386 results about "Navigation error" patented technology

An autonomous unmanned space flight system and planetary lander executes a discrete landing sequence including performing an initial velocity braking maneuver to remove velocity at altitude, coasting during which the planet surface is imaged and correlated to reference maps to estimate cross-track and along-track navigation errors and one or more lateral braking maneuvers are performed to reduce cross-track navigation error, and performing a terminal velocity braking maneuver(s) to reduce the along-track braking maneuver and remove the remainder of the velocity just prior to landing. A bi-propellant propulsion system provides a very high T/M ratio, at least 15:1 per nozzle. Short, high T/M divert maneuvers provide the capability to remove cross-track navigation error efficiently up to the maximum resolution of the reference maps. Short, high T/M terminal velocity braking maneuver(s) provide the capability to remove along-track navigation error to a similar resolution and remove the remaining velocity in a very short time window, approximately 3-15 seconds prior to touchdown. The propulsive efficiency frees up mass which can be allocated to a fuel to remove the unknown navigation errors, perform hazard avoidance and/or relocate the lander by flying it to another site or be allocated to additional payload.

A horizontal navigation system aided by a carrier phase differential Global Positioning System (GPS) receiver and a Laser-Scanner (LS) for an Autonomous Ground Vehicle (AGV). The high accuracy vehicle navigation system is highly demanded for advanced AGVs. Although high positioning accuracy is achievable by a high performance RTK-GPS receiver, the performance should be considerably degraded in a high-blockage environment due to tall buildings and other obstacles. The present navigation system is to provide decimetre-level positioning accuracy in such a severe environment for precise GPS positioning. The horizontal navigation system is composed of a low cost Fiber Optic Gyro (FOG) and a precise odometer. The navigation errors are estimated using a tightly coupled Extended Kalman Filter (EKF). The measurements of the EKF are double differenced code and carrier phase from a dual frequency GPS receiver and relative positions derived from laser scanner measurements.

The invention discloses an SINS/GPS super-compact integrated navigation system and an implementing method thereof. The method comprises the following steps: the doppler frequency assistance is provided for a GPS carrier loop by using the velocity information of a strapdown inertial navigation system, therefore, the loop equivalent bandwidth is increased, the influence of the carrier dynamic stateon the carrier loop is lowered, and the noise suppression capability is improved by reducing the bandwidth of a filter; meanwhile, in order to eliminate the correlation between the pseudo-range rate error and the inertial navigation error, a carrier loop tracking error model is obtained by establishing the relationship between the carrier tracking error and the inertial navigation speed error, andthe influence of the carrier tracking error is subduced in the measurement equation; and in addition, the carrier frequency is adjusted according to the output error estimation information, and the tracking accuracy of the carrier loop is enhanced. The invention can effectively enhance the noise suppression capability and the dynamic tracking performance of the tracking loop and enhance the tracking accuracy of a GPS receiver and the navigation accuracy of the integrated navigation system under strong interference and high dynamic circumstance.

The invention discloses a GPSINS combination navigation method and system utilizing a neural network algorithm to realize compensation and correction. The method comprises steps that a combination navigation model based on a GPS and an inertial navigation module is constructed, a Kalman filter is designed to filter combination navigation data of the combination navigation model, and navigation error data of the inertial navigation module is outputted; in a normal GPS signal state, inertial navigation data is taken as input data of a training sample of a neural network model, output data of the navigation data after filtered by the Kalman filter is taken as output data of a training sample of a neural network module, and the neural network model is trained; in a GPS signal loss state, the trained neural network model is utilized to predict an output error of the inertial navigation module, and the predicted output error is further utilized to compensate and correct inertial navigation. The method is advantaged in that the GPSINS combination navigation model is compensated and corrected through utilizing the neural network algorithm, and the inertial navigation system is assisted by the neural network algorithm to output the accurate navigation data in the GPS signal loss state.

This invention relates to a test method for error model verification of an inertial measurement system, which sets up a multifunction software platform to simulate and argue the airliner load test plan directly including: designing specific flying trails based on the error model of an inertial sensor, in which, an aircraft flies along the designed flying trail and collects inertial sensor data of specific flight path points and transmits them to a master navigation computer, then the collected data are compared with the data output by a measurement primary standard device to get the measurement error of the sensor to compute and get the error parameters in the inertial navigation error model with an airborne inertial navigation space multi-position online demarcating method to understand the inertial navigation performance and is convenient for the compensation in navigation computation.

