A method and system for space autonomous positioning and navigation of a medium-low earth orbit vehicle

Abstract

The present application relates to astronomical navigation method and system, specifically relates to a kind of low-orbit spacecraft space autonomous positioning navigation method and system, for solving the limitation of current astronomical navigation technology for accurate space positioning, so that astronomical navigation is difficult to independently complete the space coordinate positioning task of low-orbit spacecraft.The low-orbit spacecraft space autonomous positioning navigation method utilizes star sensor and photogrammetry camera to work cooperatively, obtains time information and space attitude angle when aircraft navigation positioning, calculates attitude rotation matrix, combines image point centroid pixel coordinates extracted by high-orbit satellite imaging and high-orbit satellite orbit information, establishes linear equation set using object point image point collinearity, and finally solves the space coordinates of low-orbit spacecraft in the earth-centered inertial coordinate system, to realize autonomous space positioning navigation.

Description

TECHNICAL FIELD

[0001] The present application relates to an astronomical navigation method and system, in particular to a kind of low-orbit spacecraft space autonomous positioning navigation method and system. BACKGROUND

[0002] With the rapid growth of global space exploration and satellite application demand, the navigation technology of spacecraft is facing increasing challenges. The navigation of low-orbit spacecraft usually relies on radio navigation and satellite navigation (such as GPS and Beidou, etc.), while astronomical navigation is also used as a supplement. However, with the continuous increase in the number of space missions, the limited ground station resources are under excessive pressure, and a large amount of manpower and material resources are needed to maintain the operation of these ground stations. In special periods, such as natural disasters or communication blockage, traditional ground-based radio navigation and space-based navigation methods may be limited, thereby affecting the positioning accuracy and safety of low-orbit spacecraft.

[0003] In contrast, astronomical navigation, due to its autonomous nature, has become an important means for low-orbit spacecraft to achieve autonomous navigation. Astronomical navigation can help low-orbit spacecraft achieve autonomous operation and management in orbit, reducing dependence on ground stations. However, current astronomical navigation technology is mainly applied to attitude measurement of low-orbit spacecraft, and there are limitations for accurate space positioning, which makes it difficult for astronomical navigation to independently complete the space coordinate positioning task of low-orbit spacecraft. Therefore, how to organically combine astronomical navigation with other navigation technologies to solve the problem of low-orbit spacecraft space positioning has become a problem to be solved. SUMMARY

[0004] The purpose of the present application is to solve the problem that current astronomical navigation technology has limitations for accurate space positioning, making it difficult for astronomical navigation to independently complete the space coordinate positioning task of low-orbit spacecraft, and to provide a low-orbit spacecraft space autonomous positioning navigation method and system.

[0005] In order to solve the above-mentioned deficiencies of the prior art, the present application provides the following technical solutions:

[0006] A low-orbit spacecraft space autonomous positioning navigation method, characterized in that it comprises the following steps:

[0007] Step 1, obtain the time information T0 at the time of navigation positioning, and assume that the space coordinates of the low-orbit spacecraft in the Earth-Centered Inertial coordinate system at the time of navigation positioning are (X S , Y S , Z S );

[0008] Step 2, measure the space attitude angle of the medium and low orbit vehicle in the earth-centered inertial coordinate system when navigating and positioning by the star sensor, and then obtain the attitude rotation matrix R;

[0009] Step 3, image at least two high orbit satellites by the photogrammetry camera fixedly installed with the star sensor, obtain the high orbit satellite image, and then extract the corresponding image point mass center pixel coordinates (x i , y i ) of the i-th high orbit satellite in the high orbit satellite image;

[0010] Step 4, calculate the space coordinates (X i , Y i , Z i ) of the i-th high orbit satellite in the earth-centered inertial coordinate system when navigating and positioning by using the time information T0, the programmed high orbit satellite orbit information;

[0011] Step 5, according to the collinearity of the photogrammetry object point and the image point, the following equation can be obtained:

[0012]

[0013] Wherein, f is the focal length of the optical system of the photogrammetry camera, a is the pixel size of the imaging sensor chip in the photogrammetry camera, (c x , c y ) is the principal point coordinates in the photogrammetry coordinate system, and λ is the scaling factor between the object and the image.

