985 results about "Launch vehicle" patented technology

The present invention relates to a digital accompanying flight system for a launch vehicle based on real-time telemetry data, and a method. The system comprises a ballistic simulation model, a propulsion simulation model, a mass simulation model, an aerodynamics simulation model, an environment simulation model, a timing model, and a trigger correction model. The ballistic simulation model, the propulsion simulation model, the mass simulation model, the aerodynamics simulation model, and the environment simulation model synchronously perform simulation. On the basis of a preset flight sequence and real-time telemetry data, the timing model determines, at each simulation step length, whether a key timing action has occurred, and simulates the key timing action by adopting a discrete event message. On the basis of telemetry data of a launch vehicle which is obtained in real time, the trigger correction model determines whether the ballistic simulation model needs to be corrected, and when correction is needed, sends a correction message to the ballistic simulation model.

An orbit insertion control method and system for a launch vehicle. For a guidance loop and an attitude loop, equivalent second-order error dynamics models are constructed, respectively; by making full use of the capability of fully characterizing the system design degree of freedom of a second-order fully actuated system theory, comprehensively considering multiple task requirements such as dynamic performance, control constraints, and interference suppression, and introducing an iteration idea into a discrete design framework, hybrid rapid optimization design of guidance loop and attitude loop control gains is achieved, effectively improving overall design efficiency, and design margins can be essentially released, improving the accuracy and performance of orbit insertion control; moreover, guidance loop feedback gain optimization design and attitude command solving are implemented, essentially improving the performance and robustness of orbit insertion control, and attitude loop control design is implemented, essentially ensuring the tracking performance, greatly reducing the conservative property of conventional separate design of dual-loop gains.

The invention belongs to the technical field of rocket launching, and particularly relates to a carrier rocket launching perpendicularity control method and device and a storage medium, the perpendicularity control method comprises the following steps: constructing an attitude dynamic model in a carrier rocket launching process, and the attitude dynamic model is used for describing change rules of rocket attitudes under different conditions; setting a prediction window; according to the real-time attitude state information of the rocket, predicting attitude changes of a plurality of time steps in the future in the current prediction window; designing an objective function, and optimizing the objective function through a predictive control algorithm in each time step to obtain an optimal control signal sequence; in each prediction period, executing a first signal in the obtained optimal control signal sequence, and transmitting the first signal to a rocket attitude control device; and repeating the steps. According to the invention, the launch attitude control precision of the carrier rocket can be improved, and the response speed of the system is optimized.

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 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 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 provides a carrier rocket connector shedding monitoring system and method based on multi-source data. The system comprises a multi-source data acquisition module used for acquiring initial image data, initial sensor data and a host instruction of a carrier rocket connector; the multi-source data processing module is used for receiving and processing the initial image data and the initial sensor data to obtain target image data and target sensor data; the intelligent detection module is used for receiving the target image data, the target sensor data and the host instruction, and performing fusion processing on the target image data, the target sensor data and the host instruction so as to realize comparison between an actual state and a theoretical state and obtain a comparison result; and the falling early warning module is used for receiving the comparison result and performing early warning under the condition that the comparison result does not meet the preset condition. According to the scheme, the state of the carrier rocket connector is automatically monitored, and the fault phenomena such as abnormal falling and untimely falling of the connector are early warned in time.

Implementations of the disclosed subject matter may provide an adaptive thrust structure for a space launch vehicle that may include an interface block that is fluidically coupled to a fuel supply, a fuel manifold, an oxidizer supply, and an oxidizer manifold. A first adapter plate may be coupled to a first engine assembly. First adapter lines may be fluidically coupled to the first engine assembly and the interface block, via the first adapter plate. A second engine assembly, second adapter lines, and a second adapter plate are capable of replacing the first engine assembly, first adapter lines, and first adapter plate. Implementations of the disclosed subject matter may provide a reusable and reconfigurable launch vehicle that may include a reusable second stage, and a reusable and reconfigurable transporter having a plurality of zones that are configurable to transport cargo, one or more passengers, and / or one or more satellites.

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 provides an electric pump pressure type liquid rocket engine and a carrier rocket. The electric pump pressure type liquid rocket engine comprises a gas-driven generator system, an electric pump pressure type liquid rocket engine body and an energy management system, and the gas-driven generator system is electrically connected with the energy management system; and the electric pump pressure type liquid rocket engine is electrically connected with the energy management system. According to the technical scheme, the rocket is prevented from carrying a large-mass battery through the generator system, and the highest-efficiency operation of the generator under all working conditions is realized through the energy management system. According to the technical scheme, the energy loss is small, the structural mass is light, the development and application cost of the carrier rocket is reduced, and the reliability of the carrier rocket is improved.

