34 results about "Orbiter" patented technology

The invention discloses a hydrous mineral detection method and system based on the hyperspectrum of a Mars orbiter, and relates to the technical field of hyperspectral data processing and analys.The method comprises the steps that a distribution area of hydrous minerals is obtained based on Mars hyperspectral data; performing automatic end member extraction on the distribution area to obtain an image end member; and performing spectral feature inspection on the image end members, and performing unmixing based on an image end member library to obtain a hydrous mineral abundance map. According to the method, the characteristics of the water-containing minerals can be automatically identified, the mineral distribution range can be accurately positioned, meanwhile, the de-mixing end member library can be automatically constructed, the end members can be checked, and manual participation is reduced; the end member library is automatically constructed from the image based on the qualitative recognition result, and through overlay analysis and comparison with other unmixing end member construction methods, it is shown that the method has high credibility in the aspect of reflecting the distribution of the hydrous minerals and the abundance of the minerals.

The application discloses a small-sized separated penetrator for in-situ detection of celestial bodies, and relates to the technical field of penetrator equipment, comprising an aluminum honeycomb buffer module, which comprises an external protective shell and an aluminum honeycomb, and has the beneficial effects that the application is provided with in-situ detection of high-speed penetration of celestial bodies, and discloses a small-sized separated penetrator, which comprises a planet soil penetration module and a star table detection module, and realizes autonomous separation in the impact penetration process, the star table module can realize star table detection and communication with an orbiter, the penetration module is internally provided with an energy storage device and an impact hammer, and can continue to penetrate and dive relying on the energy storage device after impact; the star table module is connected with the penetration module through flexible wires, and can supply power and signal transmission to the latter, the technology enables the penetrator with the modular characteristic to have the in-situ detection capability of planet soil profile, aims to expand the application mode of the penetrator for celestial body detection, and provides a scientific basis for the research of star table structures.

The invention discloses a self-correction state forecasting and fusion estimation method for a Mars pneumatic auxiliary orbit descending model, and belongs to the technical field of deep space exploration. The implementation method comprises the following steps: integrating various perturbation of a detector, and establishing a dynamic model suitable for long-term orbit extrapolation; the orbiter observes a natural large landmark on the surface of the Mars by using a navigation camera to obtain azimuth information of the landmark; when the detector is far away from the Mars, azimuth information of the Mars is obtained through the navigation camera; when the distance is closer to the Mars, observing a satellite surface landmark, and acquiring azimuth information of the landmark; the detector and the orbiter measure the relative distance and orientation through radio, and the measurement information obtained through fusion corrects the extrapolation result of the orbit. When the detector is located in the atmosphere, an accelerometer is used for measuring the external force acceleration of the detector, and inertial navigation is achieved. The atmospheric density is corrected by using the dynamic pressure and the heat flux density, and re-recursion is performed on the state of the detector, so that the state estimation precision of the detector in the Mars atmosphere is improved.

A rocket propulsion and satellite deployment system, comprising a rocket comprising at least a base sector, an operating sector, and a transport sector; the base sector comprises a propellant and is configured to transport the sectors to a mission target; the operating sector comprises a rotating robotic arm pivoted to the operating sector; the transport sector is configured to hold a load intended to be delivered to the target by means of a rotation of the rotating robotic arm with respect to the operating sector, accelerating the load itself which is subsequently released in a direction opposite to the speed of the rocket, increasing, due to conservation of momentum, the speed of the rocket itself. The system is based on a so-called 'useful propellant', consisting of the satellites, rovers and orbiters themselves, which, through the rotational acceleration system, are used to increase the speed of the rocket in space. As well as providing an alternative to traditional propulsion, this system provides an innovative way of putting satellites into orbit, especially in non-earth orbit, which can also be used for longer missions with more than just one target.

The invention belongs to the technical field of space flight and aviation heat dissipation, and particularly relates to a heat exchange device for a sub-orbital aircraft load under a dynamic acceleration condition, which comprises a flexible heat pipe, an evaporation section of the flexible heat pipe is coupled with a packaging structure of a heat source through a high heat conduction fixing piece, and a condensation section of the flexible heat pipe is connected with a sliding device through a heat insulation fixing piece; the flexible heat pipe is filled with working fluid. The sliding device is in sliding connection with the bottom wall of the cylindrical shell, responds to centrifugal force generated by dynamic acceleration of the aircraft and drives the condensation section of the flexible heat pipe to move to the position opposite to the acceleration direction along the bottom wall of the cylindrical shell; through the combination of the sliding device and the flexible heat pipe, the condensation section is driven by centrifugal force to move to the bottom in the acceleration direction, a high-temperature area is formed, the temperature layout in the cylindrical shell is actively optimized, the circulation efficiency of hot air rising and cold air sinking is remarkably improved, and the traditional natural convection efficiency bottleneck is broken through; and an innovative solution is provided for thermal management of the spacecraft under the high-dynamic working condition.