[0014] Substitute the attitude rotation matrix R obtained in step 2 to obtain the linear equation group about the space coordinates (X S , Y S , Z S ), and then calculate the space coordinates (X S , Y S , Z S ) to complete the space coordinate positioning of the medium and low orbit vehicle.

[0015] Further, the step 2 is specifically as follows:

[0016] Step 2.1, acquire the star image by the star sensor, extract the measured position of the star from the star image, and match with the pre-stored star table to obtain the known position of the star;

[0017] Step 2.2, calculate the space attitude angle of the medium and low orbit vehicle in the earth-centered inertial coordinate system when navigating and positioning by comparing the known position of the star with the measured position of the star;

[0018] Step 2.3, calculate the attitude rotation matrix R according to the space attitude angle obtained in step 2.2 as follows:

[0019] R=R x(φ)·R y (θ)·R z (ψ)

[0020] wherein R x (φ), R y (θ), R z (ψ) represent the rotation matrix around the x-axis, y-axis, z-axis of the medium-low orbit vehicle respectively.

[0021] Further, in step 3, the corresponding image point mass center pixel coordinates (x i , y i ) of the i-th high orbit satellite in the high orbit satellite image are extracted by a sub-pixel subdivision positioning algorithm.

[0022] Further, the step 4 is specifically as follows:

[0023] Step 4.1, six orbit parameters of each high orbit satellite are obtained according to the cataloged high orbit satellite orbit information, including semi-major axis a, eccentricity e, inclination i0, ascending node right ascension Ω, perigee argument ω, and initial mean anomaly M0 of the orbit;

[0024] Step 4.2, the spatial coordinates (X i , Y i , Z i ) of the i-th high orbit satellite are calculated according to the six orbit parameters of each high orbit satellite, and time information T0.

[0025] Further, the step 4.2 is specifically as follows:

[0026] Step 4.2.1, the perigee argument E is calculated by solving the Kepler equation through an iterative method:

[0027]

[0028] M = E - e·sin(E)

[0029] wherein M is the current mean anomaly of the orbit, n is the average angular velocity of the high orbit satellite along the orbit, μ is the standard gravity parameter of the earth, and T epoch is the epoch time;

[0030] Step 4.2.2, the two-dimensional coordinates (X o , Y o ) of the i-th high orbit satellite in the orbit plane are calculated by using the perigee argument E:

[0031] X o = a + (cosE - e)

[0032]

[0033] Step 4.2.3, convert the two-dimensional coordinates (X o , Y o ) of the i-th high-orbit satellite to the geocentric inertial coordinate system to obtain the spatial coordinates (X i , Y i , Z i );

[0034]

[0035] In the formula, R z , R x respectively represent the rotation matrix around the z-axis and the x-axis of the i-th high-orbit satellite.

[0036] Further, in step 5, the attitude rotation matrix R obtained in step 2 is substituted to obtain a linear equation group about the spatial coordinates (X S , Y S , Z S ), and the spatial coordinates (X S , Y S , Z S ) are calculated as follows:

[0037] Let the attitude rotation matrix R be:

[0038]

[0039] Further, the linear equation group about the spatial coordinates (X S , Y S , Z S ) is obtained by transformation:

[0040]

[0041] (X S , Y S , Z S ) is obtained by solving by the least square method.

[0042] Meanwhile, the application also provides a kind of low-orbit vehicle space autonomous positioning navigation system, for realizing the above-mentioned low-orbit vehicle space autonomous positioning navigation method;

[0043] Its special place is: including installation base plate, and star sensor and photogrammetry camera are arranged on installation base plate, and star sensor optical axis and photogrammetry camera optical axis are arranged in parallel.

[0044] Further, the star sensor includes a first optical lens, a first light shield, a first imaging and data processing module; the first optical lens is used for imaging, and is located on the optical axis of the star sensor and is located in the first light shield;

[0045] The photogrammetry camera comprises a second optical lens, a second light shield and a second imaging and data processing module; the second optical lens is used for imaging and is located on the optical axis of the photogrammetry camera and within the second light shield;

[0046] The first imaging and data processing module and the second imaging and data processing module are in communication connection, the first imaging and data processing module is used for calculating a posture rotation matrix R according to the spatial attitude angle of the medium and low orbit vehicle in the earth-centered inertial coordinate system and sending the spatial attitude angle of the medium and low orbit vehicle in the earth-centered inertial coordinate system to the second imaging and data processing module, and the second imaging and data processing module is used for extracting corresponding image point mass center pixel coordinates of at least two high orbit satellites in the high orbit satellite image and converting the image point mass center pixel coordinates into spatial coordinates in the earth-centered inertial coordinate system, and then combining the spatial attitude angle of the medium and low orbit vehicle in the earth-centered inertial coordinate system to calculate the spatial coordinates of the medium and low orbit vehicle in the earth-centered inertial coordinate system during navigation positioning.

