151results about "Cosmonautic ground equipments" patented technology

An aircraft has a boom, a propulsion assembly coupled to a first end of the boom, and a first wing coupled to a second end of the boom. The propulsion assembly is coupled to the boom by a rotating joint. A second wing is optionally coupled to the rotating joint. The first wing is coupled to the boom by a rotating joint. The first wing is coupled to the rotating joint by a hinge. A vehicle with roll, pitch, and yaw maneuverability able to mirror the aircraft movements may be coupled to the second end of the boom. The vehicle body may be picked up with a vehicle chassis disconnected from the vehicle body. The boom houses an energy source to power the propulsion assembly. A rudder is coupled to the second end of the boom. A paddle is disposed between the propulsion assembly and the boom.

The invention discloses a reusable carrier rocket marine landing recovery device which comprises a tower body, a driving module, a first mechanical capturing arm, a second mechanical capturing arm, a first clamping arm and a second clamping arm. The tower body comprises first to fourth stand columns and first to eighth cross beams; the driving module comprises first to twelfth driving units; each of the first to twelfth driving units comprises a driving motor, a working track, a lead screw nut, a first bearing, a second bearing, a first bearing seat and a second bearing seat. On the premise that a simple electric control system is adopted, marine recovery of the carrier can be achieved, the structure is simple, response is rapid, good reliability and maintainability are achieved, the device is suitable for recovery tasks of carrier rockets of various models, and the launching cost is effectively reduced.

The invention discloses a mobile medium-large liquid carrier rocket capturing and recycling system, and relates to the technical field of rocket recycling. According to the mobile medium-large liquid carrier rocket capturing and recycling system, through the mobile platform design that a crane is adopted to carry a connecting truss, a switching frame and a capturing arm, rocket body information is obtained in real time in combination with a camera module, a rocket position sensor and the like, and interaction with a ground center is achieved through a data transmission module; the control module accurately controls the transmission assembly to adjust the capturing arm, the buffering assembly is matched to absorb impact, and a parachute landing mode is replaced to improve the control precision. According to the movable characteristic, the dependence on a fixed land or seaborne recovery site is reduced, the site construction and maintenance cost is reduced, the influence of environmental factors on recovery is reduced, and the problems that in an existing medium-large liquid carrier rocket recovery technology, parachute landing control is difficult, the land / seaborne recovery cost is high, and the influence of the environment is large are solved.

The invention discloses a reusable carrier rocket ground active recovery system and recovery method.The system comprises a landing buffer mechanism, an omnidirectional self-driven carrying platform, a rocket body posture auxiliary stabilizing mechanism and a landing structure, the landing structure comprises bearing feet, the bearing feet are located on a rocket, and a positioning device is arranged on the omnidirectional self-driven carrying platform; the rocket body posture auxiliary stabilizing mechanism comprises two pairs of clamping rods which are arranged in parallel in a staggered mode, the two ends of each clamping rod are located on a sliding table, the clamping rods are driven by the sliding table to move, and the sliding table is located on a sliding rail; the landing buffering mechanism is located on the omni-directional self-driven carrying platform and comprises a force bearing buffering base and a flexible contact pad located on the force bearing buffering base. The rocket recovery control difficulty is reduced while the carrying capacity of the rocket is guaranteed, and the rocket recovery control device is reasonable and stable in structure, capable of being applied to recovery of rockets of multiple specifications, wide in applicability and wide in application prospect.

In a system where multiple satellites autonomously take turns acting as servers, the servers will be rotated according to instructions from a ground station. [Solution] The server, which is a satellite 200 that has obtained server rights, receives a priority notification when a ground station 300 transmits the priority notification. If the server receives the priority notification, it selects the satellite 200 specified in the priority notification as the server for the next time interval. If the server does not receive the priority notification, it selects one of the multiple satellites as the server for the next time interval based on the status information of the server and the status information of each client.