The invention provides an atmospheric parameter inversion method and system for a high-speed orbiter. The method comprises the following steps: acquiring pressure data of a plurality of measuring points arranged on the surface of the high-speed orbiter; inputting the pressure data into a pre-established mapping relation model to obtain and output corresponding atmospheric parameters; wherein the mapping relation model is a nonlinear model obtained by training by adopting a neural network fused with extended Kalman filtering according to the corresponding relation between the measuring point pressure data of the high-speed orbiter in the flight envelope and the atmospheric parameters. On the basis of the theoretical basis of the one-to-one correspondence single-valued mapping relation between the surface pressure data and the corresponding atmosphere data, the design method for inversing the atmosphere data based on surface pressure reverse modeling is provided, and the technical problem that the surrounding environment parameters of the high-speed orbiter cannot be obtained in real time through a traditional design method is solved; meanwhile, the attack angle and the sideslip angle of the orbiter are solved, and the technical problem of obtaining the high-precision attitude angle of the high-speed aircraft is solved; the applicability is high, and the calculation precision is high.

The invention relates to the technical field of spacecraft telemetry data analysis, in particular to a telemetry data analysis method for deep learning of a motor West-odd-female-Peng spacecraft, and the method comprises the steps: collecting the telemetry data of the spacecraft through a sensor and a monitoring device, carrying out the cleaning, denoising, standardization and other preprocessing operations of the original telemetry data, and obtaining the telemetry data of the motor West-odd-female-Peng spacecraft. A silicon-based memory function is used for building network nerves to simulate human brain memory; feature extraction is carried out on the preprocessed telemetry data; a model is built based on the extracted features; training is carried out by using annotation data, model parameters are optimized, and the trained model is applied to actual telemetry data; tasks of fault prediction, anomaly detection, performance evaluation and the like are achieved, task results are displayed in the forms of charts, reports and the like, support is provided for decision making, and the problem that a traditional telemetry data analysis method is difficult to meet the requirements for real-time performance, accuracy and intelligence is solved.

This invention discloses a method for predicting the high-Mach aerodynamic characteristics of a Mars orbiter based on FLUENT, belonging to the aerospace field. This invention establishes a symmetrical geometric model of the Mars orbiter and uses a dual-temperature model based on FLUENT to achieve numerical simulation of the Mars orbiter under high-Mach conditions in CO2 medium. By setting a boundary layer based on O-Block mesh coupling, it achieves unified simulation prediction for different velocities and angles of attack, reducing the computational burden of re-partitioning the mesh for different conditions, thereby improving the efficiency of aerodynamic characteristic prediction. Simultaneously, it can accurately simulate and analyze the aerodynamic characteristics of the Mars orbiter under different conditions, enabling the corresponding aerodynamic performance evaluation of the Mars orbiter. This invention can analyze and construct the hypersonic aerodynamic characteristics of the Mars orbiter under different high-Mach conditions upon entering the Martian atmosphere, assisting in the optimization of the Mars orbiter and guidance and control commands.

The sub-orbital airship comprises a single-stage rocket and an airship body, the airship body is installed at the top of the single-stage rocket and comprises a load cabin, a propelling cabin is arranged at one end of the load cabin, a consignment cavity is formed in one end of the load cabin, and a supporting cavity is formed in the other end of the load cabin. A consignment cavity is formed in the base, one end of the consignment cavity is connected with an airship nose cone, an air-drop platform is arranged in the consignment cavity, the air-drop platform comprises a first compression releasing machine, one end of the first compression releasing machine is in transmission connection with an air pressure lifting platform, and the airship nose cone comprises a landing cabin body. According to the method, safe recovery and reuse after launching can be achieved, intact recovery of airships and rockets is ensured, the launching cost is greatly reduced, the resource utilization rate is increased, and a solid technical foundation is provided for wide application of sub-orbital aircrafts in the fields of commercial spaceflight, scientific research, emergency response and the like.

