422 results about "Deep space exploration" patented technology

The invention discloses a mechanical arm digital twinning system for deep space exploration, and relates to the technical field of digital twinning simulation. The mechanical arm comprises a physical layer, a data acquisition and transmission layer, a digital twin model layer and an application layer, the physical layer comprises a mechanical arm body and a sensor network, and the sensor network is used for acquiring operation state information of the mechanical arm body; the data acquisition and transmission layer comprises data acquisition equipment and a data transmission network, wherein the data acquisition equipment is respectively connected with the sensor network and the data transmission network. According to the invention, the physical layer, the data acquisition and transmission layer, the digital twinborn model layer and the application layer of the mechanical arm are constructed, so that the actual mechanical arm body can be guided to perform related actions through virtual simulation data, and virtual and real body configuration comparison can be completed through a visual interface; and the digital twinning requirements of various spacecraft operations in the future deep space exploration field are met.

The invention discloses an indirect optimization method for a landing anti-interference trajectory of a small celestial body, and relates to the field of deep space exploration. Aiming at the problem that a small celestial body landing trajectory is sensitive to uncertain parameters, the following solving method is provided: firstly, establishing a stochastic dynamics model of small celestial body landing, and constructing an anti-interference trajectory optimization problem; secondly, establishing a relation between an anti-interference trajectory optimization problem and a gravity-energy-free optimal control problem through a performance functional, a kinetic equation and a homotopy function of process noise, and estimating a co-state initial value of the gravity-energy-free optimal control problem; and finally, gradually converting the zero-gravity energy optimal control problem into an anti-interference trajectory optimization problem by changing homotopy coefficients, and in the conversion process, converting the optimization problem corresponding to each homotopy function into a two-point boundary value problem by using a Pontryagin minimum principle, and solving the two-point boundary value problem by using a targeting method. The sensitivity of the trajectory to uncertainty is reduced, so that the anti-interference capability of the landing trajectory is improved.

The invention discloses a high-performance satellite platform based on a real-time operating system and a wireless bus and a working method of the high-performance satellite platform, and belongs to satellite communication technologies, future deep space exploration and low-orbit satellite constellations. On the basis of an original structure of a satellite platform, a real-time operating system is effectively combined with a wireless bus system, and the real-time operating system comprises a kernel module, a communication module, an equipment driving module, a file system module, an application interface module, a safety module and a fault management module. The wireless bus system comprises a physical layer, a data link layer, a network layer, a transmission layer and an application layer. According to the invention, efficient and reliable communication and task execution capabilities are provided for a satellite platform, the requirements of real-time performance and reliability of a satellite in a complex environment are effectively met through modular design and tight work of the system, and meanwhile, complex lines and interfaces brought by a wired system are avoided.

The invention discloses a scene understanding and intelligent decision-making method for a planet probe vehicle in a complex unknown environment, and relates to the technical field of deep space exploration and intelligent vehicles, and the method comprises the steps: obtaining multi-dimensional environment information; constructing a slump function based on the geometric morphology, motion characteristics and material attributes of the obstacle, and dynamically quantifying the risk of the obstacle; smoothing the potential field function by adopting Gaussian filtering, and dynamically adjusting the weight of a global potential field and the weight of a local potential field; in combination with an improved clustering algorithm, fuzzy regions of obstacles capable of being crossed and obstacles not capable of being crossed are accurately divided; screening passable areas by constructing a credibility mean square value function; carrying out multi-objective optimization by adopting a genetic algorithm, and generating an optimal path matched with the environment in cooperation with an offline data set and online reasoning; and inputting the optimal path into a bottom layer control system of the probe vehicle, controlling execution of a motor driving and steering mechanism, monitoring the state of the probe vehicle in real time, and triggering dynamic re-planning. And the autonomous obstacle avoidance capability and the path planning efficiency of the probe vehicle in a complex environment are improved.

