2065 results about "Rocket" patented technology

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 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.

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 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 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.

The invention relates to an industrial CT composite material defect reconstruction method based on an unsupervised neural network. The method comprises the following steps: carrying out industrial CT shooting to obtain a cross section two-dimensional image of a composite material component; an unsupervised model for converting an industrial CT composite material image into a defect model for reconstruction and discrimination is constructed, sampling is performed through a multi-scale sampling model, a channel attention mechanism is introduced for weight adjustment, and a multi-scale residual module is introduced for reconstruction; constructing a joint loss function for calculating the unsupervised model; training the constructed unsupervised model by using a data set, wherein a loss function is used for updating parameters; and inputting a to-be-detected industrial CT composite material image into the model, so that three-dimensional reconstruction of the defects of the composite material can be completed. According to the method, the defects of the composite material are reconstructed with high precision, nondestructive testing of the composite material is achieved, and the method can be used for structural health monitoring and defect detection of aerospace vehicles such as airplanes and rockets.

The invention discloses a carrier rocket flight load analysis method based on flight measurement data, and relates to the technical field of spaceflight, and the method comprises the steps: measuring a flight load in a rocket flight process, and obtaining actual flight strain data and actual flight trajectory data; according to the actual flight strain data and a pre-constructed load-strain corresponding relation, calculating a rocket body total load of each station on the rocket body; calculating the rocket body static load of each station in the actual flight process by using the actual flight trajectory data; subtracting the corresponding static load of the rocket body from the total load of the rocket body at each station to obtain the dynamic load of the rocket body at each station; and optimizing the flight load dynamic load theoretical calculation method by using the rocket body static load of each station and the rocket body dynamic load of each station to obtain an optimized flight load calculation method. The rocket body structure load in the actual flight process can be obtained through load measurement in the flight process, and therefore the static load of actual flight and the dynamic load of actual flight are obtained in a back-stepping mode.

The invention discloses a satellite angular rate identification method and related equipment, and relates to the technical field of rocket launching, and the method comprises the steps: obtaining a rocket telemetry video frame sequence after satellite-rocket separation; for each frame of image of the rocket telemetry video frame sequence, determining a dynamic image area of the target satellite based on the static area mask; based on a feature detection algorithm, binary descriptors are extracted from dynamic image areas of two adjacent frames of images; based on the binary descriptor, matching corresponding feature point pairs between two adjacent frames of images; estimating a rotation matrix between two adjacent frames of images through an iterative optimization algorithm based on the corresponding feature point pairs; and determining the angular rate of the target satellite according to the rotation matrix and the time interval between the two adjacent frames of images. According to the method, rapid and high-precision identification of the satellite angular rate is realized, the problem that the traditional technology depends on indirect data and is high in delay is solved, task success or failure can be judged in real time, and the launching task evaluation efficiency is improved.

The invention discloses a segmented self-destruction rain enhancement and hail suppression rocket controlled by a time schedule controller, and relates to the technical field of artificial rain enhancement and hail suppression rockets. The device comprises a catalysis bin assembly, an engine assembly and an empennage assembly, and the catalysis bin assembly, the engine assembly and the empennage assembly are connected through reducing sleeve assemblies. An explosion self-destruction device is adopted as a safe landing device, a segmented self-destruction device is arranged in a rocket body engine and a tail segment assembly in advance, a time schedule controller outputs an ignition signal according to binding information, an empennage assembly, an engine assembly and a catalytic bin assembly are exploded into floccules not larger than 100 g and fall onto the ground, self-destruction of all the assemblies is reliable, and the safety of landing is guaranteed. The rocket body wreckage preventing device can effectively reduce damage of rocket body wreckage to personnel, buildings, farmlands, livestock and the like, any one of the ignition loops of the time schedule controller does not affect performance of other loops during working, fire spurting is avoided, space is saved, and the function is reliable.

The embodiment of the utility model provides a storage tank bottom structure, a storage tank and a carrier rocket, the storage tank bottom structure comprises a shell and a top cover, the shell is of a hollow arc-shaped revolving body structure which is through in the front-back direction, the two ends of the shell are the small-diameter end and the large-diameter end respectively in the axis direction of the shell, and the small-diameter end of the shell is fixedly connected with the top cover; a propellant is arranged on the top cover; a plurality of protruding structures are evenly distributed on the outer side face of the shell in the circumferential direction in the axis direction of the shell. According to the technical scheme, the bottom structure of the thrust bearing type storage tank is simple in process, easy to manufacture in batches and beneficial to achieving light weight of a carrier rocket.

