An inertial astro-combination navigation device and method

Abstract

The application provides an inertial astronomical combined navigation device and method, which comprises an astronomical star measurement system and an inertial navigation system; a local reference inertial measurement assembly is installed on a measurement reference of the astronomical star measurement system; the local reference inertial measurement assembly is used for receiving position and speed reference information of the inertial navigation system and providing an angle measurement reference in a geographical system for the astronomical star measurement system. The local reference inertial measurement assembly is installed on the measurement reference of the astronomical star measurement system to realize local reference position and posture measurement. The inertial navigation system and the local reference inertial measurement assembly are combined through real-time relative position and speed measurement to improve the posture measurement precision of the local reference inertial measurement assembly, and then the astronomical combination of the local reference inertial measurement assembly and the astronomical star measurement system is realized, and the inertial navigation system is corrected and fed back, so that the effect of the astronomical combination of the inertial navigation system is indirectly achieved.

Description

Technical Field

[0001] This invention relates to the technical field of navigation systems, and in particular to an inertial astronomical navigation device and method. Background Technology

[0002] The inertial navigation system and the astronomical satellite measurement system together constitute an inertial-astronomical integrated navigation system. Generally, the high-precision angle measurement information of the astronomical satellite measurement system, which does not drift over time, is used to correct the sensor errors of the inertial navigation system. This can greatly improve the accuracy of the navigation system and meet the requirements of long-endurance navigation.

[0003] Since the principle is based on angle measurement correction, inconsistencies in the measurement angle references between the two navigation systems will directly affect the accuracy of the integrated navigation. Therefore, the inertial navigation system and the astronomical satellite measurement system are usually installed on a shared base, requiring necessary installation error calibration between the two systems to ensure consistency in their measurement references. However, the constraints of installing the inertial navigation system and the astronomical satellite measurement system on a shared base severely restrict the deployment and installation of the product on the carrier platform, which is detrimental to the product's application. Summary of the Invention

[0004] The purpose of this invention is to propose an inertial astronomical integrated navigation device and method to realize the distributed installation of astronomical star measurement system and inertial navigation system at different locations on the carrier.

[0005] The technical solution of the present invention: According to a first aspect of the present invention, an inertial astronomical integrated navigation device is provided, comprising an astronomical star measurement system and an inertial navigation system; a local reference inertial measurement component is installed on the measurement reference of the astronomical star measurement system; the local reference inertial measurement component is used to receive the position and velocity reference information of the inertial navigation system and provide the astronomical star measurement system with a geographic angle measurement reference.

[0006] Inertial-astronomical integrated navigation is a method of correction based on angle measurements. To ensure the consistency of angle measurement references, traditional inertial-astronomical integrated navigation devices require the astronomical angle measurement system and the inertial navigation system to share a common mounting base. However, in practical engineering applications, this rigid mounting condition is often difficult to meet, significantly limiting system layout. Especially on aircraft platforms equipped with redundant inertial navigation systems, the astronomical satellite measurement system needs to be effectively integrated with multiple inertial navigation systems. In this case, the constraint of shared mounting base further increases the complexity of system integration, raises the difficulty of installation and commissioning, and thus affects its widespread application and promotion in engineering.

[0007] This invention integrates a local reference inertial measurement unit (INS) into an astronomical satellite observation system and incorporates position and velocity information from an inertial navigation system mounted on a non-co-platform aircraft. The attitude measurement results of this local reference INS are then corrected in real time, effectively improving its attitude measurement accuracy. This method provides a high-precision pointing attitude reference for the astronomical satellite observation system. By relying solely on an information transmission mechanism, reference transfer and unification between the astronomical satellite observation system and the airborne inertial navigation system are achieved, eliminating the need for a rigid mechanical base connection. This technical solution significantly reduces the rigid constraints on the installation location and structural layout of the inertial navigation equipment and the astronomical satellite observation system, enhancing the system's deployment flexibility in practical engineering applications.

[0008] In one possible embodiment, the gyroscope accuracy of the local reference inertial measurement unit must be better than 2° / hr (1σ), and the accelerometer accuracy must be better than 2000ug (1σ).

