12 results about "Celestial navigation" patented technology

The invention provides a visual celestial navigation method based on sun centroid extraction and sea-sky-line fitting. Belongs to the application of image processing technology in the field of celestial navigation. The method mainly comprises the steps of extracting a high-precision sea-sky-line by using a textural feature segmentation and Canny-Hough transform combined algorithm, and extracting a sky region based on a sea-sky-line segmentation result. A solar facula image candidate region and a sun candidate contour are constructed through a sky region, a dual optimization strategy of region similarity screening and facula feature scoring is proposed, and sub-pixel-level sun centroid coordinate calculation is realized by combining an intensity weighted center method and nonlinear least square two-dimensional Gaussian model fitting. And further constructing a camera coordinate system space vector through the orthogonal projection point of the sea-sky-line and the sun centroid, calculating a sun elevation angle based on a vector included angle, and performing angle compensation by adopting an atmospheric refraction correction model. And finally, fusing multi-moment observation data through a sun shift line positioning method, completing autonomous calculation of the ship position, and improving the positioning precision.

The embodiment of the invention provides an interactive multi-model factor graph integrated navigation method based on tight coupling attitude constraint. In a specific embodiment, the method comprises the following steps: constructing an integrated navigation system; constructing an inertial measurement unit factor pre-integrated by the inertial navigation system; constructing a direct celestial navigation factor and a refraction celestial navigation factor which are respectively and tightly coupled with an inertial measurement unit factor based on the star sensor observation condition of the celestial navigation system; constructing a horizontal radio navigation factor and a vertical radio navigation factor based on the position information of the land-based radio navigation system; acquiring height measurement data of the barometric altimeter to construct barometric altimeter factors; dynamically adjusting the weight ratio of the height measurement data of the vertical radio navigation factor to the height measurement data of the barometric altimeter factor through an interactive multi-model algorithm; and carrying out global state estimation of the integrated navigation system on a factor graph framework formed by the factors.

The invention discloses a low-altitude high-speed flight manned flying saucer, and relates to the technical field of aviation aircrafts, the low-altitude high-speed flight manned flying saucer comprises a saucer-shaped body, a power device and an undercarriage, the saucer-shaped body is composed of a central bearing ball cabin and a peripheral reinforcing ring coaxially sleeved outside the central bearing ball cabin; the central bearing ball cabin is divided into a celestial navigation and communication cabin, a panoramic cockpit and an environment control and life support system cabin from top to bottom; the inner cabin section of the peripheral reinforcing ring forms a passenger cabin; the bearing strength and the space utilization rate of the machine body are effectively improved through the composite structure of the central spherical cabin and the peripheral reinforcing ring; due to the integrated design of the annular duct and the vector engine, the aerodynamic efficiency and the control flexibility in the vertical take-off and landing and high-speed flight states are remarkably improved; an independent environment control system ensures the safety and comfort of passengers in high-altitude flight; the overall layout gives consideration to functionality and maintenance convenience, and has high practical value.

Techniques are disclosed for improving daylight performance of a star tracker. A region of interest (ROI) within a field of view of sky is obtained. The region of interest is divided into sub-ROIs. Sub-ROIs including a candidate celestial object are identified. An angular velocity of each candidate celestial object is determined. The known celestial object is a candidate celestial object identified as having a angular velocity, during a time period, that is within a range of angular velocities about a known angular velocity associated with the known celestial object.

The invention relates to the technical field of celestial navigation, in particular to an X-ray photon detection method and device based on intensity correlation, and the method comprises the following steps: 1, employing two detection units to observe the same pulsar at the same time, recording the arrival time of X-ray photons, and completing the initial collection of X-ray pulsar signals; 2, converting the arrival time of the X-ray photons; 3, X-ray noise eliminating photons are calculated; 4, processing the de-noised X-ray signal photons to obtain an arrival time sequence of the X-ray photons with a high signal-to-noise ratio; and 5, folding the arrival time sequence of the X-ray photons with the high signal-to-noise ratio to obtain a test contour, and comparing the test contour with a standard contour to obtain high-precision TOA information. According to the method, the difference analysis of X-ray signal photons and noise photons is utilized, so that the noise photons from the same sight direction and the same energy spectrum range of the pulsar are effectively eliminated, and the signal-to-noise ratio of a detection signal is effectively improved.

A system for selecting one or more resident space objects from a number of resident space objects for targeting includes a database for storing properties of the resident space objects and one or more processors. The processors are configured to determine, for each resident space object, a utility metric representing a navigation utility of the resident space object. Determining the utility metric includes determining a projected position of the resident space object by projecting a position of the resident space object over time based on ephemeris data and covariance data associated with the resident space object, and computing the utility metric based at least in part on the projected position. The processors rank the resident space objects according to their utility metrics and select the one or more resident space objects for targeting based on the ranking.