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 provides a carrier rocket engine stage-switching motor redundancy control system, which is combined with a multi-input instruction redundancy voting method, a voting path software and hardware arbitration method and the like to complete high-reliability carrier rocket engine stage-switching motor control. According to the system, two-out-of-three voting is carried out by using a triple-redundancy input control instruction and interactive heartbeat, software and hardware joint arbitration is carried out on voting results of three modules, and an instruction control channel is determined to complete phase sequence control of the stage conversion winding. The method has strong fault capability, has no single-point failure link, and can ensure the normal operation of the engine conversion motor in the flight process of the carrier rocket.

The present invention discloses a method for attitude fault-tolerant control of a reusable launch vehicle with performance constraints, belonging to the field of automatic control. A multivariable adaptive PI fault-tolerant controller is designed to achieve performance constraints for the attitude model of the reusable launch vehicle during the powered landing phase, considering the actual engineering applications. During the entire controller design process, the complexity brought by uncertainties, external disturbances, and engine failures is considered. First, performance constraint limits are designed for the transition variables including attitude angle error and attitude angular velocity error. Secondly, new variables are derived using performance conversion techniques, and by controlling the stability of these variables, the transition variables are restricted within the performance function. Finally, a multivariable adaptive PI fault-tolerant controller is designed to stabilize these variables. The present invention can handle the performance constraint problem during the rocket powered landing process, thereby improving the success rate of the rocket's soft landing.

The invention relates to the technical field of rockets, in particular to a supporting device for a rocket and a transferring and erecting integrated launching vehicle. The supporting device comprises an erecting arm body and a plurality of auxiliary supporting mechanisms. The auxiliary supporting mechanism comprises a supporting bracket, a disc spring force limiter and at least two linear guide rails. The pre-pressure of the disc spring force limiter is adjustable, so that the rigidity of the supporting bracket is matched with the rigidity of the rocket body; the linear guide rails are arranged on the two sides of the supporting bracket respectively. According to the supporting device for the rocket and the transferring and erecting integrated launching vehicle, the supporting bracket is kept to be in a balanced state through the linear guide rail, the floating function is achieved, and the clamping stagnation phenomenon is prevented; force limiting supporting is achieved by adjusting the pre-pressure of the disc spring force limiter, the rigidity of the supporting bracket is made to be close to the rigidity of a rocket body, and therefore it is guaranteed that the additional bending moment cross section load generated by the erecting arm to the rocket is within the allowable range.

The utility model relates to an integrated light-weight launch vehicle tail section and launch vehicle, the integrated light-weight launch vehicle tail section is formed by welding two semicircular wall plates, the wall plates comprise a first arc plate and second arc plates located on the two sides of the first arc plate, the first arc plate and the second arc plates are both aluminum honeycomb plates, and the first arc plate and the second arc plates are arranged in parallel. The first arc plate comprises a first arc part and second arc parts located on the two sides of the first arc part, the first arc part and the second arc parts are tangent and opposite in direction, and the circle center of the second arc parts coincides with the circle center of the second arc plate. According to the utility model, the technical problem that the development requirements of light weight, modularization and low cost cannot be met in the prior art can be solved.

The invention provides a carrier rocket engine frame deformation compensation method. The carrier rocket engine frame deformation compensation method comprises the steps that an additional swing angle estimated value caused by frame deformation of an engine frame under the action of different thrust is measured in advance through a static test; the length of the servo mechanism is adjusted, and on the basis of the original swing angle, swing angle pre-compensation is conducted on an engine rack; according to the additional swing angle estimated value and a pre-compensation angle of swing angle pre-compensation, compensation and amplitude limiting are conducted on the engine swing angle instruction, and the compensated and amplitude-limited engine swing angle instruction is obtained; and the compensated and amplitude-limited engine swing angle instruction is sent to a servo controller. According to the method, the tilt angle compensation amount changing along with time can be provided and used for offsetting the additional tilt angle caused by deformation of the engine frame, so that the attitude control precision is improved, and then the safety during stage separation is improved, the load injection precision is improved, or the propellant consumption of an auxiliary power system is reduced.

The invention discloses a damage tolerance evaluation method and system for a friction plug repair welding joint of a carrier rocket storage tank, and the method comprises the steps: detecting an interface of the friction plug repair welding joint of the carrier rocket storage tank, and obtaining defect information, constitutive parameters of a material, and material attributes; establishing a finite element model, and accurately setting corresponding distribution interface defects; quantizing in-plane and out-plane restraint effects of the crack tip by adopting restraint parameters; processing and preparing a unilateral notch tensile sample, and measuring a local J-R curve of the tip of the precrack at the interface of the friction plug repair welding joint; correcting a real J-R curve of the friction plug repair welding joint of the rocket tank based on the influence of the distribution interface defect characteristics; then the tensile strength values of the friction plug repair welding joint in all the positions in the circumferential direction are corrected, and a corrected real J-R curve is obtained; and the limit crack size under the action of the corresponding internal pressure is obtained by utilizing the crack tip extension driving force curve, and then the critical J integral of the crack is calculated, so that the maximum crack size under the specific load or the maximum internal pressure under the crack with the specific size is obtained.