The invention discloses a microwave laser hybrid link radio star calibration device and method suitable for deep space exploration, and relates to the field of spaceflight deep space exploration. Pointing calibration of a microwave and laser hybrid link of a deep-space large-aperture antenna is completed through radio star calibration, firstly, after pointing calibration and stable tracking of the microwave link are conducted on the microwave link, fine tracking is conducted on the laser link, pointing calibration of the laser link is completed, respective pointing error correction coefficients are formed, and real-time correction of antenna pointing is achieved. Microwave laser link hybrid radio star calibration is not limited by distance and geographical conditions, and real-time calibration can be realized; the beam of the laser link is narrower than that of the microwave link, the calibration precision is higher, and the system pointing precision and the photoelectric axis consistency precision can be improved. The method is simple in design, high in generalization, low in implementation difficulty and convenient to implement and popularize.

The invention provides a deep space exploration task level demand analysis method, which comprises the steps of forming a task mission based on strategic planning, domestic and overseas trend development and engineering foundation; stage division and periodic task planning are carried out, benefit-related parties of the task are sorted, appeals of the benefit-related parties for the task are investigated, and a target system of the task is obtained based on task constraints; performing task scene analysis, task scene refinement, task capability analysis and task index analysis, and outputting a task capability system; carrying out capability clustering, capability index construction, system architecture design and system interface design to obtain a system architecture design model and a capability distribution model; and finally obtaining a deep space exploration task flow. According to the method, visual expression and analysis are carried out on deep space exploration mission and demands based on multiple views, initial design of deep space exploration mission tasks and system demands is achieved, then based on analysis of capability demands, a large system architecture is formed, initial logic design is completed, and finally a task-level demand model is formed.

The invention discloses a microwave laser composite antenna off-axis tracking method suitable for deep space exploration, and relates to the field of spaceflight deep space exploration. According to the method, Ka frequency band link angle tracking is completed firstly, and then a laser link is guided to be captured and tracked rapidly. Microwave link wave beam off-axis tracking is adopted, and simultaneous real-time target tracking of a microwave link and a laser link is achieved; the Ka beam is used for guiding the laser beam to rapidly capture and track, so that the laser beam coarse tracking robustness is improved, and the capture time is reduced; after the laser link finishes fine tracking, the system performs target tracking by using laser beam tracking, and compared with microwave Ka beam tracking, the tracking precision is higher; by using microwave link wave beam off-axis tracking, the problem that microwave link wave beams and laser link wave beams cannot be simultaneously aligned with a target due to inconsistent directions of a microwave electric axis and a laser optical axis is avoided. The scheme is simple and easy to realize, and can be directly applied to the existing integrated antenna.

The invention provides an inter-station time synchronization constrained VLBI world time measurement method, which comprises the following steps: determining at least two observation stations for observing a radio power source, and carrying out precise clock synchronization to obtain a precise inter-station clock difference; using the at least two observation stations to simultaneously observe the same radio power source to obtain observation time delays from the same radio power source to the at least two observation stations; separating the clock error in the observation time delay by using the precise inter-station clock error to obtain the observation time delay of the separated clock error; and removing ionosphere time delay, troposphere time delay and equipment time delay in the observation time delay of the separated clock error to obtain final observation time delay, and resolving a residual error between the final observation time delay and theoretical time delay to obtain world time. According to the invention, precise time synchronization between observation stations is carried out in advance, and inter-station errors are effectively separated. The problem that VLBI observation time delay and inter-station clock correction are difficult to separate is solved, the UT1 resolving precision can be effectively improved, and the requirement for high-precision UT1 in the fields of deep space exploration and the like is expected to be met.

The invention relates to the technical field of spaceflight sealing, in particular to a multi-mode pressure sealing device and method for heating and extracting extraterrestrial star soil gas. The device comprises a transposition sealing disc, a positioning base disc and a distributed driving actuator, a flow guide funnel is arranged on the outer side of the transposition sealing disc, and a sample receiving station and a sealing station are arranged on the inner side of the transposition sealing disc; an interface cylinder is arranged on the outer side of the positioning base disc, and a sample receiving station hole is formed in the inner side of the positioning base disc; the distributed driving actuator is connected with the transposition sealing disc and the positioning base disc and used for driving the transposition sealing disc to conduct axial rising, station switching and axial sealing relative to the positioning base disc in a time-sharing mode. The problem of dynamic sealing failure under the extreme temperature difference and high particle invasion environment can be effectively solved, zero-leakage collection of star soil heating gas is achieved, the reliability and operation efficiency of gas sample collection in a deep space exploration task are remarkably improved, and mechanical interference or positioning deviation caused by compound motion of a traditional linkage mechanism is also avoided.

