33 results about "Spacecraft design" patented technology

The invention discloses a spacecraft intelligent design method and system capable of autonomously learning, and belongs to the technical field of spacecraft design and simulation. The problems that in spacecraft design, natural language requirements are fuzzy and difficult to process, instruction generation is prone to ambiguity, linkage between simulation and design is not smooth, a design scheme lacks an effective iterative optimization mechanism, and traditional design is low in efficiency are solved. The method comprises the following steps: firstly, receiving a natural language demand, extracting key parameters to generate prompt words, inputting the prompt words into Design.AI, and generating a simulation instruction sequence; simulation is started through MCP Server conversion, if the result does not reach the standard, feedback correction is carried out, and iterative optimization is carried out. And repeating the steps until a design target is met, and generating a design report containing multiple items of content. Through natural language processing, real-time simulation, iterative optimization, knowledge accumulation and other technologies, the spacecraft design efficiency is greatly improved, the accuracy is enhanced, the threshold is lowered, and the flexibility is improved.

A method for automated spacecraft design includes receiving mission parameters provided by a client for spacecraft design, generating one or more concept designs based on the received mission parameters, a concept design including one or more components selected for each subsystem included in a designed spacecraft, for the concept design, automating a generation of a frame design that includes a plurality of thermal management components to be formed through a single metal-additive manufacturing process, performing a series of analyses or simulations to evaluate a performance of the frame design, and generating a digital twin representing a virtual representation of the designed spacecraft based on the generated frame design and the series of analyses and simulations.

A method for automated spacecraft design includes receiving mission parameters provided by a client for spacecraft design, generating one or more concept designs based on the received mission parameters, a concept design including one or more components selected for each subsystem included in a designed spacecraft, for the concept design, automating a generation of a frame design that accommodates the one or more components selected for each subsystem included in the designed spacecraft, the frame design being a Computer Aided Design (CAD) model that includes individual frame pieces generated, sized and fitted automatically via automated scripts compatible with a CAD system, including one or more structural design software applications; performing a series of analyses or simulations to evaluate a performance of the frame design; and generating a digital twin representing a virtual representation of the designed spacecraft based on the generated frame design and the series of analyses and simulations.

The invention relates to the technical field of spacecraft design, in particular to an aircraft information flow analysis, recheck and recalculation system, which comprises an information large graph design module used for automatically extracting design data such as an interface data list and a cable network contact table, generating a static information large graph through data check and analysis, and supporting signal management, layered display and export; the model library construction module is used for constructing a hierarchical information professional model library based on a Modelica language and covering multi-level models from a basic library, a universal library to a model customization library; and the information flow simulation verification analysis module is used for mapping the static information large graph into a simulation model, carrying out dynamic simulation verification on a system network, a communication network and a load network, and realizing multi-dimensional visual display of a result. According to the method, the information system design is rechecked and recalculated from static state to dynamic state and from manual operation to automatic operation, and the design efficiency, accuracy and verification sufficiency are effectively improved.

The invention discloses a switch ring layout method based on a communication cabin horizontal plate and a communication cabin, and belongs to the technical field of spacecraft design. Arranging microwave input equipment on the inner surface of the floor; microwave output equipment is arranged in the area above the horizontal plate on the inner surface of the south-north plate; arranging waveguides on the inner surface of the ground plate and the upper and lower surfaces of the horizontal plate to form input and output waveguide switch rings; and laser and digital processing extension equipment is arranged in the area below the south-north horizontal plate. The corresponding communication cabin adopts the partition layout structure. Through scientific partition and switch ring design, space and heat dissipation layout are optimized, meanwhile, full-system backup of the microwave channel power amplifier unit is achieved, and reliability and expandability of communication loads are remarkably improved.

This application relates to the field of orbital dynamics design and simulation technology, and provides a method and apparatus for free transformation between spatial coordinate systems for solar system spacecraft design and simulation. This method transforms the position and velocity of different coordinate systems to the celestial coordinate system for origin transformation, and then rotates to the target coordinate system. This allows position and velocity to be arbitrarily transformed between coordinate systems included in the coordinate transformation model library. Furthermore, when expanding to a new coordinate system, only the transformation function between the new coordinate system and the co-centered celestial coordinate system is needed, without having to calculate the transformation functions between the new coordinate system and all other coordinate systems separately. This achieves effective organization of scattered coordinate system transformation relationships, facilitating the design and simulation of spacecraft missions and reducing the workload generated by transformations between multiple spatial coordinate systems.

