327results about "Systems for re-entry to earth" patented technology

Transpiration-cooled systems having permeable and non-permeable portions are provided. In one aspect, a nozzle for a rocket engine includes a non-permeable outer layer, a plurality of walls, a permeable layer, and a plurality of cooling channels. The non-permeable outer layer forms a predetermined shape of the nozzle and defines an interior space. The plurality of walls extend generally orthogonal to the non-permeable outer layer and into the interior space. The permeable layer is disposed between each set of adjacent walls of the plurality of walls. The permeable layer is spaced a distance from the non-permeable outer layer. The plurality of cooling channels are formed between the permeable layer, the non-permeable outer layer, and a set of adjacent walls of the plurality of walls. The plurality of cooling channels are configured to transport a fluid.

The invention relates to the technical field of space flight and aviation, in particular to a recoverable rocket landing supporting mechanism which comprises a plurality of folding, unfolding and folding mechanisms. Each folding, unfolding and folding mechanism comprises an on-arrow locking mechanism, an auxiliary air cylinder, a telescopic assembly, a supporting assembly and a buffering device. The upper end of the telescopic assembly is hinged to the rocket body, the lower end of the telescopic assembly is hinged to the bottom end of the supporting assembly through a buffering device, the top end of the supporting assembly is hinged to the rocket body through a left hinge joint and a right hinge joint, and the left hinge joint and the right hinge joint are coaxial and located below the upper end of the telescopic assembly. The locking mechanism on the rocket is located on the rocket body and locks the edge of the supporting assembly after the supporting assembly is tightly attached to the rocket wall; one end of the auxiliary air cylinder is fixed to the arrow body, the other end of the auxiliary air cylinder extends outwards and downwards, the auxiliary air cylinder is located below the telescopic assembly, and when the supporting assembly is tightly attached to the arrow wall, the other end of the auxiliary air cylinder makes contact with the supporting assembly. The invention has the characteristics of large folding-unfolding ratio, reliable locking, high landing stability, small impact peak value and small aerodynamic resistance.

The invention relates to the technical field of cosmic sailing, and particularly discloses an offshore rocket recovery system and method, and the system comprises a floating body which floats on the sea surface; the inclination measuring unit is arranged on the floating body and is used for measuring the inclination angle between the floating body and the sea surface and generating an inclination signal; the recovery unit comprises a recovery deck, a rocket fixing device and a plurality of lifting devices, the rocket fixing device is installed on the recovery deck, the lifting devices are arranged on the periphery of the bottom of the recovery deck and installed on the floating body, and the lifting devices are used for driving the recovery deck to move up and down; and the control unit is arranged on the floating body, the control unit is in signal connection with the inclination measuring unit and the lifting device, and the control unit is configured to receive an inclination signal of the inclination measuring unit and control the lifting device to drive the recovery deck to move up and down. According to the offshore rocket recovery system and method, the rocket can be stably recovered, the recovery efficiency is improved, and the cost is reduced.

The invention discloses a rocket power landing phase guidance control method, system and device and a medium, and relates to the technical field of rocket guidance recovery. The method comprises the following steps: in a current control period, obtaining a state quantity according to a control parameter and a rocket power landing section motion model; taking the minimum guidance control comprehensive target mathematical model as a target, and updating the to-be-adjusted parameter value according to the parameter updating period and the state quantity; according to the guidance cycle, the updated partial to-be-adjusted parameter values, partial parameters in the state quantity, the flight time and the guidance model, a thrust amplitude and a channel expected attitude are obtained; and according to the attitude control period, the updated part of parameter values to be adjusted, the attitude, the attitude angular rate, the thrust amplitude, the channel expected attitude and the attitude control model, obtaining a channel engine swing angle, and entering the next control period for control. According to the method, the problems of low terminal constraint satisfaction precision, fuel consumption loss, poor attitude control quality and the like can be solved, and the reliability of recoverable rocket power landing is improved.

Container for collecting space debris. Space debris ingress through container's openable access port. A barrier deployment mechanism has a rotatable roll of flexible material on a frame assembly. A barrier translation mechanism translates the frame assembly towards or away from the access port to adjust the size of the compartment. When unwound, the roll forms a compartment for containing the space debris within the interior of the container, thus forming an obstruction between the compartment and access port.

