30 results about "Mars landing" patented technology

The invention provides an anti-interference composite online guiding method of the atmosphere entering period of a Mars lander. The method includes the following steps of a, establishing a Mars lander atmosphere entering period kinetic model, and introducing Mars atmosphere density uncertainty interference into the Mars lander atmosphere entering period kinetic model; b, establishing an interference observer, wherein the interference observer estimates the Mars atmosphere density uncertainty interference; c, designing a prediction-correction guiding rule, wherein the prediction-correction guiding rule updates to-be-determined parameters of the heel angle amplitude; d, compositing the interference observer in the step b with the prediction-correction guiding rule in the step c to obtain an anti-interference composite guiding rule, wherein the anti-interference composite guiding rule complements for the Mars atmosphere density uncertainty interference through the estimation value of the Mars atmosphere density uncertainty interference.

A method and system for anti-interference compound on-line guidance in the atmosphere-entering stage of a Mars lander is provided in the present invention. The method comprises steps of a) building a dynamics model for the atmosphere-entering phase of a Mars lander to incorporating the disturbance brought by the Mars atmosphere density uncertainty into the dynamics model for the atmosphere-entering stage of the Mars lander; b) constructing a disturbance observer to estimate the disturbance brought by the Mars atmosphere density uncertainty in the dynamics model for the atmosphere-entering phase of the Mars lander; c) building a prediction-correction guidance law, and guiding the heeling angle amplitude of the prediction-correction guidance law by continuously updating undetermined parameters; d) constructing an anti-disturbance composite guidance law to compensate the Mars atmosphere density uncertainty; and e) adjusting the deviation of the landing point of the Mars lander by compensating the Mars atmosphere uncertainty disturbance.

A method for the atmospheric entry section navigation of a mars lander based on a sampling point inheritance strategy relates to the technical field of navigation guidance and control, in particular to a method for the atmospheric entry section navigation of the mars lander. The invention aims at solving the problem that the precision is low when the status of the lander is estimated according to the conventional filtering method. The method comprises the following steps: in a filtering beginning cycle, providing a sampling point and a sampling point weight according to sampling point generating rules of unscented Kalman filtering; providing a lander status predicted value x' and a lander status variance Px according to a transmitted sampling point and a generated weight; calculating a measurement predicted value y', a variance Py and a covariance Pxy according to unscented Kalman filtering rules; updating measurement and correcting the estimate value of the lander status; estimating x'<+> and a lander status variance Px<+> according to the lander status, correcting the transmitted sampling point Xi and obtaining the sampling point Xi<+> with the corrected value. The method has higher status estimate precision and is suitable for navigation in the atmospheric entry section of the mars lander.

The invention provides a mini probe for Mars orbits of an attitude control subsystem. The mini probe comprises a cabin, a solar battery array, a thermal-control subsystem, a propelling subsystem, a power source subsystem, a communication subsystem and a comprehensive electronic subsystem. The mini probe can be directly applied to first Mars detection of China and can provide relay communication service for tasks such as Mars landing, touring, Mars sampling and returning, realize corresponding scientific Mars-surrounding detection by mounting certain miniature effective load and provide reference for design of miniature surrounding probes for other superior planets.

The invention provides a static locating method for a mars lander based on a surrounder, based on tracking and measuring a detector, the wireless distance and speed measurement between the surrounder and the lander is measured; the surrounder emits a wireless signal in a manner of double-trip measurement; and after being in transparent transmission of the lander, the wireless signal is sent to the earth to be treated by the mars measuring data. The landing locating precision obtained by multi-circle measurement of the lander can reach dozens km magnitude. The static locating method for a mars lander based on a surrounder provided in the invention can help the ground tracking and measuring and make up the defects of long underground flame distance and hard tracking and measuring difficulty of the lander; and the use ratio of the equipment on the mars can be improved without increasing extract cost by the equipment on the surrounder.

