CN114485624A - All-time optical navigation method and device based on star and satellite combination - Google Patents
All-time optical navigation method and device based on star and satellite combination Download PDFInfo
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Abstract
The invention discloses a method and a device for all-time optical navigation based on star and satellite combination, wherein the method comprises the following steps: observing the fixed star by using the star sensor to obtain a first optimal attitude matrix in a first preset time period; observing the satellite by using the infrared optical sensor to obtain a first satellite observation equation; constructing a first index function through the altitude information obtained by coupling the prior information and the barometer; obtaining a first optimal longitude and latitude; observing the sun by using a sun sensor to obtain a sun observation vector in a second preset time period; establishing an angle constraint equation of the solar observation unit vector and the satellite observation unit vector; constructing a second index function by using the altitude information obtained by coupling the prior information with the barometer to obtain a second optimal longitude and latitude; and obtaining a second optimal attitude matrix according to the satellite observation vector and the sun observation vector of the satellite observation equation. The invention realizes high-precision all-time navigation in a satellite radio rejection environment and effectively solves the problem of radio interference and deception.
Description
Technical Field
The invention relates to the field of navigation technology and aerospace technology, in particular to a full-time optical navigation method and device based on star and satellite combination.
Background
The long-term flight solar unmanned aerial vehicle such as a rainbow solar unmanned aerial vehicle of China aerospace science and technology group, an achira solar unmanned aerial vehicle of American Facebook, a western wind god solar unmanned aerial vehicle of European airmen can continuously fly for months in the atmosphere, the flying height can reach 20 kilometers, the long-term flight solar unmanned aerial vehicle is a revolutionary aircraft, can be used as an 'high-altitude pseudolite', and has wide application prospect in the fields of regional reconnaissance, environmental monitoring, search and tracking, information pickup, broadband communication, space-time Internet of things service and the like and emergency provision of external communication service in the case of major disasters. The existing long-endurance unmanned aerial vehicle navigation system depends heavily on radio satellite navigation such as Beidou and GPS, but the navigation mode based on the radio satellite is easy to be deceived and interfered.
The rapid development of the micro inertial navigation technology (MIMU) creates conditions for the realization of autonomous navigation of the unmanned system, but the MIMU has low precision, and the navigation result of the unmanned system obtained by directly using an inertial navigation algorithm can be rapidly dispersed, so that the navigation task in long-term navigation can not be completed. The star sensor can measure the attitude information of the carrier, and the geographic information can be obtained by decoupling the star sensor from the measurement result of the MIMU, but the star sensor can only be used at night, and the MIMU can cause the navigation precision to be severely limited under the dynamic condition, so that the navigation requirement cannot be met.
Disclosure of Invention
The present invention is directed to solving, at least in part, one of the technical problems in the related art.
Therefore, the invention aims to solve the problem of high-precision all-time navigation in a satellite radio rejection environment, and provides an all-time optical navigation method based on star and satellite combination.
Another objective of the present invention is to provide a whole-day optical navigation device based on star and satellite combination.
In order to achieve the above purpose, the present invention provides a method for global time optical navigation based on star and satellite combination, comprising the following steps:
observing the fixed star by using the star sensor at a first preset time interval to obtain a first optimal attitude matrix; observing a satellite by using an infrared optical sensor, and obtaining a first satellite observation equation based on the first attitude matrix; constructing a first index function based on elevation information obtained by coupling the prior information and the barometer through the first satellite observation equation; obtaining a first optimal longitude and latitude according to the first index function; observing the sun by using a sun sensor to obtain a sun observation vector in a second preset time period; observing the satellite by using the infrared optical sensor to obtain a second satellite observation equation; establishing an angle constraint equation of the solar observation unit vector and the satellite observation unit vector based on the solar observation vector and the second satellite observation equation; constructing a second index function based on the angle constraint equation and elevation information obtained by coupling the prior information with a barometer; obtaining a second optimal longitude and latitude based on the second index function; and obtaining a second optimal attitude matrix based on the second optimal longitude and latitude and according to the satellite observation vector of the satellite observation equation and the sun observation vector.