The invention is notably directed to an aircraft carrier ground vehicle (10), which comprises a chassis (102), a landing platform (104), a coupling mechanism (106), and an actuation mechanism (108). The landing platform is designed to accommodate an aircraft (20). The coupling mechanism connects the landing platform to the chassis to form a kinematic pair, in which the landing platform is rotatably mounted on the chassis. The actuation mechanism is adapted to rotate the landing platform relative to the chassis, to allow an aircraft to land on the landing platform as the vehicle moves along a landing direction of the aircraft and compensate for an inclination angle of the aircraft on landing, in operation. The aircraft carrier ground vehicles may further be designed to allow taxiing and take-off of the aircraft. The invention is further directed to related aircraft carrier ground vehicle systems, which include several vehicles that are swarmed for landing an aircraft, as well as aircraft landing and take-off methods.

A propulsion-based payload transporter (PPT) and methods of its operation are provided. The PPT may operate in a gravitational field such as that on the lunar surface. Via rocket engines that produce thrust, the PPT is configured to descend onto and over a payload to be moved. Once in position with respect to the payload, latches engage with the payload so as to secure the payload in a payload holding bay of the PPT. The rocket engines produce thrust with an exhaust plume that is angled away from the payload to avoid payload damage. The rocket engines, which are located at the sides of the payload holding bay, produce thrust to propel the PPT and the payload in a substantially horizontal trajectory with respect to gravity and toward a landing destination.

The application relates to a detachable guide rail type launch vehicle pre-launch heating and keeping warm device, which comprises a detaching mechanism and a heating and keeping warm mechanism; in a storage state, the detaching mechanism and the heating and keeping warm mechanism are independent of each other and are stored separately; in a heating working state, the heating and keeping warm mechanism is lifted by the detaching mechanism, and the heating and keeping warm mechanism heats and keeps warm the vehicle; and in a launch separation state, the detaching mechanism takes the heating and keeping warm mechanism away from the guide rail type launch device of the guide rail type launch vehicle, so that the vehicle can normally leave the rail. The application can ensure normal heating and keeping warm functions before and after the small and medium-sized vehicle is erected, and can make the heating and keeping warm device reliably detach through manual or electric ways 5 minutes before the vehicle is launched, so that the normal launch of the vehicle is not affected.

The present invention is embodied in systems and devices for transporting assets (small satellites, goods, or other payloads) in outer space. A preferred system comprises a scaffold of relay stations strategically orbiting one or more celestial bodies. Since the orbits are complementary to each other, the relay stations can transport mass objects from one station to another. In this way, the system can support various logistics networks connecting space between orbits such as low Earth orbit (LEO), geostationary orbit (GEO), and low lunar orbit (LLO), as well as space between Earth and Mars.

The invention relates to a multi-source electromagnetic matrix aircraft take-off and landing system and a waste heat utilization and building integration method thereof, and belongs to the technical field of aircraft take-off and landing. The system comprises at least three electromagnetic emission units which are arranged in a space array, and each electromagnetic emission unit comprises an independent controllable excitation coil; the central controller is used for independently adjusting current parameters of each electromagnetic emission unit; the at least three electromagnetic emission units act together to form at least four independent and controllable electromagnetic stress action points in a predetermined space area. The system comprises a building facade integration system, a waste heat recovery system and a military application system (space debris cleaning and aerospace vehicle launching). The system has the multiple advantages of high control dimension, high magnetic field intensity, low cost, mobility, building embeddability, comprehensive energy utilization, military and civilian dual-purpose, military defense value and the like, and can be used as a general take-off and landing infrastructure in the fields of urban air traffic, military aviation, spaceflight launching, military defense and the like.