The invention relates to a generalized carrier rocket trajectory calculation method which comprises the steps that a target orbit and constraint conditions corresponding to the target orbit are set, the target orbit comprises a sun-synchronous orbit or a low-dip-angle orbit, and the constraint conditions at least comprise the geocentric vector diameter, the absolute speed and the flight path angle; input parameters are simplified, wherein the input parameters comprise configuration parameters, mass parameters, pneumatic parameters and engine parameters; and based on the constraint condition of the target orbit, the corresponding control variable and the input parameter, calculating the value of the control variable through a Newton iteration method so as to obtain the trajectory of the carrier rocket. The invention provides a universal carrier rocket trajectory calculation method. The carrying capacity of a carrier rocket is ensured by optimizing a typical target orbit; according to the method, input parameters of trajectory calculation are simplified and carded, the trajectory is obtained through calculation of the Newton iteration method, the adaptability to different configuration schemes is improved on the premise that the calculation accuracy is guaranteed, rapid trajectory iteration calculation can be conducted, and generalization is achieved.

The invention provides a carrier rocket control method based on wind field characteristics, and the method comprises the steps: obtaining a wind field model generated through the fitting of actually measured wind field data; according to the speed of the carrier rocket at the current moment under the launching coordinate system, the wind speed at the current moment and the first conversion relation, the relative speed of the carrier rocket relative to the atmosphere at the current moment in the aircraft coordinate system is determined; influence parameters are calculated through the relative speed, and the aerodynamic force at the current moment is calculated according to the influence parameters; integrating the mass center motion differential equation set of the carrier rocket by combining the aerodynamic force, the preset initial value of the preset control parameter and the flight program angle expression of the carrier rocket changing along with time, and circularly calculating the preset state variable of each moment; calculating a target function of the carrier rocket according to the preset state variable at each moment; and on the basis of a sequential quadratic programming algorithm, target functions corresponding to different preset control parameters are calculated through the target functions and preset constraint conditions, so that optimal values of the preset control parameters are determined.

The invention discloses an autonomous power-up circuit and a satellite power supply system. The autonomous power-up circuit comprises a satellite-rocket separation enabling unit, a storage battery discharge switch unit and a bus switch unit; wherein the satellite-rocket separation enabling unit is connected with the storage battery discharge switch unit and the bus switch unit, and is used for controlling the storage battery discharge switch unit and the bus switch unit to be in a conducting state when a satellite and a rocket are separated and injected into an orbit; the storage battery discharging switch unit is connected with a storage battery pack of the satellite and the bus switch unit, the bus switch unit is connected with a load of the satellite, and the storage battery discharging switch unit and the bus switch unit are used for providing electric energy output by the storage battery pack for the load when being in a conducting state. According to the autonomous power-up circuit and the satellite power supply system provided by the invention, when the satellite and the rocket are separated, autonomous power-up can be carried out on the satellite through the storage battery pack, and meanwhile, the problem of uncontrollable power-up of the satellite after a fairing is thrown in an active section can be avoided.

The invention relates to the technical field of verification and evaluation, in particular to a reusable rocket attitude control system redundancy verification and fault tolerance evaluation method, which comprises the following steps: performing multi-dimensional combination on disturbance units according to disturbance types, action time periods and intensity grades to form a disturbance combination set, constructing a group of induction task profiles covering different attitude control abnormal situations; recording a switching state of an internal control path of the attitude control system, a redundant component intervention time sequence and a function takeover sequence in real time, and constructing a dynamic response path map of a redundant link; and generating a response integrity index and a fault tolerance conflict degree criterion of redundancy configuration as a fault tolerance capability evaluation result of the attitude control system. Compared with an existing method which only verifies individual static disturbance scenes, the method has the advantages that systematic design and composite excitation of disturbance conditions can be realized, scene complexity and coverage of redundancy verification are enhanced, and authenticity and representativeness of fault-tolerant testing are improved.