The invention discloses an orbit control safety shutdown time calculation method suitable for deep space exploration, and belongs to the field of deep space exploration. The method comprises the steps that firstly, the theoretical speed increment of orbit control is calculated, and a sinking engine, an orbit control engine, an attitude control engine and related engine thrust and specific impulse parameters are determined; secondly, the actual orbit control speed increment is obtained, and the orbit control nominal starting time length is determined; thirdly, calculating time error deviation, and giving out engine shutdown forbidding time and forced shutdown time in the orbit control process; if the task is a conventional orbit control task, taking the two times as final safe shutdown time; for the orbit control task related to safety, multiple groups of shutdown forbidding time and forced shutdown time under the nominal working condition and the pulling deviation working condition are given, and finally the safety shutdown time of the orbit control task related to safety is determined. According to the method, the track safety and task safety of the detector are ensured by selecting proper shutdown forbidding time and forced shutdown time.

The invention provides a complex task track thermal environment adaptability assessment method and system. The complex task track thermal environment adaptability assessment method comprises a thermal environment adaptability assessment test item making method; cold and hot alternating environment adaptability assessment test conditions and an accelerated test method; high-temperature environment adaptability assessment test conditions and an accelerated test method; the invention relates to extreme low-temperature environment adaptability assessment test conditions and an accelerated test method. According to the invention, the problem of how to carry out sufficient and effective thermal environment adaptability assessment on a product undergoing a deep space exploration extremely severe thermal environment under a complex task orbit profile is solved, and a reference can be provided for thermal environment adaptability assessment of a wood star exploration model and a subsequent deep space exploration model.

The invention discloses a lunar balance dynamic parameter high-precision resolving method and system combining LRR and LLR technologies, and belongs to the field of deep space exploration and celestial body dynamics, and the method comprises the steps: constructing an LRR and LLR combined observation data set, and carrying out the preprocessing; constructing an LRR and LLR combined moon measurement model, and deducing a partial derivative of the model for a moon rotation parameter and a tidal LOVE number to obtain a combined partial derivative matrix; iterative solution is carried out through nonlinear least square to obtain dynamic parameters of the moon balance; and evaluating a resolving result and optimizing a resolving strategy, and optimizing a resolving scheme by dynamically adjusting the weight, improving a tide model and eliminating abnormal values. According to the method, the parameter calculation error is reduced, the technical advantages of two types of data are effectively integrated, the problem of insufficient observation continuity or precision of a single data source is solved, and reliable parameter support is provided for moon rotation and internal structure research.

The invention discloses a planetary atmospheric trace gas spectrum telemetering and in-situ fusion detection system and method, and belongs to the technical field of deep space exploration. The detection system has a telemetering mode and an in-situ detection mode, and comprises a telescope module, a laser light source module, a heterodyne spectrum telemetering module, an in-situ absorption spectrum module, a gas sampling and enriching module and a signal acquisition, processing and control module. In the on-orbit stage, a heterodyne spectrum telemetering module is adopted, and a laser light source and sunlight signals are subjected to frequency mixing for atmospheric component telemetering; and after the detector enters an atmospheric descending section, the in-situ absorption spectrum module performs in-situ absorption spectrum measurement by utilizing the detection light of the same laser light source. By performing space-time fusion and joint inversion on the data acquired in the two modes, coherent detection from large-range general survey to fixed-point detailed survey is realized, the problems of load separation of telemetering and in-situ detection and difficulty in data fusion are solved, and the detection efficiency, the data consistency and the integration level of the system are greatly improved.