The all-time optical navigation method based on the combination of the fixed star and the satellite comprises a star sensor, a sun sensor and an infrared optical sensor, wherein the star sensor observes the fixed star, the sun sensor observes the sun, the infrared optical sensor observes a satellite carrying an infrared light source, the satellite number is obtained through an infrared coding light source, and the acquisition of a satellite precise ephemeris is realized, so that a combined observation equation of an infinite-distance fixed star and a finite-distance satellite is established, a method for decoupling the attitude and the position of the fixed star and the satellite at night and the attitude and the position of the sun and the satellite at daytime is provided, and the acquisition of the all-time position and the attitude of the navigation device is realized. Compared with the traditional satellite radio navigation, the method and the device have the advantages that the high-precision all-time navigation under the satellite radio rejection environment is realized by utilizing the natural celestial body and the artificial celestial body through a photoelectric measurement means, the difficult problem of radio interference deception is effectively solved, and the method and the device can be used as a standby navigation method and a standby navigation device for unmanned systems such as unmanned aircrafts and the like.
In addition, the all-time optical navigation method based on the association of the stars and the satellites according to the above embodiment of the present invention may also have the following additional technical features:
further, the first satellite observation equation is:
wherein s isECIIs a range vector of the satellite under the ECI,
for an optimal state matrix, s*Is the vector of the satellite on the infrared optical sensor, K (t) is the distance from the sensor to the satellite,
for the position of the system in ECEF coordinates,
is a transformation matrix from an ECEF coordinate system to an ECI coordinate system,
the sensor is an installation matrix of the infrared optical sensor and the star sensor.
Further, the first indicator function is:
further, in the first preset time period, observing the star by using the star sensor to obtain a first optimal attitude matrix, including:
according to the principle of the star sensor:
wherein p isECI|iIs the vector of the star in the ECI coordinate system, pST|iThe vector of the fixed star under the star sensor coordinate system is shown;
obtaining a first optimal attitude matrix from a star sensor coordinate system to an ECI coordinate system by imaging a plurality of fixed stars
Further, the second satellite observation equation is:
wherein s isECEFIs the vector of the satellite under the ECEF coordinate system, and is determined by the real-time longitude and latitude height (phi) of the satellites,λs,hs) Solving, K (t) represents the distance from the observation point to the satellite,
as a vector of observation points in the ECEF coordinate system,
and installing a matrix for the system sensor.
Further, the angle constraint equation is:
wherein the content of the first and second substances,
in order to observe the unit vector of the sun,
is a unit vector of satellite observations.
Further, the second indicator function is:
wherein d isENU(Δ x, Δ y, Δ z) represents the change in the position of the sensor in the ENU coordinate system at two predetermined times.
Further, at a preset value of t1The longitude and latitude height (phi, lambda and h) of the observation point at the moment is constructed, the following equation is constructed, and a second optimal attitude matrix is solved
Further, the expression of the star, sun and satellite observation vector is as follows:
wherein v is*Respectively representing sidereal observation vectors pSTSun observation vector r*And satellite observation vector s*，(x0,y0) Represents the midpoint coordinates of the star sensor, the sun sensor and the infrared optical sensor (x)i,yi) Representing the imaging coordinates of the star point, sun and satellite, fdevRespectively representing the focal lengths f of the star sensorsSTFocal length f of sun sensorSunFocal length f of infrared optical sensorIR。
In order to achieve the above object, another aspect of the present invention provides a full-time optical navigation device based on a star and satellite combination, comprising:
the first preset time period module is used for observing the fixed star by using the star sensor in a first preset time period to obtain a first optimal attitude matrix; observing a satellite by using an infrared optical sensor, and obtaining a first satellite observation equation based on the first attitude matrix; constructing a first index function based on elevation information obtained by coupling the prior information and the barometer through the first satellite observation equation; obtaining a first optimal longitude and latitude according to the first index function; the second preset time interval module is used for observing the sun by using the sun sensor to obtain a sun observation vector in a second preset time interval; observing the satellite by using the infrared optical sensor to obtain a second satellite observation equation; establishing an angle constraint equation of the solar observation unit vector and the satellite observation unit vector based on the solar observation vector and the second satellite observation equation; constructing a second index function based on the angle constraint equation and elevation information obtained by coupling the prior information with a barometer; obtaining a second optimal longitude and latitude based on the second index function; and obtaining a second optimal attitude matrix based on the second optimal longitude and latitude and according to the satellite observation vector of the satellite observation equation and the sun observation vector.
