1838 results about "Launch vehicle" patented technology

A practical orbit raising method and system wherein a satellite quickly escapes the Van Allen radiation belts and payload mass and mission life are maximized. A satellite is launched that contains high thrust chemical propulsion thrusters, high specific impulse electric propulsion thrusters and a solar array. The satellite quickly escapes the Van Allen radiation belts by firing the high thrust chemical propulsion thrusters at apogees of intermediate orbits, starting from the transfer orbit initiated by a launch vehicle, to successively raise the perigees until the perigee clears the Van Allen radiation belts. The payload mass and mission life are maximized by firing high specific impulse electric propulsion thrusters to raise the satellite to near synchronous orbit, while steering the thrust vector and solar array to maintain the sun's illumination on the solar array. The chemical and / or electric propulsion thrusters are then fired to achieve geosynchronous orbit.

A spacecraft such as a fly back booster or a reusable launch vehicle, or an aerospace plane has a fuselage and a set of scissors wings consisting of two main wings. Both of the main wings are rotatably mounted on the fuselage and can be yawed at opposite directions. If the spacecraft is launched vertically, both of its main wings can be yawed to be generally parallel with its fuselage so that it can connect with other vehicle or vehicles to form different launch configurations. When the spacecraft or aerospace plane is flying in the air, landing, or taking off horizontally, it can yaw both of its main wings in opposite directions to maximize its lift-to-drag ratio by optimizing the yaw angle of the main wings according to flying conditions. It can also produce desired aerodynamic characteristics such as forming a high drag configuration by adjusting the yaw angle of its main wings. The scissors wings can be used on a spacecraft that is launched vertically on the ground, or a spacecraft that is carried to the air and launched in the mid-air, or a spacecraft that takes off horizontally like an aircraft or glider. The scissors wings can also be used on an aerospace plane.

An aerospace hardware derelict item is salvaged in space, modified, and reused to provide manned facilities in orbit. The hardware packages added onto the salvaged discarded item enhances its value, and incorporates at least the subsystems required to effectively reuse portions of previously discarded launch vehicle components, and other derelict objects in space. The hardware, and technique used reduces the cost of launching comparable hardware to orbit, because of the reuse, and provides a human habitation in orbit. The salvaged items include the external tank of the space shuttle, other derelict orbital hardware, the add-on cargo pod in two forms, and the ability to convert the derelict into a cost effective reusable item. The salvaged hardware is initially capable of contributing mass, length, interior volume, strongback, rotational stability mass, interior pressurized volumes, artificial gravity, and stability with the addition of simple subsystems for salvage, and interior development packages.

A device launcher can be carried and be used to launch a device by an airborne UAV against an airborne or ground target. The target can be a moving target with the device configured to slow or disable the movement of the target. The device can be triggered manually or through an automated targeting system. The launch apparatus can be configured to launch a single device or a plurality of devices. The device launched against the target can be separated from the launch vehicle or tethered to the launch vehicle. The device can be a net that entangles the rotors of a rotor-driven vehicle. The device can include a parachute to retard movement of target vehicle whether it is falling or if it continues to move. A method of patrolling a perimeter with an airborne vehicle includes the steps of moving the airborne vehicle within the perimeter to sense unauthorized targets and launching a device against those sensed targets.

A suspension system for a vehicle is provided that allows for the vehicle body to be tilted wherein the tilt is controlled by the operator's legs in a manner that greatly enhances the stability of the vehicle while producing a fun vehicle to operate. Further, the bottom member of the suspension system is structured in a manner that provides storage capacity for installation of heavy items such as battery banks in order to preserve a lowered center of gravity for the vehicle. The tilting suspension may be used in the context of three-wheeled vehicle having a single rear drive tire and two forward wheels, but apply equally to a four wheeled vehicle as well as to water craft operating on pontoons or snow vehicles operating on skis.