A self-sensing self-adaptive deformable wheel tensioned and integrally supported by a paper folding configuration shape memory alloy wire belongs to the technical field of mechanical engineering and robots and is composed of a paper folding deformable module, a self-sensing module and a driving module, and the modules cooperate to realize the functions of wheel diameter dynamic adjustment, load self-sensing, terrain self-adaptive deformation and rigid-flexible coupling adjustment. The paper folding deformable module is designed based on an improved water bomb inlaid paper folding pattern, and an overlapping face, a small flexible piece and an additional face are additionally arranged to improve deformation flexibility. The self-sensing module comprises a tensegrity structure, continuous rigidity switching is achieved through pre-tightening force adjustment, and the load is monitored in real time through the strain-resistance effect of a shape memory alloy wire; and the driving module is directly driven by a radial electric push rod to realize wheel diameter change to adapt to terrains. By means of the structure-sensing-driving integrated design, the intelligent self-adaptive robot has the advantages of being light in weight, high in bearing capacity and intelligent and self-adaptive, and is suitable for complex terrain moving platforms such as deep space probe vehicles and disaster relief robots.

The invention discloses an intelligent prediction guidance method for Mars pneumatic auxiliary orbit descending, and belongs to the field of deep space exploration. According to the method, a high-precision kinetic model of the Mars pneumatic-assisted orbit descent is constructed, a network training data set is generated by utilizing the high-precision model, a near-fire-point state intelligent forecasting model and a far-fire-point state intelligent forecasting model of the Mars pneumatic-assisted orbit descent are constructed, and meanwhile, near-fire-point heat flow constraints are constructed; designing a predictive guidance method by taking the orbit height deviation and the control input as optimization performance indexes; and optimizing and solving the control input to control the detector until the track height of the detector is reduced to be close to the preset track height, so that the detector is reduced to the preset track under the condition of meeting heat flow and control amplitude constraints, and meanwhile, the fuel consumption of the detector is reduced. According to the method, by solving the states of the near fire point and the far fire point under the two-body dynamic model, the network is constructed to predict the deviation value between the states of the near fire point and the far fire point under the high-precision dynamic model and the prediction result of the two-body model, and the training cost of the neural network is reduced.

The invention belongs to the technical field of board games, and discloses a deep space exploration board game interactive education system based on astronomical science popularization, which adopts a galaxy, a fixed star, a planet, an asteroid belt and the like as elements of a game map to construct a macro cosmic space scene. Astronomical knowledge is decomposed into a plurality of independent modules, and the independent modules are fused into a game process, so that flexible combination and progressive education among the modules are realized. According to the game, through the combination of cards and text specifications, the science universality and education are enhanced, meanwhile, a random mechanism for simulating astronomical phenomena is introduced, and challenging and interestingness are increased for the game. A player can gradually uncover the veil of the astronomical phenomenon through the tidy puzzle and carry out deep space exploration by means of astronomical theme roles and props. According to the invention, on the basis of modular design, complicated astronomical knowledge is transmitted to the public in a visual and attractive manner through diversified interaction modes, so that the interestingness and practicability of astronomical science popularization education are improved.

The invention provides a deep space exploration simulation test system and method.The deep space exploration simulation test system comprises a thermal interface structure, a cryogenic subsystem and a heating structure, the cryogenic subsystem can refrigerate and cool the thermal interface structure, and the heating structure can heat the thermal interface structure; the deep cooling subsystem comprises a low-temperature refrigeration assembly and a gas cooling assembly, the low-temperature refrigeration assembly transmits cold energy to the thermal interface structure in a one-way mode, and the gas cooling assembly cools the thermal interface structure through cooling gas. The continuous temperature regulation in the range of 300K to 423K is realized through the combination of air cooling and electric heating, and the continuous temperature regulation in the range of 38K to 300K is realized through the combination of a G-M low-temperature refrigerator and electric heating; compared with the existing space environment simulation equipment, the temperature zone can be expanded to 38K to 423K, and the extreme temperature zone environment required by deep space exploration including the lunar surface polar zone and the Mars can be completed; meanwhile, the integrated heat exchange cryogenic subsystem based on heat conduction and convection is compact in structure, liquid nitrogen and liquid helium do not need to be used, and energy conservation and environmental protection are achieved.

The invention discloses an installation method for mounting a large-aperture microwave antenna on an optical antenna, and relates to the field of spaceflight deep space exploration. The method comprises the following steps: firstly, welding a mounting steel plate in a microwave antenna center body, forming a hole in an antenna panel, then integrally hoisting an optical antenna and an adjusting bracket into the microwave antenna center body through a hoisting device, and mounting the optical antenna and the adjusting bracket on the mounting steel plate; and finally, the angle of the optical antenna is adjusted through the adjusting bracket, and a cable is arranged. The installation mode is simple, the structure is stable, and the electric axis of the microwave antenna and the optical axis of the optical antenna can be adjusted to be consistent. According to the invention, the microwave antenna can be upgraded and reformed, and a composite structure of the optical antenna and the microwave antenna is realized.