A tiered spacecraft docking station is adapted to facilitate docking of spacecraft within outer space. A first tier includes a first frame enclosing a first area. A first net-like mesh is coupled to the first frame and fills the first area enclosed by the first frame. A second tier includes a second frame enclosing a second area. A second net-like mesh is coupled to the second frame and fills the second area enclosed by the second frame. A plurality of support beams attach the first frame to the second frame. A lander is used to slow and stop a spacecraft on a celestial body. The lander includes a first and a second webbed structure and a decelerator coupled to the first webbed structure and / or the second webbed structure. The decelerator maintains a tension in the first and / or second webbed structure below a predetermined threshold.

The invention relates to a low-cost parachute landing recovery system for a microgravity load return capsule, and aims to realize an efficient, reliable and economical recovery scheme through modular design, all modules are connected through standard interfaces, and assembly, maintenance and upgrading are facilitated. The structure module is used for assembling the parachute module and providing necessary shock resistance; the parachute module is used for pneumatic deceleration and attitude stabilization of the system and adopts a multi-stage unfolding structure; and the initiating explosive device module is used for bouncing the brake parachute, releasing the brake parachute and releasing a parachute closing opening. According to the system, the production and maintenance cost is remarkably reduced through modular design, meanwhile, the reliability and adaptability of the system are improved, and the system is suitable for various microgravity load return scenes and has wide market application prospects.

Various embodiments comprise a multi-agent spacecraft system for rendezvous and proximity operations. The spacecraft system comprises a chief spacecraft and a deputy spacecraft. The chief spacecraft maneuvers to the orbit of a Resident Space Object (RSO) and deploys a deputy spacecraft to rendezvous with the RSO. The deputy spacecraft maneuvers to the RSO and senses the RSO during the rendezvous and proximity operations maneuver. The deputy spacecraft generates a Three-Dimensional (3D) surface reconstruction of the RSO based on the sensing. The deputy spacecraft determines a roll rate and a relative trajectory of the RSO based on the sensing. The deputy spacecraft returns to the chief spacecraft.

The invention relates to the technical field of aerospace, and provides a multi-stage pneumatic telescopic mechanism and a rocket recovery landing leg device.The multi-stage pneumatic telescopic mechanism comprises a plurality of barrels which are arranged in a sliding and sleeving mode from inside to outside; the outermost cylinder is connected with an end cover, and the end cover is provided with an unfolding inflation hole; in every two adjacent barrels, a vent hole is formed in the end, away from the end cover, of the outer barrel, a locking groove is formed in the inner wall of the outer barrel, and a locking component is arranged at the end, close to the end cover, of the inner barrel. The locking component is switched from the locking state to the unlocking state through the vent hole of the outermost cylinder, the inner cylinder can slide to contract, and when the two adjacent cylinders are in the contraction state, the vent holes of the two adjacent cylinders communicate with each other, so that the locking state of the locking component is changed, and the service life of the locking component is prolonged. The corresponding locking grooves are unlocked by all the locking parts by means of the vent holes of the outermost cylinder body; the extension of the multi-stage barrel body is realized by unfolding the inflation hole, so that the extension and retraction of the multi-stage barrel body can be realized through the vent hole in the outermost barrel body and the unfolding inflation hole, and the design complexity of an air path is reduced.

The invention relates to the technical field of aerospace, and provides a multi-stage pneumatic telescopic mechanism and a rocket recovery landing leg device.The multi-stage pneumatic telescopic mechanism comprises a plurality of barrels which are arranged in a sliding and sleeving mode from inside to outside, and the barrel on the outermost side is provided with an unfolding inflation hole; in every two adjacent barrels, the outer barrel is provided with a vent hole, and the inner side of the outer barrel is provided with a locking part; a locking groove is formed in the inner side barrel body; the locking member has a locked state and an unlocked state. The locking component can be switched from the locking state to the unlocking state through the vent hole of the outermost cylinder body, so that the inner cylinder body slides to contract, and the vent holes of the two adjacent cylinder bodies are communicated under the condition that the two adjacent cylinder bodies are in the contraction state, so that the locking component is locked. The corresponding locking grooves are unlocked by all the locking parts by means of the vent holes of the outermost cylinder body; and extension of the multi-stage cylinder body is realized through the unfolding inflation hole, so that extension and retraction of the multi-stage cylinder body can be realized through the vent hole in the outermost cylinder body and the unfolding inflation hole, and the complexity of air path design is reduced.