The present invention relates to a direct-drive solar wing drive device, which relates to the field of spacecraft design technology, and includes: a stepper motor, a limit device and an output connecting shaft; the limit device and the output connecting shaft are respectively connected to the housing and the output shaft of the stepper motor by bolts; the stepper motor includes: a rotor output end flange, a housing end cover, a motor housing, a motor cable, a rotor hollow shaft, a stator coil, stator teeth, rotor teeth and a bearing; the limit device includes: a mounting seat, a limit block and a limit switch. The direct-drive solar wing drive device of the present invention simplifies the two components of the harmonic reducer and the conductive slip ring, and is only composed of a stepper motor and a limit switch; the stepper motor directly provides the torque, and the harmonic reducer's deceleration and torque increase link is omitted; the continuous rotation of the traditional SADA is converted into a reciprocating swing, and the conductive slip ring can be omitted. The direct-drive solar wing drive device of the present invention has a simple and reliable structure, is light in weight, has low cost and high reliability.

The present invention discloses a method for weighing and optimizing spacecraft design schemes based on value analysis, which belongs to the technical field of value analysis. The method comprises the following steps: establishing an evaluation index system from the functional dimension of value, and determining the weight of each index in the evaluation index system; determining the satisfaction score value of a candidate scheme for each index and the satisfaction coefficient of the candidate scheme for each index, and calculating the satisfaction value of the candidate scheme for each index; calculating the value factor of the candidate scheme for each index, and summing the value factors corresponding to all the indexes of the candidate scheme to obtain the total value of the candidate scheme; calculating the cost factor of the development of the candidate scheme; calculating the value of the candidate scheme based on the total value of the candidate scheme and the cost factor of the development; and finally selecting the candidate scheme with the highest value as the preferred one. The present invention can weigh the candidate schemes for spacecraft design from the perspective of value, propose a quantitative weighing strategy, and then select the optimal candidate scheme.

This application provides a spacecraft assembly design status integration construction system and method based on a 3D model, belonging to the field of spacecraft design and manufacturing technology, aiming to solve the problems of information silos and collaborative difficulties in spacecraft assembly design. The system includes an assembly status integration display module and a component technical requirement planning and definition module. The assembly status integration display module, based on a professional design architecture, aggregates, lightweights, and integrates multi-disciplinary layout components into a unified architecture, and parses and manages design information. The component technical requirement planning and definition module, within the whole-vehicle model environment, provides structured definitions and management of assembly technical requirements. This application breaks down information barriers, improves collaborative efficiency, and lays the foundation for digital manufacturing by constructing a unified and information-complete digital prototype.

The invention provides a gravitational wave detection spacecraft optimization method based on static full-link noise simulation, and the method comprises the steps: building a multi-physical field model of a self-gravitational field, a magnetic field, a temperature field and the like of a spacecraft platform according to a preliminary design scheme of the spacecraft; in combination with the design scheme and a multi-physical field model, a space gravitational wave detection static full-link noise model is constructed, and optical measurement noise and inspection mass residual acceleration noise simulation evaluation results are obtained through analysis; feeding back the static full-link noise simulation evaluation result to the initial design scheme of the spacecraft, and optimizing the design scheme of the spacecraft; and repeatedly executing the steps, and carrying out iterative optimization on the spacecraft design scheme. According to the method, the satisfaction degree of scientific requirements is answered on the system level, and the spacecraft design scheme is iteratively optimized on the basis. Noise and multi-physical field coupling and transmission rules can be disclosed, the requirement for evaluating the sensitivity of a spacecraft detection system can be met, and the method is feasible and has practical application significance.