The invention discloses a hard polyimide foaming structural material and a preparation method thereof. The method comprises the following steps of: mixing aromatic dianhydride and / or aromatic acid ester, high molecular alcohol, low molecular alcohol, a catalyst and a surfactant according to a ratio into a polar solvent and reacting to form polyimide foam precursor solution; reacting the precursor solution with isocyanate in a die; pressurizing and foaming to form a polyimide foam intermediate; and performing microwave radiation or / and performing heat curing by using an oven to obtain the hard polyimide foaming structural material. The preparation process route is short and the process is simple; the storage stability of the foam precursor solution is high; and the prepared hard polyimide foaming structural material has uniform aperture, hard quality, high strength, low density and high temperature resistance, is non-toxic and environmentally-friendly in the foaming process and easy to industrialize, and can be applied in the fields of electron, carrier rockets, aviation, ships, railway locomotives and the like.

An adaptor system includes a space vehicle separation plate operably coupled to a space vehicle, a launch vehicle adaptor plate operably coupled to a launch vehicle capable of carrying the space vehicle into space for release of the space vehicle from the launch vehicle, an actuator release mechanism assembly, a bend / shear restrain assembly that is non-coaxial with the actuator-release connector, and a biasing element. The actuator-release mechanism assembly may be configured to separably couple the space vehicle separation plate to the launch vehicle adaptor plate. The actuator-release actuator-release mechanism assembly may pass through the launch vehicle adaptor plate to engage the space vehicle separation plate to hold the space vehicle separation plate substantially parallel to the launch vehicle adaptor plate prior to release of the space vehicle separation plate.

A cargo carrier is disclosed for the efficient delivery of cargo to space, such as to support the International Space Station (ISS). Both pressurized and unpressurized cargo may be delivered into space on an expendable launch vehicle, such as the Delta-IV rocket. The cargo carrier may utilize a slightly modified Delta-IV second stage to provide on-orbit station keeping of the payload until it is transferred to the ISS. The cargo carrier can include an unpressurized section having a rigid central structure supporting a frame to which unpressurized cargo modules are coupled. In addition, a pressurized cargo section may be coupled to the unpressurized section. The cargo carrier may utilize existing on-orbit assets such as the European Automated Transfer Vehicle (ATV) to transfer the ISS cargo from a rendezvous orbit to the ISS.

The invention relates to a method for optimizing an orbital transfer strategy of a geostationary orbit satellite, which comprises the following steps of: 1, determining orbital transfer times, orbital transfer circle times and the controlled variable of each-time orbital transfer; and 2, determining time and a thrust direction in each-time orbital transfer. The process of launching the geostationary orbit satellite at present generally comprises the following steps of: launching the satellite into a highly elliptic transfer orbit with an inclination angle by using a carrier rocket; performing apogee / perigee orbital transfer for several times by using a self-contained liquid engine of the satellite, and transferring to a geosynchronous orbit; and correcting and rounding the inclination angle of the orbit to realize a geostationary orbit. For the satellite, operation for changing the transfer orbit into the geostationary orbit by performing apogee / perigee orbital transfer for several times is complex, so too many orbital transfer times is not suitable, and orbital transfer complexity and risk are prevented from being increased; in addition, factors such as the capacity of the liquid engine of the satellite, arc segment loss in an orbital transfer period, and the like are considered, so too few orbital transfer times is not suitable.

The invention discloses an autonomous and cooperated type aircraft cluster system and a running method. The aircraft cluster system mainly comprises at least one node aircraft cluster, wherein the node aircraft cluster has an autonomous running capability and a network information connection function, the node aircraft cluster is provided with a resource sharing interface, node aircrafts in the node aircraft cluster are connected with each other by a space self-organizing network, the node aircrafts in the aircraft cluster fly in shared navigation and cluster flight modes, the commutation is realized by the space self-organizing network, and the fleet configuration is automatically adjusted according to the task existing condition. When the aircraft cluster receives a task instruction, the resource scheduling, the task planning, the task reconfiguration and the failure regrouping are autonomously carried out according to the task types, so as to further complete the tasks. The autonomous and cooperated type aircraft cluster system is suitable for quick or maneuvering launching of small carrier rockets, the quick response of the space system is greatly improved, and the risk of the in-orbit failure of the aircrafts is reduced.