The all-time optical navigation device based on the combination of the fixed star and the satellite comprises the star sensor, the sun sensor and the infrared optical sensor, wherein the star sensor observes the fixed star, the sun sensor observes the sun and the infrared optical sensor observes the satellite carrying the infrared light source, the satellite number is obtained through the infrared coding light source, and the acquisition of the precise ephemeris of the satellite is realized, so that a combined observation equation of the infinite-distance fixed star and the finite-distance satellite is established, a method for decoupling the attitude and the position of the fixed star and the satellite at night and the attitude and the position of the sun and the satellite at daytime is provided, and the acquisition of the all-time position and the attitude of the navigation device is realized. Compared with the traditional satellite radio navigation, the method and the device have the advantages that the high-precision all-time navigation under the satellite radio rejection environment is realized by utilizing the natural celestial body and the artificial celestial body through a photoelectric measurement means, the difficult problem of radio interference deception is effectively solved, and the method and the device can be used as a standby navigation method and a standby navigation device for unmanned systems such as unmanned aircrafts and the like.
Additional aspects and advantages of the invention will be set forth in part in the description which follows and, in part, will be obvious from the description, or may be learned by practice of the invention.
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The foregoing and/or additional aspects and advantages of the present invention will become apparent and readily appreciated from the following description of the embodiments, taken in conjunction with the accompanying drawings of which:
FIG. 1 is a schematic illustration of ECI and ECEF coordinate systems according to an embodiment of the present invention;
FIG. 2 is a schematic view of the field of view and the mounting relationship of the star sensor, the sun sensor and the infrared optical sensor according to the embodiment of the invention;
FIG. 3 is a flow chart of a method of full-time optical navigation based on a star and satellite association according to an embodiment of the present invention;
FIG. 4 is a schematic diagram of a star and satellite based all-time navigation frame according to an embodiment of the present invention;
FIG. 5 is a schematic view of an imaging of a star sensor, a sun sensor and an infrared optical sensor according to an embodiment of the present invention;
FIG. 6 is a schematic diagram of a night navigation method according to an embodiment of the present invention;
FIG. 7 is a schematic diagram of a daytime navigation method according to an embodiment of the present invention;
fig. 8 is a schematic structural diagram of a full-time optical navigation device based on the association of stars and satellites according to an embodiment of the present invention.
Detailed Description
It should be noted that the embodiments and features of the embodiments in the present application may be combined with each other without conflict. The present invention will be described in detail below with reference to the embodiments with reference to the attached drawings.
In order to make the technical solutions of the present invention better understood, the technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be obtained by a person skilled in the art without making any creative effort based on the embodiments in the present invention, shall fall within the protection scope of the present invention.
The following describes a method and an apparatus for full-time optical navigation based on the association of stars and satellites according to an embodiment of the present invention with reference to the accompanying drawings, and first, a method for full-time optical navigation based on the association of stars and satellites according to an embodiment of the present invention will be described with reference to the accompanying drawings.
It should be understood that the coordinate systems mentioned in the present invention need to be explained before specifically describing the method for star and satellite based joint all-time optical navigation according to the embodiments of the present invention. As shown in fig. 1, ECI represents the equatorial coordinate system of the celestial sphere, also called the equatorial coordinate system of the geocentric, and is based on the J2000.0 coordinate system, i.e. the coordinate system is an inertial coordinate system established in the 1 st geodynamics time of the public yuan 2000, the ECI coordinate system takes the geocentric as the origin, the X axis points to the vernal equinox, the Z axis points to the north polar celestial earth, and the Y axis is in the equatorial plane and determined by the right-hand rule; ECEF represents a ground-fixed coordinate system with the geocentric as the origin, the X-axis pointing to a point with zero latitude and longitude, the Z-axis pointing to the north and the Y-axis in the equatorial plane and determined by the right-hand rule. The ENU represents a geographic coordinate system, i.e., a northeast coordinate system, with the center of mass of the carrier as the origin, the X-axis pointing east, the Y-axis pointing north, and the Z-axis pointing vertically towards the sky.