The invention discloses liquid rocket system polarity test equipment and a test method, and the method comprises the steps: carrying out the analysis of the working principles of a rocket system polarity test servo system, a final correction attitude control correction system and an engine power system, selecting a corresponding data collection device, and installing the data collection device at a to-be-detected position. During testing, a corresponding data acquisition device is used for carrying out data acquisition on parts with relative position changes or electromagnetic valve actions in the servo system, the final correction attitude control correction system and the engine power system. The data transmission system transmits the collected test data to the control and data processing system. And the control and data processing system compares the received test data with the instruction data to realize automatic interpretation of the system polarity test. The liquid rocket polarity test data and the system data are effectively and deeply fused, the accurate closed loop of the data is realized, the requirements of systematization and automation of carrier rocket test are met, the product reliability is improved, and the production efficiency is also improved.

The invention provides a carrier rocket ground test system and method supporting visual process configuration. The system mainly comprises: (1) a visual process configuration module, which configures a test process file through a graphical interface; (2) a process analysis and scheduling module for analyzing a process file and converting the process file into a task sequence; the test execution engine module (3) is used for driving bottom hardware and executing a test instruction sequence; (4) the test resource library module is used for storing and managing various information of test elements and various information of test equipment; the execution monitoring and data management module (5) is used for displaying a test execution state in real time and recording whole-process data; according to the method, the test logic and the execution program are separated, and a user can quickly construct, modify and reuse an automatic test process in a graphical mode, so that the test flexibility is greatly improved, and the test adaptability and efficiency are remarkably improved.

The invention discloses an automatic generation method and device for a dynamic model of a carrier rocket structure and a storage medium, and belongs to the technical field of carrier rocket model generation, and the method comprises the steps: describing the carrier rocket structure through carrier rocket parameters; reading and storing the edited parameter information of the carrier rocket; and automatically generating a carrier rocket structure dynamic model by using the stored carrier rocket parameters. According to the method, the carrier rocket is subjected to height parameterization description, the method can be combined with a secondary development function of commercial finite element software, rapid and automatic generation and rapid analysis of the model are achieved, and the method has the advantage of being easy to debug.

The invention discloses a carrier rocket avionics system based on a TSN (Time Sensitive Network) and a carrier rocket avionics data transmission method, and belongs to the technical field of carrier rocket control. A network is physically or logically divided into a periodic time-sensitive data stream, a high-speed time-sensitive data stream and a best effort data stream, the three data streams are subjected to isolation management, and transparent interconnection between a 1553B bus network and a TSN switching network is realized on a protocol layer through an intelligent protocol conversion gateway. Meanwhile, in a network management layer, through an instruction response type configuration protocol, common Ethernet equipment is seamlessly incorporated into the global scheduling and control category of the TSN network. According to the invention, by constructing the avionics network system with three-layer partition and dual-network fusion, multi-network seamless fusion on the rocket can be realized, the real-time performance of the network is improved, and the key requirements of rocket recovery are met.

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

Stackable satellites with a large diameter antenna that are configured to stack within each other like fully overlapping shallow bowls and to fit tightly within the area in a pay load fairing. The bowl-shaped satellites allow the satellites to nest within each other so that many satellites can be stacked within the height limits of the launch vehicle pay load fairing. A cradle is used to hold the stack, and tensions cables keeps the satellites in compression during launch. The satellite is an antenna with the bus integrated into the antenna structure. The antenna occupies the entire diameter / footprint of the satellite and thus can be matched up with the cross section of the fairing. This allows the largest possible antenna size to be launched without a complex deployment mechanism.

The invention relates to a digital supercharging control method and device based on a fitting algorithm, and belongs to the field of carrier rocket gas flow monitoring. According to the digital pressurization control method based on the fitting algorithm, the gas flow adjusting efficiency and the response speed are improved through the steps of signal recording, data rationality judgment and fitting algorithm processing, the pressurization control algorithm and a two-out-of-three control mode are adopted, control decision making is conducted, and an auxiliary electromagnetic valve is controlled. As the two-out-of-three control mode is adopted in the three paths of redundant signals, two paths of most reliable signals are selected for control decision making, and the reliability of the control method is further improved. The design of double-path independent short-circuit protection, double EMI filters, double DC / DC modules, redundant power output and the like is adopted in the power supply module in the digital boost control equipment, high-reliability power supply from a primary power supply to a secondary power supply is achieved, the fault-tolerant capability of the system is improved, and the stability and reliability of the power supply module are remarkably enhanced.

