172 results about "Exploration of Mars" patented technology

The invention relates to a mars soft landing guidance method, in particular to a mars probe landing guidance method based on fuel minimization, and belongs to the technical field of deep space exploration. The mars probe landing guidance method based on fuel minimization solves the problems that fuel minimization guidance can not be achieved through an existing explicit guidance law, and the optimal guidance law must store an entire track, so that a large storage space is occupied. The technical scheme includes the steps that a fuel optimal solution of a probe is acquired; path points are set and a path point library is established; a linear feedback guidance law is designed, if initial state information of the probe is matched with path point information of the path point library, landing is executed, and otherwise, landing is executed after the path points are fitted. Through the guidance strategy of the path points and the linear guidance law, the method can achieve the fuel optimal guidance of a power descent section of the mars probe based on a small storage space. The mars probe landing guidance method based on fuel minimization is suitable for the guidance law of the power descent section of the mars probe.

The invention relates to an X-ray pulsar navigation ground test system, which comprises a signal control sub-system, an orbital data generation sub-system, a navigation database sub-system, a pulse contour extraction sub-system, a navigation parameter estimation sub-system, a large-scale dynamic effect sub-system, an X-ray pulsar signal simulation sub-system, a detector sub-system and a navigation demonstration sub-system, wherein the orbital data generation sub-system comprises a Walker constellation satellite orbit module, a lunar exploration module and a Mars exploration orbit module; the large-scale dynamic effect sub-system comprises a large-scale space effect simulation module and an in-orbit dynamic effect simulation module; and the X-ray pulsar signal simulation sub-system comprises an X-ray generator and a vacuum channel. By realistically simulating the radiation characteristics of pulsar X-ray photon signals in a large-scale dynamic environment, the performance of a pulsar navigation probe is subjected to a closed loop test, the pulsar navigation algorithm is examined, and the feasibility of the scheme design of the pulsar navigation system is verified.

A Mars exploration landing buffer device comprises three or four sets of landing legs symmetrically distributed on the periphery of a Mars exploration lander body mechanism. Each set of landing legs is composed of a foot pad (1), a main leg (2), an auxiliary leg (3) and an unfolding locking mechanism (4), wherein the foot pad (1), the main leg (2), the auxiliary leg (3) and the unfolding locking mechanism (4) of the same set of landing legs are connected with one another through universal joints or rotating shafts, and the main leg (2), the auxiliary leg (3), the unfolding locking mechanism (4) of each set of landing legs and the lander body mechanism form a four-bar mechanism so that the landing buffer device can be overturned outwards up and down. The upturning folding structure enables the mass center of the lander to move forwards, the movement stability of the lander in an entering stage is improved, in addition, energy is absorbed through own movement deformation of buffer cellular structures in the main legs (2) and the auxiliary legs (3), soft landing of the lander is achieved, and the device can be applied to design of soft landing buffer mechanisms for landing detection on a planet or a satellite with atmosphere.

The invention relates to a deep space detector astronomy / radio combination navigation method based on ephemeris correction. The method comprises the following steps: establishing a Mars detector state model and astronomy astronomy navigation and radio navigation subsystem measurement models, acquiring the measured values of astronomy and radio navigation subsystems, and carrying out filtering estimation to obtain the position of the detector in an inertia coordinate system treating a target heavenly body as a center and the speed of the detector; and establishing the ephemeris error state model and the measurement model of the target heavenly body based on the position and the speed, obtaining the ephemeris error measurement value of the target heavenly body according to the estimation states of the two astronomy and radio navigation subsystems, utilizing a Kalman filtering method estimate the ephemeris error of the target heavenly body, feeding back to a navigation system model, and carrying out information fusion. The deep space detector astronomy / radio combination navigation method belongs to the technical field of spaceflight and navigation, can realize the online estimation of the heavenly body ephemeris error and the correction of the navigation system model error, and is suitable for the capture segment of a detector.