As an example, as shown in fig. 2, the all-time navigation device is composed of a 20 ° × 20 ° field-of-view star sensor, a 120 ° × 120 ° field-of-view sun sensor, and a 40 ° × 40 ° infrared optical sensor. The star sensor and the sun sensor are respectively used for observing the fixed star and the sun, and the infrared optical sensor is used for observing the satellite. The all-day navigation is divided into two conditions of night and day, and the star sensor and the infrared optical sensor work in a combined mode at night; in the daytime, the sun sensor and the infrared optical sensor work in combination. The working principle of the star sensor is that the star point position is calculated according to the shot star image and is compared with a known star table to further obtain the attitude information of the star sensor relative to an ECI coordinate system. The sun sensor can obtain the vector of the sun in the sensor, and further calculate the direction angle and the altitude angle of the sun. Once the satellite carrying the infrared light source enters the view field of the infrared sensor, the navigation equipment can observe the satellite light-emitting signal all the day, obtain the satellite observation vector and further obtain the ECEF coordinate of the satellite.
It should be noted that the all-time optical navigation is based on an infinite star and a finite distance satellite, and the decoupling of the attitude and the geographic position is realized. In order to realize the observation of the satellite all the day, the artificial satellite needs to be provided with an infrared band light source, and the precise position of the satellite in an ECEF coordinate system is provided by a precise ephemeris. Under the condition that the UTC time is known, once the satellite is observed, the ground device obtains the satellite number through decoding, and the position of the satellite in an ECEF coordinate system can be obtained in real time when the satellite is observed through the precise ephemeris at the infrared sensor.
As an example, the invention requires an artificial satellite carrying an infrared optical beacon, which has a central spectrum band of 1550nm and a divergence angle of at least + -10 deg. in order to allow the staring light source to pass through the atmosphere and not be absorbed by atmospheric molecules. The satellite orbit is 400 to 600 km high. The satellite infrared light source flickers at a certain frequency, an infrared optical sensor in the navigation equipment determines the satellite number according to the flickering frequency, and then the all-day navigation is carried out under the assistance of the satellite precise ephemeris.
The present invention will be specifically described for the case of night and day, respectively. The attitude under the night condition can be directly obtained through the star sensor, and the position needs to be solved by the combination of the star sensor and the infrared optical sensor; the attitude and position under the daytime condition need to be solved by combining the multi-time sun sensor and the infrared optical sensor.
FIG. 3 is a flow chart of a method for star and satellite based joint all-time optical navigation according to an embodiment of the present invention.
As shown in fig. 3, the method for global time optical navigation based on the association of stars and satellites comprises the following steps:
step S1, observing the fixed star by using the star sensor at a first preset time interval to obtain a first optimal attitude matrix; observing a satellite by using an infrared optical sensor, and obtaining a first satellite observation equation based on a first attitude matrix; constructing a first index function based on elevation information obtained by coupling prior information and a barometer on the basis of a first satellite observation equation; and obtaining a first optimal longitude and latitude according to the first index function.
It should be understood that this step is illustrative of the night case.
Firstly, a star sensor is used for observing a fixed star to obtain a first optimal attitude matrix.
Specifically, as shown in fig. 4 and 6, at night, the navigation device navigates by using a star and a satellite, wherein the star sensor observes the star and the infrared optical sensor observes the satellite.
Specifically, a large number of stars in visible light (400nm-780nm) wave bands exist at night, and because the distance from the stars to the earth is infinite, the star sensor is used for observing and calculating the stars, so that the attitude information of the sensor can be obtained, as shown in fig. 5, according to the principle of the star sensor:
wherein p isECI|iIs the vector of the star in the ECI coordinate system, pST|iThe vector of the fixed star under the star sensor coordinate system has the expression:
wherein (x)0,y0) Is the midpoint coordinate of the star sensor, (x)i,yi) Imaging coordinates for the stars, fSTIs the focal length of the star sensor;
by imaging a plurality of fixed stars, the optimal attitude matrix from the star sensor coordinate system to the ECI coordinate system can be obtained by methods such as QUEST and the like
And then, observing the satellite by using the infrared optical sensor, and obtaining a first satellite observation equation based on the first attitude matrix.