[0009] According to a second aspect of the present invention, an inertial astronomical integrated navigation method is proposed, employing the aforementioned inertial astronomical integrated navigation device, comprising the following steps: Step 1: The inertial navigation system and the local reference inertial measurement unit perform position, velocity, and attitude calculations respectively; Step 2: Design the integrated navigation Kalman filter. The filter states include the error states of the inertial navigation system, the error states of the local reference inertial measurement unit, and the error states of the astronomical satellite measurement system. Step 3: Use the position and velocity calculated by the inertial navigation system and the relative position and relative velocity calculated by the local reference inertial measurement unit as measurement 1 of the Kalman filter; at the same time, use the altitude difference and azimuth difference of the astronomical star measurement system as measurement 2. Step 4: Combine the initialization parameters of the Kalman filter with Measurement 1 and Measurement 2 to perform error estimation calculation through the Kalman filter.

[0010] In one possible embodiment, in step 2, the estimated state variables of the Kalman filter are specifically: in: This refers to the longitude error of the inertial navigation system. This refers to the latitude error of the inertial navigation system. For the eastward velocity error of the inertial navigation system; This refers to the northbound velocity error of the inertial navigation system. The eastward platform deflection angle of the inertial navigation system; The northward platform deflection angle of the inertial navigation system; The tilt angle of the inertial navigation system towards the platform; For X-axis gyroscope drift in inertial navigation systems; For Y-axis gyroscope drift in inertial navigation systems; For Z-axis gyroscope drift in inertial navigation systems; Zero bias for the X-axis accelerometer of the inertial navigation system; Zero bias for the Y-axis accelerometer of the inertial navigation system; Zero bias for the Z-axis accelerometer of the inertial navigation system; This refers to the longitude error of the local reference inertial measurement unit; This refers to the latitude error of the local reference inertial measurement unit; The eastward velocity error of the local reference inertial measurement unit; The northward velocity error of the local reference inertial measurement unit; The eastward platform deflection angle of the local reference inertial measurement unit; The northward platform deflection angle of the local reference inertial measurement unit; The tilt angle of the azimuth platform for the local reference inertial measurement unit; X-axis gyroscope drift for the local reference inertial measurement unit; For the Y-axis gyroscope drift of the local reference inertial measurement unit; Z-axis gyroscope drift for the local reference inertial measurement unit; The X-axis accelerometer of the local reference inertial measurement unit has zero bias; The Y-axis accelerometer of the local reference inertial measurement unit has zero bias; The Z-axis accelerometer of the local reference inertial measurement unit has zero bias; This refers to the azimuth axis system error of the astronomical star measurement system; This refers to the elevation axis error of the astronomical astronomical measurement system.

[0011] In one possible embodiment, in step 3, measurement 1 is calculated according to the following formula:

[0012] in: , The longitudes calculated for the inertial navigation system and the local reference inertial measurement unit, respectively; , The latitudes are calculated by the inertial navigation system and the local reference inertial measurement unit, respectively. , The eastward velocities calculated by the inertial navigation system and the local reference inertial measurement unit are respectively. , The northbound velocities are calculated by the inertial navigation system and the local reference inertial measurement unit, respectively.

[0013] In one possible embodiment, the Kalman filter observation matrix corresponding to measurement 1 is: .

[0014] In one possible embodiment, in step 3, measurement 2 is calculated according to the following process: Calculate the pointing vector of the measured celestial body:

[0015] in:

[0016] The altitude angle of the measured celestial body. The horizontal azimuth of the celestial body being measured; Attitude matrix calculated for the local reference inertial measurement unit: ; , , These are the pitch, roll, and heading measurements of the local reference inertial measurement unit, respectively. Calculate the measured pointing vector

[0017] in: , The azimuth and elevation angles of the measured celestial body in the reference coordinate system of the astronomical star measurement system; Calculation measurement

[0018] in: , This refers to the lateral and longitudinal miss distances of the measured celestial body on the imaging target surface of the astronomical astrometry system.