The invention belongs to the technical field of spacecraft design and thermal management, and provides a spacecraft temperature field modeling method and system based on device point cloud data and 3D U-Net, and the method comprises the steps: carrying out the preprocessing of the point cloud data, and determining a feature tensor; then, according to the feature tensor and a preset 3D U-Net network, predicting a three-dimensional temperature field as an initial temperature field; and finally, according to the initial temperature field, adjusting the center coordinate of the device through reinforcement learning, minimizing the volume of the high-temperature region, and generating a new temperature field. A high-resolution temperature field can be directly generated from point cloud data of a plurality of devices in a spacecraft by utilizing the powerful three-dimensional convolutional neural network nonlinear fitting capability of the 3D U-Net network, and the parallel computing capability of deep learning is combined, so that the computing time is remarkably shortened, and the prediction precision under a complex layout is improved; meanwhile, the coordinates of the device are efficiently adjusted through reinforcement learning, the highest temperature is minimized, and the risk of heat concentration is reduced.

The application discloses a kind of super low orbit spacecraft aerodynamic configuration optimization method, device, equipment and medium, it is related to spacecraft design field.The method comprises: obtaining target parameter;According to target parameter and NRLMSISE-00 model, calculate real-time atmospheric environment parameter;Based on real-time atmospheric environment parameter, construct multi-objective optimization function;With NSGA-III algorithm, with the minimum drag coefficient, the maximum working medium capture efficiency, the minimum structure weight and the optimal atomic oxygen erosion resistance performance as target, the multi-objective optimization function is solved, and the core geometric parameter of super low orbit spacecraft is obtained;Using constraint condition, core geometric parameter, derivative constraint parameter and material and coating parameter are filtered, and the optimal aerodynamic configuration of super low orbit spacecraft is obtained.The application can realize the rapid optimization of super low orbit spacecraft aerodynamic configuration, and has high reliability and long on-orbit life.

The invention discloses a low-cost constrained multidisciplinary robust optimization design method, and belongs to the field of spacecraft manufacturing and application. The implementation method comprises the following steps: constructing an aircraft multidisciplinary model; a numerical simulation model and a finite element analysis model are respectively constructed for pneumatic and structural subjects, and a proxy model is constructed for the pneumatic and structural subjects in combination with Bayesian learning. And constructing an aircraft manufacturing cost prediction model based on knowledge distillation. Based on a global sensitivity equation GSE of a deterministic domain coupling multidisciplinary system, input random input and model cognition uncertainty are introduced, expressions of mean values and covariances of subject coupling variables and system responses are derived respectively, and a rapid semi-analytical multidisciplinary random and model cognition hybrid uncertainty propagation method is established. Multidisciplinary uncertainty analysis is achieved, and aircraft multidisciplinary system performance response is obtained. And constructing a multi-objective multidisciplinary optimization model with optimal performance and lowest cost, and solving by adopting an NSGA-II algorithm to obtain a final aircraft design scheme.

The application discloses a kind of bilateral swingable scanning ultra-wide coverage SAR satellite system, scanning method and device, and relates to spacecraft design field.The application solves the problem that traditional spaceborne SAR is difficult to realize high-resolution wide swath imaging.The application solves the problem that the existing SAR satellite scheme is mostly equipped with a single SAR antenna, and the field of view is expanded by the attitude maneuver of satellite platform.Besides, due to the natural contradiction between resolution and swath, traditional spaceborne SAR is difficult to realize high-resolution wide swath imaging.The system comprises a satellite, a rotating mechanism and two SAR antennas.The two SAR antennas are symmetrically distributed on both sides of the satellite by the rotating mechanism.The method is realized based on the above-mentioned bilateral swingable scanning ultra-wide coverage SAR satellite system, comprising: dividing the scanning area into N distance direction and M azimuth direction sub-bands;the SAR antenna scans according to the divided M sub-bands, and splicing is carried out.The application is applied to the field of satellite monitoring.

The application relates to the technical field of spacecraft assembly, and provides a spacecraft assembly process generation method based on multi-model cooperation. The method comprises the following steps: receiving a current process generation task prompt instruction, analyzing and content extracting a target spacecraft design file and a process flow framework template file provided, so as to generate a first edition process flow of the target spacecraft; generating an assembly process sequence of the target spacecraft by using a process generation model; generating a second edition process flow and a third edition process flow by using two pairs of generation models and a check model; and performing process evaluation on the third edition process flow. The application reduces the excessive dependence of process flow design and checking on manpower, reduces the probability of human error, improves process compilation efficiency, and realizes the automation of the spacecraft assembly process flow.

