2775 results about "Rocket" patented technology

An attitude control simulation method for a launch vehicle comprises: receiving a pulse accumulation sent by a laser strapdown inertial measurement unit (11); performing calculation processing according to the pulse accumulation to obtain target calculation results (12), wherein the target calculation results comprise a navigation calculation result, a guidance calculation result and an attitude control calculation result; and sending the navigation calculation result, the guidance calculation result and the attitude control calculation result to a simulation server side (13). By means of the attitude control simulation method for a launch vehicle, the problems whereby a real environment is difficult to be simulated, a verification method is not sufficiently intuitive and the safety risk is high when a launch vehicle is tested are solved, and the simulation and optimization of the launch vehicle on a computer under different working conditions are realized without experimental verification, thereby saving on the time and cost, and accurately predicting attitude responses and control effects of the launch vehicle under different working conditions.

The invention relates to the field of aerospace data transmission and monitoring, and discloses a carrier rocket power supply data transmission and monitoring method and device and a storage medium, and the method comprises the following steps: obtaining historical data for analyzing a network state, and predicting network transmission abnormality based on a long short-term memory network model; dynamically adjusting a threshold value through a sliding window, and triggering an abnormal response mechanism; a reliability enhancement mechanism is adopted, data transmission is optimized through a TCP exponential backoff algorithm and a UDP redundancy coding technology, and data loss and retransmission are reduced; a Q-Learning algorithm in reinforcement learning is utilized to dynamically adjust transmission parameters according to a network state, the transmission parameters comprise a window size, a retransmission interval and a data packet size, and self-adaptive optimization is realized. The method is suitable for reliable transmission of key data and instructions in launch tasks of carrier rockets, and can be widely applied to data monitoring of spaceflight launch, high-concurrency network transmission and complex scenes.

The invention discloses an online efficient load shedding control method based on a forecast wind field. The method comprises the following steps: firstly, acquiring a load attack angle and a sideslip angle which are calculated in real time on a rocket; amplitude limiting processing is carried out on the load attack angle and the sideslip angle, an attack angle control instruction is generated after gain control network calculation is carried out, and an attitude control system carries out direct attack angle control in a load sensitive area in combination with the load shedding requirement and the guidance deviation correction capacity; and pitching and yaw feedback control instructions are generated, active load shedding control is implemented, and errors generated by attack angle control are compensated. The on-line active and passive load shedding mode covers the real deviation or interference influence during flight, the generated load shedding instruction is better matched with the actual flight, the universality is good, the implementation is easy, and the load shedding effect is better.

The present disclosure discloses a method for designing a terminal guidance law based on deep reinforcement learning, and relates to the field of missile and rocket guidance. The method includes: establishing a relative kinematics equation between a missile and a target in a longitudinal plane of a target interception terminal guidance section of the missile; to adapt to a research paradigm of reinforcement learning, abstracting a research problem and modeling as a Markov decision process; building an algorithm network, and setting algorithm parameters, where a selected deep reinforcement learning algorithm is a deep Q-network (DQN); in a terminal guidance process of each round, obtaining a sufficient number of training samples through Q-learning, training a neural network and updating a target network at fixed frequencies respectively, and continuously repeating the above process before set learning rounds are reached.

The invention discloses an RFID-based weather modification rocket projectile operation full-process monitoring and management method and system, and the method comprises the steps: assembling a high-temperature-resistant and impact-resistant RFID tag for each rocket projectile, and writing identification information and operation parameters; the storage position and the transportation track of the ammunition are monitored in real time through a fixed reader-writer network; the matching relation between the rocket projectile and the launcher is verified before launching, and an interception instruction is generated if not; and generating an ammunition management report based on the whole-process collected data and marking an exception. According to the invention, the whole life cycle accurate tracing of the rocket projectile from production to recovery is realized, the operation safety is ensured through a multiple verification mechanism, the meteorological data is fused to construct an intelligent decision system, and the management efficiency and scientificity of the figure operation are significantly improved. The system has the characteristics of quick response, complete data, convenient operation and the like, and provides an innovative solution for modern management of figure operation.