The application provides a rocket erecting frame overturning docking device and a docking method thereof, and the device comprises an overturning docking base, overturning docking blocks, a rotating assembly, an erecting oil cylinder assembly and a hydraulic chuck mechanism; the overturning docking base is a main torsion-resistant mechanism during the overturning process of the erecting frame, and high-strength steel is selected as the material; the overturning docking blocks are arranged on the left and right sides of the overturning docking device and are symmetrically arranged, the bottom of the overturning docking block is connected with the overturning docking base; the upper surface of the overturning docking block is reserved with connecting hole positions of the erecting frame and the hydraulic chuck mechanism, and the inner cavity is provided with an interface connected with the upper fulcrum of the erecting oil cylinder. The adjustment and positioning of the erecting frame are completed through the expansion of the hydraulic chuck. The docking method can realize the accurate posture positioning of the erecting frame in a short time, increases the reliability and efficiency of the docking, reduces the docking time and the participation degree of manpower, and reduces the safety risk to a certain extent.

The application relates to a modular quick-assembly and opening type launching station aircraft protection device scheme and belongs to the technical field of launching protection devices. The device is composed of a plurality of basic modules, each basic module is designed by using quick-guiding positioning, quick hoisting positioning can be realized, a plurality of groups of quick-assembly mechanisms are arranged to be connected with quick locking / unlocking, quick fixing between the modules can be realized after installation, an integrated protection device is formed, and the protection device is convenient for quick assembly and opening.

Orbital Elevator with Magnetic Rail Propulsion System: The invention relates to an orbital elevator designed to transport payloads between Earth and geostationary orbit. The system uses an ultra-strong, multi-strand cable made of carbon nanotubes to withstand gravitational and centrifugal forces. A transport module (climber) equipped with magnetic rails is in continuous contact with the cable, enabling stable, frictionless propulsion and precise control of movement along the cable. An integrated emergency braking system allows the climber to be stopped if necessary, thus enhancing transport safety. The ground base, which anchors the cable, includes a control center to monitor the movement of the modules and maintain structural tension. An orbital counterweight completes the system, stabilizing the cable through dynamic equilibrium.

The invention provides a multi-load observation task planning method, which comprises the following steps of: after any load on a satellite triggers an observation task, converting the position of an observation target into a position in a reference load view field, then calculating a constraint condition according to the position of the observation target in the reference load view field, comparing the constraint condition with a preset condition, and if the preset condition is met, determining that the observation target is a target. If not, the task is cancelled. By setting the reference load and calculating the constraint condition based on the reference load, it can be ensured that all loads in the observation task can observe the target, and then the task success rate and the observation accuracy are ensured.

The invention provides a supporting device. The supporting device is used for a carrier rocket sub-stage power system test run and comprises a supporting frame, a first supporting frame and a second supporting frame, the multiple stand columns are detachably connected to the lower portion of the first platform assembly. The first platform assembly comprises a first platform and a bearing frame; a first through groove is formed in the center of the first platform, and the bearing frame is detachably connected with the inner circumferential wall of the first through groove; the lower surface of the transfer frame is detachably connected with the bearing frame, and the upper surface of the transfer frame is configured to be detachably connected with the bottom end of the carrier rocket sub-stage power system; one end of the mooring tool is detachably connected with the first platform, and the other end of the mooring tool is constructed to be detachably connected with the top end of the carrier rocket sub-stage power system; therefore, the supporting device can be quickly disassembled and assembled and is convenient to transport. Moreover, the mooring tool provides a downward pulling force, so that the phenomenon that the carrier rocket sub-stage is out of control and flies off due to accidents in the test run process of the carrier rocket sub-stage power system can be avoided.

The invention discloses an equipment six-degree-of-freedom automatic posture-adjusting docking device applying a binocular vision system, belongs to the technical field of cabin docking, and solves the problems that the mounting effect of a cabin docking device in the prior art depends on the experience of operators, and the mounting stability is poor. The invention discloses an equipment six-degree-of-freedom automatic posture adjusting and docking device applying a binocular vision system. The equipment six-degree-of-freedom automatic posture adjusting and docking device comprises a horizontal walking mechanism, a binocular vision camera, a first docking section posture adjusting mechanism and a second docking section posture adjusting mechanism. The binocular vision camera can shoot the postures of the butt joint sections on the first butt joint section posture adjusting mechanism and the second butt joint section posture adjusting mechanism in real time. The positions of the first butt-joint section and the second butt-joint section are recognized through the binocular vision camera, butt joint of the first butt-joint section and the second butt-joint section is achieved by controlling the horizontal walking mechanism, the first butt-joint section posture adjusting mechanism and the second butt-joint section posture adjusting mechanism, and dependence on experience and level of operators is avoided; and the butted cabin sections have good stability and consistency.