The invention discloses an aerodynamic layout structure of a Mars landing patrolling device, and relates to the technical field of aerodynamic layout of patrolling devices, the aerodynamic layout structure comprises a cabin body, the cabin body is of a revolving body structure, and the cabin body comprises a head part, a rear body, a transition body and a tail part which are sequentially arranged from bottom to top; the mass center of the cabin body is arranged on one side of the axis of the revolving body structure; a balancing wing which is telescopically embedded in the rear body, the balancing wing is arranged on the other side of the axis of the revolving body structure, and when the balancing wing extends out, the attack angle of the Mars landing patrolling device can be changed into 0 degree; according to the pneumatic layout structure of the Mars landing patrolling device, the mass center of the cabin body is offset, the patrolling device can enter at an attack angle without unfolding the balancing wings, and the balancing wings extend out before parachute opening, so that the attack angle of the patrolling device returns to 0 degree, parachute opening safety is guaranteed, the service time of the balancing wings is greatly shortened, and the requirements on an actuating mechanism and a heatproof system of the balancing wing are reduced.

The invention discloses a Mars dust storm space-time probability extraction and landing safety evaluation method, comprising the following steps: 1, identifying a dust storm object based on an RGB color remote sensing image; 2, performing Mars dust storm time probability analysis; 3, performing Mars dust storm spatial probability analysis; and 4, evaluating the safety of the Mars landing area. TheMars dust storm space-time probability extraction and landing safety evaluation method has the advantages of accurately recognizing the Mars dust storm object, and performing area extraction; solvingthe problem that the regularity of repeated dust storms of multiple Mars is ignored in the prior art; solving the problem that dust storm spatial distribution rule characteristics are ignored in theprior art; and selecting proper landing time and a proper safe area for a Mars task.

The invention discloses a heat insulation component for heat control of a Mars Landing machine. The heat insulation component comprises a thin layer structure, structure support pieces, first aerogel blocks, second aerogel blocks, honeycombs and skins, wherein the second aerogel blocks are filled inside the honeycombs and clamped between the upper skin and the lower skin so as to form a honeycomb plate, the first aerogel blocks and the structure support pieces are arranged on the first surface of the honeycomb plate, the thin layer structure is covered on the outer surface of the first aerogel blocks, and both the first aerogel blocks and the second aerogel blocks are composite SiO2 aerogel materials. The invention further discloses a manufacturing method for the heat insulation component. By utilizing the heat insulation component based on SiO2 aerogel, the problem that the aerogel material is difficult to use due to poor mechanical property is solved, and the heat insulation component is excellent in heat insulation performance and capable of effectively meeting requirement on heat insulation under the environment on the surface of the Mars, has obvious effect of light weight, can reduce the weight of a Mars lander, and is resource saving.

The invention relates to a distance and speed measurement sensor beam pointing determination method for Mars landing, which comprehensively considers the configuration layout of a detector, and considers the measurement precision requirements of long and short distance sections and heterogeneous backup of two types of distance and speed measurement sensors. The pointing directions of four fixed pointing beams of the microwave distance and speed measuring sensor and nine fixed beam pointing directions of the phased array sensor are determined. In actual use, the phased array sensor calculates and appoints four wave beams through a spaceborne computer in each measurement period to provide distance measurement and speed measurement information. Multi-beam information fusion processing is carried out on relative distance and speed information obtained in eight beam directions of the microwave distance and speed measuring sensor and the phased array sensor, so that the relative distance, speed and other information of the detector relative to the surface of the Mars can be solved.

The invention relates to a high-fault-tolerance initiating explosive device control method for a Mars landing process, aiming at a large dynamic and strong impact environment and a strict time sequence requirement when Mars atmosphere enters the landing process, on the basis of triggering in a bus communication mode, the dual-redundancy initiating explosive device control method is presented, wherein a spaceborne computer hard wire is adopted to directly control the initiating explosive device, and an interface control unit is notified to control the initiating explosive device through a bus message, so that the spaceborne computer and the interface control unit can timely send out an initiating explosive device control instruction according to sequential logic under normal conditions; therefore, the dual redundancy control logic of the initiating explosive device related instructions is realized, and the test verification of the initiating explosive device instructions is realized through the combined triggering use of the time sequence test ports of the spaceborne computer.