It should be noted that the low-earth orbit satellites (400km to 1000km) have orbits of a certain height, which only needs about 95 minutes to make a circle around the earth, and the observation positions are different when the satellites are observed, which results in different observation satellite vectors.
Specifically, the low-earth orbit satellite carries an infrared light source load, the infrared optical sensor can realize observation of the low-earth orbit satellite, and an observation equation can be expressed as follows:
wherein s isECIIs a range vector of the satellite under the ECI,
for the optimal state matrix, s*The vector of the satellite on the infrared optical sensor is shown in fig. 5, and the calculation method is similar to the calculation of the star point coordinates of the star sensor. K (t) is the distance from the sensor to the satellite,
the position of the system under the ECEF coordinate is represented by (phi, lambda and h) latitude, longitude and elevation respectively, and represents longitude and latitude height information required to be solved.
Is a transformation matrix from an ECEF coordinate system to an ECI coordinate system,
the sensor is an installation matrix of the infrared optical sensor and the star sensor.
Specifically, k (t) is obtained by the following formula:
as an example, in the WGS-84 coordinate system,
can be expressed as:
wherein e represents the Earth's ellipsoidal eccentricity, R, defined by the WGS-84 coordinate systemE(phi) represents the curvature radius of the earth-made unitary fourth of twelve earthly branches, and the calculation formula is respectively as follows:
wherein R ispRepresenting the major semi-axis of the earth ellipse, RoThe minor half axis is indicated.
It can be known that, before the system operates, the time needs to be aligned with the standard UTC time in advance, and when the system works, the high-precision crystal oscillator is used for time service of the system. Thus, UTC time is available according to the system
Obtaining an installation matrix of the infrared optical sensor and the star sensor through calibration according to the system installation relation
sECIIs the vector of the satellite in the ECI coordinate system, which can be determined according to the longitude and latitude height (phi) of the satellite in the ECEF coordinate systems,λs,hs) And
calculation, as follows:
further, a first index function is constructed on the basis of elevation information obtained by coupling the prior information and the barometer of a first satellite observation equation.
Specifically, the elevation information h of the sensor is obtained in a mode of coupling the prior information and the barometer, and then the following index function is constructed:
further, a first optimal longitude and latitude is obtained according to the first index function.
Specifically, the following formula is minimized by a nonlinear least square method, and the optimal longitude and latitude (phi, lambda) can be obtained:
min‖G(φ,λ)‖2
step S2, observing the sun by using the sun sensor to obtain a sun observation vector in a second preset time period; observing the satellite by using the infrared optical sensor to obtain a second satellite observation equation; establishing an angle constraint equation of the solar observation unit vector and the satellite observation unit vector based on the solar observation vector and the second satellite observation equation; constructing a second index function based on the angle constraint equation and elevation information obtained by coupling the prior information and the barometer; obtaining a second optimal longitude and latitude based on a second index function; and obtaining a second optimal attitude matrix based on a second optimal longitude and latitude and according to the satellite observation vector of the satellite observation equation and the sun observation vector.
It is understood that this step is illustrative of the case of daytime.
In the daytime, due to the radiation of the atmospheric background, the starry light can be covered by the atmospheric reflected light, and the only star capable of being used for navigation only has the sun, but the attitude under an inertial system cannot be obtained under the condition of only one vector, and the geographic position cannot be solved. Therefore, as shown in fig. 4 and 7, in one embodiment of the present invention, a finite distance satellite is introduced, a satellite vector is continuously observed by using the high-speed motion characteristic of the satellite and the daytime observability of the infrared light source, a pose coupling equation is established with the sun vector, multi-frame iterative decoupling is performed, and the solution of the pose and the geographic position is realized. The equipment used for daytime navigation is a sun sensor and an infrared optical sensor, the sun sensor observes the sun to obtain a sun vector, and the infrared optical sensor observes the satellite. The specific steps of the calculation are as follows:
firstly, in a second preset time period, the sun is observed by using a sun sensor to obtain a sun observation vector. The second preset time period is a daytime time period.