[0019] In one possible embodiment, the Kalman filter observation matrix corresponding to measurement 2 is: in:

[0020]

[0021]

[0022]

[0023]

[0024]

[0025] .

[0026] In one possible embodiment, in step 4, the error parameters of the inertial navigation system are calculated using a Kalman filter, and the combined navigation result is obtained by compensating for them according to the following formula;

[0027]

[0028]

[0029] .

[0030] The beneficial effects of this invention are as follows: 1. By adding a local reference inertial measurement unit to the measurement reference of the astronomical star-measuring system, local reference attitude measurement is achieved. The inertial navigation system and the local reference inertial measurement unit improve the attitude measurement accuracy of the local reference inertial measurement unit through real-time relative position velocity measurement and combination, thereby achieving astronomical combination of the local reference inertial measurement unit and the astronomical star-measuring system. Feedback is then used to correct the inertial navigation system, indirectly achieving the effect of astronomical combination with the inertial navigation system. 2. This invention eliminates the need for physical co-mounting of the astronomical star-measuring system and the inertial navigation system to transfer attitude references. Instead, it uses relative position and relative velocity information transmission, achieved through combined navigation processing, transforming physical transmission into information transmission. This eliminates the need to calibrate the installation errors of the inertial navigation system and the astronomical star-measuring system. It also facilitates flexible layout of the inertial navigation system and the astronomical star-measuring system on the carrier. Attached Figure Description

[0031] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. The drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0032] Figure 1 This is a schematic diagram of a preferred embodiment of the present invention. Detailed Implementation

[0033] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0034] The features and illustrative embodiments of various aspects of the present invention will now be described in detail. Numerous specific details are set forth in the following detailed description to provide a thorough understanding of the invention. However, it will be apparent to those skilled in the art that the invention may be practiced without requiring some of these specific details. The following description of embodiments is merely intended to provide a better understanding of the invention by illustrating examples of the invention. The invention is by no means limited to any specific setups and methods set forth below, but covers any improvements, substitutions, and modifications to structures, methods, and devices without departing from the spirit of the invention. Well-known structures and techniques are not shown in the drawings and the following description to avoid unnecessarily obscuring the invention.

[0035] In the description of this invention, it should be noted that the directions or positional relationships indicated by terms such as "center," "upper," "lower," "left," "right," "vertical," "horizontal," "inner," and "outer" are based on the directions or positional relationships shown in the accompanying drawings and are only for the convenience of describing and simplifying the invention, and should not be construed as limiting the invention. Furthermore, the use of ordinal numbers (e.g., "first and second," etc.) is for distinguishing objects and is not limited to this order, and should not be construed as indicating or implying relative importance.

[0036] In the description of this invention, it should be noted that, unless otherwise explicitly specified and limited, the terms "installation," "connection," and "linking" should be interpreted broadly, encompassing both direct connection and indirect connection via an intermediate medium. Those skilled in the art can understand the specific meaning of these terms in this invention based on the specific circumstances.

[0037] It should be noted that, unless otherwise specified, the embodiments of the present invention and the features thereof can be combined with each other, and the various embodiments can be referenced and cited in each other. The present invention will now be described in detail with reference to the accompanying drawings and embodiments.

[0038] The present invention will be further described in detail below with reference to the embodiments and accompanying drawings, but the embodiments of the present invention are not limited thereto.

[0039] See implementation plan Figure 1 1. Install a local reference inertial measurement unit (gyroscope accuracy not exceeding 1° / hr (1σ), accelerometer accuracy not exceeding 1000ug (1σ)) on the measurement reference of the astronomical star measurement system, and configure an inertial navigation system at the same time; in this embodiment, the initial longitude of the two systems is set to 108.82835°, the initial latitude is set to 34.19834°, the initial altitude is set to 400m, and the initial velocity is set to 0. Based on this, position, velocity and attitude calculations are started and the system enters the star measurement state; 2. Design a Kalman filter for integrated navigation. The filter status should include the error status of the inertial navigation system, the error status of the local reference inertial measurement unit, and the error status of the astronomical satellite measurement system.