This disclosure proposes a method and system for determining a Mars landing site based on the yield rate of Martian dust devils, including: acquiring high-resolution Mars remote sensing time-series image pairs in the same area at different times; identifying the acquired Mars remote sensing time-series image pairs to obtain Mars landing newly generated dust devil trajectory in the area; based on the above trajectory, obtain the overlapping area of ​​the Mars landing area, the observation time interval, and the number of newly generated dust devil objects observed in the image pair to calculate the dust devil yield rate; calculate the landing area image Centering, the coverage area of ​​the dust devil object track newly generated in each Martian day, and then calculate the number of Martian days required to cover the entire landing area area on the basis of the newly generated dust devil track area in each Martian day, namely Mars rover solar panel cleaning cycle; determine the Mars landing site based on the dust devil production rate and the Mars rover solar panel cleaning cycle. Select the landing and inspection area with the optimal cleaning cycle for the subsequent Mars rover inspection mission after a successful Mars landing.

The invention provides a device for centering postures of a Mars landing capsule. The device comprises a base plate (1), a support (2), a connecting shaft (3), a torsional spring (4), a centering ring (5) and a compressing and releasing device (6). Under the conditions that the size, the weight and the energy supply of the device are limited strictly, an initiating explosive device acts to release the compacted torsional spring, and therefore the centering ring is driven to center the postures of the Mars landing capsule. The device for centering the postures of the Mars landing capsule has the advantages of being simple in structure, safe, reliable, small in size and light in weight. As the effective means for guaranteeing the correct postures of the landing capsule on the Mars, the device for centering the postures of the Mars landing capsule can be directly applied to development for a landing capsule in the first Chinese Mars landing detecting task, and can also provide reference for design of Chinese follow-up small-scale planet landing capsules.

The invention provides a Mars landing point determination method and system based on Mars dust wind yield. The method comprises the following steps: acquiring high-resolution Mars remote sensing time sequence image pairs at different times in the same region; identifying the obtained Mars remote sensing time sequence image pair to obtain a newly-generated dust entrainment wind track in the Mars landing area; calculating the dust entrainment wind yield according to the area of the overlapping range of the Mars landing area, the observation time interval and the number of newly-generated dust entrainment wind objects observed in the image pair obtained on the basis of the trajectory; calculating the coverage area of the newly generated dust entrainment object trajectory of each Mars day in the landing area image pair, and then calculating the number of Mars day days required for covering the whole landing area range on the basis of the area of the newly generated dust entrainment object trajectory of each Mars day, namely the cleaning period of the solar panel of the Mars rover; and determining the Mars landing point is determined based on the dust entrainment wind yield and the Mars over solar panel cleaning period. A landing and patrolling area with the optimal cleaning period is selected for a follow-up Mars rover patrolling task after successful Mars landing.

The invention discloses a method for simulating the surface morphology of Mars in an earth environment. The method comprises the following steps: constructing a basic terrain for simulating the surface morphology of Mars; based on the basic terrain, constructing a first simulated Mars surface morphology with different Mars pit distribution characteristics; based on the first simulated Mars surface morphology, constructing a second simulated Mars surface morphology with different Mars pit distribution characteristics and Mars rock distribution characteristics; based on the second simulated Mars surface morphology, constructing a third Mars surface morphology with different Mars pit distribution characteristics, Mars rock distribution characteristics and fire slope distribution characteristics; and based on the third Mars surface morphology, constructing a simulated Mars surface morphology which has different Mars pit distribution characteristics, Mars rock distribution characteristics and fire slope distribution characteristics and can reflect Mars surface microwave and aurora reflection characteristics. The invention aims to realize the real representation of the Mars surface topography and landform, and ensures that the Mars lander can identify the Mars surface morphology (Mars pit, Mars rock, Mars slope and the like) with high precision in the Mars surface landing process.

The invention discloses a simulated Mars soil preparation method which comprises the following steps: sequentially drying, removing impurities and crushing volcanic ash to obtain volcanic ash particle aggregates with different particle size ranges, and obtaining corresponding simulated Mars soil standard gradation according to a preset soil mechanical property target value range. The simulated Mars soil can be produced in batches according to the standard grading; the invention further discloses a construction method of the simulated Mars soil test bed, the simulated Mars soil laying area is arranged in the foundation pit, layered filling of the simulated Mars soil is carried out to obtain the test bed, and after filling of each layer is completed, the physical and mechanical parameters of the test bed are controlled according to the physical and mechanical property requirements; a severe landing environment is built on the surface of the test bed. The Mars surface fire soil landform and physical and mechanical properties can be truly reproduced, it is ensured that the Mars lander can recognize the severe environment of the Mars surface with high precision in the Mars surface soft landing process, and meanwhile accidents such as foot pad subsidence are avoided.