To solve the problem that there is a case in which it is difficult to recover a flying object.SOLUTION: A recovery device has an engine part which burns supplied propellant to obtain a thrust in a predetermined direction, and is configured to be mounted on an airframe of a recovery object such that the recovery device can obtain, with the engine part, the thrust in a direction opposite to a direction in which the airframe of the recovery object obtains a thrust.SELECTED DRAWING: Figure 1

Systems and methods for a fully reusable upper stage for a multi-stage launch vehicle are provided. The reusable upper stage uses an aerospike engine for main propulsion and for vertical landing. A heat shield can include a plurality of scarfed nozzles embedded radially around a semi-spherical surface of the heat shield, wherein inboard surfaces of the plurality of scarfed nozzles collectively define an aerospike contour. The heat shield can be actively cooled to dissipate heat encountered during reentry of the upper stage.

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.

A system and method are disclosed for softly capturing, protecting, and landing a spacecraft returning from space using a VTOL (vertical takeoff and landing) aircraft including drones and helicopters. The spacecraft is decelerated by a parachute. One or more VTOL drones transport a waterproof pocket that catches and captures the descending spacecraft in the air. In this way, the spacecraft is protected inside the pocket and, after continuing to descend, makes a soft landing in a body of water. In another embodiment, a recovery helicopter, which is a type of VTOL aircraft with heavy-lift capabilities, is used to directly catch the returning spacecraft. One or more VTOL drones are attached to the lower end of a recovery cable suspended from the helicopter. These drones carry a clutch that directly, quickly, and accurately catches the descending spacecraft without interfering with the parachute. In this way, the spacecraft is caught and protected by the helicopter while maintaining the floating function of the parachute.

Tether-based reaction control systems for spacecraft undergoing atmospheric entry or reducing orbital energy, including systems that use atmospheric drag on one or more tethers to alter the angle of attack of the spacecraft during its atmospheric transit to reduce the thermal and mechanical loads on the spacecraft during the reentry process. Systems may include tethers that together generate a force that functionally acts on the spacecraft at a point that is offset from the spacecraft's center of mass. Such systems may be either actively or passively controlled.

A composite article comprises a body comprising a radial region, a tapered region, and a cavity region. The cavity region comprises a cylindrical shape and an inner diameter of the cylindrical shape decreases between the radial region and the tapered region. A nosecone region is adjacent to the tapered region of the body and at an opposing end of the body to the cavity region. A thermal protection structure is on the body and comprises a base layer, an insulating layer is on the base layer, and an erosion resistant layer is on the insulating layer. The insulating layer exhibits a specific gravity of from about 0.01 to about 1.0. Methods of forming the thermal protection structure and of thermally protecting an article are also disclosed.

The invention discloses a rocket sublevel active rotor pneumatic deceleration recovery system and a control method, and aims to solve the problems of low recovery precision, large fuel consumption, large landing impact and the like. The recovery process is divided into five stages of reentry, rotor wing unfolding, main pneumatic deceleration, accurate approaching and soft landing. Fine adjustment of the descending speed and the horizontal position is achieved through cooperative control of aerodynamic resistance of the body and the rotor wing system. And a multi-source sensor and a trajectory tracking algorithm are fused, so that stable soft landing at the sub-level zero horizontal speed and the minimum vertical speed is finally ensured. According to the invention, the recovery precision is obviously improved, the fuel consumption is reduced and the landing impact is reduced.

To enable making a rocket ultra-small-sized, a robot body light-weighted, a rocket manufacturing cost significant reduced, so as to facilitate capture of space debris.SOLUTION: An entire outer wall of a rocket body, including a fairing at a head of a space rocket, is made of wood. In order to use wood material for the entire outer wall of the rocket body, the wood material must be treated to be tougher in terms of heat resistance than conventional aluminum veneer. In an inner wall of the rocket body, all of inner wall surfaces are lined with fireproof fabric that can withstand fires up to 1700°C, or heat insulating fabric, and toughened wood technology is used to ensure that rocket inner wall will not have any problems with heat resistance until the rocket ascends and reaches outer space. Further, in order to reduce weight of the rocket, the wood materials to be used is to be selected from those having a low specific gravity, such as cedar or hinoki.SELECTED DRAWING: None

The invention provides a system device for core-level offshore recovery of a carrier rocket and a control method. The system device comprises an offshore recovery platform body, a core-level recovery capture rotary locking device, a dumping thrust reverser, an anti-dumping damping device, an auxiliary holding device and a locking device moving platform. The locking device moving platform can move to any position on the offshore recovery platform body; the core-level recycling, capturing and rotating locking device is used for receiving core-level position signals and supporting vertically descending cores; the dumping thrust reverser is used for holding the core; the anti-toppling damping device is matched with the toppling thrust reverser to enable the core to topple to a horizontal state; the auxiliary clasping device is used for clasping a core in a horizontal state. According to the invention, the carrier rocket core-level can be converted from a vertical fixed state to a horizontal fixed state after being recovered on the offshore recovery platform, the core-level mass center is reduced, the overturning risk of the core-level and the offshore recovery platform body is greatly reduced, and the harbor returning safety after core-level offshore recovery is improved.