The invention discloses a SiO2-coated ZnO white dustproof coating prepared based on a step-by-step sol-gel method and a preparation method of the SiO2-coated ZnO white dustproof coating, and belongs to the field of spatial special functional materials. According to the invention, stable performance of the thermal control coating under lunar dust pollution is realized. The preparation method comprises the following steps: introducing SiO2 particles on the surface of zinc oxide through a sol-gel method to construct a micro-level rough surface structure, adding a low-surface-energy modifier into a mixed solution of zinc oxide and silica sol, uniformly mixing, and drying to obtain multi-scale hydrophobic SiO2 (at) ZnO particles. The preparation method comprises the following steps: putting a SiO2 coated ZnO powder material into silica sol prepared in advance, adding a surfactant, uniformly mixing to prepare white slurry, finally covering the slurry on an alloy substrate through a blade coating technology, and curing to form the SiO2 coated ZnO white dustproof coating. The method is applied to the field of deep space exploration and has a very wide application prospect.

The present invention discloses a robust optimization method for autonomous optical navigation landmarks for planetary landing, which belongs to the field of deep space exploration technology. The implementation method of the present invention is as follows: using the height information of the lander to determine the scale of the image plane projected onto the planetary surface, and deriving the quadrilateral vertex expression of the field of view based on the geometric relationship; using the relationship between the field of view vertex deviation and the lander posture error, converting the posture uncertainty to the field of view vertex uncertainty, and using the error ellipse to describe the field of view vertex uncertainty distribution; screening the quadrilateral area formed by the intersection of common tangents, which is the field of view intersection under the lander posture uncertainty; using the nominal trajectory and camera parameters to estimate the field of view sequence, combining the field of view intersection acquisition method to obtain the field of view intersection sequence, the field of view intersection sequence is obtained by offline estimation, and the amount of onboard calculation of the lander is reduced; using the observability of the navigation system to establish the landmark optimization index, optimize the selection of the landmarks in the field of view intersection sequence, and reduce the influence of posture uncertainty on the selected landmarks.

The invention discloses an FPGA (Field Programmable Gate Array) hardware system based on an intelligent sensing algorithm and a deployment method thereof, and relates to the technical field of intelligent video image processing, and the FPGA hardware system comprises a reconfigurable sensing module, a compatible FPGA module, an expandable external storage module and a customizable information processing module, through the technical means of a configurable hardware system scheme, a pipelined deployment process method, an extensible external storage module, a customized information processing module and the like, the efficient, flexible and extensible FPGA hardware system based on the intelligent sensing algorithm is provided. The system can meet the deployment requirements of different intelligent sensing algorithms, improves the performance and processing efficiency of the algorithms, and can be widely applied to the fields of industry, aerospace, deep space exploration and the like.

The invention discloses a lunar scientific research station micro-grid system based on a multi-source access electric energy router, and belongs to the technical field of deep space exploration and micro-grid energy supply systems. The micro-grid system comprises a medium-large micro-grid and a plurality of small micro-grids, and the medium-large micro-grid takes a high-voltage alternating-current backbone network as a framework and is connected with the small micro-grids; each micro-grid comprises a plurality of electric energy routers, and the electric energy routers are connected in parallel through high-voltage alternating-current ports; the electric energy router is connected with various distributed power supplies through a plurality of ports, supplies power to a load, and is responsible for scheduling and managing multi-source energy; and each small micro-grid performs micro-source configuration and port and capacity configuration of the electric energy router according to the load and in-situ resource characteristics of the power supply area of the small micro-grid. A quasi Z source converter and high-frequency transformer technology is adopted to realize cross-voltage-level energy efficient conversion, the flexibility and robustness of an energy system are remarkably improved, and theoretical support and engineering reference are provided for all-weather autonomous operation of a lunar scientific research station.