[0047] Compared with the prior art, the present application has the following beneficial effects:

[0048] (1) The present application is a medium and low orbit vehicle spatial autonomous positioning and navigation method, which utilizes the cooperation of a star sensor and a photogrammetry camera, calculates a posture rotation matrix through the acquisition of time information and spatial attitude angle during vehicle navigation positioning, combines the image point mass center pixel coordinates extracted by high orbit satellite imaging and high orbit satellite orbit information, uses the collinearity of object points and image points to establish a linear equation group, and finally solves the spatial coordinates of the medium and low orbit vehicle in the earth-centered inertial coordinate system to realize autonomous spatial positioning and navigation; the present application combines high orbit satellite observation and medium and low orbit vehicle sensor data, and does not require ground auxiliary equipment, thus having the characteristics of high efficiency and independence.

[0049] (2) The present application uses a star sensor for spatial attitude positioning, has the characteristics of high precision (measurement precision better than 5"), high refresh frequency (better than 10 Hz) and mature technology, and can meet the rapid attitude adjustment requirements of a vehicle.

[0050] (3) The present application uses photogrammetry technology to use the known three-dimensional coordinates of high orbit satellites for positioning, and only two high orbit satellites are required to realize high-precision positioning of a medium and low orbit vehicle, with a positioning precision better than 50 meters, and the algorithm has good robustness.

[0051] (4) The present application is a medium and low orbit vehicle spatial autonomous positioning and navigation system, which comprises a star sensor and a photogrammetry camera, has a simple design, is light in weight and small in size, and is suitable for application scenarios of vehicles with strict requirements on load weight and size.

[0052] (5) The attitude of the star sensor measurement coordinate system and the photographic imaging camera measurement coordinate system in the application can be flexibly designed according to installation requirements, and the attitude relationship between the two is determined through calibration, which is suitable for various environments and installation requirements, and has strong adaptability. BRIEF DESCRIPTION OF DRAWINGS

[0053] Figure 1 It is a structural schematic diagram of an embodiment of a low-orbit aircraft space autonomous positioning and navigation system of the application.

[0054] Figure 2 It is a working principle schematic diagram of a photographic survey camera in the embodiment of the application.

[0055] The reference signs are explained as follows: 1, star sensor optical axis; 2, first light shield; 3, first optical lens; 4, first imaging and data processing module; 5, photographic survey camera optical axis; 6, second light shield; 7, second optical lens; 8, second imaging and data processing module; 9, mounting base plate; 10, imaging sensor chip. DETAILED DESCRIPTION

[0056] The application will be further described below in combination with the drawings and exemplary embodiments.

[0057] Reference Figure 1 A low-orbit aircraft space autonomous positioning and navigation system, comprising a mounting base plate 9, and a star sensor and a photographic survey camera arranged on the mounting base plate 9; the star sensor optical axis 1 and the photographic survey camera optical axis 5 are arranged in parallel, which facilitates the calibration of the three-dimensional attitude relationship between the star sensor measurement coordinate system and the photographic survey coordinate system.

[0058] The star sensor is a star sensor, which is used to measure the space attitude angle of the low-orbit aircraft in the earth-centered inertial coordinate system, and then obtain the attitude rotation matrix R; the star sensor comprises a first optical lens 3, a first light shield 2, and a first imaging and data processing module 4; the first optical lens 3 is used for imaging, and is located on the star sensor optical axis 1 and in the first light shield 2; the first light shield 2 is used to reduce the influence of stray light on the attitude measurement accuracy, and ensure the attitude measurement accuracy.

[0059] Reference Figures 1-2 The photographic survey camera is used to image at least two high-orbit satellites, and extract corresponding image point centroid pixel coordinates; for example Figure 2 In the application, the photographic survey camera uses a central perspective projection model to image two high-orbit satellites S1 and S2 in space, and projects them on the imaging sensor chip 10 to form corresponding image points S1' and S2'.