[0040] in: This refers to the longitude error of the inertial navigation system. This refers to the latitude error of the inertial navigation system. For the eastward velocity error of the inertial navigation system; This refers to the northbound velocity error of the inertial navigation system. The eastward platform deflection angle of the inertial navigation system; The northward platform deflection angle of the inertial navigation system; The tilt angle of the inertial navigation system towards the platform; For X-axis gyroscope drift in inertial navigation systems; For Y-axis gyroscope drift in inertial navigation systems; For Z-axis gyroscope drift in inertial navigation systems; Zero bias for the X-axis accelerometer of the inertial navigation system; Zero bias for the Y-axis accelerometer of the inertial navigation system; Zero bias for the Z-axis accelerometer of the inertial navigation system; This refers to the longitude error of the local reference inertial measurement unit; This refers to the latitude error of the local reference inertial measurement unit; The eastward velocity error of the local reference inertial measurement unit; The northward velocity error of the local reference inertial measurement unit; The eastward platform deflection angle of the local reference inertial measurement unit; The northward platform deflection angle of the local reference inertial measurement unit; The tilt angle of the azimuth platform for the local reference inertial measurement unit; X-axis gyroscope drift for the local reference inertial measurement unit; For the Y-axis gyroscope drift of the local reference inertial measurement unit; Z-axis gyroscope drift for the local reference inertial measurement unit; The X-axis accelerometer of the local reference inertial measurement unit has zero bias; The Y-axis accelerometer of the local reference inertial measurement unit has zero bias; The Z-axis accelerometer of the local reference inertial measurement unit has zero bias; This refers to the azimuth axis system error of the astronomical star measurement system; This refers to the elevation axis system error of the astronomical astronomical measurement system. The initial value of the filter state is initialized to 0.

[0041] In this embodiment, the variance matrix of the state estimation error of the integrated navigation Kalman filter is... It should include the variance of the inertial navigation system estimation error. Variance of estimation error for local reference inertial measurement unit , Estimation error variance of astronomical star measurement system ;

[0042] in: To estimate the error variance of the inertial navigation system, the off-diagonal elements are initialized to 0, and the initial values ​​of the diagonal elements are set as follows: {0, 0, 0, 0, 9.4e-11, 9.4e-11, 9.4e-11, 2.11e-16, 2.11e-16,2.11e-16, 8.6e-8, 8.6e-8, 8.6e-8}; To estimate the error variance of the local reference inertial measurement unit, the off-diagonal elements are initialized to 0, and the initial values ​​of the diagonal elements are set as follows: {2.45e-12, 2.45e-12, 0.01, 0.01, 1.2e-3, 3.0e-2,5.876e-14, 5.876e-14, 5.876e-14, 3.82e-6, 3.82e-6, 3.82e-6}; To estimate the variance of the astronomical stellar measurement system, the off-diagonal elements are initialized to 0, and the initial values ​​of the diagonal elements are set as follows: {0.0, 7.62e-7}.

[0043] In this embodiment, the noise variance matrix of the integrated navigation Kalman filter system It should include the inertial navigation system noise variance. Local reference inertial measurement unit noise variance astronomical star-measuring system noise variance ;

[0044] in: For the noise variance of the inertial navigation system, the off-diagonal elements are initialized to 0, and the initial values ​​of the diagonal elements are set as follows: {0, 0, 9.56e-9, 9.56e-9, 8.46e-14, 8.46e-14, 8.46e-14, 0, 0, 0,0, 0, 0}; To determine the noise variance of the local reference inertial measurement unit system, off-diagonal elements are initialized to 0, and the initial values ​​of diagonal elements are set as follows: {0, 0, 1.17e-7, 1.17e-7, 8.46e-12, 8.46e-12,8.46e-12, 0, 0, 0, 0, 0, 0}; To represent the noise variance of the astronomical stellar measurement system, off-diagonal elements are initialized to 0, and diagonal elements are initialized to the following values: {0, 0}.