The invention discloses a water unmanned ship local hierarchical path planning method based on the navigation error constraint. The method mainly includes the steps of local static path planning based on the navigation error constraint and combined with a genetic algorithm; local dynamic path planning fused with the maritime affair rule. The hierarchical thinking is adopted, local path planning is divided into two levels of static path planning and dynamic path planning, the static barrier avoiding problem is solved, and the dynamic barrier avoiding problem is also solved. According to static path planning based on the navigation error constraint and combined with the genetic algorithm, adverse influences of navigation errors on path selection are reduced, and the safety of path planning is improved; according to dynamic path planning based on the marine affair rule, the dynamics constraint of a water unmanned ship is considered, and the feasibility of path planning is improved.

An improved system and method are disclosed for facilitating target aiming and aircraft control by flight crews using electronic displays. For example, an aircraft display system is disclosed that includes a database for storing target location and boundary data, a processing unit, a position and direction determination unit for use in determining the position of an aiming symbol on a display, a graphics display generator, and a visual display. The processing unit determines whether or not a selected target to be displayed is obscured or covered by the aiming symbol to be displayed. If any portion of the target is obscured or covered by the aiming symbol, then the processing unit directs the graphics display generator to generate graphic control signals for the visual display, which increase the transparency of the portions of the aiming symbol that obscure or cover the target. Notably, the transparency of any symbol (or portion of a symbol) can be increased if that symbol obscures or covers a target being displayed or viewed. Thus, the visibility of targets viewed using onboard visual displays is significantly enhanced, whereby the targets are obscured by the symbols used to aim and control the aircraft. Advantageously, target aiming and aircraft control are facilitated, which reduces pilot workload and navigation errors, and thus results in increased aircraft safety.

The invention relates to a navigation method realized by means of loose combination of visual navigation/inertial navigation. The main research content is to periodically correct inertial navigation parameters by using the position of the visual navigation and velocity information. The navigation method solves the problem that inertial navigation error accumulates over time. According to the navigation method, an error equation of an inertial navigation system is used as a state equation of a filter, the difference between the position calculated by means of the visual navigation and the position solved by using the inertial navigation system is used as measurement of Kalman filtering, and the parameters of the inertial navigation system and the error of a sensor are estimated, so that the inertial navigation parameters are corrected. Compared with the prior art, by adopting the scheme, the navigation method can enable the inertial navigation system to be maintained to have high accuracy for a long time, and the accuracy of the inertial navigation system is not affected due to failure of visual navigation calculation.

The invention relates to a GPS navigation map accurate matching method of a pilotless automobile, and a method thereof. The map accurate matching method comprises the following steps: 1, obtaining road information; 2, determining a start point; 3, obtaining vehicle positioning information; 4, matching and screening the information; and 5, repeating step 1-4 until successful matching. The method and the system can reduce the navigation error to 2m or less, can timely adjust when the navigation error exceeds 2m, and greatly reduces the navigation error.

The invention discloses a CKF filtering-based vehicle dynamic model auxiliary inertial navigation combined navigation method. The CKF filtering-based vehicle dynamic model auxiliary inertial navigation combined navigation method comprises the following steps: calculating posture, speed and position of a vehicle according to angle increment and specific force output by a micro-inertia device and by an inertial navigation numerical value updating algorithm; establishing a three-degree-of-freedom vehicle dynamic model, and calculating the speed of a carrier by taking a steering wheel angle and a longitudinal force as control input quantity and by a fourth order Ronge-Kutta method in real time; designing a CKF filter by taking an inertial navigation equation as a state equation and speed difference between a dynamical model and inertial navigation calculation to perform state estimation on a combined navigation system; performing output correction on strapdown inertial navigation calculation result by the position the speed and the posture error obtained by CKF estimation, and performing feedback correction on the inertial navigation through peg-top and adding error. The method aims at the problems that the inertial navigation error is accumulated along with time and navigation precision cannot be maintained for a long time, and the accuracy and the reliability of a vehicle navigation system can be improved.

The invention discloses an information fusion method for an airborne inertia/Doppler radar integrated navigation system. The method comprises the following six major steps of: I, establishing an error model of an inertia device and a Doppler radar; II, determining a state variable based on the application of inertia/Doppler integrated navigation on an airborne fixed-wing aircraft; III, establishing an inertia/Doppler radar integrated navigation state equation; IV, establishing an inertia/Doppler radar integrated navigation measurement equation; V, discretizing a Kalman filtering continuous system; and VI, performing Kalman filtering data fusion resolving, obtaining the sub-optimal estimation related to navigation parameters through state estimation, and correcting the inertial navigation and Doppler radar in a closed-loop feedback mode at the same time. The method disclosed by the invention makes full use of the Doppler information and inertial navigation information, puts forward a data fusion scheme for airborne inertia/Doppler radar integrated navigation, effectively solves the problem of a navigation error caused by an imperfect error model, and improves the system precision.