[0060] The photogrammetry camera comprises a second optical lens 7, a second light shield 6 and a second imaging and data processing module. The second optical lens 7 is used for imaging and is located on the optical axis 5 of the photogrammetry camera and in the second light shield 6; the focal length of the second optical lens 7 is determined by the flight height of the medium-low orbit vehicle, the flight height of the high-orbit satellite to be detected and the pixel size of the imaging sensor chip 10; the second light shield 6 is used to reduce the detection imaging of the high-orbit satellite by stray light. The second imaging and data processing module is arranged on the mounting base plate 9, and the imaging sensor chip 10 is arranged in the second imaging and data processing module.

[0061] The first imaging and data processing module 4 and the second imaging and data processing module 8 are in communication connection, the first imaging and data processing module 4 is used for calculating the attitude rotation matrix R according to the spatial attitude angle of the medium-low orbit vehicle in the earth-centered inertial coordinate system, and sending the spatial attitude angle of the medium-low orbit vehicle in the earth-centered inertial coordinate system to the second imaging and data processing module 8, and the second imaging and data processing module 8 is used for extracting the corresponding image point mass center pixel coordinates of at least two high-orbit satellites in the high-orbit satellite image, and converting them into spatial coordinates in the earth-centered inertial coordinate system, and then calculating the spatial coordinates of the medium-low orbit vehicle in the earth-centered inertial coordinate system during navigation and positioning in combination with the spatial attitude angle of the medium-low orbit vehicle in the earth-centered inertial coordinate system.

[0062] A medium-low orbit vehicle space autonomous positioning and navigation method adopts the above-mentioned medium-low orbit vehicle space autonomous positioning and navigation system, and comprises the following steps:

[0063] Step 1, obtaining time information T0 during navigation and positioning, and assuming that the spatial coordinates of the medium-low orbit vehicle in the earth-centered inertial coordinate system during navigation and positioning are (X S , Y S , Z S );

[0064] Step 2, measuring the spatial attitude angle of the medium-low orbit vehicle during navigation and positioning in the earth-centered inertial coordinate system by the star sensor, and then obtaining the attitude rotation matrix R;

[0065] Step 2.1, collecting a star image by the star sensor, extracting the measured positions of stars from the star image, and matching the measured positions of stars with the pre-stored star table to obtain the known positions of stars;

[0066] Step 2.2, calculating the spatial attitude angle (roll angle φ, pitch angle θ and yaw angle ψ) of the medium-low orbit vehicle during navigation and positioning in the earth-centered inertial coordinate system by comparing the known positions of stars with the measured positions of stars;

[0067] Step 2.3, calculating the attitude rotation matrix R according to the spatial attitude angle obtained in step 2.2 as follows:

[0068]

[0069] wherein R x (φ), R y (θ), R z (ψ) represent the rotation matrix around the x-axis, y-axis, z-axis of the medium-low orbit vehicle respectively;

[0070] Step 3, imaging at least two high-orbit satellites by a photogrammetry camera to obtain high-orbit satellite images, and the corresponding image point mass pixel coordinates of the i-th high-orbit satellite in the high-orbit satellite images are (x i , y i ); and then the image point mass pixel coordinates (x i , y i ) are extracted by a sub-pixel subdivision positioning algorithm;

[0071] Step 4, calculating the spatial coordinates (X i , Y i , Z i ) of the i-th high-orbit satellite in the Earth-Centered Inertial coordinate system at the time of navigation positioning by using the time information T0 and the programmed high-orbit satellite orbit information;

[0072] Step 4.1, obtaining six orbit parameters of each high-orbit satellite according to the programmed high-orbit satellite orbit information, including semi-major axis a, eccentricity e, inclination i0, ascending node right ascension Ω, perigee argument ω, and orbit initial mean anomaly M0;

[0073] Step 4.2, calculating the spatial coordinates (X i , Y i , Z i ) of the i-th high-orbit satellite according to the six orbit parameters of each high-orbit satellite and the time information T0;

[0074] Step 4.2.1, calculating the perigee anomaly E by solving the Kepler equation by an iterative method:

[0075]

[0076] M = E - e·sin(E)