The present application relates to a method for designing a transfer orbit for a HALO orbiter to fly over a given ground point. The method includes: finding the moment when the lunar declination value is equal to the latitude value of the given ground point within the given mission time range, and selecting the time point closest to the initial moment as the initial value of the iteration for the orbit transfer moment; calculating the flight duration of the orbit transfer point relative to the starting point using the starting time and the initial value of the iteration for the orbit transfer moment; calculating the normalized time by dividing the flight duration by the HALO orbit period; calculating the magnitude of the orbit transfer velocity increment based on the normalized time using the polynomial fitting coefficients; establishing a mapping relationship between the magnitude of the orbit transfer velocity increment and the initial value of the iteration for the orbit transfer moment through numerical integration based on a high-precision force model; and completing the solution of the parameter to be solved using the Newton iteration method based on the mapping relationship. The method adopted has a short response time, high calculation result accuracy, and fast iteration convergence.

A planetary rotation system for coating includes a solar orbiter plate configured to rotate about a solar axis. A lunar rotation system is mounted to the solar orbiter plate. The lunar rotation system includes a lunar orbiter plate configured to rotate about a planetary axis. A plurality of lunar substrate holders is mounted on the lunar orbiter plate. Respective lunar substrate holders include a lunar axis extending centrally therethrough. The respective lunar axes are configured to rotate about the planetary axis. The respective ones of the lunar substrate holders are configured to rotate about their respective lunar axis and to hold a substrate for coating.

The invention provides a dynamic evolution analysis method for a sub-orbital fragment risk area. The method comprises the following steps: acquiring fragment group data before and after a sub-orbital aircraft is disassembled; constructing a fragment dynamic model, and establishing a state function including position, speed and acceleration; the fragment group is grouped and processed, and an optimal grouping scheme is determined by adopting a Freedman-Diacois rule; performing state prediction and updating on each group by combining a Runge-Kutta method and a self-adaptive unscented Kalman filtering algorithm; constructing a probability distribution model of the fragment group based on a grouping processing result, and representing by adopting multivariate t distribution; and calculating the risk probability of the fragment group in the specific space region through probability density function integration. Through the adaptive grouping processing strategy, the limitation of overall approximate processing of the fragment group in a traditional method is broken through, the spatial distribution characteristics and the evolution law of the large-scale fragment group can be accurately described, and the accuracy of group behavior analysis is remarkably improved.

The application provides a folding and unfolding rod side swing continuous transfer celestial orbiter transfer device and method, and belongs to the field of celestial orbiter transfer. The transfer device comprises a swing rod, a boom, a slow-release rope, an attitude control rope and a fixed-length rope. One end of the swing rod is connected with one end of the boom through an attitude control hinge, and the other end is connected with a first cam through a hinge. The other end of the boom is connected with the celestial orbiter through a separation mechanism. The first cam is connected with a swing motor. The swing motor controls the swing rod to swing along the horizontal direction. One end of the slow-release rope is wound around a slow-release rope wheel, and the other end is connected with the separation mechanism. The slow-release rope wheel is connected with a slow-release motor. One end of the attitude control rope is arranged on the first cam, and the other end is connected with the boom. One end of the fixed-length rope is wound around a fixed-length rope wheel, and the other end is connected with the swing rod. The slow-release rope wheel, the fixed-length rope wheel, the slow-release motor, the swing motor and the first cam are arranged on a lander. The device is mainly used for the transfer of the celestial orbiter.

The application relates to a kind of generalized control constraint-based joint adjustment method of Mars multi-source orbiters image, comprising: for low-resolution reference DOM and Mars orbiters optical image, extracting plane control point, and matching to obtain connecting point;Based on the fitting residual of Mars orbiters image rational function model, establish the image side compensation model of superimposed polynomial and cubic spline function and obtain initial compensation parameter;According to plane control point, connecting point and initial compensation parameter, combined with plane and elevation control constraint, the adjustment model is constructed, and the adjustment model is solved using adaptive weighting strategy, and the geometric positioning parameter of image is obtained.Compared with the prior art, the application can effectively improve the absolute and relative positioning accuracy of Mars multi-source orbiters image by carrying out multi-source image joint beam adjustment in cooperation with plane weak control constraint and elevation control constraint, and realizes the accurate registration of multi-source orbiters image.

