of the propulsion mounting oil cylinder is hinged to the front end of a tailstock fixed to the top of the transport vehicle; a positioning guide rail is arranged at the top of the reversal docking device; the universal swing mechanisms, which are matched with the bottom end of the positioning guide rail and are inverted, are fixedly connected to the bottom at the front end of the erecting frame; a positioning support is fixedly connected to the front end of the erecting frame; a positioning pin which is matched with the positioning support is fixedly connected to left end of the positioning guide rail; and a positioning switch is mounted in a position close to the positioning pin. Through the adoption of the self-adaptive propulsion mounting system disclosed by the invention, the propulsion mounting cost of the launch vehicle is effectively reduced, and the accuracy and efficiency are improved.

The invention discloses a service life ground test method for an ion thruster. By using the ground test method, a ground service life test for the iron thruster can be performed for a long time. According to the ground test method, service life ground test equipment is built first of all, the performance of the ion thruster after working for a long time is tested on the basis of the simulation space condition of the test equipment, and accordingly actual results of the service life of the ion thruster can be obtained, and real and reliable test data are provided for evaluation of the service life and the reliability of the thruster. Main performance parameters of the ion thruster in the whole service life can be obtained through substantive tests in the test process, and the change rule of the performance parameters of the thruster along with accumulative working time is further obtained; the basis for long-service-life working and application of the ion thruster is provided, and essential data can be obtained for more effective acceleration tests of the iron thruster in future. The test method is effective and reliable, and the economic cost and the time cost of the service life test are greatly lowered.

A system and a related method are disclosed for matching space launch vehicle, satellite, and mission hosted payload opportunities utilizing multi-level, priority algorithms. The potential matches include numerous mission-matching scenarios, e.g., 1) matching hosted payload and host spacecraft, and 2) matching spacecraft(s) and launch opportunities, to include rideshare opportunities. The matching opportunities can be pursued by any interested project team, e.g., hosted payload, spacecraft, and space launch vehicle.

A rocket launch tower is provided, including: a vertical support structure including two or more guide towers defining a vertical shaft between the two or more guide towers, each guide tower including one or more pulleys engaging one or more cables; a platform located within the vertical shaft and connected to the one or more cables; a drive mechanism that applies a force to the one or more cables to accelerate the platform along a trajectory within the vertical shaft; one or more sensors collecting data regarding the position of the platform along the trajectory within the vertical shaft and communicating the platform position data to a controller, the controller in communication with an acceleration control system including one or more brakes acting on the platform; wherein the controller causes the acceleration control system to actively correct the platform acceleration towards an intended platform acceleration.

The method provided by the invention is based on the designed counterweight system. According to the counterweight system, a controller serves as the center; a mass measurement table is adopted for measuring the mass and the mass center position of the spacecraft before filling of the propellant is carried out each time; a controller determines an object in need of filling according to the mass center position; the controller adopts torque balance calculation to obtain the mass of the filling object according to the filling object, the mass center position of the spacecraft and the measured mass center position of each tank; the controller controls the corresponding filler to fill the propellant to the corresponding tank via a tank interface according to the mass obtained after calculation and the filling object; and if the mass center position measured by the mass measurement table can not meet the design demands, the above process is repeated by the control table until the mass center position of the spacecraft meets the design demands.

The invention discloses a carrier rocket sub-stage recycling system and method. The carrier rocket sub-stage recycling system comprises a recycling box body, a probe device, a signal transceiver, a plurality of airbags, an inflation device, moving mechanisms and a ground control system. The recycling box body is of a U-shaped box body structure with the top end being not closed; the probe device,the signal transceiver and the inflation device are arranged on the recycling box body; the multiple airbags are arranged at the bottom and on the side faces of the recycling box body correspondingly;the inflation device is connected with the multiple airbags; the moving mechanisms are arranged at the bottom of the recycling box body; and the ground control system makes remote communication withthe carrier rocket sub-stage recycling system through the signal transceiver. Based on a carrier rocket in active service, recycling of a sub-stage part is completed while the carrier rocket sub-stagepart is less changed.

