CN106382927B - An autonomous navigation method for star sensors based on satellite identification - Google Patents

Abstract

The invention provides an autonomous navigation method for a star sensor based on satellite identification, and belongs to the technical field of autonomous navigation methods for a star sensor. The invention can realize full autonomous navigation based on the star sensor, and provide the carrier with high-precision attitude and position information that does not diverge with time. The method of the invention is to create a satellite star map and create a fusion star map, and obtain all information and attitude information of stars and satellites according to the matching of all celestial bodies in the field of view with the star map. According to the obtained star and planet information, the improved starlight angular distance method is used for high-precision positioning, and then the conversion of the carrier attitude information is completed, realizing the true full autonomous navigation method based on the star sensor. Advantages of the present invention: 1) adopt satellite information, have good adaptability and flexibility; 2) adopt satellite positioning, have good stability and precision; 3) adopt redundant multi-satellite solution, anti-jamming; 4) adopt Fusion star map matching increases information reliability.

Description

An autonomous navigation method for star sensors based on satellite identification

technical field

The invention relates to a star sensor autonomous navigation method based on satellite identification, and belongs to the technical field of star sensor autonomous navigation methods.

Background technique

The star sensors are able to work independently and provide extremely high-precision inertial attitude information that has no cumulative error and does not diverge over time. The inertial attitude information output by the star sensor includes the implicated motion component introduced by the earth's rotation and the error components such as precession, nutation, and pole shift caused by the earth's perturbation, so the inertial attitude output of the star sensor cannot be directly used as a navigation parameter. Solve. In order to eliminate the influence of the above factors, the inertial attitude information output by the star sensor needs to be transformed into attitude information in the navigation coordinate system through coordinate transformation and compensation. In the process of attitude transformation, position information needs to be introduced at the same time to realize the decomposition and conversion of attitude, so the accuracy of position information will directly affect the accuracy of the final attitude output.

The position information in the attitude calculation of the star sensor, the vast majority of position information acquisition methods are provided by external information sources, such as inertial navigation systems, but the position information of inertial navigation systems will produce error accumulation due to the working principle of inertial navigation. That is, the navigation accuracy diverges with time, so the position information provided by the inertial navigation system cannot meet the high-precision requirements of the star sensor for attitude determination and positioning; Parameters or adding specific maneuvers, such as using a sextant to directly sense the horizon, such as increasing atmospheric density parameters to measure starlight refraction and indirectly sensitive horizons, such as using specific maneuvers to keep the platform at a specific attitude, etc. The advantage of using the star sensor to directly solve the position is that the position information obtained by the star sensor has no cumulative error and does not diverge with time. However, when the star sensor direct positioning method is used, although the system can ensure a certain positioning capability, it has the following problems: more errors are introduced while adding more parameters, and the maneuverability of the carrier is reduced while performing specific maneuvers When adding more auxiliary equipment, the passivity and autonomy of the star sensor are reduced, so it is urgent to propose a new method to realize the autonomous positioning of the star sensor.

At present, the autonomous positioning methods of star sensors can be generally divided into direct sensitive horizon, indirect sensitive horizon and pure astronomical geometry methods, among which the pure astronomical geometry method has more advantages, because it does not need to introduce additional auxiliary parameters, equipment or maneuver to simultaneously solve Simpler and faster. In the pure astrogeometric positioning method, the starlight angular distance method can use the angle and position relationship between the star and other non-stellar celestial bodies in the field of view of the star sensor to perform positioning. The size is proportional to and inversely proportional to the distance between the observation point and the measured celestial body, that is, the larger the starlight angle and the closer the celestial body distance, the higher the positioning accuracy. Therefore, in the positioning process, it is hoped to use near-Earth objects closer to the earth, such as satellites, which can effectively improve the positioning accuracy.

In the satellite positioning process, the position information of the satellite needs to be used, which requires identifying the satellite information currently involved in the calculation. Usually satellite identification is carried out when the position and attitude of the carrier are known, but in the autonomous navigation of the star sensor, the position and attitude of the carrier are unknown, and there is no research to propose an effective satellite identification method, there is no corresponding star sensor autonomous attitude determination, navigation method for positioning.

Contents of the invention

The purpose of the present invention is to solve the problems existing in the above-mentioned prior art, that is, the existing research cannot realize satellite identification under the condition of unknown navigation parameters, and then cannot independently use star sensors for high-precision attitude determination and positioning to realize star sensors The problem of autonomous navigation. Furthermore, an autonomous navigation method of the star sensor based on satellite identification is provided.

