1260 results about "Launch vehicle" patented technology

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 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.

The invention relates to the field of intelligent control and software management, and discloses a carrier rocket software control method and decision making device based on SAC, and the method comprises the following steps: obtaining the state information of a carrier rocket software system, and constructing a unified expression; an optimal action is generated based on an SAC algorithm; executing the action and collecting feedback information after the action is executed; utilizing feedback information to optimize an actor network and a commentator network in the SAC model, and continuously updating a decision strategy; based on the optimized strategy, the task priority, the running state of the software module and the abnormal recovery operation are dynamically adjusted, and the system state is collected again to optimize the model in a closed-loop mode. The system operation state can be adjusted in real time, the task scheduling efficiency, the resource utilization rate and the exception handling capacity are improved, intelligent management of a carrier rocket software system is achieved, and the method is suitable for complex task environments and high-dynamic operation scenes.

An inlet cover mechanism for an air launched vehicle includes a quick release pin, a chassis, a collar, and a piston pin. The inlet cover mechanism facilitates attachment and deployment of an inlet cover with the air launched vehicle. The quick release pin is connected to the inlet cover. The chassis is connected to the air launched vehicle. The collar is connected to the chassis and is releasably engageable with the quick release pin. The piston pin is slidably engaged within the chassis and is configured to contact the quick release pin.

The invention provides a carrier rocket box bottom and a forming method thereof. The forming method comprises the following steps: S1, calculating box bottom stress distribution under the action of internal pressure by adopting a film stress method and a moment-free theory; s2, determining a latitudinal tension-compression stress boundary according to box bottom stress distribution under the action of internal pressure; s3, the top is subjected to forming machining, and the bottom is subjected to forming machining; and S4, the machined top and the machined bottom are subjected to tailor welding according to the latitudinal tension and compression stress boundary, and therefore the finally-formed box bottom is obtained. According to the carrier rocket box bottom forming method, the problems that under the action of internal pressure, the bottom of the box bottom is prone to wrinkling and elastic buckling, and the top of the box bottom is prone to cracking and instability are solved.

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 provides an integral box bottom mirror image milling program generation method and system. The integral box bottom mirror image milling program generation method comprises the steps of 1, compiling a profile scanning path according to pattern / model input; step 2, compiling a profile scanning path; 3, generating a profile scanning numerical control program; 4, obtaining an actual programming entity; step 5, extracting processing features; step 6, dividing a rough machining area; 7, planning an initial rough machining path; 8, a finish machining area is divided; step 9, planning an initial finish machining path; step 10, generating a milling processing path; and step 11, generating a machining numerical control program. According to the integral box bottom mirror image milling program generation method, the integral box bottom mirror image milling tool path generation steps and method are stipulated, the normalization of integral box bottom mirror image milling program generation is improved, and the machining efficiency and quality of an integral box bottom are further improved; and the bearing capacity and reliability of the carrier rocket are met.

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 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 discloses a remote supportable rocket test launch control system based on a cloud platform. The remote supportable rocket test launch control system comprises a rear-end remote test module, a remote test network and a front-end remote test module, wherein communication and data interaction between the rear-end remote test module and the front-end remote test module are realized through a remote test network; the front-end remote test module transmits the collected test data of the launching site and the workshop to the rear-end remote test module through a remote test network; and the rear-end remote test module realizes communication and data interaction with carrier rocket test launch control equipment and post personnel in various regions through a remote test network, and realizes remote auxiliary test, comprehensive management and remote command of the whole rocket launch process. The problems that a traditional rocket test launching mode is low in efficiency, and front-rear cooperation cannot be achieved are solved.