The invention provides an offshore launch rocket stabilizing device which comprises a first anti-overturning device and a second anti-overturning device, the first anti-overturning device and the second anti-overturning device are fixedly connected with a launching pad, and the first anti-overturning device and the second anti-overturning device are connected to the two sides of a rocket in a supporting mode correspondingly; the first anti-overturning device and the second anti-overturning device are the same in structure; each of the first anti-overturning device and the second anti-overturning device comprises an anti-overturning device underframe, an anti-overturning arm and a bracket; the bottom end of the anti-overturning arm is installed on the anti-overturning device bottom frame, the bracket is fixedly connected to the top end of the anti-overturning arm, and the bracket is used for being fixedly connected to the side portion of a rocket body of a rocket in a clasping mode. The end, close to the rocket, of the bottom of the anti-overturning arm is rotationally connected with the anti-overturning device underframe, and the other end of the anti-overturning arm is supported on the anti-overturning device underframe through a supporting and rapid separating mechanism. After the rocket is erected and parked on the launching platform, the rocket body is stabilized, the rocket is prevented from toppling over, and the stability and safety of the rocket are improved.

The invention provides a reusable rocket landing stage starting height prediction method, a guidance system and a processor, and aims to dynamically and accurately determine the ignition time of an engine by constructing a real-time online starting height prediction model based on an energy conservation principle and combining multiple safety thresholds (including an allowable starting height upper limit and a standby starting height lower limit) so as to predict the starting height of the rocket in a landing stage. And the safety and fuel optimality in the whole flight envelope are ensured, so that the reliability, accuracy and economical efficiency of the reusable rocket landing are remarkably improved. The problems that in the prior art, engine starting height prediction precision and robustness are insufficient, online adaptation to uncertainty of a complex flight environment cannot be achieved, and safety margin consideration is insufficient are solved, and potential safety risks are thoroughly eliminated.

Disclosed are systems and methods for a distributed space transportation network. Satellite launches to orbit are more efficiently performed by large rockets. Modern satellites are in smaller form factor, leaving the large launch rockets with excess capacity. Small satellite operators can use ride-shares, but do not have efficient options for delivering their satellites to their desired destination and may be forced to operate their satellites in compromise orbits. The disclosed distributed space transportation network maintains a fleet of space tugs, which can dock with satellites in space at an initial arrival destination and transport the satellites to their final destinations. In one embodiment, the space tugs can dock with satellite depots to obtain fuel and repairs.

The invention relates to a non-pyrotechnic secondary satellite and rocket locking and separating device which comprises a shape memory alloy primary separating assembly, a pressing and releasing assembly and a turntable four-claw secondary clamping and separating assembly, and the shape memory alloy primary separating assembly is arranged in a box body of the pressing and releasing assembly; one end, extending out of the shell, of a grooving bolt shaft of the shape memory alloy first-stage separation assembly penetrates through a through hole of the separation block to be connected with a nut, and the shape memory alloy rod provides driving force for the shape memory alloy first-stage separation assembly to jointly form a first-stage unlocking mechanism; a clamping jaw guide groove of the rotary table four-jaw two-stage clamping and separating assembly is fixed to the bottom plate, a rotary table controls a clamping jaw set to be locked, and clamping jaws are connected with a separating block of the pressing and releasing assembly to jointly form a two-stage unlocking mechanism. When the device is locked, after being locked by the shape memory alloy first-stage separation assembly, the device is locked by the turntable four-claw second-stage clamping separation assembly; when the device is separated, the device is unlocked by the turntable four-claw secondary clamping and separating assembly, and then is unlocked by the shape memory alloy primary separating assembly.

Ejection seat systems include an ejection seat and a propellant assembly mounted to the ejection seat. A nozzle assembly having an adjustable nozzle is moveably attached to the ejection seat and configured to receive propellant exhaust from the propellant assembly and exhaust through the adjustable nozzle. A first linear electric actuator is operably connected to the adjustable nozzle and arranged to adjust an orientation of the adjustable nozzle in a first direction and a second linear electric actuator is operably connected to the adjustable nozzle and arranged to adjust an orientation of the adjustable nozzle in a second direction. The second direction includes a directional component that is normal to the first direction. A control module is configured to control operation of the first linear electric actuator and the second linear electric actuator.

The utility model discloses a rocket type catching net launching device which comprises gun barrels, gas bombs, a percussion mechanism, a catching net, a cartridge receiver and a folding telescopic rod, the gas bombs are arranged in the corresponding gun barrels, the percussion mechanism is arranged at the tail ends of the gun barrels, the catching net is connected with the gas bombs, and a fixing part is fixedly installed at the front end of an inner cavity of the cartridge receiver. The tail end of the gun barrel is detachably installed on the front side of the fixing component, and the folding telescopic rod is arranged at the front end of the gun barrel. The high-pressure gas cylinder is adopted as launching power, compared with power sources such as ammunition and fuel, higher safety is achieved, materials are easy to obtain, cost is lower, the unfolding angle of the gun barrel can be freely adjusted according to actual requirements, the larger the unfolding angle is, the higher the unfolding speed of the catching net is, and the closer the flying distance is. The method is suitable for close-range combat; otherwise, the smaller the unfolding angle is, the slower the unfolding speed of the catching net is, the farther the flying distance is, and the catching net is suitable for long-distance combat.