A space transportation system (1) intended to be mounted on a space launcher (100) comprising a re-ignitable propulsion device (2), steerable flaps (3) arranged at a second end (12), and a monitoring unit (4), each flap (3) comprising an actuation means configured to change the orientation of said flap (3), said monitoring unit (4) being configured to: · a launch step in which the propulsion device (2) is off and the flaps (3) are in the retracted position; a return step in which the monitoring unit (4) deploys the flaps (3) and individually controls the orientation of the flaps (3) to slow down the system (1), while the propulsion device (2) is off; A landing step in which the monitoring unit (4) controls the propulsion device (2) and the actuation means to orient and slow down the flaps (3) and roll over the system (1). In order to carry out the above, the activation of the propulsion device (2) is controlled and the orientation of the flap (3) is individually controlled.

The present invention is a recovery system for recovering a spacecraft that drops a fuselage while generating thrust when landing. The recovery system includes: a plurality of cable members connected to the fuselage and extending toward the outside of the fuselage at the time of fuselage dropping; a plurality of unmanned aircraft that respectively seize the cable members and fly; and a recovery structure that is provided at a landing location of the spacecraft, and that supports the spacecraft in a suspended state as a result of the seized / released plurality of cable members being latched to the unmanned aircraft. At the time of fuselage dropping, the recovery system executes: a step in which the plurality of unmanned aircraft latch the seized plurality of cable members to the recovery structure; a step in which the plurality of unmanned aircraft releases the seizing of the cable members; and a step in which, after thrust stoppage of the spacecraft, the recovery structure suspends and supports the spacecraft.

The invention provides a flexible solar wing unfolding test device and test method, and relates to the technical field of spacecrafts, the flexible solar wing unfolding test device comprises a one-dimensional unfolding unloading mechanism and a two-dimensional unfolding unloading mechanism; in the one-dimensional unfolding unloading mechanism, a first air floating sliding block is arranged on a transverse guide rail in a sliding mode, and the first air floating sliding block is connected with the flexible solar wing through a lifting assembly. In the two-dimensional unfolding unloading mechanism, a second air floating sliding block of a first tackle assembly is arranged on a first longitudinal guide rail in a sliding mode. The first connecting pieces of the first tackle assembly are correspondingly connected with the flexible solar wings respectively; a sliding block of the second pulley assembly is slidably arranged on the second longitudinal guide rail through a rolling piece. And the second connecting pieces of the second pulley assembly are correspondingly connected with the flexible solar wings respectively. According to the flexible solar wing unfolding test device, continuous verification of the solar wing in one-dimensional and two-dimensional unfolding processes can be realized, so that the unfolding reliability and safety of the solar wing are ensured.

The invention provides an offshore efficient maritime launching and recycling system and method for a liquid carrier rocket. The system comprises a wharf plant, a maritime transportation launching platform, a measurement and control command ship, a propellant filling ship, a maritime recycling transportation platform, a rocket horizontal transfer system, a rocket incendiary agent filling system and a rocket oxidizing agent filling system. The method comprises the steps that firstly, final assembly and butt joint of a rocket and a satellite are completed at a wharf; then the satellite-rocket combination body is horizontally rolled to an offshore launching platform and navigated to a preset sea area together with the filling ship; on the sea, the rocket is transferred to a launching position and erected, an oxidizing agent and an incendiary agent are injected through the platform and the injection ship respectively after personnel evacuate, and then ignition and launching are conducted. The whole process is compact and efficient, the whole sea transportation mode is adopted, the rocket body size limitation of land transportation is avoided, meanwhile, the propellant filling ship is independently arranged, and the launching safety is improved.