The invention discloses a low-impact non-pyrotechnic separation device, which is used for the locking and release between stages of a rocket, between a satellite and a launch vehicle, and of a satellite solar panel substrate, an antenna and the like. The low-impact non-pyrotechnic separation device is formed by using a shape memory alloy as a triggering switch and being provided with components and parts of springs, a locking sleeve, a split nut and the like in a matching manner. Based on the shape memory effect of the shape memory alloy, stretching and pre-deformation are firstly carried out on the shape memory alloy during assembly; when separation and unlocking are needed, an alloy material is electrified and heated to be at a phase-transition temperature of the alloy material; the alloy material is recovered to be in a shape before the pre-deformation; thus, a triggering block is driven to act; the locking sleeve is disengaged from the split nut; the release of loads of the satellite and the like is realized. After the one-time separation is completed, a trigger part inside a shell can be conventionally taken out; a shape memory alloy wire is loaded in the low-impact non-pyrotechnic separation device after the pre-deformation is carried out on the shape memory alloy wire once again by adopting a special tool; thus, the low-impact non-pyrotechnic separation device can be repeatedly used. The separation device provided by the invention is low in impact, can be repeatedly used, has no pollution, is low in cost, is safe to use, and is especially applicable to an occasion which is high in the requirement on a separation impact.

Affordable commercial service platforms in space focusing on customer affordability, quality microgravity services, innovation and combining five emerging space technologies. The invention provides an integrated approach to microgravity services in orbit. First, the reusable launch vehicle (RLV) offers affordable transportation services, platform reboost, and eliminates platform subsystems of propulsion, RCS, liquid storage / resupply, and the like. Second, deployable structures are not heavy for the launch, but expand in orbit. Third, orbital phone networks offer customers control of unmanned experiments. Fourth, an enhanced robotic system transfers payloads. Fifth, manufactured thin film solar cells in orbit offer advantages including weight / cost reductions. The orbital service platform has a low initial cost, expands as the market demands, is repairable, offers quality unmanned microgravity, leads to production facilities using similar hardware and offers numerous affordable commercial services.

A capsule contains fluid and a solid inertial mass that is free to move within the capsule. The capsule is embedded in a foam panel. A plurality of such foam panels are attached to the internal wall of a launch vehicle fairing. This device augments acoustic energy dissipation with damping the resonant frequency of the fairing to reduce the amount of energy that is transmitted into the acoustic volume contained within the wall. Incorporating a plurality of capsules respectively tuned to many frequencies provides broadband structural attenuation. This abstract is provided to comply with the rules requiring an abstract, and is intended to allow a searcher or other reader to quickly ascertain the subject matter of the technical disclosure. It is submitted with the understanding that it will not be used to interpret or limit the scope or meaning of the claims.

A high temperature sealing system is operative to seal an interface between adjacent hot structures and to minimize parasitic flow between such structures that move relative to one another in-plane or out-of-plane. The sealing system may be used to seal thrust-directing ramp structures of a reusable launch vehicle and includes a channel and a plurality of movable segmented sealing elements. Adjacent ramp structures include edge walls which extend within the channel. The sealing elements are positioned along the sides of the channel and are biased to engage with the inner surfaces of the ramp structures. The segmented sealing elements are movable to correspond to the contour of the thrust-directing ramp structures. The sealing system is operative to prevent high temperature thrust gases that flow along the ramp structures from infiltrating into the interior of the vehicle.

An active satellite dispenser is preferably attachable to a reusable launch vehicle for deployment of one or more satellites into one or more desired orbits. The active satellite dispenser includes a center mast that releasably receives the satellite(s), a liquid propellant rocket, and an orbital control system on an avionics pallet. In the preferred embodiment, a pressurized gas selectively pressurizes the propellant tanks (which may include fuel and oxidizer tanks), to provide propellant to the rocket. In operation, the launch vehicle releases the satellite dispenser in a first deployment orbit. The active dispenser rocket and orbital control system then transport the active dispenser and satellite(s) into the final deployment orbit. In the preferred embodiment the active dispenser can operate multiple times to place individual satellites in different orbits.