This patent discloses a powerplant for an aerial vehicle comprised of Perovskite-Silicon tandem photovoltaic solar cells covering the wings and fuselage, a lithium-sulfur battery, a high-pressure unitized regenerative proton exchange membrane (“PEM”) device, and hydrogen tanks. The PEM device has a fuel-cell mode and an electrolysis mode. During level flight, the PEM device operates in fuel-cell mode, converting hydrogen into electricity. The electricity is used to run a plurality of pairs of permanent magnet synchronous motors, coupled to propellers, and mounted in a coaxial rotor configuration. During level flight, the array of solar cells re-charges the Li—S battery pack. During takeoff and landing, the Li—S battery pack supplements the electricity generated by the PEM device in fuel-cell mode. On the ground, the solar cells provides electricity to the PEM device, which operates in electrolysis mode, converting water into hydrogen gas, which is then stored in the hydrogen tanks.

The invention discloses a conical deceleration airbag and a folding and unfolding supporting mechanism thereof, and belongs to the technical field of aerospace air-drop. A connecting ring with a section of sleeve, a short connecting rod hinged to the connecting ring, a long connecting rod hinged to the short connecting rod and a base hinged to the long connecting rod are sequentially arranged on the outer side of the structure from top to bottom; the inner side structure is sequentially provided with an air chamber from top to bottom, a spherical lock clamping groove formed in the connecting ring is formed below the air chamber, a sleeve below the clamping groove is connected with a center rod with a spherical lock locking device, and an air pipe with the top end fixedly connected with the air chamber is arranged in the whole columnar space. The foldable supporting mechanism is added to the air bag structure, the structure can effectively support normal unfolding of the shape of the air bag and provide a space matched with the folded state of the air bag when the air bag is folded, meanwhile, due to the fact that the structure does not interfere with the air bag, the air bag can be inflated at various opportunities, and the design of more air bags can be met.

Connecting device for establishing a mechanical connection between a first body (11) and at least one second body (12), comprising a first connecting element (20) and a second connecting element (22) cooperating therewith, wherein the second connecting element (22) has an engagement portion (27) for a coupling portion (28) of the first connecting element (20), and wherein the first connecting element (20) and the second connecting element (22) are designed to form a positive connection when connected to one another, characterized in that the engagement portion (27) is formed on a holding component (29) which is connected to a core of the second connecting element (22), wherein the holding component (29) and / or the second connecting element (22) consists of or comprises a thermoplastic material, and in that at least one spring (3) is provided,which applies a force which repels the two connecting elements (20, 22) and / or the two bodies (11, 12) from each other and moves the two connecting elements (20, 22) away from each other above a predetermined temperature threshold (T1).

The invention relates to a method for planning a satellite maneuver, in particular for planning an orbital change of a satellite orbit after the completion of a mission (240), with the satellite (220) having a drive unit (150) with a ground-based power supply, at least comprising the steps of: - configuring at least one ground-based laser system (100) for producing a laser beam, in particular laser pulses; - configuring at least one laser ground station (200) including the at least one laser system (100); - designing an optical transmission path (120) for transmitting the laser pulses from the at least one ground station (200) to the satellite (220); - configuring the satellite-based drive unit (150), in particular an ablation drive where mass is removed from the drive unit (150) by means of the laser pulses; - determining a trajectory attainable by means of the drive unit (150), with the steps being run through in an iteration until the desired orbital change has been achieved.

A system and method for decelerating space objects are provided. The system comprises at least one enclosure that is inflated by a medium, such as a gas, liquid, solid particles, or any mixture thereof. The enclosure can be transported to a predetermined orbital location in the path of the space object. Upon penetrating the enclosure membrane, the object passes through the medium and is thereby decelerated leading to decay of its orbit and / or disintegration thereof.