The invention provides a launching pad and a carrier rocket launching method. The launching pad comprises a launching pad body provided with a through flow guide opening, a first heat-proof plate, a second heat-proof plate and a driving structure. The second heat-proof plate detachably covers the inner side wall of the second edge side of the flow guide opening, the first heat-proof plate is rotationally arranged on the inner side wall of the first edge side of the flow guide opening, and the driving structure is arranged at the bottom of the launching pad body and used for driving the first heat-proof plate to rotate to be in an unfolded state or a folded state. And the first heat-proof plate is horizontally arranged at the top of the flow guide opening. During maintenance, the second heat-proof plate can be quickly disassembled so as to facilitate maintenance or replacement, meanwhile, the first heat-proof plate can be rotated to a horizontal unfolding state through the driving structure, a worker stands on the first heat-proof plate to maintain the heat-proof coating, high-altitude operation is not needed, and the safety and convenience of maintenance are improved; and the heat-proof coating is coated in the horizontal state, so that adhesion and uniform distribution of the heat-proof coating are facilitated.

The invention relates to the field of rocket recovery, and discloses a rocket capture arm and a recovery tower. The rocket capturing arm comprises a main body; the supporting piece is arranged on the main body in a liftable mode through a first connecting rod assembly and a second connecting rod assembly and used for supporting a rocket; the first connecting rod assembly comprises at least two first connecting rods arranged in the vertical direction, one end of each first connecting rod is rotationally connected with the body, and the other end of each first connecting rod is rotationally connected with the supporting piece. The second connecting rod assembly comprises at least two second connecting rods arranged in the vertical direction, one end of each second connecting rod is rotationally connected with the body, the other end of each second connecting rod is rotationally connected with the supporting piece, and the plane where the second connecting rod assembly is located and the plane where the first connecting rod assembly is located form an angle. The buffering assembly is connected between the main body and the supporting piece and used for providing resistance in the descending process of the supporting piece. The non-coplanar layout of the two sets of connecting rod assemblies constructs a composite supporting structure in a three-dimensional space, so that the degree of freedom of an original single-plane four-connecting-rod is counteracted by rigid constraint of crossed connecting rods.

The invention provides a free boundary supporting system for a horizontal modal test of a large rocket. The free boundary supporting system at least comprises a rack car, a pressure container, a free membrane type air spring and a control assembly. The rack car at least comprises a bottom truss and a top truss, and the bottom truss is provided with a moving device; the pressure container is specifically a gas storage container located between the bottom truss and the top truss, and the top of the pressure container is provided with a sealing flange connector. The free membrane type air spring is arranged on the upper end face of the top truss, the bottom of the free membrane type air spring communicates with the pressure container through the sealing flange connector, and a bracket used for supporting a rocket is arranged on the top of the free membrane type air spring. The control assembly at least comprises an acquisition module, a controller and an execution module which are in communication connection in sequence; freedom degrees of translation, rotation and torsion of the rocket in a modal test are provided by the free membrane type air spring.

The invention discloses an offshore rocket erection state anomaly detection method based on sparse representation and adaptive filtering, and the method comprises the following steps: obtaining state data in a rocket erection process through multiple sensors, constructing a time sequence, setting a sliding window with change capability in each time period, and extracting sparse features for state prediction. A feedback adjustment signal is formed by calculating the difference between a predicted value and an observed value, and parameters in the state estimation process and the external disturbance modeling process are adjusted respectively. A double-path structure is adopted to independently estimate the state and disturbance, and a final estimation result is obtained through a fusion strategy. The system continuously updates parameters according to error conditions, continuous self-adaptive adjustment and anomaly recognition are achieved, and the recognition precision and the response capability of the anomaly trend are improved. The method is suitable for the recognition processing of the abnormal state in the erection process of the rocket in the marine environment, and has high dynamic adaptive capacity and time sequence anomaly detection capacity.