The invention discloses a mars ultimate approach segment autonomous navigation method based on multi-source information fusion, relates to a mars ultimate approach segment autonomous navigation method and belongs to the technical field of deep space exploration. The method is characterized in that optical measurement, radio measurement based on a mars orbiter and X-ray pulsar measurement information are combined based on a mars ultimate approach segment dynamical model, are effectively fused in the application range based on a federated filtering structure and are standby for one another, and the advantages of various navigation systems are fully exerted; on the basis of establishing a mars ultimate approach segment state model and a mars ultimate approach segment autonomous navigation model, detector real-time navigation status information is resolved by the utilization of an integrated navigation filtering algorithm of a federation structure to realize the mars ultimate approach segment autonomous navigation. According to the mars ultimate approach segment autonomous navigation method, the precision and the real time of the mars ultimate approach segment autonomous navigation can be improved, and the reliability and the fault tolerance of a navigation system can be improved. The method can provide a technical support for the design of a future mars exploration task autonomous navigation solution.

The invention relates to a dual-camera-based motion estimation method of a light stream detection device for an aircraft. First, the light stream detection device for the aircraft, which consists of a forward-looking camera and a down-looking camera, is used for collecting forward-looking images and down-looking images in real time, and the light stream values, which are represented by a first formula shown in the description and a second formula shown in the description, of corresponding multiple characteristic points between the current frames and the last frames of the forward-looking images and the down-looking images are calculated respectively; secondly, the angular speed of the aircraft is calculated by utilizing the distribution characteristics of the translational light streams of the forward-looking images and the decoupling thoughts of the light streams, and an attitude angle is updated; thirdly, the translational light streams of the down-looking images are obtained by utilizing the obtained angular speed and the second formula shown in the description. By utilizing the obtained attitude angle, each characteristic point is deeply represented to be the function of the terrain clearance of the aircraft; a novel translational motion estimation equation is established; the speed v of the aircraft is solved by using a UKF (unscented Kalman filter). Finally, the six-degree-of-freedom motion estimation of the aircraft is gained. The dual-camera-based motion estimation method of the light stream detection device for the aircraft is applicable to the navigation of indoor, low-altitude or deep-space aircraft such as a mars exploration aircraft. The forward-looking camera to which the dual-camera-based motion estimation method of the light stream detection device for the aircraft relates can be also used for completing the calculation of an attack angle and a sideslip angle and other navigation tasks of evading an obstruction, and the like.

The invention discloses a Mars capture process autonomous management method which comprises a Mars detector rover Mars capture former management series, a Mars detector Mars capture process management series and a Mars detector Mars capture later management series which are used for managing an attitude control subsystem, a power supply subsystem, an integrated electronic subsystem and a measurement and control data transmission subsystem. Through the rational planning of detector management events before capture, in the capture and after the capture, through the multidisciplinary design optimization of energy, communication, attitude, lighting and the like, a completed capture process autonomous management series is established, and the autonomous management of the detector capture process is realized.

The invention relates to a deep space explorer acquisition phase celestial navigation method based on target object ephemeris correction. According to the method, firstly, a Martian explorer state model, a startlight angle navigation sub system measuring model and an X-ray pulsar navigation sub system measuring model are built; then, the startlight angle and the X-ray pulsar quantity measurements are respectively obtained; filtering estimation is performed to obtain the position and the speed of a detector in a heliocentric inertia coordinate system and a target object center inertia coordinate system; on the basis, a state model and a measuring model of target object ephemeris error are built; the quantity measurement about the target object ephemeris error is obtained through estimation state vectors of the startlight angle navigation sub system and the X-ray pulsar navigation sub system; a Kalman filtering method is used for estimating the target object ephemeris error; the target object ephemeris error is fed back into a navigation system model; the position of a target object in the navigation model is corrected. The deep space explorer acquisition phase celestial navigation method belongs to the technical field of aerospace navigation, can be used for estimating the object ephemeris error and correcting the model error of the navigation system on line, and is applicable to the explorer acquisition phase.