The invention discloses a universal liquid carrier rocket launching pad technology. A satellite general assembly test subsystem realizes recovery and detection of a satellite and forms a star cover assembly; the rocket horizontal final assembly test subsystem realizes subsystem test, matching test and total inspection test of a first-stage rocket body and a second-stage rocket body in a horizontal state, and the first-stage rocket body, the second-stage rocket body and a star cover combination body form a whole rocket in the horizontal state; the rocket horizontal transfer subsystem transfers the whole rocket to the fixed launching platform in a horizontal state; the rocket launching subsystem erects the whole rocket and achieves testing, filling and launching of the whole rocket. According to the universal liquid carrier rocket launching pad technology, by adopting a three-plane test launching mode of horizontal final assembly testing, horizontal transferring and vertical launching, the construction cost of a rocket launching site is effectively reduced, the operation process is simplified, the rocket launching efficiency and frequency are improved, and the requirement for rapid and batch launching is met.

The invention provides an online fault detection method for a reusable liquid rocket engine, and aims to realize real-time online adjustment of a fault threshold according to the change of the working condition of the engine under the condition of large-range variable thrust of the reusable engine so as to form a red line shutdown strategy based on an online dynamic threshold. Faults occurring in the working process of the engine are pre-warned and judged, so that effective measures are taken in time, and the safety of the carrier rocket and the load is protected.

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 belongs to the technical field of friction stir welding of carrier rocket tank body circular seams, and particularly relates to an automatic leveling and tightening device for a cylinder section and a common bottom lock bottom circular seam fixing inner support. The fixed base assembly comprises a fixed base and linear guide rails, the fixed base is an annular welding structural part with a trapezoidal section, the linear guide rails which are uniformly distributed are laid on the upper surface of the fixed base and used for supporting, positioning and accurately guiding the radial outer supporting block assembly and the radial inner supporting block assembly, and the axial wedge block assembly is fixedly installed on the back of the fixed base; every two sets of axial wedge block assemblies fix one set of radial inner supporting block assembly in a wedge-caulking mode, every two adjacent sets of radial inner supporting block assemblies wedge the radial outer supporting block assembly between the two adjacent sets of radial inner supporting block assemblies, and the outer sides of every two adjacent sets of axial wedge block assemblies are connected with a common-bottom leveling assembly. The device provides strong internal supporting force for the welding state of the common-bottom assembly and the cylinder section lock bottom girth, it is guaranteed that concession does not occur under the action of strong upsetting force of a welding execution mechanism, and the stable and reliable storage tank girth welding process is achieved.

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 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 fairing pulsating pressure scaling wind tunnel test method capable of reducing space-ground difference. The method comprises the following steps: determining initial parameters; based on the determined initial parameters, carrying out a transonic speed scaling manual transition pulsating pressure wind tunnel test, arranging pulsating pressure measuring points, and carrying out transonic speed state fairing pulsating pressure environment measurement to obtain a measurement result; performing parameter optimization according to a comparison result of a measurement result and a standard parameter under a real flight test condition to obtain an optimal parameter of manual transition; and based on the optimal parameters, carrying out a transonic speed scaling manual transition pulsating pressure wind tunnel test to obtain test results of the configuration in different free incoming flow Mach number, attack angle and sideslip angle states. According to the method, the free incoming flow local shock wave-boundary layer interference state and the boundary layer separation state of the ground wind tunnel test and the flight test are ensured to be closer through a manual transition mode, and the ground test prediction precision of the carrier rocket fairing pulsating pressure environment is improved.

The invention discloses a carrier rocket application-oriented adaptive augmented controller frequency domain analysis method, which comprises the following steps of: further simplifying a nonlinear damping module of an improved AAC controller to obtain a frequency domain model of the AAC controller, inputting a sine characteristic signal into the frequency domain model to obtain an analytical solution of an output value of the AAC controller, and removing high-order harmonics from the analytical solution to obtain a frequency domain model of the AAC controller. Obtaining an output signal with the same frequency band as the input signal; frequency domain response of an integrator at a low frequency band is used as an integral filter to simplify an output signal, and a linear part of a result is used as a frequency domain analysis model of the AAC controller; the precision and the application range of the AAC controller frequency domain analysis model are evaluated through a frequency sweep test, and the application range of the AAC controller is given; the obtained AAC controller frequency domain analysis model is integrated into a carrier rocket control model, and the frequency domain margin of a carrier rocket control system is improved at high frequency and low frequency by adjusting control parameters.