The invention discloses a Mars lander parachute avoidance method based on parachute drop point offline analysis, relates to the technical field of guidance control, and aims to solve the problem that a parachute back cover assembly covers a lander when the lander falls to the ground after the lander is separated from the parachute back cover assembly in the prior art. According to the application, the drop point dispersion of the parachute back cover assembly is determined off line through a Monte Carlo simulation method, and the shortest and longest space time of the parachute and different wind direction combinations are ensured to be included in simulation by using aerodynamic parameter pulling deviation in the simulation analysis process, so that the dispersion ellipse of the parachute covers various extreme working conditions; and the result is more conservative and reliable. In the landing process, the landing point of the lander is predicted based on polynomial guidance, maneuvering avoidance is only applied to the situation that the lander falls in a parachute landing point scattering area, and a maneuvering target enables the lander to move by the shortest distance to reach the position outside the parachute scattering area.

The invention discloses a method for estimating the interruption of Mars atmosphere entry capacity based on multiple models, relates to a method for estimating the interruption of Mars atmosphere entry capacity, and belongs to the technical field of deep space exploration. The implementation steps of the present invention are as follows: Step 1: Establish the dynamic equation of the Martian atmosphere entry section. Step 2: Establish the radio measurement model and the radio measurement interruption model of the Martian atmosphere entry section, which are used in step 3 to correct the position and velocity information of the probe. Step 3 combines the radio model, interruption model and filtering algorithm of step 2 to solve the state of the detector, and estimate the state of the detector. The invention can effectively suppress the divergence problem of the state estimation of the detector when the radio measurement is interrupted, and improve the reliability and state estimation accuracy of the autonomous navigation system of the detector in the Mars atmosphere entry section. The invention can provide technical support and reference for the design of the radio navigation scheme adopted in the atmospheric entry section in the future Mars landing detection.

The invention discloses a system anti-impact method suitable for a Mars landing task. The system anti-impact method comprises relay anti-impact, balancing wing state autonomous diagnosis and processing and grounding switch state autonomous diagnosis and processing. The electromagnetic coil in the relay is continuously powered up from the time before an initiating explosive device is detonated to the time after the initiating explosive device is detonated, so that the relay works normally; the balancing wing state autonomous diagnosis and processing comprises the following steps: before the initiating explosive device is detonated, if a signal that a certain path of balancing wing is unfolded in place is effective, considering that a switch of the path is in fault, and not using the signal of the path subsequently; after the initiating explosive device is detonated, if one or more paths of balancing wing unfolding in-place signals are collected to be effective, considering that the balancing wing is unfolded in place, and otherwise, considering that the balancing wing is not unfolded; the grounding switch state autonomous diagnosis and processing comprises the following steps: before grounding, if a certain grounding switch is in a grounding state, considering that the grounding switch is in a fault state, and no longer using the signal in the subsequent process; and after grounding, judging whether grounding exists or not according to the states of the remaining grounding switches.

The invention discloses an aerodynamic layout structure of a Mars landing patrol device, relates to the technical field of aerodynamic layout of the patrol device, and comprises: a cabin body, which is a revolving body structure, and the cabin body includes a head and a rear body arranged sequentially from bottom to top , transition body and tail; the center of mass of the cabin is set on one side of the axis of the rotary structure; the trim wing is telescopically embedded in the rear body, and the trim wing is set on the other side of the axis of the rotary structure. When it is out, the angle of attack of the Mars landing rover can be changed to 0°; the aerodynamic layout structure of the Mars landing rover offsets the center of mass of the cabin, and the rover can enter at an angle of attack without deploying the trim wings. Extend the trim wing in front of the parachute, so that the angle of attack of the patrol device returns to 0°, ensuring the safety of the parachute, which greatly shortens the use time of the trim wing and reduces the requirements for the action mechanism and heat protection system of the trim wing.