The invention relates to a designing method for a guidance law applicable to the vertical recovery phase of a rocket. The designing method comprises the following steps: acquiring the full-distance standard trajectory of the rocket; according to the full-distance standard trajectory, acquiring the latitude and longitude of a target point, the motion parameters of the rocket at the starting point of an aerodynamic deceleration phase, and the latitude and longitude of a point below the rocket, and constructing an auxiliary coordinate system; based on the auxiliary coordinate system, additionallysetting terminal angular constraints on the basis of a proportional guidance process so as to obtain a biased proportional guidance law; and applying the biased proportional guidance law in the aerodynamic deceleration phase and applying the biased proportional guidance law with an additional main power deceleration algorithm in a vertical descending phase. The improved biased proportional guidance method in the invention can effectively solve problems in guidance control under the condition of multiple constraints by location, velocity and attitude during rocket recovery; the guidance law for a vertical landing phase employs aerodynamic deceleration and tail-phase short-time main-power deceleration, so no impact is exerted on the carrying capacity of the rocket in an in-orbit flying phase; and at the same time, proportional guidance has better adaptability to various non-guidance deviations.

A suspension system for a vehicle is provided that allows for the vehicle body to be tilted wherein the tilt is controlled by the operator's legs in a manner that greatly enhances the stability of the vehicle while producing a fun vehicle to operate. Generally a hub is provided for use ad the vertical members in a four-bar suspension linkage that includes an inboard side and an outboard side such that the tilting pivots are positioned inboard and the steering pivots are positioned outboard. This arrangement allows for the optimum placement of the pivot axis for steering as close to the wheel as possible, even directly within the rim of the wheel, while the pivot point of the suspension linkage is optimally positioned a few inches inboard relative to the wheels thereby increasing the lever arm used to control the relative tilt of the wheel.

The invention relates to a spherical three-degree-of-freedom parallel mechanism antenna structure system, which is formed by connecting an antenna reflector and an antenna base frame. The antenna base frame is formed by connecting an upper platform and a lower platform through three branch mechanisms with the same structure; each branch is formed by connecting an upper platform, an upper circulararc-shaped bracket, a lower circular arc-shaped bracket and a lower platform through three rotating pairs, wherein one rotating pair is connected with a motor; the upper platform is fixedly connected with the antenna reflector; the lower platform is connected with a foundation frame or an antenna carrier frame; and the motor drives the lower circular arc-shaped bracket, the upper circular arc-shaped bracket, the upper platform and the antenna reflector to rotate spherically with three degrees of freedom directly or by a transmission. The spherical three-degrees-of-freedom parallel mechanism antenna structure system can perfectly acquire remote sensing and remote detecting signals and data of satellites, carrier rockets and other aircrafts, and continuously track in overhead airspace of the antenna so as to reach hemispherical working airspace and over-semisphere working airspace.

A low cost, on demand, dedicated launch system is provided for placing micro satellites or space-based instruments at orbital and sub-orbital altitudes and velocities. The invention describes a space launch vehicle (SLV) that incorporates a single, integrated guidance, navigation, and control unit (GNCU) that performs all guidance and control for the SLV from main stage ignition to orbital insertion. The GNCU can remain with the payload after orbital insertion to provide satellite station keeping and orbital maneuvering capability. The use of a single integrated avionics unit for all guidance, navigation, and control simplifies the SLV, reducing weight and significantly reducing cost. In addition, this architecture allows for a combined launch and satellite bus system as the GNCU can also be used as a satellite bus. This further reduces cost and increases the payload capacity to orbit by optimizing the use of launch vehicle and satellite bus subsystems and reducing non-instrument mass delivered to orbit. All support functions are provided by the IDMV. This approach represents a significant improvement over conventional systems, especially with respect to the orbital launch of payloads less than about 100 kg.