An inflatable docking station for a Mars roving vehicle (Mars rover) provides a recharging station as well as a storage facility for the Mars rover during adverse environmental conditions. The docking station / garage allows for extended operational life for the Mars Rover, allows for the Mars rover to operate at the polar regions of Mars, where extreme cold conditions restrict rover operation, and provides for a warmer habitat for the Mars rover. Conventional planetary roving vehicles only be deployed from the warmer equator region of the planet surface. However, these roving vehicles may never explore the polar regions because the roving vehicle may not have enough stored power. The docking station / garage provides a remote outpost for, among other things, storing and recharging the Mars Rover to allow it to reach more remote areas.

An autonomous astronomical navigation method of a spark detector based on an improved MMUPF filtering method relates to an autonomous navigation method of a spark detector. The method comprises the following steps: accurately modeling a deep space detector positioned on a transfer orbit; then measuring with starlight angular distances as quantities; and finally, carrying out optimal estimation ofnavigation parameters by adopting the improved MMUPF filtering method, thereby solving the problem that the planet gravitation size and an orbit dynamics model are changed continuously when the distances are different, and obviously improving the navigation precision. The autonomous astronomical navigation method can be used for determining the navigation parameters of the spark detector or a planet detector positioned on the transfer orbit.

The invention discloses a Mars self-orientating method of a large elliptical orbit Mars probe, to enable a Mars probe to accomplish three-axis stable control to Mars orientation under the condition that the Mars probe does not completely depend on ground monitoring and control and self-orientation sensors are not available. The method comprises the following steps of: carrying out corresponding short-term orbit recursion calculation according to the number of initial orbits and a reference coordinate system for describing the number of the orbits; adopting a simplified analysis method for Mars self-recursion, and adopting a high precision numerical method for orbit recursion on the ground; establishing a Mars orientation reference coordinate system by taking a Mars center-probe position vector as the Z-axis direction; and turning to a posture working mode for Mars-orientation after the Mars receives remote control instructions. Compared with the prior art, the Mars self-orientating method has the beneficial effects that under the condition that the ground communication has time delay and the self-orientation sensors cannot achieve engineering application, the Mars self-orientation control of the Mars probe is reliably realized through the orbit recursion and the establishment of the orientation reference coordinate system.

The invention discloses a gyroscope-free independent space attitude maneuver control method of a Martian probe. The gyroscope-free independent space attitude maneuver control method is applied to attitude maneuver control of the probe among different oriented targets, and can implement attitude maneuver in any direction by virtue of data of a star sensor without gyroscope data. The method comprises the following steps of: determining four elements of a target attitude before implementing the attitude maneuver of the probe, and performing closed-coop control on a star body by a flywheel to implement the attitude maneuver around an Euler axis within a shortest route. Compared with the prior art, the method has the advantages that the attitude maneuver in any orientation can be completed within the shortest route with the configuration of minimum hardware, namely the star sensor and the flywheel, and the method is practicable.

The invention provides a mars acquisition phase optical autonomous navigation semi-physical simulation method and a system. the method comprises steps: a track generator module acquires a theoretical position and a theoretical speed of a mars detector in a mars J2000 coordinate system, a simulation image of the mars is acquired according to the theoretical position, and the image is projected on a screen; a camera circuit of an optical system navigation module acquires projection on the screen via an optical lens so as to acquire an analog image of the mars; an image pre-processing unit uses an Sobel operator to carry out edge detection on the analog image of the mars, and an image centroid extraction unit acquires the circle center of the analog image of the mars according to the edge detection result and by using a least square method; and an autonomous navigation module acquires the initial position value and the initial speed value of the mars detector, and according to the initial position value, the initial speed value and the circle center of the analog image of the mars, a UKF filter method is used for acquiring the filter solution position and the filter solution speed of the mars detector. Thus, the navigation precision can be improved.

The invention belongs to an on-board high-precision extrapolation algorithm of a Mars probe, and particularly relates to the extrapolation for an on-board high-precision rail of the Mars probe, which is used for forecasting the rail of the probe and the position and the speed of the probe relative to the Mars in real time. The method has the advantages that: no Mars probe is emitted in China, and no relative report about the on-board algorithm is provided, so the algorithm aiming at the on-board rail extrapolation of the Mars probe meets the requirement of engineering practice and keeps higher precision.