The invention relates to a dynamic modeling method for an attitude control system under a dynamic fault. The method comprises the following steps: determining a modeling hypothesis of a nonlinear dynamic model; establishing a coordinate system used for modeling; rocket configuration parameters and engine related parameters are obtained; building displacement and acceleration descriptions of the three types of structures on the rocket; establishing a rigid body translation equation, a rigid body rotation equation, a rocket body elastic vibration equation, a propellant shaking equation and a sensitive element measurement equation; summarizing the five established equations to form a nonlinear dynamic equation set; and performing small-disturbance linearization expansion on the nonlinear dynamic equation set on the basis of the nominal reconstructed trajectory to obtain a linearized small-deviation equation set suitable for controller design and stability analysis. According to the method, the influence of the thrust drop fault is comprehensively considered, and the model can be used for carrying out relatively accurate dynamic response analysis and reconstruction scheme design under the fault.

The invention belongs to the field of rocket recovery, particularly relates to an offshore recovery rocket sealing device and a using method thereof, and provides the following scheme that the offshore recovery rocket sealing device comprises a rocket sealing conveying mechanism movably mounted on an offshore landing platform, and the top end of the outer side wall of the offshore landing platform is rotationally sleeved with a rotating platform with a rotating driving mechanism; a rocket sealing transportation mechanism is movably mounted on one side of the rotating platform through two fixing blocks and a locking bolt, and a weight box is mounted on the other side of the rotating platform; the rocket sealing transportation mechanism comprises a loading cylinder with a sealing gate plate at an opening in one side, and an auxiliary fixing mechanism I with a plurality of clamping mechanisms is mounted in the loading cylinder. According to the device, rockets landing on any landing point on the offshore landing platform can be rapidly clamped and fixed in an auxiliary mode, then the rockets are loaded into the loading cylinder to be transported, the problem that the rockets topple over in the harbor returning process of a rocket recovery platform is solved, and the transportation safety is high.

An improved liquid-propellant engine utilizing a combination of a gas turbine, a generator, an electric pump fed engine, and a power management system. When in use, the gas turbine and generator are configured for providing power to an electric motor for operating a fuel pump and an oxidizer pump to pump fuel and oxidizer from individual tanks into a combustion chamber for mixture and combustion. The combustion of the fuel and oxidizer is configured for providing thrust from the base of a rocket to propel it. The generator of the present invention provides continuous electrical power to the electric pump(s) to continue the supply of fuel and oxidizer to the combustion chamber. The power management system of the present invention is electrically connected to the generator, the electric motor(s) configured for operating the fuel and oxidizer pumps, a flight computer, and fuel and oxidizer valves.

The invention provides a high-precision optical pose measurement method based on tower recovery, which adopts a close-range bidirectional optical cooperation mode, measures a mark point of a recovery tower through an optical detector on a rocket, and determines the distance and the pose of the rocket relative to the recovery tower; and the optical measurement system of the recovery tower synchronously detects and identifies the rocket and controls a capture arm and a buffer mechanism of the recovery tower to move, and optical measurement data of the recovery tower are fed back to the rocket and are integrated with optical measurement data on the rocket to control landing of the rocket. According to the invention, optical measurement and navigation are introduced into rocket tower type recovery, and the method has the advantages of high measurement precision, high data frequency, interference resistance and the like. The relative relation between the rocket and the recovery tower can be directly obtained through optical pose measurement, compared with inertial navigation, the coordinate conversion process can be reduced, and the advantage of no accumulated error is achieved. Compared with satellite navigation, optical pose measurement does not depend on external signals, and higher independence and electromagnetic resistance are achieved.

The invention relates to a working platform in the bottom of a rocket storage tank and a mounting method of the working platform. The bottom of the rocket storage tank at least comprises a scalloped segment and a manhole formed in the bottom of the rocket storage tank. The working platform comprises a supporting mechanism, a mounting mechanism and a plurality of standing mechanisms, one end of the supporting mechanism is used for extending into the bottom of the rocket storage tank through a manhole, and when one end of the supporting mechanism is located in the bottom of the rocket storage tank, the mounting mechanism is detachably mounted at the end, located in the bottom of the rocket storage tank, of the supporting mechanism; the multiple standing mechanisms are detachably mounted on the mounting mechanism in the circumferential direction of the mounting mechanism in sequence, buffer layers are mounted at the ends, away from the mounting mechanism, of the standing mechanisms, and the tail ends of the buffer layers are used for abutting against the side walls of the ends, located in the bottom of the rocket storage tank, of the scalloped segments. The occupied area of the rocket storage tank can be reduced, the product quality of the rocket storage tank is guaranteed, the working efficiency is improved, and danger is reduced.