The purpose of the present invention is achieved through the following technical solutions:

A method for autonomous navigation of a star sensor based on satellite identification, comprising the following steps:

Step 1. Inertial attitude measurement: The star sensor is matched with the created all-sky star map Map _star according to the star S _i information captured in the field of view, and the attitude information Att of the current star sensor relative to the inertial coordinate system is obtained;

Step 2. Star information identification: According to the matching attitude information, identify all star information in the current field of view, obtain right ascension and declination information, and calculate the current star sensor optical axis pointing unit vector L _i ;

Step 3. Determine the observation plane: the star sensor adopts the form of a multi-eye star sensor, assume that the optical axes of all star sensors intersect at a point in the calculation space, and any two star sensor optical axes determine the inertia of the optical axis of a star sensor System plane P _{optical axis} ;

Step 4. Satellite map creation: Create a satellite map map _{satellite based on the inertial coordinate system based on all satellites in orbit in space. Map satellite is} related to the world time UTC, that is, it changes with time and rotates around the center of the earth;

Step 5. Satellite map segmentation: the plane P _{optical axis} and the satellite map Map _satellite intersect to form a circle C _P×Map ; take C _P×Map as the center, and take half of the width of the field of view Width of the star sensor as the bandwidth to divide the satellite star The map Map _satellite is extended to both sides and divided into an annular zone F _width/2 parallel to the P _{optical axis} , S _width/2 has a total of I=2π/width parts, and F _i //F _j {i,j∈I,i≠ j};

Step 6. Divide the star map projection: project the ring zone F _width/2 onto the plane P _⊥oa perpendicular to the optical axis pointing to L _i , this projection is to project the spherical satellite space into the plane to form a certain relative position distortion Rectangular plane; the belt-shaped star map formed by Map _satellite projection is Map _projection , and the projection is related to UTC time and attitude Att;

Step 7. Constellation satellite map fusion: Map _star and Map _projection are synthesized to form a fusion star map Map _integrated (UTC, Att), and Map _integrated is a function of UTC and Att;

Step 8, fusion star map matching: match all celestial bodies in the field of view, including stars and satellites, with the star map _integrated to obtain the information R _i of the currently observed satellite;

Step 9. Starlight angular distance positioning: use the starlight vector information S _i of the stars, the starlight angle information IA _ij between the stars and satellites, the relative distance information ρ _i of the satellites, and the inertial position information R _i of the satellites to perform positioning and calculate the position of the carrier for r;

Step 10. Improve satellite selection: In each calculation, stars and satellites that exceed the minimum number of satellites required for calculation are introduced for positioning to improve calculation accuracy; if the number of stars is N and the number of satellites is K, choose 2 out of N Calculate the angle IA between the starlight vectors of stars, that is, IA _M , Choose 3 starlight angles in M to calculate L _i , that is, L _i , Similarly satellite R _j ,

Step 11. Improve positioning calculation: When the influence of the carrier’s deviation from the origin of the celestial sphere on the attitude measurement is not considered, the star sensor directly gives the carrier’s attitude angle Att _i relative to the inertial space through star sky observation and star map matching. The relationship in the sensor image plane coordinate system is R(α,β), then the satellite direction vector can be calculated by L _i =R(α,β)·Att _i ;

Step 12. Limit the selection of fixed satellites: limit the number of introduced stars M and satellites K. Since there will be certain measurement errors in the process of star observation, using too many stars and satellites to participate in the calculation will instead introduce too many errors, and then Reduce the accuracy of positioning and navigation; follow in Hengwei selection:

1. The selected stars and satellites should be as close as possible to the center of the field of view;

2. The starlight vector angle between stars and stars, between stars and satellites is greater than the set threshold δ _threshold ;

3. The satellite is a new satellite, the orbital perturbation is weak, and the orbital parameters are more accurate;

4. The single measurement error of the starlight angle is within σ;

Step 13, Xingmin autonomous navigation: calculate the position of the carrier, and convert it into position information in the navigation coordinate system to complete autonomous positioning, perform attitude conversion on the basis of known position, time and earth parameters to complete autonomous attitude determination, comprehensive Finally, autonomous navigation is completed.

The present invention has the following beneficial effects: the use of satellite identification method can directly use satellite information, which greatly improves the flexibility and applicability of positioning; the use of fusion star map matching can obtain more accurate satellite identification results and improve system positioning. Reliability of information sources; using satellite positioning method, because the accuracy of satellite position information is higher, the positioning is more accurate than planetary positioning method, and it is less affected by the perturbation of observed celestial bodies. At the same time, the number of satellites is large, which improves the stability of continuous positioning The use of redundant multi-satellite measurement and calculation can effectively eliminate the calculation error caused by inaccurate star and satellite observation information, and improve the final positioning accuracy. The present invention can realize attitude determination and positioning of a high-precision star sensor, and the navigation result does not diverge with time, the positioning accuracy is better than 50m, and the attitude determination accuracy error caused by the positioning error does not exceed 1".