The invention belongs to the field of geomagnetism auxiliary navigation location, and in particular relates to an inertial navigation error correction method based on geomagnetism modulus gradient and particle filter. The method comprises the following steps: solving the position on which a carrier is arranged by adopting an inertial navigation system according to the information of an accelerometer on a submarine; predicting the position error of the carrier according to a state equation of a geomagnetism modulus gradient/inertial combination navigation system; acquiring the geomagnetism modulus gradient measurement value of the submarine on a real position in real time by utilizing a geomagnetism modulus gradient measuring device when the submarine sails underwater; obtaining a difference value between the predicted modulus gradient value and the observed modulus gradient value; estimating the system state by utilizing a particle filter estimation algorithm based on the mass point dynamics physicomimetics optimization, and carrying out the error compensation on the inertial navigation system. According to the method, the sailing track of the carrier is corrected according to an estimation result, and the drifting of the inertial navigation gyroscope is estimated and compensated. An ideal way is provided for the underwater carrier to realize the precise autonomous navigation.

The invention relates to a satellite starlight refraction navigation error determination and compensation method. The method comprises the following steps: firstly, generating data of a satellite orbit by virtue of STK software; establishing a satellite attitude planning model; determining an actual observation visual field, and simulating an observation star map comprising refraction stars and non-refraction stars; calculating to obtain the tangential height h of the refraction stars; calculating to obtain the tangential height h'' of the refraction stars with errors of refraction angles and errors of an atmospheric model, wherein tangential height errors are mainly caused by the measurement precision errors of the refraction angles and the errors of the atmospheric model; calculating to obtain the position of a satellite under a geocentric inertial coordinate system; and performing filtration by an extended Kalman filtration method, and outputting a starlight refraction navigation estimated position and position errors. The method provided by the invention can realize accurate prediction of navigation precision of a satellite starlight refraction navigation system, and is comprehensive in error analysis and accurate and reliable in results.

A position estimation system including a first arrangement for providing an image with a known target in a known reference frame. A second arrangement correlates the image with a stored image. The correlation is used to compute an error with respect to a position estimate. In a specific embodiment, the error is referenced with respect to first (x), second (y) and third (z) directions. A target location error is computed with respect to a stored image provided by a target image catalog. The target image catalog includes target geo-locations and digital terrain elevation data. In an illustrative application, the image data is provided by synthetic aperture radar and forward-looking infrared systems. An observation model and a measure noise matrix are Kalman filtered to ascertain a position error in navigation data generated by an integrated inertial navigation and Global Positioning system. In the illustrative application, geo-registered SAR/FLIR imagery is used to track targets and to determine a target location error (TLE). This TLE information is a set of error equations that describe the relationship between vehicle navigation information and target data. In accordance with the invention, this relationship is used to form an observation model for vehicle navigation with respect to target locations. Using Kalman filtering and the observation model, vehicle navigation errors can be bound and the navigation accuracy of the vehicle can be improved.

The invention relates to a combined navigation method based on polarization information. The combined navigation method comprises the following steps: firstly, checking astronomical ephemeris according to carrier positions and moments so as to obtain sun position information; secondly, measuring by using a polarization sensor so as to obtain polarization information, and measuring speed information by using a light stream sensor; thirdly, on the basis of polarization and light stream, establishing a filtering measuring model; and finally, estimating navigation error information by using a kalman filter. By adopting the combined navigation method, a linear measuring model based on single-polarization sensor is established, and the model is simple, small in filtering calculation amount, high in precision and applicable to autonomous navigation of the carrier for a relatively long time.

The invention provides an intelligent vehicle road junction passing method and system. Log data recorded by a log system in the driving process by human beings at the earlier stage is used for assisting an intelligent vehicle in automatic driving later. Compared with the prior art, the invention provides a road junction passing method and system with low cost and high performances of an intelligent vehicle by utilizing road junction memory bars, so that an unmanned driving vehicle can pass correctly at different specific road junctions under the conditions of precise map deficiency, great navigation error as well as universal road junctions, and the intelligent vehicle road junction passing method and system can be used for auxiliary driving and unmanned driving of vehicles.

The invention discloses a method for realizing integrated navigation through a ship's inertial navigation system (SINS) and a celestial navigation system (SNS) and belongs to the field of integrated navigation. The invention uses the SINS error mathematic platform as CNS level base and creates a deep combined mathematical model according to the coupling SINS attitude error celestial positioning principle. The method of the invention comprises the following steps: the first step, creating the celestial navigation positioning model based on the SINS mathematic platform; the second step, creating an SINS and CNS combined system model; the third step, performing discretization process through the system state equation and the measuring equation of the created integrated navigation and using a Kalman filter to optimally estimate and verify the inertial navigation error. The inertial/celestial integrated navigation system has excellent autonomy and anti-interference capability. Through the method of the invention, no special level base is needed on the aircraft so as to simplify the design of the SINS/CNS integrated navigation system under an onboard condition.