A tailor-welding, spinning and variable-wall-thickness forming method for a large-sized extension section of a liquid propellant rocket engine, which belongs to the technical field of metal forming. The method comprises: performing laser numerical control machining on a GH3044 superalloy to obtain two chord-tangent circles; performing electron beam welding to obtain a complete circle; performing a vacuum solid solution treatment; performing pre-forming to obtain a shallow dish shape; performing a vacuum solid solution treatment; assembling a variable-wall-thickness flat blank and a spinning mandrel of a first spinning die, and then performing spinning on a spinning machine to form a curved generatrix variable-wall-thickness blank part; performing a vacuum solid solution treatment; and assembling the curved generatrix variable-wall-thickness blank part and a spinning mandrel of a second spinning die, and then performing spinning on the spinning machine to form a curved generatrix variable-wall-thickness part. The present invention is mainly used for the manufacturing of an ultra-large-sized GH3044 extension section of the liquid propellant rocket engine, which has a diameter of 1200 mm and a length of 1500 mm, and the product precision reaches a wall thickness tolerance of + / -0.05 mm for each point and a surface profile of ≤3 mm.

The utility model provides a supporting and enclasping structure. The supporting and enclasping structure comprises an enclasping mechanism, the enclasping mechanisms are rotatably mounted on the two sides of the transverse moving mechanism, the supporting mechanism is mounted in the middle of the transverse moving mechanism, and the supporting mechanism and the enclasping mechanisms on the two sides form a circular hoop shape; the enclasping mechanism and the transverse moving mechanism are connected through a transverse moving support in a structural part hinging mode, an electric push rod is connected with the enclasping mechanism and the transverse moving mechanism, and the electric push rod drives the enclasping mechanism to rotate relative to the transverse moving mechanism through stretching out and drawing back; the supporting mechanism is fixed to the transverse moving mechanism through bolts. The multi-direction adjusting device can adapt to multi-direction adjustment of the rocket, and can effectively support the loading rocket; the holding arm mechanism can realize holding in the rocket transportation and erection process, the opening angle of the holding arm mechanism is large in the installation process, a larger space is provided for lifting of the rocket, dead point self-locking can be realized, and the reliability is improved; the structure is compact and the integration level is high.

The invention provides a rocket power anchor. The rocket power anchor comprises a bullet of a spiral structure. The anchoring structure is detachably connected with the warhead, and the anchoring structure can stretch; the second-stage engine is detachably connected with the anchoring structure, a second-stage delay igniter capable of driving the anchoring structure to stretch is arranged at the joint of the second-stage engine and the anchoring structure, and the second-stage engine is filled with a second-stage propelling grain; the first-stage engine is detachably connected with the second-stage engine, a middle partition bin is arranged at the joint of the second-stage engine and the first-stage engine, a first-stage delay igniter is arranged on the middle partition bin, and the first-stage engine is filled with a first-stage propelling grain; the exhaust nozzle can be obliquely arranged at a certain angle, so that the rocket power anchor can rotate during movement; the lifting ring is rotationally connected to the first-stage engine; the total mass of the first-stage propelling grain is smaller than that of the second-stage propelling grain; and a pressure-bearing diaphragm is arranged on the middle partition bin.

The invention discloses a space-based telemetering transmission system for a carrier rocket, and relates to the technical field of rocket measurement and control communication, and the system comprises a rocket-borne network camera which is used for collecting video image data of a preset monitoring point position of the rocket, compressing the video image data, and outputting the compressed video image data as telemetering data; the telemetering transmission equipment is respectively connected with the rocket-borne network camera and the ground equipment, performs telemetering framing on the telemetering data, outputs first telemetering data and outputs second telemetering data to the ground equipment; the space-based power amplifier unit equipment is connected with the telemetering transmission equipment, receives the first telemetering data, amplifies the first telemetering data and then outputs the first telemetering data through radio frequency; the space-based telemetering antenna is connected with the space-based power amplifier unit equipment, receives the first telemetering data and then transmits the first telemetering data to a relay satellite, so that the satellite sends the first telemetering data to a ground measurement and control station; and the ground equipment is used for receiving the second telemetry data and then analyzing and uploading configuration parameter information. The problem that the coverage blind area of a telemetering system mainly based on ground-based measurement and control is large in the prior art is solved.