The invention discloses a high-precision quantum sensing environment physical quantity real-time monitoring system and method, and belongs to the technical field of environment monitoring. The system comprises a quantum sensing front end, a signal excitation and regulation module, a quantum classical hybrid computing engine, an edge cloud collaborative transmission module and a terminal application module. The quantum sensing front end adopts a quantum entanglement sensing array integrated by NV color center diamond and a rubidium atom gas chamber to synchronously sense a magnetic field, temperature, pressure and vibration; the signal excitation and regulation module optimizes regulation parameters and inhibits noise through a quantum optimal control algorithm; a quantum-classical hybrid calculation engine is combined with a quantum neural network and classical multi-algorithm collaboration to realize efficient resolving and data fusion of quantum state features. According to the invention, the monitoring precision, the real-time response speed and the anti-interference capability are obviously improved, synchronous monitoring of multiple physical quantities is supported, and the system is suitable for various scenes such as industry, scientific research and deep space exploration, and has wide application value.

The invention discloses a point cloud registration method based on target sailboard features. In order to solve the problem of registration error accumulation when low-overlapping-rate sparse point clouds are processed under the condition of resource limitation in the prior art, the invention provides the following improvement schemes: firstly, designing a preprocessing module in which multi-scale statistical outlier detection and radius filtering local smoothing processing are coordinated, so that the noise point filtering rate is improved to 92% or above; secondly, a weighted PCA coordinate system construction method constrained by an axial bounding box (AABB) is innovatively provided, and the fitting precision of the coordinate system is improved by 40% through a weighted fusion strategy of a covariance matrix; and finally, constructing a registration evaluation function for verifying geometric consistency based on key points, and effectively eliminating mismatching caused by symmetric ambiguity in combination with an improved RANSAC algorithm. Experiments show that the method can still keep the registration success rate of 89.7% when the point cloud overlapping rate is lower than 30%, and the registration error is reduced by 62.5% compared with a traditional ICP algorithm. The method is especially suitable for non-cooperative target three-dimensional reconstruction in resource limited scenes such as deep space exploration, and provides reliable technical support for on-orbit service of spacecrafts.

The invention discloses a method for calculating the requirement of signal capture among lunar orbit formation satellites on satellite orbit measurement accuracy, and belongs to the technical field of deep space exploration. The method comprises the following steps: under a ground test condition, measuring an antenna pattern, transmitting signal power and receiving sensitivity of a master satellite, measuring an antenna pattern, receiving sensitivity and a transmitting EIRP value of a slave satellite, measuring an installation error and an antenna mechanism deformation error of a high-gain directional antenna of the master satellite, and measuring attitude control precision of the master satellite. According to the main satellite antenna pointing total error requirement, the antenna installation error, the antenna mechanism deformation error and the beam pointing error caused by the attitude control precision of the main satellite, the position error requirement of the slave satellite relative to the main satellite is calculated; and determining orbit measurement precision requirements of the master satellite and the slave satellite based on the position error. According to the method, the orbit measurement requirement can be put forward to a ground measurement and control system in advance, or the proper inter-satellite distance can be selected according to the orbit measurement precision provided by the ground, and the success probability of establishing the inter-satellite link is improved.

The invention relates to the technical field of deep space exploration, in particular to a cryogenic propellant on-orbit low-loss storage system and method for a deep space probe. After the low-temperature propellant is injected into the storage tank through the injection pipe, along with continuous accumulation of external leaked heat, the first pressure sensor detects that the pressure in the storage tank is continuously increased, and when the numerical value of the first pressure sensor exceeds the preset pressure, the cooling assembly is started to cool the low-temperature propellant in the storage tank, so that the external leaked heat is consumed; before the engine needs to supply the low-temperature propellant, the low-temperature propellant in the storage tank is in a supercritical state, and the storage temperature is increased in a high-pressure storage mode, so that the temperature difference between the storage tank and the outside and the heat leakage amount are reduced, and the requirement for the refrigerating capacity of the cooling assembly is reduced; when the low-temperature propellant is conveyed, due to the fact that the low-temperature propellant is in a supercritical state and free of gas-liquid phase separation, the interior of the whole storage box is in a homogeneous state, the stable conveying flow of the low-temperature propellant can be guaranteed, and stable output of an engine is also facilitated.