The invention discloses a multi-source information fused Mars entering integrated navigation method and system. The method comprises the following steps: acquiring multi-source measurement information of the Mars entering an aircraft in an entering stage, and performing time synchronization and coordinate conversion on the multi-source measurement information to obtain initial observation data; establishing a Mars entry dynamical model based on Mars rotation characteristics, and performing state prediction on the dynamical model by using the initial observation data to obtain a prediction state vector; constructing an observation model, associating the prediction state vector with the distance measurement and speed measurement information of the orbiter and the beacon to form a combined observed quantity, and performing estimation updating on the combined observed quantity by using a nonlinear filtering algorithm to obtain an updated state estimation result; and calculating the condition number of the observation matrix according to the updated state estimation result, and when the condition number is smaller than a preset threshold value, determining that the system is completely observable, and outputting a final navigation calculation result. The technical problem of inaccurate navigation in the Mars entering stage is solved.

The present invention discloses a small penetrator for planetary subsurface penetration detection and its working method. The penetrator structure includes: a penetration buffer module, a spring power module, an unlocking and positioning module, and a communication antenna spiral expansion module. When locked, the bottom protrusion of the antenna frame is located in the T-slot; when unlocked, the antenna frame is driven by the motor to rotate out of the T-slot, and under the action of the spring power module, the unlocking and positioning module is driven to rise along the inner wall of the penetrator. When it reaches the specified position, the positioning pin realizes positioning. The present invention relies on a lander or orbiter for precise delivery and performs in-situ detection of the planetary subsurface soil. It has the advantages of compact structure, light weight, low cost, high reliability, etc., and meets the task requirements of planetary subsurface soil penetration detection.

The invention belongs to the cross technical field of satellite gravity mechanics, geodetics, orbit perturbation, space science and the like. The invention particularly relates to a planet gravity field inversion method. The planet gravity field inversion method comprises the following steps: S1, acquiring shadow region duration of a planet in an orbital period through orbital device load data; s2, acquiring an included angle beta (t) between a sun-planet vector and an orbiter plane through a conical earth shadow model or a cylindrical earth shadow model according to the shadow region duration time tau (t), and further acquiring an ascending node right ascension advancing amount omega (t); s3, solving the ascending node right ascension advancement amount omega (t) to obtain a long-term item and a long-period item; and S4, obtaining a low-order even-number term harmonic coefficient and a low-order odd-number term harmonic coefficient of the gravity field by using a linear perturbation method, an average perturbation method or an analytical method. According to the invention, the low-order gravity field can be measured only by using the sun sensitive load carried by the orbiter without radio measurement, and long-distance measurement can be realized.

The present invention provides a method and system for static closed-loop verification of the separation strategy for Mars exploration, including the following steps: Step S1: Verify the separation process of the orbiter and the lander; Step S2: Perform static equivalent separation of the orbiter and the lander according to the verified separation process. The present invention can deduce and verify the separation process between spacecraft during the development stage of Mars probes. The two spacecraft are statically placed on their respective parking racks. By operating the separation device between the two spacecraft and using the closed-loop verification system to collect separation data, the correctness of the separation control strategy between spacecraft can be examined, the coordination of the joint operation of the two spacecraft during the separation process and the rationality of the flight program design can be evaluated, possible fault events and design defects can be investigated, and it is ensured that the Mars probe can successfully execute the separation process between spacecraft according to the designed program after being launched into orbit.

This invention relates to the field of thermal protection structures for low Earth orbit (LEO) vehicles, specifically to a high-density thermoelectric conversion device and its design method. The high-density thermoelectric conversion device includes: a high-density thermal layer disposed on the inner side of the vehicle surface; a heat exchanger surrounding the inner side of the high-density thermal layer; and a thermoelectric converter disposed between the high-density thermal layer and the heat exchanger. The heat-absorbing end of the thermoelectric converter is connected to the high-density thermal layer, and the heat-releasing end of the thermoelectric converter is connected to the wall of the heat exchanger. Embodiments of this invention, by setting a high-density thermal layer and a heat exchanger on the surface of the vehicle, create a large temperature difference between the heat-absorbing and heat-releasing ends of the thermoelectric converter on the vehicle surface. This reduces the surface temperature of the vehicle and provides power to the vehicle, thereby solving the thermal protection and power supply problems of LEO vehicles.