The invention discloses an independent earth orienting control method of a Martian probe. The method can be used for completing full-attitude earth capturing and earth orienting when the Martian probe is at any attitude, and only takes a star sensor for measuring and a counteractive flywheel as an executing component for providing control moment. According to the independent earth orienting control method, the measuring value of the star sensor can be directly utilized, the process of determining the attitude by a filtering algorithm is eliminated, and thus a gyroscope-free safe mode can be realized. Compared with the prior art, the independent earth orienting control method is independent from an earth sensor and a gyroscope, can complete the full-attitude earth capturing and earth orienting independently, and has high practicability and popularization values.

The invention provides a Mars probe vehicle with an obstacle crossing function. The Mars probe vehicle comprises two base plates, first servo motors, walking units, a drilling fixing unit, a camera, a first cylinder, first hinge seats and fifth cylinders. The Mars probe vehicle is characterized in that the base plates are rectangular flat plates, the first servo motors are mounted above the base plates correspondingly, the first hinge seats are mounted on motor shafts of the first servo motors, the two first hinge seats are connected through the first cylinder, and the fifth cylinders are mounted on the inner sides of the first hinge seats correspondingly, wherein a piston rod of one of the fifth cylinders is hinged to the lower side of a first cylinder body; according to the Mars probe vehicle with the obstacle crossing function, the left and right wheel tread can be controlled through two separated base plates, and the walking stability of the probe vehicle on the Mars is increased; and each walking unit can walk and steer independently, and has a shock absorption function, good operability and flexibility.

The invention discloses a nonlinear non-Gaussian ranking filtering method for a Martian atmosphere entering section. The method comprises four steps of 1, constructing an engineering actual equation based on a Martian atmosphere entering section detector; 2, supplying initial values: x0, P0 and statistical properties of system noise wk and vk, namely a mean value and a variance; 3, performing nonlinear non-Gaussian ranking method; 4, setting k=k+1, returning to the step 3 to continue the circulation until k is equal to the moment T corresponding to the end of the time of the Martian atmosphere entering section, namely the time when a supersonic parachute is opened, thus completing the nonlinear non-Gaussian ranking filtering method for the Martian atmosphere entering section. By the adoption of the nonlinear non-Gaussian ranking filtering method, the detector can reduce the influence, caused by non-Gaussian noise, on the estimation of a system state in the Martian atmosphere entering section, so that the state estimation precision can be improved, and the requirement of a Mars detection task in the future on the landing precision can be well met.

The invention provides an integrated antenna shell for a Mars exploration landing rover and a preparation method of the integrated antenna shell. The integrated antenna shell is formed through the steps of shell molding mold manufacturing, carbon fiber / polyimide prepreg preparation, shell laying-up, imidization, curing, demolding, machining, metal plating on the inner surface of the shell and thelike. According to the integrated antenna shell for the Mars exploration landing rover, by the adoption of an auxiliary molding manner of a female and male combined mold and a high-temperature-resistant silicone rubber soft mold, one-time integrated molding of an autoclave molding process is adopted, and the prepared shell has the multi-function integrated characteristics of the light mass, resistance of the high temperature of 370 DEG C, space environment resistance and structure bearing; and meanwhile a coupling technology of a nanometer silver particle modified shell base body and an innersurface vacuum sputtering metallization layer is adopted, an electromagnetic shielding function of the antenna shell is achieved, the containing structure space occupied by a traditional antenna connecting cable is omitted, the interface connection is further reduced, and the structure manner of the shell is simplified while the space utilization rate is increased.

The invention provides a rover tethered airship system for Mars exploration and a working method of the rover tethered airship system, and belongs to the technical field of Mars flight vehicle exploration. The rover tethered airship system comprises a tethered airship, a temperature control system, a power system, a high-resolution camera, a communication platform, sensing elements, a tethering collection-release control system and a rover. The temperature control system and the power system guarantee normal operation of electronic equipment, the sensing elements monitor working environment, flight conditions and other information, the high-resolution camera shoots the complex terrain of the Mars, and messages are sent to the rover through the communication platform to facilitate route planning and avoid dangerous zones. The tethered system can expand the exploration range, can carry out real-time data transmission, greatly increases the exploration range of the rover, assists the rover in completing exploration tasks and promotes the efficiency of Mars exploration. In addition, the tethered system can serve as a communication relay for communication between multiple rovers and spacecrafts.