Detailed ways

The present invention will be described in further detail below: the present embodiment is implemented on the premise of the technical solution of the present invention, and detailed implementation is provided, but the protection scope of the present invention is not limited to the following examples.

A kind of star sensor autonomous navigation method based on satellite identification involved in the present embodiment, comprises the following steps:

Step 1. Inertial attitude measurement: The star sensor is matched with the created all-sky star map Map _star according to the star S _i information captured in the field of view, and the attitude information Att of the current star sensor relative to the inertial coordinate system is obtained;

Step 2. Star information identification: According to the matching attitude information, identify all star information in the current field of view, obtain right ascension and declination information, and calculate the current star sensor optical axis pointing unit vector L _i ;

Step 3. Determine the observation plane: the star sensor adopts the form of a multi-eye star sensor, assume that the optical axes of all star sensors intersect at a point in the calculation space, and any two star sensor optical axes determine the inertia of the optical axis of a star sensor System plane P _{optical axis} ;

Step 4. Satellite map creation: Create a satellite map Map _satellite based on the inertial coordinate system based on all satellites in orbit in space, and new satellites and satellites with smaller perturbation errors have a higher selection priority. Map _satellite It is related to the world time UTC, that is, it changes with time and rotates around the center of the earth;

Step 5. Satellite map segmentation: The plane P _{optical axis} and the satellite map Map _satellite intersect to form a circle C _P×Map ; with C _P×Map as the center and half the width of the field of view Width of the star sensor as the bandwidth, the satellite star The map Map _satellite is extended to both sides and divided into an annular zone F _width/2 parallel to the P _{optical axis} , S _width/2 has a total of I=2π/width parts, and F _i //F _j {i,j∈I,i≠ j};

Step 6. Divide the star map projection: project the ring zone F _width/2 onto the plane P _⊥oa perpendicular to the optical axis pointing to L _i , this projection is to project the spherical satellite space into the plane to form a certain relative position distortion Rectangular plane; the belt-shaped star map formed by the Map _satellite projection is Map _projection , and the projection is related to UTC time and attitude Att; in order to eliminate the distortion caused by the projection as much as possible, it can be properly matched within the field of view of the maintenance unit Reduce the width Width of F _width/2 under the condition that the number of satellites is sufficient;

Step 7. Constellation satellite map fusion: Map _star and Map _projection are synthesized to form a fusion star map Map _integrated (UTC, Att), and Map _integrated is a function of UTC and Att;

Step 8, fusion star map matching: match all celestial bodies in the field of view, including stars and satellites, with the star map _integrated to obtain the information R _i of the currently observed satellite;

Step 9. Starlight angular distance positioning: use the starlight vector information S _i of the stars, the starlight angle information IA _ij between the stars and satellites, the relative distance information ρ _i of the satellites, and the inertial position information R _i of the satellites to perform positioning and calculate the position of the carrier for r;

Step 10. Improve satellite selection: In each calculation, stars and satellites that exceed the minimum number of satellites required for calculation are introduced for positioning to improve calculation accuracy; if the number of stars is N and the number of satellites is K, choose 2 out of N Calculate the angle IA between the starlight vectors of stars, that is, IA _M , Choose 3 starlight angles in M to calculate L _i , that is, L _i , Similarly satellite R _j ,

Step 11. Improve positioning calculation: When the influence of the carrier’s deviation from the origin of the celestial sphere on the attitude measurement is not considered, the star sensor directly gives the carrier’s attitude angle Att _i relative to the inertial space through star sky observation and star map matching. The relationship in the sensor image plane coordinate system is R(α,β), then the satellite direction vector can be calculated by L _i =R(α,β)·Att _i ;

Step 12. Limit the selection of fixed satellites: limit the number of introduced stars M and satellites K. Since there will be certain measurement errors in the process of star observation, using too many stars and satellites to participate in the calculation will instead introduce too many errors, and then Reduce the accuracy of positioning and navigation; follow in Hengwei selection:

1. The selected stars and satellites should be as close as possible to the center of the field of view;

2. The starlight vector angle between stars and stars, between stars and satellites is greater than the set threshold δ _threshold ;

3. The satellite is a new satellite, the orbital perturbation is weak, and the orbital parameters are more accurate;

4. The single measurement error of the starlight angle is within σ;

Step 13, Xingmin autonomous navigation: calculate the position of the carrier, and convert it into position information in the navigation coordinate system to complete autonomous positioning, perform attitude conversion on the basis of known position, time and earth parameters to complete autonomous attitude determination, comprehensive Finally, autonomous navigation is completed.