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.

The invention discloses a carrier rocket composite attitude control method, which comprises the following steps: establishing a swing nozzle correction model and a side jet flow correction model of a carrier rocket, sending a rocket body attitude angle deviation value to the two models at the same time, calculating an output quantity, comparing the output quantity of the side jet flow correction model with an attitude control nozzle opening and closing threshold, and calculating the attitude control nozzle opening and closing threshold; and when the output quantity of the side jet flow correction network reaches the opening and closing door limit of the attitude control jet pipe, the output of the side jet flow correction network can actually act on the rocket body attitude control process of the carrier rocket. Through common application of the swing spray pipe correction model and the side jet flow correction model, the jitter attitude angle deviation value generated during operation of the swing spray pipe correction system in a severe environment is eliminated, the attitude control precision of the rocket body is improved, and the output quantity of the side jet flow correction model is compared with the attitude control spray pipe opening and closing threshold, so that the attitude control precision is improved. It is guaranteed that the side jet flow correction model controls the attitude of the rocket body only if necessary, and therefore the fuel utilization rate is increased.

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 present application discloses a control method and system for a booster system of a launch vehicle engine, wherein the control method includes the following steps: collecting pressure data; making a rational judgment on the collected pressure data; if the rational judgment is passed, making a validity judgment on the pressure data; making a comprehensive calculation on the pressure data according to the validity judgment result and the rational judgment result; judging whether the result after the comprehensive calculation is valid; if the result after the comprehensive calculation is valid, performing a closed boost control; if the result after the comprehensive calculation is invalid, performing a boost valve opening and closing control according to a preset valve action sequence. The present application also makes a reasonable threshold judgment and data fusion on the data of redundant sensors, and when one sensor fails, it can ensure the correctness of the boost calculation and the correct opening and closing of the boost valve. Under extreme conditions, the boost system can be guaranteed to work within the normal range by fixing the boost timing. Thereby improving the reliability of the boost control system of the launch vehicle.

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 the technical field of satellites, in particular to a deployment method of a low-orbit large-scale satellite constellation, which comprises the following steps: S1, establishing a static network for launching deployment by taking each orbit and a launching site as nodes and taking a satellite transportation path between the nodes as an edge; the network side comprises non-time dimension information of each node; s2, on the basis of the static network, traversing all possible paths according to launch interval time of the vehicle and orbit dynamics to obtain a time expansion network; s3, a parking track and a target track are obtained, and a two-channel logistics network model is constructed according to the parking track and the target track; s4, based on the two-channel logistics network model, establishing a multivariable mixed integer optimal mathematical equation according to information contained in nodes and edges in the time expansion network; and S5, solving the two-channel logistics network model according to the mathematical equation to obtain an optimal launching deployment scheme of the satellite constellation.

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 discloses a liquid recoverable carrier rocket control system and control method, and belongs to the technical field of rocket recovery control. Comprising a flight control combination, a recovery control combination, an inertia measurement combination, a battery, a safety control combination, an engine controller, an optical target camera, a laser range finder, a comprehensive collecting and editing device, a relay measurement and control terminal and a space-based telemetering phased array antenna. The flight control combination and the recovery control combination are arranged in the second rocket sublevel and the first rocket sublevel respectively, partial hot backup is achieved through the two control combinations, and the reliability of the system is improved. In the first-stage flight section with the worst and most dangerous mechanical environment, the recovery control combination can serve as hot backup execution work of the flight control combination, and normal recovery of the first sub-stage cannot be affected even if flight failure is caused by abnormity of the second-stage rocket. Effective control over launch and recovery of the carrier rocket can be met, and a necessary control system is provided for recovery of the 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.