The present invention relates to a method for collaborative obstacle avoidance for multi-node landers under non-uniform constraints, comprising: controlling the multi-node lander to descend to a preset altitude and acquiring an environmental map of a star catalog; calculating the trajectory of the multi-node lander's virtual center and the position of each node; determining the real-time distance of each node relative to the virtual center and calculating each node's own trajectory based on flexible transformation constraints such as the multi-node lander's configuration potential field, obstacle avoidance potential field, and collision avoidance potential field; and controlling each node to track its own trajectory to achieve collaborative obstacle avoidance. This method enables future asteroid probes to robustly and accurately adhere to landing sites of high scientific value during landing and sample return missions, and to cope with various hazardous terrains, such as steep slopes and narrow valleys.

The application relates to a parachute opening failure discrimination method and device and a recovery landing point judgment method. An enhanced data set is obtained by performing enhancement processing on an obtained data set, and the generalization capability and robustness of a training model during training can be improved by using the enhanced data set; a threshold setting method is used to mark the data set to obtain a labeled data set, and the labeled data set comprises a first parachute image and a second parachute image; the actual projection area is corrected by reconstructing the opening area of the second parachute image, so that the actual opening area of the parachute can be more accurately obtained, and the accuracy of model training is improved; in the model construction stage, a lightweight failure discrimination model is used, the accuracy and operation efficiency are considered, the hardware requirement is reduced, the model has the characteristics of lightweight portability, and is more easily deployed in a small spacecraft recovery project.

The present application relates to a kind of big buffering capacity arrow body vertical damping buffer mechanism in small space, belong to the field of reusable launch vehicle;The vertical damping buffer mechanism includes unfolding piece, locking piece, connecting piece, buffer device, landing piece and unlocking piece, with light quality, easy operation, easy replacement by ingenious structure design, can effectively improve mechanism installation, replacement, maintenance efficiency, reduce processing cost, and the vertical damping buffer mechanism is locked in the tail section inside arrow body in rocket ascent segment, not protruding from the outside of arrow body, without additional setting device such as pneumatic cover, without occupying the external space of arrow body, maintain the aerodynamic shape of arrow body, the internal space of arrow body occupied when retracted is also smaller;And preferably adopt layered crush buffering method can reduce structure size under the premise of guaranteeing bearing and buffering performance.

The invention relates to the technical field of aerospace craft water recovery, and provides a buoyancy air bag assembly for aerospace craft recovery, a recovery system and an aerospace craft, the aerospace craft comprises the recovery system, the recovery system comprises a buoyancy air bag assembly and a mounting part, and the buoyancy air bag assembly comprises a buoyancy air bag and a push-out air bag; the push-out air bag penetrates through and is fixedly connected to the installation opening of the buoyancy air bag, the first end of the push-out air bag is located in the buoyancy air bag, the second end of the push-out air bag is located outside the buoyancy air bag, and the push-out air bag is used for pushing the buoyancy air bag in the first direction during inflation. When the aerospace craft is recycled, the push-out air bag in the buoyancy air bag assembly pushes the buoyancy air bag and the mounting piece in the inflating process, the buoyancy air bag is pushed out of the mounting piece, the buoyancy air bag can be smoothly and safely popped out, the buoyancy air bag is prevented from being clamped, the buoyancy air bag is rapidly inflated under the condition that the buoyancy air bag is smoothly and safely popped out, and the safety of the aerospace craft is improved. The bursting risk of the buoyancy air bag is reduced, and smooth work of the buoyancy air bag is guaranteed.

The present invention relates to the field of spacecraft return landing technology, and in particular to a cold air reverse thrust and air cushion landing device for a manned spacecraft, comprising a return capsule, a parachute installed in the top of the return capsule, a jet assembly installed in the bottom of the return capsule, a metal bottom fixedly connected to the bottom of the return capsule, a cavity formed between the inner wall of the metal bottom and the bottom surface of the return capsule; the jet assembly comprises an air supply portion arranged in the bottom of the return capsule, the air supply portion is connected to a plurality of nozzles, and the nozzles are arranged in the cavity. The present invention ejects the gas in the air supply portion through the nozzle, so that the gas ejected by the nozzle can not only generate reverse thrust, but also form a supporting force in the cavity to perform air cushion cushioning on the return capsule when the metal bottom is close to the ground, which can effectively consume the impulse of the return capsule, so that the return capsule can be buffered and land smoothly.