It can be understood that the expression of the star, sun and satellite observation vector of the present invention is:
wherein v is*Respectively representing sidereal observation vectors pSTSun observation vector r*And satellite observation vector s*，(x0,y0) Represents the midpoint coordinates of the star sensor, the sun sensor and the infrared optical sensor (x)i,yi) Representing the imaging coordinates of the star point, sun and satellite, fdevRespectively representing the focal lengths f of the star sensorsSTFocal length f of sun sensorSunFocal length f of infrared optical sensorIR。
Specifically, as shown in fig. 5, the imaging principle of the sun sensor is similar to that of the star sensor, and the sun observation unit vector can be expressed as:
wherein (x)0,y0) Represents the midpoint coordinate of the sun sensor (x)i,yi) Imaging coordinates representing the sun point, fSunIndicating the focal length of the sun sensor.
Under the condition of knowing the UTC time t, the solar vector r in the ECEF coordinate system at any time is obtained according to the solar ephemerisECEFFurther, the following relationship is obtained:
wherein the content of the first and second substances,
and the attitude matrix from the sun sensor coordinate system to the ECEF coordinate system is represented.
Further, a second satellite observation equation is obtained by observing the satellite by using the infrared optical sensor.
It can be known that the star sensor cannot observe the fixed star in the daytime, so that the attitude relationship cannot be directly solved. In the daytime, the satellite can be observed by the infrared optical sensor through the infrared light source load on the satellite, so that a satellite observation equation is obtained:
wherein s isECEFThe vector of the satellite in the ECEF coordinate system can be represented by the real-time longitude and latitude height (phi) of the satellites,λs,hs) Solution of s*The unit vector of the satellite on the infrared optical sensor is shown and is consistent with the star point solved by the star sensor. K (t) represents the distance from the observation point to the satellite.
Representing the vector of the observation point in the ECEF coordinate system. Wherein the content of the first and second substances,
is calculated in the same manner as above, sECEF、Ks(t) is calculated by the following formula:
further, an angle constraint equation of the sun observation unit vector and the satellite observation unit vector is established based on the sun observation vector and the second satellite observation equation.
Specifically, by comparing the sun observation vector of the sun sensor at two moments with the satellite observation vector of the infrared optical sensor at two moments, the following relationship can be obtained:
the measured values of the sun sensor and the infrared sensor are unit vectors, so that a sun observation unit vector is established
Unit vector of observation with satellite
The angle constraint equation of (c) can be expressed as:
system sensor mounting matrix
Are calibrated in advance. Therefore, the vector included angle theta can be obtained by observing the sun and the satellite at the same timet. After the relative accurate elevation information h of the sensor is known, two unknown quantities of longitude and latitude also exist. Because the motion speed of the near-earth satellite is high, the vector included angle between the satellite and the sun changes at the next moment. Therefore, at two moments, two equations exist, and the optimal longitude and latitude can be solved.
Further, a second index function is constructed based on the angle constraint equation and elevation information obtained by coupling the prior information and the barometer, and a second optimal longitude and latitude is obtained.
Specifically, while navigating, motion of the device may occur. In order to compensate for the change of the geographic position of the navigation equipment, the initial speed and the acceleration of the sensor under an ENU coordinate system are introduced. The following equation is established.
Wherein d isENU(Deltax, Deltay, Deltaz) represents the change in position of the sensor in the ENU coordinate system at two moments, measured by the moving sensor of the carrier,
the rotation matrix representing the observation point p from the ENU coordinate system to the ECEF coordinate system can be expressed as:
respectively obtaining the included angle of the solar unit vector and the satellite unit vector and the displacement change condition between two moments, and constructing the following index function
Multiple iterations may be performed by a nonlinear least squares method to minimize the following.
min‖L(φ,λ)‖2
Through multiple iterations, the longitude and latitude information of the sensor in the daytime can be solved.