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 provides an adjustable launcher for a rocket-boosted unmanned aerial vehicle. The adjustable launcher is mainly used for solving the technical problems that the pitching angle of an existing launcher for the rocket-boosted unmanned aerial vehicle cannot be adjusted, the launching direction cannot be changed by 360 degrees according to the wind direction, the requirement for ground flatness is high, and flexibility is poor. The platform comprises a pitching platform, an azimuth platform and a fixed platform. The fixed platform is arranged on the ground through ground nails, and the gradient of the ground can be horizontally adjusted; a slewing bearing is arranged between the azimuth platform and the fixed platform, the azimuth platform is driven by the slewing bearing to rotate, rotation adjustment of the azimuth platform is facilitated, and the unmanned aerial vehicle takes off along the wind; a second screw elevator is arranged between the pitching platform and the azimuth platform and used for adjusting the pitching angle when the unmanned aerial vehicle is launched, and the take-off effect of the unmanned aerial vehicle is guaranteed. The take-off device is only composed of a mechanical structure, the structure is simple, and a guarantee is provided for take-off of the rocket-assisted unmanned aerial vehicle.

The invention discloses a carrier rocket ascending section control method, system and equipment, and relates to the field of rocket flight control. Constructing a carrier rocket ascending section rigid body dynamic model and a carrier rocket ascending section elastic vibration model; forming a dynamic model of the rising stage of the carrier rocket; rewriting into a speed coordinate system and carrying out linearization processing to obtain a small-disturbance linearization model; obtaining an extended state observer of the rocket system according to the small disturbance linearization model; performing stability derivation on the extended state observer to obtain a first ESO open-loop transfer function and a second ESO open-loop transfer function; adjusting the first preset parameter to obtain a first control result of the rising stage of the carrier rocket; adjusting a second preset parameter to obtain a second control result of the rising stage of the carrier rocket; the carrier rocket ascending section first control result and the carrier rocket ascending section second control result represent the carrier rocket ascending section control result. According to the invention, the precision of attitude stable tracking and load shedding control in the flight process of the rising stage of the carrier rocket is improved.

The invention discloses a rocket seaborne alignment method based on federated filtering and a system thereof. The rocket seaborne alignment method based on federated filtering comprises the following steps of performing equipment power-up and filter filtering parameter initialization; after equipment power-up and initialization of filtering parameters of the filter are completed, whether coarse alignment calculation can be carried out or not is judged; if the coarse alignment calculation can be carried out, carrying out coarse alignment calculation to obtain an attitude angle; and after coarse alignment calculation is completed and an attitude angle is obtained, fine alignment calculation is performed. According to the method, data of the multiple sensors are fused through the federated filtering algorithm, and the alignment precision of the rocket during offshore launching is remarkably improved. Meanwhile, the influence of the complex marine environment on rocket alignment can be effectively dealt with, and the stability of the launching process is improved. Moreover, by optimizing the structure of the federated filter, the fault tolerance performance of the system is improved, and even if a certain sensor breaks down, accurate alignment of the rocket can be ensured.

The invention provides multichannel spinning and turning integrated equipment for a cylindrical part of a carrier rocket and a machining method. The multichannel spinning and turning integrated equipment comprises a lathe bed, a main spindle box, a closing-in spindle box, a spinning roller frame, a tail top, a heating system, a control system and a hydraulic system. The main spindle box, the closing-in spindle box, the spinning roller frame and the tail top are installed on the lathe bed. The main spindle box is connected with the closing-in spindle box and drives the closing-in spindle box to work, a cylindrical part blank is clamped between the closing-in spindle box and the tail top, the spinning roller frames are located on the two sides of the cylindrical part blank, and the heating system is installed on the spinning roller frames; the control system is connected with the main spindle box, the closing-in spindle box, the spinning roller frame, the heating system and the hydraulic system. The hydraulic system is connected with the main spindle box, the closing-in spindle box, the spinning roller frame and the tail top. Through cooperation of the main spindle box, the closing-in spindle box, the spinning roller frame, the tail top and the like, the in-situ machining function combining forced spinning machining, closing-in machining and turning machining of cylindrical parts, conical parts and curved-bus-shaped parts of the carrier rocket is achieved.

The system and method for decelerating space objects moving at high velocities are provided. The system comprises at least one closed volume containing a combination of gas, liquid, solid particles, or a mixture thereof. The volume can be transported to the targeted location using a range of means, including a chemical gun, light gas gun, electromagnetic coil gun, superconducting quench gun, a rocket, or a combination thereof. The volume is strategically positioned on the trajectory of the moving object. Upon penetrating the walls of the volume, the object passes through it, experiencing deceleration.