[0077] wherein M is the current mean anomaly of the orbit, n is the average angular velocity of the high-orbit satellite along the orbit, μ is the standard gravitational parameter of the Earth, and T epoch is the epoch time;

[0078] Step 4.2.2, calculating the two-dimensional coordinates (X o , Y o ) of the i-th high-orbit satellite in the orbit plane by using the perigee anomaly E:

[0079] X o= a + (cos E - e)

[0080]

[0081] Step 4.2.3, convert the two-dimensional coordinates (X o , Y o ) of the i-th high-orbit satellite to the geocentric inertial coordinate system to obtain the spatial coordinates (X i , Y i , Z i );

[0082]

[0083] In the formula, R z , R x respectively represent the rotation matrix around the z-axis and the x-axis of the i-th high-orbit satellite;

[0084] Step 5, since the star sensor and the photogrammetric camera are fixed together, the spatial attitude of the medium-low-orbit vehicle obtained by the star sensor, that is, the self attitude of the photogrammetric camera;

[0085] Assuming that the focal length of the optical system of the photogrammetric camera is f, the pixel size of the imaging sensor chip 10 is a, and the principal point coordinates in the photogrammetric coordinate system are (c x , c y ), according to the collinearity relationship of the photogrammetric object point and the image point, the following formula can be obtained:

[0086]

[0087] In the above formula, the focal length f, the pixel size a, and the principal point (c x , c y ) of the photogrammetric camera are internal parameters and are known quantities; the spatial coordinates (X i , Y i , Z i ) of the i-th high-orbit satellite and the attitude rotation matrix R of the medium-low-orbit vehicle are known quantities; the photographic center, that is, the coordinates (X S , Y S , Z S ) of the medium-low-orbit vehicle in the spatial coordinate system, are to be solved objects, and λ is the scaling factor between the object and the image;

[0088] Let the attitude rotation matrix R be:

[0089]

[0090] Then through deformation, a linear equation group about the spatial coordinates (X S , Y S , Z S ) can be obtained:

[0091]

[0092] When two or more high-orbit satellites are simultaneously imaged as cooperative targets, the equation set is an over-determined equation set, and (X S , Y S , Z S ) can be solved by least square method to complete the spatial coordinate positioning of the medium-low orbit vehicle.

Claims

1. A method for space autonomous positioning and navigation of a medium-low earth orbit vehicle, characterized in that, The method comprises the following steps: Step 1, obtain time information T0 when navigation positioning, and assume that the space coordinates of the medium and low orbit vehicle in the earth-centered inertial coordinate system at the time of navigation positioning are (X S , Y S , Z S ) ; Step 2, measuring the space attitude angle of the medium and low orbit vehicle in the inertial coordinate system during navigation positioning by a star sensor, and then obtaining an attitude rotation matrix R; Step 3: Use a photogrammetric camera fixedly mounted on the star sensor to image at least two high-orbit satellites to obtain high-orbit satellite images, and then extract the pixel coordinates (x i ,y i ); Step 4, using the time information T0, the cataloged high-orbit satellite orbit information, the space coordinates (X i , Y i , Z i ) of the i-th high-orbit satellite in the earth-centered inertial coordinate system are calculated when the high-orbit satellite is navigated and positioned; Step 5, according to the collinear relationship of the photogrammetry object points and the image points, the following can be obtained: wherein f is the focal length of the optical system of the photogrammetric camera, a is the size of a pixel of the imaging sensor chip (10) in the photogrammetric camera, (c x , y ) is the principal point coordinate in the photogrammetric coordinate system, and λ is the scaling factor between object and image; Substitute the attitude rotation matrix R obtained in step 2 to obtain a linear equation system about the space coordinates (X S , Y S , Z S ), and then calculate the space coordinates (X S , Y S , Z S ) to complete the space coordinate positioning of the medium and low orbit vehicle.

2. The method according to claim 1, wherein, The step 2 is specifically as follows: Step 2.1, acquiring a star image by the star sensor, extracting a measured position of a star from the star image, and matching the measured position with a pre-stored star table to obtain a known position of the star; Step 2.2, calculating the space attitude angle of the medium and low orbit vehicle in the inertial coordinate system during navigation positioning by comparing the known position of the star with the measured position; Step 2.3, calculating the attitude rotation matrix R according to the space attitude angle obtained in step 2.2 as follows: R = R x (φ) · R y (θ) · R z (ψ) wherein R x (φ), R y (θ), R z (ψ) represent the rotation matrices about the x-axis, y-axis, z-axis of the medium-low earth orbit vehicle, respectively.