The invention relates to a spacecraft structure damage on-line and off-line detection capability evaluation index determination method. The invention relates to the technical field of spacecraft structure health monitoring, and aims at identifying spacecraft damage on line, performing off-line detection, identifying a damage position and quantifying a damage degree. The method is suitable for long-term on-orbit and multi-time reciprocating spacecrafts when constructing a spacecraft structure damage off-line detection technical capability evaluation index system, and provides technical support for improving the spacecraft on-orbit damage real-time early warning capability and ground maintenance decision-making efficiency, reducing leak detection risk and maintenance cost, and realizing high-reliability spacecraft design and service life extension.

The present invention discloses a method for calculating collision debris parameters through sudden changes in spacecraft attitude and orbit parameters. The method calculates the velocity increment of the spacecraft impacted by debris through changes in the spacecraft orbit, and obtains the momentum change of the spacecraft when the debris impacts the spacecraft in combination with the spacecraft mass. The torque change caused by the collision is obtained based on the angular velocity change of the gyroscope, and the vector of the collision point in the fixed coordinate system of the spacecraft body is obtained in combination with the momentum change. The expected mass of the debris is directly obtained based on the statistical average velocity distribution and momentum change of the space debris relative to the spacecraft. The present invention can directly obtain the spacecraft momentum change and angular momentum change, the expected mass of the debris, and the maximum and minimum masses of the debris corresponding to sudden changes in the spacecraft attitude and orbit parameters, which helps to estimate the long-term impact and statistical laws of the debris collision on the spacecraft on the orbit, helps to trace the source of space emergencies and diagnose and design disaster plans in the long-term management of the spacecraft, and assists in the effective implementation of anti-collision reinforcement in the spacecraft design process.

The invention relates to the technical field of spacecraft design, in particular to a rapid scheme demonstration system for an aircraft. Comprising a task analysis module, a constraint analysis module, a function demand analysis module, a special design module, a design integration module, a demonstration analysis evaluation module, a system modeling simulation module, a model data management module, a demonstration management module, a collaborative demonstration support module and an aircraft multi-field unified model library. According to the method, decomposition, transmission and structured association of different levels from tasks to functions, from functions to indexes and the like are achieved through all the modules, rapid aircraft scheme argumentation from top to bottom is achieved through simulation analysis and rechecking, the complexity and cost of scheme argumentation are reduced, and the design efficiency is improved.

This invention provides a method for analyzing the on-orbit utilization of propellant in a spacecraft surface tension propellant tank, under conditions where pressurized gas cannot be introduced by the tank management device. The method includes: obtaining the residual propellant dose V attached to the inner wall of the tank. tc ; Obtain the residual propulsion dose V in the pipeline gc ; Obtain the propellant dose V in the tank management device tg ; Obtain the safety margin V for propellant in spacecraft design a Obtain the total volume V of the storage tank. t According to the residual propellant dose V on the inner wall of the tank tc Residual propellant dose V in pipeline gc Propulsion dose V in tank management device tg Spacecraft design with a safety margin for propellant V a Total volume V of the storage tank t The propellant utilization rate was determined to be 1-(V). tc +V gc +V tg +V a ) / V t By integrating the working mechanism of the surface tension tank, the propellant consumption process, the overall load conditions, and the on-orbit mission profile, accurate measurement of the on-orbit utilization rate of propellant, which is less affected by on-orbit environmental factors, can be achieved.

The application provides a spacecraft multi-loop energy balance design method and system, relates to the technical field of spacecraft design and simulation, and comprises solar cell array series number design, solar cell array parallel number design, storage battery group series number design and storage battery group parallel number design. The energy balance design method is suitable for the multi-loop energy balance state of a spacecraft. The application effectively solves the problem that the volume and weight of the power supply and distribution subsystem are over-limited when the overall spacecraft load power is too large, provides a feasible scheme for the power supply and distribution subsystem design of the multi-loop energy balance of a spacecraft, and expands the application range of the energy balance design method.