[0045] In this embodiment, the integrated navigation Kalman filter measures the noise variance matrix. It should include noise from relative position and relative velocity measurements. Measurement noise of astronomical star-measuring systems ; To reduce noise in relative velocity and position measurements, off-diagonal elements are initialized to 0, and diagonal elements are initialized to the following values: {3.54e-12, 3.54e-12, 0.25, 0.25}; To measure noise in the astronomical stellar measurement system, off-diagonal elements are initialized to 0, and diagonal elements are initialized to the following values: {2.35e-9, 2.35e-9}.

[0046] In this embodiment, the state transition matrix of the integrated navigation Kalman filter system It should include the state transition matrix of the inertial navigation system. State transition matrix of local reference inertial measurement unit State transition matrix of astronomical star measurement system ;

[0047] in: and This is the system state transition matrix for the inertial navigation system and the local reference inertial measurement unit, which is the same as the standard strapdown inertial navigation system state transition matrix; Let be the system state transition matrix of the astronomical star measurement system, which is a second-order zero matrix.

[0048] 3. The relative quantities (relative position and relative velocity) of the position and velocity calculated by the inertial navigation system and the position and velocity calculated by the local reference inertial measurement component in the astronomical satellite measurement system are used as measurement 1 of the Kalman filter; at the same time, the altitude difference and azimuth difference of the astronomical satellite measurement system are used as measurement 2. (1) Calculation method for measurement 1:

[0049] in: , The longitudes calculated for the inertial navigation system and the local reference inertial measurement unit, respectively; , The latitudes are calculated by the inertial navigation system and the local reference inertial measurement unit, respectively. , The eastward velocities calculated by the inertial navigation system and the local reference inertial measurement unit are respectively. , The northbound velocities calculated by the inertial navigation system and the local reference inertial measurement unit are respectively. The Kalman filter observation matrix corresponding to measurement 1 is:

[0050] (2) Calculation method for measurement 2: i. Calculate the pointing vector of the measured celestial body.

[0051] in:

[0052] The altitude angle of the measured celestial body. The horizontal azimuth of the celestial body being measured; Attitude matrix calculated for the local reference inertial measurement unit:

[0053] , , These are the pitch, roll, and heading measurements of the local reference inertial measurement unit, respectively.

[0054] ii. Calculate the measured pointing vector

[0055] in: , The azimuth and elevation angles of the measured celestial body in the reference coordinate system of the astronomical star measurement system.

[0056] iii. Calculation and measurement

[0057] in: , This refers to the lateral and longitudinal miss distances of the measured celestial body on the imaging target surface of the astronomical astrometry system.

[0058] The Kalman filter observation matrix corresponding to measurement 2 is:

[0059] in:

[0060]

[0061]

[0062]

[0063]

[0064]

[0065]

[0066] 4. Error estimation calculation The error parameters of the inertial navigation system can be obtained by calculating the Kalman filter, and the combined navigation result can be obtained by compensating for the error parameters.

[0067]

[0068]

[0069]

[0070] .

[0071] The above detailed embodiments are a description of the present invention. It should not be considered that the specific embodiments of the present invention are limited to these descriptions. For those skilled in the art, several simple deductions and substitutions can be made without departing from the concept of the present invention, and all of these should be considered to fall within the protection scope of the present invention.

Claims

1. An inertial astronomical navigation device, characterized in that, It includes an astronomical satellite measurement system and an inertial navigation system; a local reference inertial measurement component is installed on the measurement reference of the astronomical satellite measurement system; the local reference inertial measurement component is used to receive the position and velocity reference information of the inertial navigation system and provide the astronomical satellite measurement system with a geographic angle measurement reference.

2. The inertial astronomical navigation device according to claim 1, characterized in that, The gyroscope accuracy of the local reference inertial measurement unit must be better than 2° / hr (1σ), and the accelerometer accuracy must be better than 2000ug (1σ).