The invention discloses a multi-level embedded integrated navigation system and a navigation method, belonging to the technical field of integrated navigation. The method comprises the following steps: firstly establishing an aviation airborne inertial navigation system (INS) error state value equation and an inertia/satellite integrated measurement model, then defining a navigation error state value of an embedded inertia/satellite integrated system and establishing an inertia/star sensor integrated measurement model; on that basis, carrying out embedded inertia/satellite integrated system internal integrated navigation resolving and close-loop filtering correction, and finally carrying out external inertia/star sensor integrated navigation resolving and open-loop filtering correction to obtain multi-information fusion navigation results. According to the invention, the complete independent integrated structure of the embedded inertia/satellite integrated system can be kept, and simultaneously the multi-level fusion of inertia/satellite/astronomy asynchronous multi-sensor information and the step-by-step correction of the inertial navigation system are realized, the integrated navigation performance is effectively raised, and the invention is suitable for engineering application.

The invention discloses an inertial navigation-fingerprint location-combined positioning error correction method. The method comprises A, determining a reference position, B, in movement, building positioning road signs at intervals by an inertial navigation method and recording the information of the positioning road signs, wherein the information comprises space coordinates, a wireless signal set and inertial navigation errors of a travel distance relative to the reference position, and C, according to the wireless signal set measured in real time, determining if the wireless signal matches with the recorded positioning road sign, and when the recorded positioning road sign is matched, based on coordinates of the matched positioning road sign, calibrating and correcting coordinates determined by the current inertial navigation technology. When the on-site positioning error is small, fingerprint is acquired and is used as a road sign, and through fingerprint coupling, when a user passes through the road sign again, the current position coordinates are calibrated by the small-error position coordinates so that an error estimation range is reduced and positioning precision is improved.

The invention belongs to the technical field of guidance, navigation and control of deep sky detectors, and particularly discloses a high-precision navigation method for a landed or attached deep sky celestial body detector. The method comprises the following steps: (1) determining the three-dimensional speed of a coordinate system of a detector body; (2) determining initial value of the center position of a celestial body in the coordinate system of the detector body; (3) determining the position and the direction of the celestial body in the body system; (4) determining the attitude angle and the speed of the detector relative to the coordinate system of an orbit; (5) determining the height of the detector relative to the surface of the celestial body; and (6) determining the angular speed of the detector relative to the coordinate system of the orbit. The method is not affected by the situation that inertial navigation error is constantly increased along the time, thereby effectively improving the precision of key navigation parameters for autonomous navigation of the landed or attached detector, and meeting the demands on navigation and control of the high-precision landed or attached deep sky celestial body detector.

The invention discloses a formation cooperation navigation method for an unmanned aerial vehicle. During formation of the unmanned aerial vehicle, a lead aircraft unmanned aerial vehicle node is carried with an inertial navigation system and a GPS receiver, the posture, speed and position are estimated through airborne inertial navigation/GPS fusion filtering, and high-precision posture, speed andposition navigation information of the lead aircraft unmanned aerial vehicle node is obtained; other unmanned aerial vehicle nodes are carried with the inertial navigation system, a vision navigationdevice and a light stream sensor, through the vision navigation device, the relative distance of the unmanned aerial vehicle and a lead aircraft, a pitch angle and an azimuth angle are obtained, absolute position information of the adjacent lead aircraft unmanned aerial vehicle nodes is mathematically solved, speed information of the unmanned aerial vehicle is obtained through solution of the light stream sensor, and the airborne inertial navigation error is corrected through Kalamn filtering. In addition, the relative distance between the unmanned aerial vehicle node, further away from the lead aircraft, in formation and other unmanned aerial vehicle nodes is solved through the vision navigation device, and based on the relative distance, the absolute position information is solved.

This invention relates to a calibration method for ultra-short baseline acoustic positioning system, which comprises a main processor, a transceiver, a control unit, an ultra-short base line sound head, an underwater target transponder, a GPS system and IMU device both arranged on survey ship body above water surface. The calibration method comprises: a. stores GPS data, ship attitude data and the geometrical distance from sound head to the transponder; b. inputs to compute the transponder position and the position error from the sound head to GPS antenna; c. calls the computation to compute the navigation error for sound head; d. outputs new computer program for high-precise positioning.