A Hold Down Release Mechanism, HDRM, interface for attachment of a spacecraft to an adjacent structure of a launch vehicle or another spacecraft, wherein the HDRM interface is configured for forming part of a single or multi-point releasable attachment of the spacecraft to said adjacent structure. The HDRM interface includes first and second connector parts, wherein one of the first and second connector parts is fastened to said adjacent structure, and other is fastened to the spacecraft; wherein the first connector part has a tapered projection with a non-circular external surface; wherein the second connector part has a tapered recess with a non-circular interior surface configured for form-lockingly receiving the tapered projection, for enabling transfer of torsion and shear load between the first and second connector parts, when the tapered projection is inserted in the tapered recess and the first and second connector parts are pressed together.

The invention provides a launch device of a carrier rocket, a use method and a launch system. The launching device is used for executing a launching task or a test run task, and comprises a fixed launching pad which is fixed at a launching site, is in butt joint with a carrier rocket during execution of the launching task, is in butt joint with a movable launching pad during execution of the test run task, and comprises a traction assembly; the bottom end of the movable launching pad is butted with the fixed launching pad, and the top end is butted with a carrier rocket sub-stage during a test run task execution period; one end of the mooring tool is used for being detachably connected with the top end of the carrier rocket sub-stage during the test run task execution period, and the other end of the mooring tool is in transmission connection with the traction assembly in a detachable mode; and the control system is in communication connection with the traction assembly and the mooring tool during execution of the test run task, and controls the traction assembly to adjust the pre-tightening force of the traction assembly according to the pre-tightening force information fed back by the mooring tool. According to the launching device, manpower, material resources and time cost can be greatly saved, and the structural integrity of the rocket body is guaranteed.

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

The invention provides a simulation method and system for the launch process of a carrier rocket and electronic equipment, and relates to the technical field of rocket launch simulation. The method comprises the steps that a launching scheme of the carrier rocket is determined, and the launching scheme comprises launching parameters of the carrier rocket; modeling and verifying the launching scheme; and performing simulation test on the launch process of the carrier rocket based on the verified launch scheme, and responding and correcting faults in the simulation test process. According to the invention, the problem that the safety of the launch task of the carrier rocket is threatened due to the fact that key design defects are not found in time due to multiple technical bottlenecks existing in the distributed simulation technology of the reusable carrier rocket is solved.

The embodiment of the invention provides a satellite-rocket separation scheme generation method and device, electronic equipment and a computer readable storage medium, and relates to the technical field of carrier rocket trajectories. The method comprises the following steps: acquiring parameter information of satellite-rocket separation; constructing an orbital coordinate system, and distributing a plurality of first alternative rocket postures in the orbital coordinate system for each satellite-rocket separation; carrying out calculation operation on each first alternative rocket attitude, and calculating to obtain included angle data of the rocket relative to the beam directions of the antenna chain satellite and the space-based antenna after the satellite and the rocket are separated; when the included angle data meets a preset space-based antenna beam angle maximum constraint condition, taking the first alternative rocket attitude corresponding to the included angle data as a second alternative rocket attitude; and generating a satellite-rocket separation scheme based on the second alternative rocket attitude. According to the embodiment of the invention, the measurement and control continuity of a satellite-rocket separation scheme is effectively enhanced, the design period of the satellite-rocket separation scheme is shortened, and the design efficiency is improved.

The invention discloses a carrier rocket propellant management device which comprises a porous partition plate, a screen structure, a screen support, a storage box, a conveying pipe and a pressurizing pipe. The porous partition plate is installed on the upper portion of the storage box and used for restraining the propellant from shaking in the long-time small negative overload section, and the low-temperature liquid propellant is prevented from being mixed with the air pillow. The porous partition plate is a disc formed by splicing fan-shaped pore plates and is mounted at the upper part of the storage box; the screen structure is positioned at the lower part of the storage box; the screen structure is formed by splicing a plurality of fan-shaped components into a circle and comprises an outer ring, a middle ring and an inner ring; the screen support comprises a plurality of vertical partition plates and a framework and is used for supporting the screen structure and playing an anti-vortex role at the same time; the framework is fixed on the vertical partition plate; the storage box is communicated to a rocket engine through a conveying pipe; and the pressurizing pipe is used for conveying pressurizing gas to the storage box. The device has the functions of gas isolation and liquid storage, can meet the matching requirements of a sinking engine, is beneficial to secondary weight reduction, and improves the effective load.