The invention relates to a star sensor star measurement probability evaluation method, and belongs to the technical field of airborne celestial navigation. The method comprises the following steps: generating a screening star library according to a limit star magnitude design value of a star sensor; calculating the optical axis direction of the star sensor according to the star measurement time, place and attitude; screening out an observable star set again from the screening star library according to the star sensor design parameters and the optical axis direction; setting satellite measurement time, place and attitude sequences, performing continuous satellite measurement analogue simulation, and recording the number of satellites in an observable satellite set; and counting the satellite measurement probability and the satellite measurement distribution condition, and evaluating the availability of the design parameters. According to the method, the preliminary analogue simulation verification of the airborne star sensor can be realized, a reference is provided for the parameter iterative design of the star sensor, and too much trial and error investment in the aspect of real objects in the earlier stage of the design is avoided.

The application discloses an aerial calibration method of an aviation celestial navigation device, combines a flight route in a flight task of an airplane, integrates aerial calibration work of a celestial navigation system into the flight route, and synchronously carries out tracking and positioning work of a star tracker and a navigation processor in the celestial navigation device after the airplane enters a calibration task area; according to a task route plan, the airplane enters each side of a route path once, and the celestial navigation system completes aerial calibration once under the condition that a heading determination meets a requirement; after completing calibration of the last side, a calibration result is output, and the aerial calibration is completed. The application shifts the environment of calibration work of the celestial navigation device from the ground to the top of the clouds in the air, and synchronously carries out the calibration work with a task, avoids important influence of change of the clouds in the air on a calibration success rate, improves the calibration success rate, saves a large amount of input of manpower and material resources, and can be popularized to the fields of spaceflight and navigation, and has a good guiding role for application, popularization and development of the celestial navigation device.

The application discloses a star sensor, which comprises a detection lens, a star sensor body frame, a rotating mechanism, a circuit board, a heat-conducting pad, a heating film and a PEEK shell; the detection lens is installed on the star sensor body frame; the rotating mechanism is connected with the star sensor body frame to make the star sensor body frame rotate axially; the circuit board is arranged on the star sensor body frame; the heat-conducting pad is arranged between the circuit board and the star sensor body frame to transfer the heat generated by the chip of the circuit board to the star sensor body frame; the outer side of the circuit board is sequentially provided with the heating film and the PEEK shell; the heating film can heat the circuit board; and the PEEK shell can isolate the radiant heat of the high-temperature environment. The application further discloses an astronomical navigation system. The application can adapt to high and low temperatures, strong radiation and other harsh environments, and ensures the reliability of the equipment.

The invention relates to a laboratory simulation satellite measurement navigation device and method, and belongs to the technical field of airborne celestial navigation. The device comprises a target simulation servo table, a star target simulator and an acquisition processing computer, the star target simulator is mounted above the target simulation servo table, the target simulation servo table is provided with a load platform double-degree-of-freedom rotating shaft and a star target simulation single-degree-of-freedom rotating shaft, and the acquisition processing computer is cross-linked with the target simulation servo table. The method comprises the following steps: compiling a servo table transposition program according to requirements; calculating a simulated star vector under each transposition, and storing the simulated star vector into a simulated star library; the target simulation servo table rotates in place, and a simulation star calculation vector under the current transposition is calculated; the simulated star is observed, and the actual measurement direction of the simulated star is calculated; calculating combined navigation measurement; the combined navigation calculation of the current transposition is completed; and traversing all translocations to finish the laboratory simulation satellite measurement navigation task. According to the method, the limitation of weather factors on satellite measurement navigation tests is broken through, most satellite measurement navigation debugging and testing requirements in the research and development process can be met, the product research and development speed is increased, and the research and development period is shortened.

A navigational initialization system that includes a reference oscillator, altitude reference system, a celestial navigation system (CNS), and a controller is provided. The reference oscillator is used to generate timing signals. The altitude reference system is used to generate altitude information. The CNS includes a star tracker and inertial sensor assembly (ISA). The star tracker is configured to determine orientation of the star tracker with respect to an earth centered inertial frame. The ISA is used to determine at least an attitude with respect to a local vertical frame. The controller is configured to generate an initialization signal using the timing signals generated by the reference oscillator, the altitude information generated from the altitude reference system, and at least one output of the CNS. The initialization signal is configured to initialize a navigation system.

The invention discloses a multi-source integrated navigation method, system and equipment and a storage medium. The method comprises the following steps: establishing a Kalman filtering state equation based on an inertial navigation speed error equation, a position error equation and an attitude error equation; and establishing a Kalman filtering observation equation by taking multi-source measurement information as observed quantity, carrying out filtering iterative operation, and estimating and compensating an error of an inertial device as well as navigation resolving speed, position, attitude and course errors. By adopting integrated navigation of inertial navigation, celestial navigation, satellite navigation and other measurement sources, multi-source integrated navigation adopting different types of measurement sources can be flexibly configured according to requirements, so as to meet the navigation requirements in various scenes.