Device (1) for compressing a gaseous fluid, in particular a scroll compressor, comprising a housing (2) with a wall (12) and a compression mechanism with a stationary stator (3) and a movable orbiter (4), which is driven via a drive shaft (6, 6a, 6b, 6b-1, 6b-2, 6b-3), wherein - the wall (12) and the orbiter (4) are designed to at least partially enclose a counterpressure region (13) and the wall (12) is designed between the counterpressure region (13) and a suction region (14), delimiting the counterpressure region (13) from the suction region (14), - the drive shaft (6, 6a, 6b, 6b-1, 6b-2, 6b-3) is arranged so as to protrude through an opening formed within the wall (12), wherein a shaft sealing element (15, 15a, 15b, 15b-1, 15b-2, 15b-3, 15b-4, 15c) is arranged between the drive shaft (6, 6a, 6b, 6b-1, 6b-2, 6b-3) and the wall (12) for sealing the counterpressure region (13) from the intake region (14), wherein between the drive shaft (6, 6a, 6b, 6b-1, 6b-2, 6b-3) and the shaft sealing element (15, 15a, 15b, 15b-1, 15b-2, 15b-3, 15b-4, 15c) a sealing surface (23) is formed with a through opening (24a, 24b, 24c) interrupting the sealing surface (23) and hydraulically connecting the counterpressure region (13) and the suction region (14).

The invention relates to a transfer orbit design method for a HALO orbital vehicle to fly to a given ground point. The method comprises the following steps: finding a moment when a lunar declination value is equal to a given ground point latitude value according to the lunar declination value in a given task time range, and selecting a time point closest to an initial moment as an orbit transfer moment iteration initial value; calculating the flight duration of the orbital transfer point relative to the starting point by using the starting point time and the iteration initial value of the orbital transfer time; calculating normalization time according to the quotient of the flight duration and the HALO orbital period; according to a polynomial fitting coefficient, calculating the increment of the orbital transfer speed for the normalization time; based on a high-precision force model, establishing a mapping relation between the increment of the orbital transfer speed and the iteration initial value at the orbital transfer time through numerical integration; and solving the to-be-solved parameters by adopting a Newton iteration method based on the mapping relation. The method is short in response time, high in calculation result precision and fast in iteration convergence.

The invention discloses a Mars fine terrain reconstruction method, system and device based on an atmospheric correction shadow method and a medium, and belongs to the technical field of terrain reconstruction. The Mars fine terrain reconstruction method comprises the steps that Mars orbiter remote sensing images and a low-resolution digital elevation model under different illumination conditions are acquired to serve as input data; based on the input data, constructing a terrain-atmosphere coupling reflection model to correct interference caused by Mars atmospheric scattering and earth surface albedo in image brightness; and estimating atmospheric parameters in the terrain-atmosphere coupling reflection model through the brightness information of the input multiple images and the radiation transmission model so as to realize parameter adaptive estimation without external atmosphere data. According to the method, by constructing the terrain and atmosphere coupling reflection model, the coupling effect of atmospheric scattering and earth surface albedo in the Mars image can be accurately quantified and corrected, the problem of terrain calculation misalignment caused by neglecting the atmospheric influence in a traditional method is solved, and thus the accuracy of a terrain reconstruction result is remarkably improved.

The present invention provides an autonomous management method and system for a Mars orbiter without uplink failures in orbit, including: resetting or powering off and restarting the remote control terminal; resetting or powering off and restarting the transponder; switching the cross-backup channels of the transponder and the receiving antenna; performing main-backup switching processing on the crystal oscillator; autonomously switching the TT&C to the low-gain or medium-gain antenna mode, and at the same time, the spacecraft is in the +X sun-pointing mode and rotates 180° around the +X axis; in the +X sun-pointing mode, switching the receiving antenna back to the omnidirectional low-gain receiving mode; in the omnidirectional low-gain receiving mode, the spacecraft slowly rotates around the +X axis facing the sun; after the above-mentioned processing, if the long-term uplink-free autonomous management process still does not jump out on the spacecraft, power on the UHF relay communication machine to the single-carrier transmission mode; use the locked state of other Mars orbiters or Mars rovers to determine the life and death of the orbiter. The present invention can solve the problem that it is impossible to process through the ground when a failure occurs in the uplink channel.

This invention belongs to the field of lunar exploration and in-situ resource utilization, and specifically discloses a method and system for 3D modeling of lunar lava tubes. The method comprises: fusing multimodal data of lunar lava tubes to obtain a basic 3D model; the multimodal data includes point cloud data collected by a lidar, point cloud data collected by an orbiter penetrating radar, 3D images collected by a stereo camera, and thermal images collected by an infrared thermal imager; inputting the complete regions of the basic 3D model into generative models based on CGAN and VAE, respectively, to obtain two corresponding 3D completion models; and performing a weighted fusion of the two 3D completion models to obtain a complete 3D model of the lunar lava tubes. This method can improve the integrity and accuracy of 3D models of lunar lava tubes, providing technical support for in-depth exploration of lunar lava tubes.

