330results about "Cosmonautic ground equipments" patented technology

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.

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.

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.).

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.

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 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.

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.

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.