The above are only preferred specific implementations of the present invention. These specific implementations are all based on different implementations under the overall concept of the present invention, and the scope of protection of the present invention is not limited thereto. Anyone familiar with the technical field Within the technical scope disclosed in the present invention, any changes or substitutions that can be easily conceived by a skilled person shall fall within the protection scope of the present invention. Therefore, the protection scope of the present invention should be determined by the protection scope of the claims.

Claims (2)

1. a kind of star sensor autonomous navigation method based on satellite identification, which is characterized in that

Step 1: inertial attitude measures: star sensor is according to the fixed star S captured in visual field_uInformation, with the whole day ball created Star chart Map_starIt is matched, obtains posture information Att of the current star sensor relative to inertial coodinate system；

Step 2: fixed star information identifies: according to matched posture information, identifying fixed star information all in current field, obtain The information of right ascension, declination, while calculating current star sensor optical axis and being directed toward unit vector L_i；

Step 3: plane of vision determines: star sensor uses more mesh star sensor forms, if all star sensor optical axis meet at meter It calculates in space a bit, and any two star sensor optical axis determines inertial system plane where a star sensor optical axis P_{optical axis}；

Step 4: satellite star chart creates: creating the satellite based on inertial coodinate system according to the satellite of operation on orbit all in space Star chart Map_satellite, Map_satelliteIt is related to zebra time UTC, that is, it changes over time, and rotated around earth the earth's core；

Step 5: satellite star-fields segmentation: plane P_{optical axis}With satellite star chart Map_satelliteIt is crossed to form round C_P×Map；With C_P×MapCentered on, using the half of star sensor visual field width Width as bandwidth, by satellite star chart Map_satelliteExtend to both sides It is divided into and is parallel to P_{optical axis}Ring belt F_width/2, F_width/2I=2 π/width parts in total, and F_p//F_q{p,q∈I,p ≠q}；

Step 6: segmentation star chart projection: by ring belt F_width/2L is directed toward to perpendicular to optical axis_iPlane P_⊥oaUpper projection, the projection Spherical surface satellite spatial is projeced into plane, forms the rectangle plane with certain relative position distortion；By Map_satelliteIt throws The band-like star chart that shadow is formed is Map_projection, and the projection is related to UTC time and posture Att；

Step 7: permanent satellite mapping fusion: by Map_starAnd Map_projectionIt is comprehensive to form fusion star chart Map_integrated(UTC, Att), Map_integratedFor the function of UTC and Att；

Step 8: fusion star pattern matching: by celestial body all in visual field, including fixed star and satellite and star chart Map_integratedIt carries out Matching obtains the inertial position information R for being currently observed satellite_j；

Step 9: starlight angular distance positions: utilizing the starlight vector information S of fixed star_u, starlight angle information between fixed star satellite IA_uv, satellite relative distance information ρ_vAnd the inertial position information R of satellite_jThe position for carrying out location Calculation carrier is r；

Chosen Step 10: improving Heng Wei: introduced in calculating every time be more than minimum perseverance needed for calculating defend quantity fixed star and satellite into Row positioning, to improve computational accuracy；If number of stars are N, number of satellite is K, 2 fixed stars calculate starlight vector in optional N Angle IA_M, i.e.,3 starlight vector angle calcu-lation L in optional M_i, i.e.,The similarly inertia position of satellite Confidence breath

Step 11: improving location Calculation: when not considering that carrier deviates influence of the celestial sphere origin to attitude measurement, star sensor is logical Cross the attitude angle Att that Starry sky observation and star pattern matching directly give carrier relative to inertial space_w, satellite and fixed star are in star sensitivity Relationship in device photo coordinate system is R (α, β), then satellite direction vector can pass through Q_w=R (α, β) Att_wIt calculates；

Step 12: limitation Heng Wei chooses: limitation introduces the quantity of fixed star M and satellite K, due to that can have in celestial body observation process There is certain measurement error, excessive error can be introduced instead by participating in calculating using excessive fixed star and satellite, and then is reduced and determined The accuracy of position and navigation；It is followed in Heng Wei selection:

1), the fixed star, satellite of selection are as far as possible close to field of view center；

2), the starlight vector angle between fixed star and fixed star, fixed star and satellite is greater than the threshold value δ of setting_threshold；

3), satellite is new satellite, orbit perturbation is weaker, orbit parameter is more accurate；

4), the single measurement error of starlight angle is within σ；

Step 13: the quick independent navigation of star: calculating the position of carrier, and be converted to the completion of the location information under navigational coordinate system Autonomous positioning carries out posture on the basis of known location, time and earth parameter and converts autonomous attitude determination, to sum up final complete At independent navigation.

2. the star sensor autonomous navigation method according to claim 1 based on satellite identification, which is characterized in that the step It in rapid six, is distorted to eliminate as far as possible by projection bring, is keeping matching the enough items of number of satellite in unit visual field F is reduced under part_width/2Width Width.