The application provides a multi-point approximation method based on an approximate Hessian matrix for structure optimization, comprising the following steps: step 1, establishing a finite element model of a spacecraft based on an initial design of a spacecraft structure; step 2, performing mechanical prediction of the spacecraft, and the analysis mode comprises static analysis, modal analysis, buckling analysis and transient response; step 3, establishing a structure optimization mathematical model, namely, defining an optimization target, a constraint function and a design variable; step 4, establishing a sequence approximation problem considering second-order information to approximate the original problem; step 5, estimating the Hessian matrix of the multi-point approximation function by using a quasi-Newton method; step 6, selecting known points by using a neighborhood selection strategy based on the Euclidean distance; step 7, establishing convergence discrimination and an automatic optimization system; step 8, method comparison, analysis and iteration explicitness; and setting an optimization example to illustrate the flow from spacecraft design to analysis and then to optimization, and the effectiveness is verified by comparing the results of the optimization method with those of a commercial software.

The invention relates to a method for acquiring dynamic characteristics of a stacked satellite assembly, which belongs to the field of spacecraft design and kinetic analysis, and comprises the following steps: firstly, establishing a single-layer satellite kinetic model, then acquiring actually measured modal parameters of a single-layer satellite through a ground modal test, and correcting the single-layer satellite kinetic model by using a modal confidence criterion; establishing an interface model, and obtaining an interface equivalent dynamic model through theoretical calculation and a simulation method; simulating an axial pretightening force, a shear load and a bending moment condition under an actual launching working condition, and correcting the interface equivalent dynamic model; based on a modal synthesis method, obtaining an N-layer stacked satellite combination dynamic model; acquiring actual modal parameters and dynamic response data of the stacked satellite assembly, and performing reverse correction according to the actual modal parameters and the dynamic response data; and assembling the combination dynamical model and the rocket dynamical model, carrying out satellite-rocket integrated joint simulation, and optimizing the N-layer stacked satellite combination dynamical model according to a simulation result.

The application provides a SAR satellite structure and a design method suitable for multi-satellite launching, relates to the technical field of spacecraft design, and comprises a satellite main body, a solar wing, a radiation unit, an antenna substrate, a through-plate connector, a TR assembly and a semi-rigid connecting assembly. The SAR satellite structure adopts a trapezoidal cross-section configuration, fully utilizes the internal space of a rocket fairing, improves satellite networking speed, sets heat conduction units and heating units between the radiation unit and the antenna substrate and between the antenna substrate and the TR assembly, increases heat conduction between single machines while ensuring the radiation array plane flatness of the radiation unit and the installation flatness of the TR assembly, improves antenna temperature consistency, the semi-rigid connecting assembly can reduce the antenna response in the active stage, release the antenna thermal deformation, reduce the flatness change, improve the radiation array precision, and solves the problems of low space utilization rate of the existing SAR satellite, the antenna flatness being easily affected by thermal deformation and the difficulty in antenna temperature control.

The invention relates to a direct-drive solar wing driving device, which relates to the technical field of spacecraft design, and comprises a stepping motor, a limiting device and an output connecting shaft, the limiting device and the output connecting shaft are connected to a shell of the stepping motor and the output shaft through bolts respectively. The stepping motor comprises a rotor output end flange, a shell end cover, a motor shell, a motor cable, a rotor hollow shaft, a stator coil, stator teeth, rotor teeth and a bearing. The limiting device comprises a mounting seat, a limiting block and a travel switch. According to the direct-drive type solar wing driving device, the harmonic reducer and the conductive slip ring are simplified, and the direct-drive type solar wing driving device is only composed of the stepping motor and the travel switch; the stepping motor directly provides rotating torque, and the link of speed reduction and torque increase of a harmonic reducer is omitted; continuous rotation of traditional SADA is changed into reciprocating swing, and a conductive slip ring can be omitted. The direct-drive type solar wing driving device is simple and reliable in structure, light in weight, low in cost and high in reliability.

The invention belongs to the technical field of spacecraft design and three-dimensional thermal field modeling, and provides a thermal field modeling method and system based on point cloud feature extraction and physical dimensionality reduction, original point cloud data of a spacecraft are directly processed by adopting Point CNN, local spatial features are coded through transformation weighting, information loss caused by gridding is avoided, and the method and the system have the advantages that the method and the system are simple and convenient to operate, and the cost is low. The modeling precision of the irregular geometric structure of the spacecraft is obviously improved; a principal component analysis dimension reduction method of physical constraint is introduced, key physical parameters such as heat conduction coefficients in thermal field distribution can be explicitly decoupled, and an interpretable mapping relation between low-dimensional physical characteristics and a temperature field is constructed; the technical barriers that traditional finite element analysis is low in calculation efficiency and a data-driven model is weak in physical interpretability are broken through, three-dimensional temperature field proxy modeling is completed, and a new solution is provided for thermal design and optimization of spacecrafts.