The invention relates to an octagonal battery-equipped array upright post micro satellite configuration. A satellite is designed into a non-equilateral symmetrical octagonal upright post structure; simultaneously, the octagonal configuration and the internal space design of the satellite are optimized; the satellite is partitioned into a platform cabin and a load cabin; a cabin section between a top plate and a middle plate is the platform cabin; and a cabin section between the middle plate and a bottom plate is the load cabin. Each cabin section is partitioned into three regions through a +Y clapboard and a -Y clapboard; according to equipment enveloping size, reasonable layout and arrangement are performed, the design of a main force transfer path is optimized and cable direction and the insertion space of an electric connector are fully considered, so that a compact and reasonable structural space can be ensured; simultaneously, an enveloping space provided for carrying rocket is fully utilized and the solar cell installation area of a satellite body is enlarged to the greatest extent.

The invention provides configuration of a satellite loaded with a large deployable antenna with double reflecting surfaces. The relative position relation of a feed source array, a main reflecting surface and an auxiliary reflecting surface in an in-orbit flight state is guaranteed, main force transfer path design is optimized through design of a payload capsule in a truss structure style and reasonable layout and placement while the constraint conditions about enveloping of a fairing of a carrier rocket and correct deployment of the large deployable antenna are met, compactness and reasonableness of structure space of the large deployable antenna in a folded state can be guaranteed, meanwhile, the deploy path requirement of the large deployable antenna is met, and the area of a satellite solar cell array is guaranteed by fully using the enveloping space provided by the carrier rocket.

A high capacity broadband service is provided from an Earth orbiting satellite having a payload that includes multiple large antennas, the satellite having an inverted configuration when deployed on orbit. The satellite has an aft surface proximate to a launch vehicle structural interface, a forward surface opposite to said aft surface, and a structure disposed therebetween. In a launch configuration, two or more deployable antennas are disposed, undeployed, on or proximate to and forward of the forward surface; on-orbit, the payload is operable when the satellite is disposed with the aft surface substantially Earth facing while the forward surface is substantially facing in an anti-Earth direction. Each deployable antenna is disposed, so as to be Earth facing, when deployed, in a position substantially forward and outboard of the forward surface.

An intelligent control method for vertical recovery of carrier rockets based on deep reinforcement learning is disclosed. A method of autonomous intelligent control for carrier rockets is studied. Theinvention mainly studies how to realize attitude control and path planning for vertical recovery of carrier rockets by using intelligent control. For the aerospace industry, the autonomous intellectualization of spacecrafts is undoubtedly of great significance whether in the saving of labor cost or in the reduction of human errors. A carrier rocket vertical recovery simulation model is established, and a corresponding Markov decision-making process, including a state space, an action space, a state transition equation and a return function, is established. The mapping relationship between environment and agent behavior is fitted by using a neural network, and the neural network is trained so that a carrier rocket can be recovered autonomously and controllably by using the trained neural network. The project not only can provide technical support for the spacecraft orbit intelligent planning technology, but also can provide a simulation and verification platform for attack-defense confrontation between spacecrafts based on deep reinforcement learning.

The invention disclosed here is a pressurization system for pressure-fed launch vehicles. The system employs a hot pressurant gas to minimize the mass of pressurant gas required. It also employs a set of baffles within the propellant tank to reduce heat transfer between the propellant and the hot pressurant gas. The baffles keep the pressurant gas flowing uniformly in one direction as the propellant is expelled, and inhibit mixing of the cold gas near the propellant with the hot gas being blown into the tank. The baffles also prevent large-scale sloshing of the propellant. The metal baffles are rigidly attached to one end of the tank and attached with a mount which allows travel in the longitudinal direction but not in the two lateral directions which minimizes stess on the pressure vessel from differences in thermal expansion.

An apparatus comprises a core having a first surface configured for attachment to a surface of a structure, a face sheet located over a second surface of the core, a number of cavities within an interior of the core, and a number of ports for the number of cavities. The number of ports provides communication between the number of cavities within the interior of the core and the exterior of the core. The number of cavities and the number of ports are configured to reduce noise traveling through the core.