The invention discloses a ground emitting control device of a solid-liquid powered aircraft; the ground emitting control device consists of a front end device and a rear end device. The front end device is a ground control box, wherein the output port of the front end device is connected with the solid-liquid powered aircraft; the rear end device comprises a host computer and a remote control box, and the remote control on the solid-powered liquid-powered aircraft is realized. The ground emitting control device of the solid-liquid powered aircraft mainly finishes two functions; the first function is the igniting emitting function which relates to pressure increase of storage boxes, ignition of igniter and opening of liquid-path electromagnetic valve, and the other function is emitting safety ensuring function which relates to high-pressure gas release of high-pressure bottles, pressure release of storage boxes and oxidant discharge of storage boxes. The ground emitting control device solves problems of complicated system, poor portability and high cost of traditional ground emitting control devices, the precise control of igniting emitting timing sequence aiming at the characteristics of solid-powered liquid-powered aircraft is realized, and the programming can be carried out in an on-line manner so as to modify igniting emitting timing sequence simply and conveniently.

The invention discloses a flexible self-adaptive dynamic balancing device. The device includes a lower base plate, a moving platform, motor screw guide rail modules, first branch chains and a second branch chain, the lower base plate is fixed on the ground, each motor screw guide rail module is fixed on the lower base plate, one end of each first branch chain is connected with a slide block of thecorresponding motor screw guide rail module through a spherical pair, and the other end of each first branch chain is hinged with the moving platform; one end of the second branch chain is connectedwith the lower base plate through a spherical pair, and the other end of the second branch chain is hinged with the moving platform; the number of the first branch chains is five, the number of the corresponding motor screw guide rail modules is five, and the number of the second branch chain is one. The advantages are fully used that a parallel mechanism is high in carrying capacity, compact in structure, flexible to move, small in inertia and convenient to control in time, and the whole symmetrical structure design has the advantages of simple and compact structure, reasonable design and good processing and assembly technologies.

The invention discloses a rotational centrifugal launcher which comprises a power device, a control device, an accelerator and a base. The power device is composed of a driving motor (1) and a speed changing mechanism (15); the control device (3) adopts a single chip microcomputer or a computer to control the speed changing mechanism (15) and the launching initial speeds and launching angles of launching objects (4); the accelerator comprises a rotating device (5), multiple launching canisters (16) and electromagnetic releasing mechanisms (7), the launching canisters (16) are fixed on the edge of the whole circumference of the rotating device (5) and driven by the power device to rotate to reach certain revolutions, and therefore the launching objects (4) in the launching canisters (16) can acquire certain initial speed and kinetic energy. The rotational centrifugal launcher is small in size and weight and high in invisibility and launching precision, one launcher can launch warheads of different warhead diameters, and single launching, continuous launching and mass launching can be all achieved; heavy objects can be safely sent to the far place at low cost, even the outer space.

The invention discloses a deep space exploration large-caliber antenna calibration ground test system based on suspension device. The deep space exploration large-caliber antenna calibration ground test system based on the suspension device comprises onboard equipment and ground equipment; the onboard equipment comprises a large-caliber antenna, an onboard low gain receiving antenna, antenna driving control equipment and onboard communication equipment; and the ground equipment comprises the suspension device, a ground signal simulator, a ground low-gain transmitting antenna, a theodolite anda frequency spectrograph. By using the suspension device, gravity force effects of the large-caliber antenna in a static state and a driving state are effectively balanced; by using a ground signal simulation source, the low-gain transmitting antenna and the large-caliber antenna, in-orbit communication state to the ground is simulated; through reasonable position layout, the signal intensity of the large-caliber antenna is inversely calculated by using the low-gain receiving antenna; by using the theodolite, the pointing accuracy of a mechanical shaft of the large-caliber antenna is calibrated, and the radio frequency calibration accuracy of the large-caliber antenna is inverted. According to the deep space exploration large-caliber antenna calibration ground test system based on the suspension device, the in-orbit calibration process of the large-caliber antenna can be effectively simulated, and the reasonability and the effectiveness of the in-orbit calibration method are guaranteed.