The invention provides a device and a method for horizontal assembly in a sealed cabin of a large spacecraft. The device is characterized in that the device comprises a front-end support (1), a rear-end support (2), a longitudinal beam (3), a suspended trolley (4) and a rotating platform (5), a cabin body of a sealed cabin (8) is horizontally parked on the rotating platform (5), and the rotating platform (5) drives the cabin body to rotate along a longitudinal axis and is used for enabling the position of equipment to be installed on the cabin body to be aligned with the ground; the front-end support (1) and the rear-end support (2) are of a gantry type, the longitudinal beam (3) enters a cabin body from a front hatch (6) and penetrates out of a rear hatch (7), and the two ends of the longitudinal beam (3) are erected on the front-end support (1) and the rear-end support (2) respectively. The hanging trolley (4) is installed on the longitudinal beam (3), and equipment (9) to be installed in the cabin is hoisted through a lifting rope and transversely moves along the cabin body. The device can work continuously and is high in working efficiency; the device has the advantages of low degree of freedom, simple function and structure, low capital investment and low cost.

The invention provides a novel satellite-rocket electrical interface system, which comprises an interface assembly arranged on the outer surface of a satellite deck, and is characterized in that the interface assembly comprises an X1F connector, an X2F connector and an X10B connector which are arranged side by side; one end of the ground test cable assembly is provided with a first connector used for being connected with satellite ground test equipment, and the other end of the ground test cable assembly is provided with a second connector used for being in butt joint with the X1F connector and the X2F connector; an operation opening is formed in the position, corresponding to the interface assembly, of the rocket fairing, and the operation opening is configured to allow the ground test cable assembly to penetrate through the operation opening to be connected with a satellite connector of the interface assembly or to be manually disconnected and withdrawn before launching. Therefore, the satellite-rocket interface relation is simplified, the reliability of satellite-rocket separation is improved, the carrying and launching capability is improved, and the system development cost is reduced.

A propulsion-based payload transporter (PPT) and methods of its operation are provided. The PPT may operate in a gravitational field such as that on the lunar surface. Via rocket engines that produce thrust, the PPT is configured to descend onto and over a payload to be moved. Once in position with respect to the payload, latches engage with the payload so as to secure the payload in a payload holding bay of the PPT. The rocket engines produce thrust with an exhaust plume that is angled away from the payload to avoid payload damage. The rocket engines, which are located at the sides of the payload holding bay, produce thrust to propel the PPT and the payload in a substantially horizontal trajectory with respect to gravity and toward a landing destination.

The invention relates to a method for testing a power supply system under satellite shadow entering and exiting conditions, and belongs to the field of satellite testing. The method comprises the following steps of: simulating an in-orbit solar wing in-out shadow power supply current change condition through output change of a solar cell simulator of a whole-satellite-level test system, and collecting telemetering data such as charge and discharge current of a power manager and a storage battery of an on-satellite power system; simulation of electrical performance compatibility analysis of a satellite power supply system and a load subsystem under the in-orbit in-out shadow working condition is facilitated, and effective feedback is formed for power supply compatibility between satellite equipment and in-orbit working condition design.

The application discloses a kind of transport, erecting and self-adaptive leveling integrated equipment for small spacecraft, including carrying chassis, erecting mechanism, self-adaptive leveling platform and intelligent control system.Carrying chassis is used to execute transport and parking locking;Intelligent control system is based on launch site surveying and mapping and ideal vertical attitude model, constructs optimal kinematics solution matrix;Data fed back using multidimensional inclination sensor are used to carry out forward reasoning, drive self-adaptive leveling platform to complete coarse leveling;In the process that spacecraft is erected to vertical state by erecting mechanism, based on control error function and PID algorithm closed loop adjustment platform parameter, realize dynamic accurate leveling.The application solves the problems of low efficiency, easy to produce coupling interference and poor environmental adaptability of traditional split type equipment, realizes high-precision automatic collaborative control for complex field terrain, and greatly shortens the launch preparation period.