3. An inertial astronomical integrated navigation method, characterized in that, The inertial astronomical navigation device according to any one of claims 1-2 includes the following steps: Step 1: The inertial navigation system and the local reference inertial measurement unit perform position, velocity, and attitude calculations respectively; Step 2: Design the integrated navigation Kalman filter. The filter states include the error states of the inertial navigation system, the error states of the local reference inertial measurement unit, and the error states of the astronomical satellite measurement system. Step 3: Use the position and velocity calculated by the inertial navigation system and the relative position and relative velocity calculated by the local reference inertial measurement unit as measurement 1 of the Kalman filter; at the same time, use the altitude difference and azimuth difference of the astronomical star measurement system as measurement 2. Step 4: Combine the initialization parameters of the Kalman filter with Measurement 1 and Measurement 2 to perform error estimation calculation through the Kalman filter.

4. The inertial astronomical integrated navigation method according to claim 3, characterized in that, In step 2, the estimated state variables of the Kalman filter are specifically as follows: in: This refers to the longitude error of the inertial navigation system. This refers to the latitude error of the inertial navigation system. For the eastward velocity error of the inertial navigation system; This refers to the northbound velocity error of the inertial navigation system. The eastward platform deflection angle of the inertial navigation system; The northward platform deflection angle of the inertial navigation system; The tilt angle of the inertial navigation system towards the platform; For X-axis gyroscope drift in inertial navigation systems; For Y-axis gyroscope drift in inertial navigation systems; For Z-axis gyroscope drift in inertial navigation systems; Zero bias for the X-axis accelerometer of the inertial navigation system; Zero bias for the Y-axis accelerometer of the inertial navigation system; Zero bias for the Z-axis accelerometer of the inertial navigation system; This refers to the longitude error of the local reference inertial measurement unit; This refers to the latitude error of the local reference inertial measurement unit; The eastward velocity error of the local reference inertial measurement unit; The northward velocity error of the local reference inertial measurement unit; The eastward platform deflection angle of the local reference inertial measurement unit; The northward platform deflection angle of the local reference inertial measurement unit; The tilt angle of the azimuth platform for the local reference inertial measurement unit; X-axis gyroscope drift for the local reference inertial measurement unit; For the Y-axis gyroscope drift of the local reference inertial measurement unit; Z-axis gyroscope drift for the local reference inertial measurement unit; The X-axis accelerometer of the local reference inertial measurement unit has zero bias; The Y-axis accelerometer of the local reference inertial measurement unit has zero bias; The Z-axis accelerometer of the local reference inertial measurement unit has zero bias; This refers to the azimuth axis system error of the astronomical star-measuring system; This refers to the elevation axis error of the astronomical astronomical measurement system.

5. The inertial astronomical integrated navigation method according to claim 3, characterized in that, In step 3, measurement 1 is calculated according to the following formula: in: , The longitudes calculated for the inertial navigation system and the local reference inertial measurement unit, respectively; , The latitudes are calculated by the inertial navigation system and the local reference inertial measurement unit, respectively. , The eastward velocities calculated by the inertial navigation system and the local reference inertial measurement unit are respectively. , The northbound velocities are calculated by the inertial navigation system and the local reference inertial measurement unit, respectively.

6. The inertial astronomical integrated navigation method according to claim 3, characterized in that, The Kalman filter observation matrix corresponding to measurement 1 is: 。 7. The inertial astronomical integrated navigation method according to claim 3, characterized in that, In step 3, measurement 2 is calculated according to the following process: Calculate the pointing vector of the measured celestial body: in: The altitude angle of the measured celestial body. The horizontal azimuth of the celestial body being measured; Attitude matrix calculated for the local reference inertial measurement unit: ； , , These are the pitch, roll, and heading measurements of the local reference inertial measurement unit, respectively. Calculate the measured pointing vector in: , The azimuth and elevation angles of the measured celestial body in the reference coordinate system of the astronomical star measurement system; Calculation measurement in: , This refers to the lateral and longitudinal miss distances of the measured celestial body on the imaging target surface of the astronomical astrometry system.

8. The inertial astronomical integrated navigation method according to claim 3, characterized in that, The Kalman filter observation matrix corresponding to measurement 2 is: in: 。 9. The inertial astronomical integrated navigation method according to claim 3, characterized in that, In step 4, the error parameters of the inertial navigation system are calculated using a Kalman filter, and the combined navigation result is obtained by compensating for them according to the following formula. 。

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