The application discloses a reconfigurable ground test system and method with a triangular prism as a module, belongs to the technical field of reconfigurable system design, and is based on reconfigurable system technology, and a reconfigurable ground system with a triangular prism as a module is designed. Since the module is a triangular prism and the rotating surface is an inclined surface, the module can form various configurations to cope with different task scene requirements, and meanwhile, various reconfiguration algorithms can be verified, thereby laying a foundation for real spacecraft design. In addition, the module is subjected to ground tests to ensure the feasibility of the model, provides a new idea for the field of spaceflight, and can form a research framework in which theory and tests are mutually examined.

The invention provides a relay antenna layout method and device and electronic equipment, and relates to the technical field of spacecraft design, and the method comprises the steps: configuring a designated relay antenna on a first cabin in a multi-cabin spacecraft as an antenna facing the sky; determining an installation azimuth angle of each specified relay antenna in the remaining relay antennas based on the configuration of the remaining cabin sections; and determining a target zenith angle of the remaining relay antennas by taking the condition that the joint relay coverage rate of the antenna and the remaining relay antennas is greater than or equal to a first threshold value as an optimization target. According to the method, the advantages and disadvantages of the relay antenna layout are measured by using the joint relay coverage rate, so that the relay antenna layout method is converted from qualitative design to quantitative design, the overall coverage capability of a plurality of relay antennas can be fully exerted, and repeated redundancy and resource waste of the relay antennas of each cabin section are avoided.

The present invention provides a method and apparatus for determining a spacecraft's transfer orbit, belonging to the field of orbital transfer technology. The method provided in an embodiment of the present application is applied in a scenario where a first spacecraft is pursuing a second spacecraft. When the first spacecraft lags behind the second spacecraft, the method determines a maximum phase difference based on the first spacecraft's current position and the radius of a central celestial body. The transfer time of the first spacecraft is determined based on the maximum phase difference and the current phase difference between the first and second spacecraft, thereby determining the start and end times of the first spacecraft's orbital transfer. Furthermore, the trajectory information of the target transfer orbit is determined based on the Lambert orbital transfer model. The target transfer orbit determined by this method has an orbital altitude greater than the radius of the central celestial body. A reasonable transfer orbit can be designed for the spacecraft, thereby smoothly adjusting the spacecraft's motion state to enable the spacecraft to complete the pursuit mission.

The invention relates to a design method of a vertical rail scanning and along-rail splicing optical satellite. The invention relates to the technical field of spacecraft design, and the method comprises the steps: carrying out the satellite task demand analysis, and designing the satellite image quality; and designing a core subsystem and other subsystems to form a satellite overall scheme. Through the design method provided by the invention, the functional performance difference between a vertical orbit scanning and along-orbit splicing imaging new system optical satellite and other traditional push-scan system satellites can be accurately identified, and a core subsystem required for realizing the functional performance and the design requirements of the core subsystem can be determined according to brand new functional performance requirements; through hierarchical design of the core subsystem and other subsystems, the satellite can be guided to pay attention to key core technology identification and attack in the early stage and pay attention to engineering and reliability design in the later stage, and the design efficiency is improved.

The invention relates to the technical field of spacecraft structure design and manufacturing integration, in particular to a man-machine collaborative spaceflight design method, which comprises the following steps of: constructing a design contract and a design hypergraph, and optimizing configuration parameters in a microphysical model; an accessibility cone field and a flutter critical frequency field are obtained through analytic geometry and dynamic modes, and a feasibility certificate is generated through square sum form verification; executing continuous relaxation planning and constraint projection based on a type system on the process resource hypergraph, and generating a safe trajectory and a measurement path by combining a control barrier function and a harmonic field; and gathering into a patch based on the minimum hit of the counter-example set, writing back, and triggering local resolution and reverification. According to the invention, consistent convergence of contract satisfaction, manufacturing feasibility and evidence auditing is realized.