And further, obtaining a second optimal attitude matrix based on the second optimal longitude and latitude and according to the satellite observation vector and the sun observation vector of the satellite observation equation.
In particular, to solve the attitude matrix
With t1For example, the following equation is constructed:
by means of the methods such as QUEST and the like, under the condition that the longitude and latitude height (phi, lambda and h) of an observation point is known, the optimal attitude moment can be solved according to the satellite observation vector and the sun observation vectorMatrix of
Through the steps, high-precision all-time navigation in a satellite radio rejection environment is realized through a photoelectric measurement means, the problem of radio interference deception is effectively solved, and the method and the device can be used as a standby navigation method and a standby navigation device for unmanned systems such as unmanned aircrafts and the like.
In order to implement the above embodiment, as shown in fig. 8, the present embodiment further provides an all-time optical navigation device 10 based on a star and satellite combination, where the device 10 includes: a first preset time period module 100 and a second preset time period module 200.
A first preset time period module 100, configured to observe the star by using the star sensor at a first preset time period to obtain a first optimal attitude matrix; observing a satellite by using an infrared optical sensor, and obtaining a first satellite observation equation based on a first attitude matrix; constructing a first index function based on elevation information obtained by coupling prior information and a barometer through a first satellite observation equation; obtaining a first optimal longitude and latitude according to a first index function;
a second preset time interval module 200, configured to observe the sun by using the sun sensor to obtain a sun observation vector at a second preset time interval; observing the satellite by using the infrared optical sensor to obtain a second satellite observation equation; establishing an angle constraint equation of the solar observation unit vector and the satellite observation unit vector based on the solar observation vector and the second satellite observation equation; constructing a second index function based on the angle constraint equation and elevation information obtained by coupling the prior information and the barometer; obtaining a second optimal longitude and latitude based on a second index function; and obtaining a second optimal attitude matrix based on the second optimal longitude and latitude and according to the satellite observation vector and the sun observation vector of the satellite observation equation.
According to the all-time optical navigation device based on the star and satellite combination, high-precision all-time navigation under the satellite radio rejection environment is realized through a photoelectric measurement means, the problem of radio interference deception is effectively solved, and the all-time optical navigation device can be used as a standby navigation method and device for unmanned systems such as unmanned aerial vehicles and the like.
It should be noted that the foregoing explanation of the embodiment of the all-day optical navigation method based on star and satellite combination is also applicable to the all-day optical navigation device based on star and satellite combination in this embodiment, and is not repeated here.
Furthermore, the terms "first", "second" and "first" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include at least one of the feature. In the description of the present invention, "a plurality" means at least two, e.g., two, three, etc., unless specifically limited otherwise.
In the description herein, references to the description of the term "one embodiment," "some embodiments," "an example," "a specific example," or "some examples," etc., mean that a particular feature, structure, material, or characteristic described in connection with the embodiment or example is included in at least one embodiment or example of the invention. In this specification, the schematic representations of the terms used above are not necessarily intended to refer to the same embodiment or example. Furthermore, the particular features, structures, materials, or characteristics described may be combined in any suitable manner in any one or more embodiments or examples. Furthermore, various embodiments or examples and features of different embodiments or examples described in this specification can be combined and combined by one skilled in the art without contradiction.
Although embodiments of the present invention have been shown and described above, it is understood that the above embodiments are exemplary and should not be construed as limiting the present invention, and that variations, modifications, substitutions and alterations can be made to the above embodiments by those of ordinary skill in the art within the scope of the present invention.
Claims (10)
1. A full-time optical navigation method based on star and satellite combination is characterized by comprising the following steps:
observing the fixed star by using the star sensor at a first preset time interval to obtain a first optimal attitude matrix; observing a satellite by using an infrared optical sensor, and obtaining a first satellite observation equation based on the first attitude matrix; constructing a first index function based on elevation information obtained by coupling the prior information and the barometer through the first satellite observation equation; obtaining a first optimal longitude and latitude according to the first index function;
observing the sun by using a sun sensor to obtain a sun observation vector in a second preset time period; observing the satellite by using the infrared optical sensor to obtain a second satellite observation equation; establishing an angle constraint equation of the solar observation unit vector and the satellite observation unit vector based on the solar observation vector and the second satellite observation equation; constructing a second index function based on the angle constraint equation and elevation information obtained by coupling the prior information with a barometer; obtaining a second optimal longitude and latitude based on the second index function; and obtaining a second optimal attitude matrix based on the second optimal longitude and latitude and according to the satellite observation vector of the satellite observation equation and the sun observation vector.