3. The method of claim 1, wherein: In step 3, the corresponding image point centroid pixel coordinates (x i , y i ) of the i-th high-orbit satellite in the high-orbit satellite image are extracted by a sub-pixel subdivision positioning algorithm.

4. The method according to any one of claims 1 to 3, wherein, The step 4 is specifically as follows: Step 4.1, obtaining six orbit parameters of each high orbit satellite according to the programmed high orbit satellite orbit information, including a semi-major axis a, an eccentricity e, an inclination i0, an ascending node right ascension Ω, a perigee amplitude ω, and an initial perigee argument M0; Step 4.2, calculate the spatial coordinates (X i , Y i , Z i ) of the i-th high orbit satellite according to the six orbit parameters of each high orbit satellite, time information T0.

5. The method of claim 4, wherein, The step 4.2 is specifically as follows: Step 4.2.1, calculating a perigee argument E by solving a Kepler equation through an iterative method: M = E - e·sin(E) where M is the current mean anomaly of the orbit, n is the average angular velocity of the high orbit satellite along the orbit, μ is the standard gravitational parameter of the Earth, T epoch is the epoch time; Step 4.2.

2. Calculate the 2D coordinates (X o , Y o ) of the i-th high orbit satellite in the orbital plane using the eccentric anomaly E: E = 2arctan((sqrt(1 - e2) * tan(π / 4 - M / 2)) / (1 + sqrt(1 - e2) * tan(π / 4 - M / 2))) (1) where M = E - e sin E (2) X o = a + (cos E - e) Step 4.2.3, convert the two-dimensional coordinates (X o , Y o ) of the i-th high-orbit satellite into the geocentric inertial coordinate system to obtain the spatial coordinates (X i , Y i , Z i ); wherein R z , R x respectively represent the rotation matrix around the z-axis, x-axis of the i-th high orbit satellite.

6. The method of claim 5, wherein, In step 5, the pose rotation matrix R obtained in step 2 is substituted to obtain a linear equation system about the space coordinates (X S , Y S , Z S ), and the space coordinates (X S , Y S , Z S ) are calculated in detail as follows: Let the attitude rotation matrix R be: The linear equations about the spatial coordinates (X S , Y S , Z S ) are obtained by deformation again: (X S , Y S , Z S ) are obtained by solving the least square method.

7. A medium and low orbit vehicle space autonomous positioning and navigation system for realizing the medium and low orbit vehicle space autonomous positioning and navigation method of claim 1. characterized in that The system comprises a mounting base plate (9), and a star sensor and a photogrammetry camera arranged on the mounting base plate (9), wherein an optical axis (1) of the star sensor is arranged in parallel with an optical axis (5) of the photogrammetry camera.

8. The medium and low orbit vehicle space autonomous positioning and navigation system according to claim 7, characterized in that: The star sensor comprises a first optical lens (3), a first light shield (2), and a first imaging and data processing module (4); the first optical lens (3) is used for imaging and is located on the optical axis (1) of the star sensor and in the first light shield (2); The photogrammetry camera comprises a second optical lens (7), a second light shield (6), and a second imaging and data processing module (8); the second optical lens (7) is used for imaging and is located on the optical axis (5) of the photogrammetry camera and in the second light shield (6); The first imaging and data processing module (4) and the second imaging and data processing module (8) are in communication connection; the first imaging and data processing module (4) is used for calculating the attitude rotation matrix R according to the space attitude angle of the medium and low orbit vehicle in the inertial coordinate system, and sending the space attitude angle of the medium and low orbit vehicle in the inertial coordinate system to the second imaging and data processing module (8); the second imaging and data processing module (8) is used for extracting corresponding image point mass center pixel coordinates of at least two high orbit satellites in the high orbit satellite image, and converting the image point mass center pixel coordinates into space coordinates in the inertial coordinate system, and then calculating the space coordinates of the medium and low orbit vehicle in the inertial coordinate system during navigation positioning in combination with the space attitude angle of the medium and low orbit vehicle in the inertial coordinate system.

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