The invention relates to a redundancy switching circuit used for a ground test launch and control system of a carrier rocket, which mainly comprises an FPGA module, a singlechip, a switching relay and a manual transfer switch, wherein the FPGA module realizes an automatic switching function and an Ethernet switching function, and the switching priority level of the Ethernet is high; the singlechip receives the interface signal of the Ethernet, transmits the signal to the FPGA module, meanwhile receives the master-slave machine working condition from the FPGA module and carries out real-time display; an amplifying circuit receives the output signal of the FPGA module and carries out power amplification; the relay receives the drive signal of the amplifying circuit and executes action; and the manual transfer switch is used for the manual switching involved by people under an emergency state, and the priority level of manual switching is the highest. The invention ensures that the test launch and control system is not disturbed when in outputting, thereby ensuring the reliability and the safety of the whole test launch and control system. The invention solves the problem that a rocket is unattended before being launched because of low reliability of general test launch and control equipment, thereby meeting the demands of high reliability and safety of the carrier rocket on the ground test launch and control system.

A suspension system for a vehicle is provided that allows for the vehicle body to be tilted wherein the tilt is controlled by the operator's legs in a manner that greatly enhances the stability of the vehicle while producing a fun vehicle to operate. Further, the bottom member of the suspension system is structured in a manner that provides storage capacity for installation of heavy items such as battery banks in order to preserve a lowered center of gravity for the vehicle. The tilting suspension may be used in the context of three-wheeled vehicle having a single rear drive tire and two forward wheels, but apply equally to a four wheeled vehicle as well as to water craft operating on pontoons or snow vehicles operating on skis.

The invention provides a device and a method for repairing a friction stir welding weld defect of a storage tank for a carrier rocket. A repair welding pin tool comprises a tool pin clamping structure and a tool pin, the end face of a shaft shoulder portion of the tool pin clamping structure is a recessed slope provided with two concentric-circle annular grooves, the front of the tool pin is of a conical thread structure, and a conical thread is circumferentially provided with three spiral grooves at intervals. The method includes the steps: firstly, manually excavating the tool pin broke at a friction stir welding weld; secondly, performing solid filling for the weld; thirdly, utilizing the repair welding pin tool to start friction stir welding from a center of the broke position of the tool pin filled with welding wires or materials to a set welding end point; and finally, continuing utilizing the repair welding pin tool to perform friction stir welding for the weld again from a welding start point to the welding end point. The method and the device are capable of guaranteeing the quality of the storage tank for the carrier rocket and decreasing scrapping of the storage tank for the carrier rocket due to the key hole type defect.

The invention belongs to a thrust vector control servo system for a manned space carrier rocket or missile which has a demand for high reliability to provide thrust for an engine or a steering engine on the rocket or the missile, and particularly relates to a novel multi-redundancy electromechanical hydrostatic servo mechanism. The novel multi-redundancy electromechanical hydrostatic servo mechanism has the advantages that since a servo motor is directly used for driving a bidirectional rotating constant plunger pump, the defect that a valve-controlled electro-hydraulic servo system is apt to be blocked is overcome and the requirement of high reliability is satisfied; since a controller, a motor driver, a servo electric pump and a displacement sensor adopt multi-redundancy configuration, under the situation of at most two-redundancy faults, the system can still work normally and the feasibility of high reliability is guaranteed from structural design; since peak power is provided for a load through the maximum overload of an alternating-current permanent magnet motor and the alternating-current permanent magnet motor is designed according to load average power, compared with an alternating-current permanent magnet motor which is designed according to load maximum power, the size is small, the weight is light and the requirements on long-term use, light weight and small size are better satisfied.

A suspension system for a vehicle is provided that allows for the vehicle body to be tilted wherein the tilt is controlled by the operator's legs in a manner that greatly enhances the stability of the vehicle while producing a fun vehicle to operate. Generally a hub is provided for use ad the vertical members in a four-bar suspension linkage that includes an inboard side and an outboard side such that the tilting pivots are positioned inboard and the steering pivots are positioned outboard. This arrangement allows for the optimum placement of the pivot axis for steering as close to the wheel as possible, even directly within the rim of the wheel, while the pivot point of the suspension linkage is optimally positioned a few inches inboard relative to the wheels thereby increasing the lever arm used to control the relative tilt of the wheel.