The invention relates to the technical field of deep space exploration visual three-dimensional terrain reconstruction error analysis, in particular to a visual three-dimensional terrain reconstruction error modeling method and system for a Mars patroller. In view of the limitation of the Mars surface environment and the existing error analysis method, the purpose of improving the three-dimensional reconstruction precision is achieved. Technically, the method comprises the following steps: sparsely reconstructing a covariance propagation model based on bundle adjustment, processing a camera imaging equation, carrying out re-projection error calculation and optimization, carrying out nonlinear least square algorithm iterative solution and the like to obtain a camera attitude, a three-dimensional point position parameter and a covariance matrix of the three-dimensional point position parameter; densely reconstructing a covariance propagation model based on linear triangularization, and calculating a three-dimensional point coordinate covariance matrix through equation conversion, least square solution and covariance propagation formula derivation. The experiment simulates the Mars environment to obtain simulation data for dense reconstruction error analysis, and the result shows that the method can effectively analyze errors, significantly improve the three-dimensional terrain reconstruction precision, and powerfully guarantee Mars detection and navigation tasks.

The invention discloses a lunar soil penetration type shearing mechanical in-situ testing device, and relates to the field of testing devices for deep space exploration. The blade conical head is suspended; the penetration and recovery mechanism is arranged between the bracket and the blade conical head; the first driving mechanism is arranged between the bracket and the injection and recovery mechanism; the first driving mechanism is used for driving the penetrating and recycling mechanism to ascend and descend so as to drive the blade conical head to penetrate and recycle. And the second driving mechanism is arranged between the penetration and recovery mechanism and the blade conical head so as to drive the blade conical head to rotate and shear. Through cooperative driving control of the penetration and recovery mechanism, the blade conical head, the first driving mechanism and the second driving mechanism, automatic penetration, shearing and recovery of the blade conical head can be achieved, and manual penetration and recovery are not needed; the device is more suitable for being installed on a lunar vehicle for unmanned lunar soil in-situ mechanical testing in a lunar exploration task, and the exploration efficiency is improved.

The invention discloses a deep space exploration low-power-consumption AI navigation algorithm processing system based on a storage and calculation integrated architecture, the system adopts the storage and calculation integrated architecture based on RRAM cross arrays and calculates an AI navigation algorithm, the system comprises an RRAM storage and calculation core, a dynamic configuration PE array, an anti-radiation controller and an on-chip storage interaction unit, the RRAM storage and calculation core comprises a plurality of independent RRAM cross arrays, in-situ execution of matrix-vector multiplication and vector-matrix multiplication is supported; the dynamic configuration PE array is used for processing nonlinear operation, and dynamic switching of two parallel granularity modes of 16 * 16 and 4 * 4 is supported through a configuration register; the anti-radiation controller is used for repairing radiation induction errors in real time and coordinating task scheduling of the storage and calculation core and the PE array; and the on-chip storage interaction unit adopts an SRAM (Static Random Access Memory) as an intermediate cache and is used for temporarily storing intermediate data. The problem of high power consumption of an AI navigation system for extreme environment detection can be solved.

The invention relates to the technical field of deep space detection equipment, in particular to an extraterrestrial celestial body rock coring drill and a coring and breaking method thereof. The coring drill comprises an attachment mechanism, a rotary impact sliding table mechanism and a coring and core breaking mechanism which are sequentially connected, the attachment mechanism is used for being anchored to the surface of celestial rock, and the rotary impact sliding table mechanism is used for providing drilling rotation, impact power and axial feeding motion; the coring and breaking mechanism comprises an outer drill bushing, an eccentric coring pipe assembly mounted in the outer drill bushing and a drill bit connected to the front end of the outer drill bushing; the eccentric coring pipe assembly comprises a coring pipe and a driving shaft for driving the coring pipe to rotate, and the coring pipe and the driving shaft are eccentrically arranged; the driving shaft can be controlled to synchronously rotate in the same direction with the outer drill bushing or independently rotate in the opposite direction relative to the static outer drill bushing. Active anchoring and multiple buffering are combined, the sample integrity, the operation success rate and the automation degree of extraterrestrial hard rock coring are remarkably improved, and lossless capture and mechanical reliable breaking of a rock core are achieved.