The invention provides a calculation method for optical navigation data of a mars approaching section based on installation matrix correction. The method comprises the following steps of step I, enabling a mars probe to adjust the posture of the probe in a process of approaching mars so that an optical navigation sensor points to mars; step II, enabling the optical navigation sensor to image mars,phobos and deimos; and step III, processing a navigation image by use of a center point extraction technology, determining a navigation target source center and acquiring a sight line direction of anavigation target source relative to the mars probe. With the calculation method for the optical navigation data of the mars approaching section based on installation matrix correction, the autonomousnavigation task requirement of Mars exploration in future can be satisfied and good technical means is provided for high-precision autonomous navigation of a deep space probe.

The invention discloses a mars probe onboard autonomous orbit calculation method, which is a calculation method based on radial basis function neural network curve fitting. The method includes the following steps: building a three-layer radial basis function network model; getting orbit prediction data of a mars probe within a period of time according to ground orbit determination, and training the radial basis function network model by taking the data as a training sample; and finally, uploading the trained radial basis function neural network model as an onboard orbit prediction model to the mars probe. There is neither need for establishment of a complex dynamic model on a satellite nor need for ephemeris calculation. The prediction precision is almost equal to the ground orbit prediction precision. The method is also suitable for engineering implementation. The method not only can satisfy the engineering precision constraints of the recursion result, but also can satisfy the constraints of limited resources of onboard computers.

The embodiment of the invention provides a Mars exploration method and a Mars probe. The Mars exploration method comprises the steps that the Mars probe is launched from the earth, lifts off and enters an earth Mars transfer orbit, orbit trimming is conducted, and the Mars probe enters a Mars atmospheric layer; pneumatic braking is conducted through the Mars atmospheric layer, and the Mars probe enters and parks on a large elliptical orbit around the Mars; when the Mars probe parks on the large elliptical orbit around the Mars, dust storm on the Mars is monitored; and when the situation that the dust storm on the Mars is formed is determined, the Mars probe enters the Mars atmospheric layer to sample the dust storm on the Mars. According to the scheme, the cost can be reduced during Mars exploration.

The invention provides a ground navigation and probe autonomous navigation switching method during a Mars capturing and approaching phase, which comprises the following steps: step 1: turning on an optical navigation sensor when a cruise section is 10 million kilometers to 100,000 kilometers from Mars in the process of a Mars probe approaching Mars; step 2: analyzing radio navigation and optical navigation accuracy based on radio navigation and optical navigation measurement data; step 3: determining a navigation switching strategy such as ground radio navigation absolute priority policy / probeoptical autonomous navigation absolute priority strategy / combined navigation information fusion strategy; and step 4: turning off the optical navigation sensor and switching to a ground radio navigation when the distance to Mars is less than 100,000 kilometers. According to the ground navigation and probe autonomous navigation switching method during the Mars capturing and approaching phase, thedemand for future Mars exploration and navigation missions can be met, and a good technical method for high-precision navigation of deep space probes is provided.

The invention discloses a manned Mars detection system and method based on an earth-Mars circulating orbit. A manned Mars detection aircraft large system comprises an earth-Mars circulating track orbit space station, a manned aircraft, a freight aircraft and other space aircrafts, and realizes manned detection on the Mars and earth-Mars space with the corporation of relevant systems, such as a cosmonaut system, a carrier rocket, a measurement and control system and the like. The system is applied to long-term, persistent and multi-batch manned Mars detection tasks, can realize manned detection tasks on the earth, Mars space and the Mars, and meanwhile provides an ideal research platform for the scientific research on earth and Mars space.