2. The method of claim 1, wherein the first satellite observation equation is:
wherein s isECIIs a range vector of the satellite under the ECI,
for an optimal state matrix, s*Is the vector of the satellite on the infrared optical sensor, K (t) is the distance from the sensor to the satellite,
for the position of the system in ECEF coordinates,
is a transformation matrix from an ECEF coordinate system to an ECI coordinate system,
the sensor is an installation matrix of the infrared optical sensor and the star sensor.
3. The method of claim 1, wherein the first indicator function is:
4. the method for global-time optical navigation based on star and satellite combination according to claim 1, wherein the observing the star by using the star sensor during the first preset time period to obtain the first optimal attitude matrix comprises:
according to the principle of the star sensor:
wherein p isECI|iIs the vector of the star in the ECI coordinate system, pST|iThe vector of the fixed star under the star sensor coordinate system is shown;
obtaining a first optimal attitude matrix from a star sensor coordinate system to an ECI coordinate system by imaging a plurality of fixed stars
5. The method of claim 1, wherein the second satellite observation equation is:
wherein s isECEFIs the vector of the satellite under the ECEF coordinate system, and is determined by the real-time longitude and latitude height (phi) of the satellites，λs，hs) Solving, K (t) represents the distance from the observation point to the satellite,
as a vector of observation points in the ECEF coordinate system,
and installing a matrix for the system sensor.
6. The method of claim 1, wherein the angle constraint equation is:
wherein, the first and the second end of the pipe are connected with each other,
in order to observe the unit vector of the sun,
is a unit vector of satellite observations.
7. The method of claim 1, wherein the second indicator function is:
wherein d isENU (Δ x, Δ y, Δ z) represents the position change of the sensor in the ENU coordinate system within a preset two moments.
8. The method for full-time optical navigation based on star and satellite combination as claimed in claim 1, wherein the following equation is constructed for solving the second optimal attitude matrix at the longitude and latitude height (phi, lambda, h) of the observation point at the preset time t1
9. The sidereal and satellite based all-time optical navigation method according to claim 1, wherein the expression of the sidereal, solar and satellite observation vector is as follows:
wherein v is*Respectively representing sidereal observation vectors pSTSun observation vector r*And satellite observation vector s*，(x0，y0) Represents the midpoint coordinates of the star sensor, the sun sensor and the infrared optical sensor (x)i，yi) Representing the imaging coordinates of the star point, sun and satellite, fdevRespectively representing the focal lengths f of the star sensorsSTFocal length f of sun sensorSunFocal length f of infrared optical sensorIR。
10. An all-time optical navigation device based on star and satellite combination, comprising:
the first preset module is used for observing the fixed star by using the star sensor at a first preset time period to obtain a first optimal attitude matrix; observing a satellite by using an infrared optical sensor, and obtaining a first satellite observation equation based on the first attitude matrix; constructing a first index function based on elevation information obtained by coupling the prior information and the barometer through the first satellite observation equation; obtaining a first optimal longitude and latitude according to the first index function;
the second preset module is used for observing the sun by using the sun sensor to obtain a sun observation vector in a second preset time period; observing the satellite by using the infrared optical sensor to obtain a second satellite observation equation; establishing an angle constraint equation of the solar observation unit vector and the satellite observation unit vector based on the solar observation vector and the second satellite observation equation; constructing a second index function based on the angle constraint equation and elevation information obtained by coupling the prior information with a barometer; obtaining a second optimal longitude and latitude based on the second index function; and obtaining a second optimal attitude matrix based on the second optimal longitude and latitude and according to the satellite observation vector of the satellite observation equation and the sun observation vector.
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