This invention discloses an experimental apparatus and method for quantifying the uncertainty of the assembly process of a large flexible module. The apparatus includes a large flexible module, a free-boundary suspension system, a non-contact motion capture system, and a disassembly / removal robotic arm. The large flexible module is the test object. The free-boundary suspension system provides boundary conditions for the free movement of the large flexible module, featuring adjustable static position and attitude, and dynamic free boundaries. The non-contact motion capture system is used to measure the position and attitude responses during the assembly process. The disassembly / removal robotic arm is used for assembling and disassembling multiple large flexible modules. The apparatus of this invention provides a ground-based experimental scheme for studying the assembly and disassembly process of large flexible modules in an on-orbit microgravity environment, and can repeat the assembly and disassembly process multiple times. The method of this invention can identify the dynamic characteristics of the structure using the vibration response generated by the large flexible module during the assembly process, and then quantify the uncertainty of the structural dynamic parameters using data from multiple assembly and disassembly processes.

The motorized launch system includes a vehicle having at least one electrically conductive coil arranged to interact with at least one ultra-high conductivity material track in the guide rail when the vehicle is moved through the guide rail. The drive device drives the vehicle at a speed at which a current is induced in the track, the current being sufficient to suspend the vehicle and to inhibit the vehicle from colliding with the guide rail and the track by electric suspension when the vehicle is moving along the track, and drives the vehicle at a firing speed to launch the vehicle from the guide rail. The firing speed is at least as high as the first speed. The controller controls the release mechanism to release the payload from the vehicle to launch the vehicle from the rail while the at least one drive device drives the vehicle at the launch speed.

This invention discloses a propellant venting connector docking and locking device for venting propellant from a rocket being recovered and suspended from a robotic arm on a launch tower. The docking and locking device includes an installation platform; an installation plate driven by a first drive mechanism to rotate vertically on the installation platform; multiple rocket-holding mechanisms evenly distributed circumferentially on the installation plate, driven by a second drive mechanism to synchronously move radially toward each other towards the center and synchronously move radially away from each other towards the center; multiple rocket-bearing mechanisms evenly distributed circumferentially on the installation plate to support the rocket; an industrial robot is installed on the installation plate, and a connector is provided at the end effector of the industrial robot. The connector is driven by the industrial robot to dock with the rocket's venting port and is used to vent the propellant. This invention can effectively reduce the safety risks to personnel involved in the handling of rocket propellant after recovery and improve the efficiency of propellant venting and handling.

The invention provides a carrier rocket ground electrical control system and method, and belongs to the technical field of carrier rocket electrical control. The system comprises a front-end subsystem and a rear-end subsystem which is in communication connection with the front-end subsystem through a redundant optical fiber link, the rear-end subsystem is used for generating a first monitoring instruction, a second monitoring instruction and at least one third monitoring instruction; the front-end subsystem is used for acquiring power supply state information of front-end power supply equipment of the carrier rocket, operation state information of front-end power equipment and other state information according to the first monitoring instruction, the second monitoring instruction and at least one third monitoring instruction; and the front-end subsystem sends the power supply state information, the operation state information and other state information to the rear-end subsystem through the redundant optical fiber link, so that the rear-end subsystem monitors each device of the front-end subsystem in real time. According to the scheme, the communication reliability of the rocket ground electrical system is improved, maintenance is easy and convenient, and expansion is flexible.

Systems and methods disclosed herein are for determining a pre-launch center of gravity (CG) of at least one component of a launch vehicle which may be subject to shifts or changes in CG after launch. The systems and methods may include a tool with an adapter base which may be configured to attach to the component and may include counterweights or segmented weights. The tool may be hoisted with the component for first CG measurements and may be hoisted independently of the component for second CG measurements. The first and second CG measurements may be used to determine the CG of the component of the launch vehicle.