According to the high-reliability composite power supply system applied to the Mars landing patrol task provided by the invention, a high-specific-energy primary battery with the characteristics of low self-discharge rate, no maintenance, wide temperature adaptation range and the like is selected for an entering cabin power supply subsystem only needing short-time power supply, so that a charging path entering a cabin is saved, and the weight of the whole power supply system is reduced; for a Mars rover power supply subsystem needing long-time power supply, a rechargeable battery is selected, and the purposes of adapting to long-time storage and wide-temperature-range detection tasks and realizing power supply path multiplexing backup at the most critical EDL section are achieved by adopting different types of energy storage batteries and corresponding topological design. In other words, the backup power supply path can independently complete the EDL task and can match the main power supply path to complete the EDL task, and the system reliability is improved; therefore, the invention is particularly suitable for detectors which need to be subjected to long-term flight, can be stored for a long time, have relatively independent cabin tasks and have multiplexing requirements, such as Mars detection tasks and interplanetary detection tasks.

The invention discloses a Mars EDL process large dynamic navigation test verification system and method, and belongs to the technical field of navigation. According to the verification system, a mechanical rotary table is adopted to carry the IMU sensor to simulate the real attitude motion of the Mars lander in the EDL process through an attitude mapping method, and the key technology and related performance of inertial navigation can be effectively verified and evaluated by using the actual output data of the mechanical rotary table and the data acquired by the IMU; the system has the characteristics of simple test method, easy implementation, high reliability, strong pertinence and the like.

The invention discloses a high-gain movable spot beam antenna suitable for Mars landing and patrol detection. According to the invention, a feed source is supported by a waveguide D-waveguide E and other two supporting members in a three-point manner and is suspended in the center of a reflecting surface, so that the influence of the supporting structure on the antenna performance is effectively reduced; and meanwhile, the waveguide calibers of the waveguide D, the waveguide E and the feed source are changed, so that the miniaturization of the feed source is effectively realized, and the shielding influence of the feed source is reduced; meanwhile, a non-compact antenna assembly pressing structure is adopted, and a pressing mounting surface with a height difference and a mounting surface provided by an unfolding supporting structure are adopted, so that the working requirement after the antenna is unfolded is met, and meanwhile, the maximum additional envelope after the antenna is installed can reach the maximum height size of an unfolding driving shaft assembly part to the minimum, so the additional envelope is greatly reduced, and the application envelope requirement of the Mars rover is met.

The invention provides a space transportation system based on manned Mars detection and a detection method. The space transportation system comprises a nuclear propulsion stage, a storage tank stage and a payload stage, the nuclear propulsion stage is located at the tail end of the transportation system, comprises a main engine containing a nuclear reactor and a first-stage liquid hydrogen storage tank, and is used for propelling the whole transportation system to enter a fire transfer track, track control and near-fire and near-ground braking; the storage tank stage is located at the front end of the nuclear propulsion stage, comprises a plurality of secondary liquid hydrogen storage tanks and is used for storing propellants and providing the propellants for a main engine of the nuclear propulsion stage when the transportation system enters a ground fire transfer track; the payload level comprises a manned payload and a cargo-carrying payload, the manned payload comprises a manned probe aircraft and a transfer dwelling cabin, and the cargo-carrying payload comprises a Mars landing and rising probe. According to the space transportation system based on manned Mars detection and the detection method provided by the invention, reference in great significance is provided for manned Mars detection.

A method for the atmospheric entry section navigation of a mars lander based on a sampling point inheritance strategy relates to the technical field of navigation guidance and control, in particular to a method for the atmospheric entry section navigation of the mars lander. The invention aims at solving the problem that the precision is low when the status of the lander is estimated according to the conventional filtering method. The method comprises the following steps: in a filtering beginning cycle, providing a sampling point and a sampling point weight according to sampling point generating rules of unscented Kalman filtering; providing a lander status predicted value x' and a lander status variance Px according to a transmitted sampling point and a generated weight; calculating a measurement predicted value y', a variance Py and a covariance Pxy according to unscented Kalman filtering rules; updating measurement and correcting the estimate value of the lander status; estimating x'<+> and a lander status variance Px<+> according to the lander status, correcting the transmitted sampling point Xi and obtaining the sampling point Xi<+> with the corrected value. The method has higher status estimate precision and is suitable for navigation in the atmospheric entry section of the mars lander.