The invention discloses a multifunctional integrated type mechanical arm for a planet exploration vehicle. The mechanical arm is provided with four rotational joints; and a three-dimensional camera with a two-dimensional rotary table is arranged at an elbow joint of the mechanical arm. By a triangle stability principle, the mechanical arm and a locking device form a triangular stable structure, and the mechanical arm and a mast for lifting the three-dimensional camera are combined into a whole, so that the mast for lifting the three-dimensional camera is eliminated, a mechanism is simplified,the weight is reduced, electric energy consumption required by the mast is eliminated, and the energy efficiency of a Mars exploration vehicle is greatly improved.

The invention provides a Mars probe secondary brake capture control method based on Newton iteration. During capturing and braking, after a Mars detector breaks down, an on-board computer is reset, then, the Mars probe is subjected to secondary braking by adopting a Newton iteration-based capturing and braking optimization method, so that the detector can smoothly enter a target orbit in the Marscapturing and braking process, and a secondary capturing and braking strategy can be autonomously designed for faults in the capturing and braking period. Besides, a safety zone for capturing and braking is also arranged and comprises the shortest ignition duration and the longest ignition duration of braking and capturing; when the on-board time is within the interval of the shortest ignition duration and the longest ignition duration, the speed increment reaches the nominal speed increment or the on-board time reaches the maximum ignition duration, the engine shutdown operation is executed,so that the Mars detector can form an annular fire track during braking and capturing and cannot impact Mars due to overlong ignition time, and the safety of Mars braking and capturing is guaranteed.

The invention discloses a comprehensive environment heat balance test system for a Mars detector single-machine product. The system comprises a Mars environment simulation chamber, a solar irradiationsimulation system, a Mars atmosphere simulation system, a Mars surface wind simulation system, a temperature control system and a test piece support system, wherein the Mars environment simulation chamber is a horizontal and capsule-shaped cabin body, container gates which can be opened are arranged at the two ends of the cabin body, and the Mars environment simulation chamber is mainly used forforming a closed space without obvious leakage and maintaining a special atmospheric environment in the cabin. The system can provide comprehensive stress environment tests of five parameters including gas atmosphere, wind speed, pressure, solar irradiation and environment temperature on the surface of Mars, and is easy to operate and good in generalizability.

The invention discloses a Mars navigation sensor installation error in-orbit calibration method, which comprises that 1, before autonomous optical navigation is performed, moon or asteroid celestial body is screened as a calibration celestial body according to a Mars exploration orbit design result, and attitude maneuver is performed at an appropriate moment, such that a Mars navigation sensor isaligned with the calibration celestial body; 2, a sequence star map containing the target celestial body is subjected to in-orbit shooting by using the Mars navigation sensor, the mass center of the calibration celestial body is extracted by using a star map recognition algorithm and a mass center algorithm, and the mass center coordinate of the star map is subjected to multi-frame average in-orbit processing so as to reduce random error; and 3, calculation is performed according to a reference ephemeris, ground station navigation data and the current attitude of the Mars detector to obtain the expected mass center parameter and the like of the calibration celestial body. According to the present invention, with the Mars navigation sensor installation error in-orbit calibration method, theparameter error of the navigation sensor can be determined, and the reliability, the stability and the precision of the optical autonomous navigation can be improved.

The invention relates to a simulation method of a Mars object for deep space exploration, which belongs to the field of deep space exploration application. In order to solve the problems in the prior art of lacking of a method for simulating the location, size, contour, imaging orientation and imaging gray level of Mars and existing technical gaps in engineering practices, the simulation method of the Mars object for deep space exploration is proposed. The simulation method includes transforming a center coordinate of Mars from a heliocentric ecliptic coordinate system to a display plane coordinate system and a projector coordinate system, calculating an imaging size of Mars according to an imaging field of view of a Mars explorer and the relative distance between Mars and the Mars explorer, simulating illuminated areas of Mars which are illuminated by the sun and unilluminated areas which are not illuminated by the sun, calculating an imaging orientation of the contour of Mars according to the relation of relative positions among the sun, Mars and the Mars explorer, and mapping a magnitude of Mars to the gray level of a computer to display on an interface. The simulation method of the Mars object for deep space exploration is applicable to the simulation software of deep space exploration.