A multiply-redundant system that protects fueling of rockets, aircraft and other vehicles using Liquefied Natural Gas (LNG) along with an oxidizer such as Liquefied Oxygen. One or more sensors in combination with one or more optional microswitches combine to detect any leaks, fire or explosion hazards quickly locking out further fueling. For different levels of safety, different combinations of sensors can be used.

The invention discloses a modular assembly small satellite platform with adjustable quality characteristic, and relates to the field of precise attitude and orbit control of satellite; the modular assembly small satellite platform comprises an attitude and orbit control module, two solar panel modules, a vertebral column module, a load module, a communication module and n centroid adjusting modules; the vertebral column module is of a square column structure; the vertebral column module is vertically arranged; the attitude and orbit control module is fixedly arranged on the upper surface of the vertebral column module; two solar panel modules are symmetrically and fixedly arranged on the two sides of the top end of the vertebral column module; the load module is fixedly arranged at the bottom of the vertebral column module; the communication module is fixedly arranged at one side of the load module; the n centroid adjusting modules are fixedly arranged inside the vertebral column module or the other side of the load module; and n is an integer greater than or equal to 1. The existing design process of design, test and balance weight of the quality characteristics of the small satellite are changed, development cycle is shortened, and the quality characteristics control of the modular assembly small satellite on orbit can be realized at the same time.

A gateway segment assembly line that allows for the construction of space stations while in orbit by fabricating individual segments in space. An assembly housing provides an open ended structure through which materials are processed in order to construct a segment for a space station. The materials are loaded into a plurality of workstations positioned along the assembly housing through the use of a plurality of external manipulators adjacently connected to the assembly housing. Each of the plurality of workstations provides the equipment for sequentially loading materials into the assembly housing. An assembly line conveyor, positioned throughout the plurality of workstations, guides materials through the assembly housing as the materials are mated to form the segment. Upon completion of the segment, a plurality of segment transport units transports the segment to an orbital construction site, wherein the segment is mated with subsequent segments to form the space station.

The invention relates to a zero-second dropping-off normal-temperature propellant adding and discharging connector for a carrier rocket and a method. The zero-second dropping-off normal-temperature propellant adding and discharging connector for the carrier rocket comprises a connector ground part, a connector on-rocket part, a first pulley set, a second pulley set, a third pulley set, a recoveryrope, a traction rope, a counterweight, an air cylinder and a ball lock; the connector on-rocket part is mounted on a propellant storage tank of the carrier rocket, the second pulley set and the air cylinder are mounted on a launching tower, the connector ground part is connected with the connector on-rocket part through the ball lock, the counterweight is connected with the connector ground partthrough the second pulley set and the recovery rope, and the air cylinder is connected with the ball lock through the third pulley set, the traction rope and the first pulley set. The connector between a ground normal-temperature propellant filling pipeline and the on-rocket propellant storage tank can drop off reliably after zero-second ignition for taking off of the carrier rocket.

A method, operable in the presence of ambient cosmic rays, is provided for braking a craft upon approach to a planet, moon or other space body, e.g. in preparation for landing. Deuterium-containing particle fuel material is projected in a specified direction outward of the craft, which interacts with both the cosmic rays and their principal decay product muons to generate energetic micro-fusion products that produce a braking thrust on the craft for a specified trajectory. The micro-fusion products may push directly against the craft, e.g. upon a pressure plate, or upon a sail or parachute connected to the craft, to decelerate the craft. A prepositioned automated landing system at a landing site may project the fuel material toward the craft based on telemetry tracking of an incoming craft and likewise directly disperse the material cloud to form a braking cushion at the landing site. The micro-fusion landing system may be part of a site-to-site transport, where the craft was launched using either conventional chemical rockets or micro-fusion for accelerating thrust.

Systems and methods for launching space vehicles into orbit involve placing a space vehicle into a protective container. The protective container is placed on a sled in a maglev tunnel and then launched into the atmosphere. Once the protective container reaches a certain height the space vehicle is released from the protective container and the space vehicle then proceeds to the desired orbital position.