A multi-strategy obstacle avoidance logic and control method for Mars landing adopts a cooperative control strategy of coarse obstacle avoidance and parachute-back cover combination avoidance for a new requirement of unified avoidance of a parachute-back cover combination and landform obstacles in the aspect of control strategy. In a hovering fine obstacle avoidance stage, on the basis of laser three-dimensional imaging, a redundancy strategy of binocular fine obstacle avoidance by using an optical obstacle avoidance sensor and a multifunctional obstacle avoidance sensor optical obstacle avoidance module is added. According to the strong time sequence requirements of mars in the entering, descending and landing processes, working time sequences of data transmission, information interaction and the like among an entering and descending control unit, a multifunctional obstacle avoidance sensor and a optical obstacle avoidance sensor are provided, and it is guaranteed that the obstacle avoidance task is completed autonomously and rapidly in a short time.

The invention discloses an extraterrestrial celestial body landing multi-dimensional all-condition IMU performance verification and evaluation method. The method comprises the following steps: 1, obtaining a test matrix of an IMU product according to task requirements of a Tianzhi No.1 Mars landing patroller; 2, testing the IMU product according to the test matrix of the IMU product; and 3, carrying out mechanical environment adaptability evaluation, thermal environment adaptability evaluation, navigation working condition adaptability evaluation and inertial parameter measurement capability evaluation on the tested IMU product. According to the invention, the high requirement on reliability of the IMU in a landing task and the high requirement on inertial navigation precision of the IMU under multi-dimensional and full working conditions are met.

The invention relates to an attitude correction method using a speed measurement sensor in a Mars landing process. The method comprises the following steps of: adopting a parallel attitude determination frame, using traditional inertial navigation extrapolation at a foreground to obtain an inertial position, a speed and an attitude, and obtaining a projection of a gravitational force direction under a body coordinate system by using the inertial position and the attitude; carrying out benchmark reconstruction in the background by using a speed measurement sensor and an inertial measurement unit to obtain the projection of another gravity direction in the system; and monitoring and comparing the gravitational force direction of the background and the gravitational force direction of the foreground in real time, once the gravitational force directions exceed a certain threshold value, rotating the gravitational force direction of the foreground according to the Euler rotation principle, so that the gravitational force direction of the foreground can coincide with the gravitational force direction of the background so as to achieve the purpose of posture correction. According to the method, the problem of filtering divergence caused by incomplete observability of speed measurement attitude correction is avoided, the inertial navigation attitude precision in a large dynamic flight environment can be improved, and the landing index is improved.

The invention relates to a Mars lander anti-interference controller under the harmonic interference at an unknown frequency. Firstly, an attitude kinematics and dynamics model for a Mars lander is built up. Secondly, the harmonic influence caused by structural uncertainty during the severe friction with the atmosphere is determined for a lander engine, and the disturbing influence of the unknown-frequency harmonic on an interference observer is estimated. Furthermore, the compensation is conducted by a feed-forward path. Thirdly, a gesture instruction is tracked by a sliding-mode controller. Finally, the interference observer for the unknown-frequency harmonic and the sliding-mode controller are combined to construct an anti-interference controller. The anti-interference controller is strong in anti-interference property. Compared with a conventional Mars lander controller, the above Mars lander anti-interference controller is better in adaptability and anti-interference capability.

The present invention relates to a Mars landing track optimization control method based on convex optimization. The method comprises the following steps of: (1) establishing a Mars landing dynamical model and performing convex optimization; (2) taking fuel optimization as a target to perform single off-line optimization to obtain a nominal track; and (3) employing a moving horizon model to performprediction and control to perform track of the nominal track to complete landing control. Compared to the prior art, the convex optimization and the moving horizon model prediction and control are effectively combined to effectively reduce the accumulated errors caused by modeling errors and interferences and improve the landing precision.