The invention relates to the technical field of stage equipment, and discloses a stage orbital aircraft safety control system and method, and the system comprises an acquisition module which obtains target data; the simulation training module is used for forming contrast data and judgment data through a data training strategy; the judgment module is used for forming regulation and control mode judgment data through a direction relation analysis strategy; the control module is used for forming regulation and control data through a regulation and control strategy; the execution module is used for performing safety control on the aircraft; according to the relation between the current aircraft translation direction and the wind direction, the adopted regulation and control mode is judged to be regulation and control of the aircraft translation mechanism or regulation and control of the aircraft lifting mechanism, the stability of the aircraft and the stability of a target person are kept, the stage performance safety is improved, meanwhile, the current stage performance requirement is met as much as possible, and the practicability is high. And a stage performance effect is provided.

The invention relates to the technical field of rubber pad processing, in particular to a special cutter for processing a satellite orbiter gas cylinder rubber pad part, which comprises a satellite orbiter gas cylinder rubber pad, a plurality of groups of processing holes and a first milling cutter, a plurality of sets of machining holes are machined in the middle of the satellite orbiter gas cylinder rubber pad through the first milling cutter, a stepped layered cutting structure is formed through the first milling cutter and the second milling cutter, and the design of a cutting edge with the outer diameter being 10 + / -0.1 mm and the inner diameter being 9.6 + / -0.1 mm, a protruding block and a guide groove is matched. The problems that a satellite orbiter gas cylinder rubber pad is poor in machining surface smoothness, the size precision is difficult to control, part edges are prone to curling, and materials are layered or torn are effectively solved, machine tool vibration and rubber elastic deformation are restrained, and the use performance and assembly reliability of parts are guaranteed.

The application relates to a DRO orbit aircraft orbit design method for falling on the moon at a given angle and coordinates. The method comprises the following steps: DRO initial position and speed value integration, output of L2 coordinate system positions of the moon at different moments in a period, and mapping of the moon falling speed and transfer orbit and white plane intersection position. The mapping and DRO period position are combined to reversely coarsely calculate the moon falling moment, initial speed and transfer time length. The three are taken as iteration initial values and constraints are set to reversely accurately calculate the integral end moment, speed and orbit flight time length. The end moment is taken as the orbit transfer moment initial value, the end speed and the DRO speed difference value are taken as the orbit transfer speed increment initial value, the orbit flight time length is taken as the transfer time length initial value, the forward accurate calculation is carried out by setting the terminal constraint and the integral termination condition, and finally the transfer orbit transfer moment and speed increment projection are obtained. The method can be suitable for different task times and different DRO orbit sizes, the calculation result is high in precision, and the iteration is fast and reliable.

The application discloses a Mars remote sensing image enhancement method and device, which comprises the following steps: acquiring a plurality of clear Mars remote sensing images shot by a Mars orbiter; using a sand dust rendering method to generate a plurality of corresponding images with sand dust for each clear Mars remote sensing image; training a deep neural network by using each plurality of clear sand dust images; inputting a real image with sand dust into the trained neural network, reducing the spatial resolution of the real image with sand dust and expanding the channel dimension through a convolutional neural network layer and two two-fold downsampling layers to obtain a feature map; sequentially passing the feature map through three dense deep convolution modules to extract features; and passing the extracted features through twice upsampling operations and once convolution operation. The application can be applied to a Mars orbiter computer vision system and Mars image ground analysis, and provides convenience for ground data analysis and Mars observation work.

The invention relates to an orbital design method for a DRO orbital aircraft falling into the moon according to a given angle and coordinates. The method comprises the following steps: outputting earth-moon L2 coordinate system positions at different moments in one period according to DRO initial position speed numerical integration, and constructing mapping between the moon falling speed and the intersection point position of a transfer orbit and a white track plane; and performing reverse rough calculation by combining the mapping with the DRO periodic position to obtain the falling month moment, the initial speed and the transfer duration. Taking the three as iteration initial values and setting constraints, and performing reverse actuarial calculation to obtain integral end point time, speed and orbit flight duration. And by taking the end point moment as an orbit transfer moment initial value, the difference value between the end point speed and the DRO speed as an orbit transfer speed increment initial value and the orbit flight duration as a transfer duration initial value, forward actuarial calculation is carried out by setting a terminal constraint and an integral termination condition, and the transfer orbit transfer moment and speed increment projection is finally obtained. The method can adapt to different task time and different DRO orbit sizes, the calculation result is high in precision, and iteration convergence is rapid and reliable.