Spacecraft integrated navigation system and method based on INS / GNSS / XPNAV

Abstract

The invention provides a spacecraft integrated navigation system and method based on an inertial navigation system (INS), a global navigation satellite system (GNSS) and an X-ray pulsar navigation system (XPNAV), aiming at the navigation requirement of a spacecraft, and the spacecraft integrated navigation system comprises an INS, a global navigation satellite system (GNSS) and an X-ray pulsar navigation system (XPNAV). A federated filtering structure based on double-layer filtering is used, navigation data of three navigation systems are fused, multi-source information fusion is achieved, compared with a single navigation mode, the navigation precision is improved, the overall navigation stability is enhanced, compared with an existing INS / XPNAV integrated navigation method, GNSS navigation information is fused, and the navigation efficiency is improved. In the data fusion method, double-layer filtering is used for replacing standard Kalman filtering for sub-filtering in federated filtering, so that the robustness of the integrated navigation system in an unstable environment is enhanced.

Description

Technical Field

[0001] The present invention belongs to the field of deep space navigation technology, and relates to a spacecraft integrated navigation system and method, and specifically to a spacecraft integrated navigation system and method based on INS / GNSS / XPNAV. Background Art

[0002] There are a large number of orbital resources at the Earth-Moon libration point, which can provide ideal orbital positions for spacecraft. Spacecraft in this scenario can be used for deep space exploration, lunar resource development, and astrophysical research. However, carrying out space missions requires navigation of spacecraft to determine their position relative to Earth.

[0003] Spacecraft navigation can be categorized as autonomous and non-autonomous. Autonomous navigation independently outputs the spacecraft's position without relying on external information, while inertial navigation systems (INS) can independently output the spacecraft's position and are therefore often used as a navigation method. INS provides all navigation parameters relative to a specific reference coordinate system. By automatically integrating the INS output, the spacecraft's velocity and position data are obtained. However, this method suffers from the accumulation of errors over time, necessitating the use of other navigation technologies to correct for these errors in deep space. The existing typical spacecraft autonomous navigation is the INS / XPNAV combined navigation method, which overcomes the defects of a single inertial navigation INS. For example, the patent application with application publication number CN117191050A, entitled "Spacecraft Navigation Method Based on XPNAV and INS Combination", discloses a spacecraft combined navigation system and method based on inertial navigation INS and X-ray pulsar navigation XPNAV. The invention calculates the pulse arrival time difference between the pulsar observation pulse profile and the standard pulse profile in two observation time periods respectively, and calculates the pulse arrival time difference between the pulsar observation pulse profile and the standard pulse profile in the two observation times through the inertial navigation system, and then solves the navigation observation equation to obtain the position and velocity of the spacecraft, thereby effectively improving the navigation accuracy. However, since the photon arrival time measured by XPNAV may have errors, the further improvement of navigation accuracy is affected. Summary of the Invention

[0004] The purpose of the present invention is to address the above-mentioned deficiencies in the prior art and to propose a spacecraft navigation method based on an INS / GNSS / XPNAV combination to solve the technical problem of low navigation accuracy in the prior art.

[0005] To achieve the above object, the technical solution adopted by the present invention is:

[0006] An INS / GNSS / XPNAV-based spacecraft integrated navigation system includes an inertial navigation system INS and an X-ray pulsar navigation system XPNAV carried on a spacecraft in a Halo orbit of the Earth-Moon libration point, and The Global Navigation Satellite System (GNSS) is composed of 100 satellites. .

[0007] A spacecraft integrated navigation method based on INS / GNSS / XPNAV includes the following steps:

[0008] (1) Initialization parameters:

[0009] Initialize the spacecraft in the body coordinate system measured by the inertial navigation system INS Next The angular increment of the spacecraft at time , Included in the body coordinate system The angle increments of the pitch, roll and yaw angles of the spacecraft 、 and ,in , represents the time when the spacecraft navigation solution is performed; the non-gravitational acceleration is , the length of the adjacent time interval is , the Earth's inertial coordinate system measured by the Global Navigation Satellite System GNSS Next The pseudo-range between the spacecraft and each satellite at time , Earth's inertial coordinate system measured by the X-ray pulsar navigation system XPNAV Next The distance between the spacecraft and the spacecraft in the direction of each pulsar unit vector at the moment is , , represents the number of pulsars, 3:

[0010] (2) Calculate the predicted position error of the spacecraft at each moment in the Earth's inertial coordinate system:

[0011] By angle increment Constructing the Earth's inertial coordinate system Next The direction cosine matrix of the spacecraft at time , and through non-gravitational acceleration and Calculating the Earth's inertial coordinate system Next Real-time acceleration of the spacecraft at this moment , then by and Calculating the Earth's inertial coordinate system Next The predicted position error of the spacecraft at time ;

[0012] (3) Calculate the observed position error of the spacecraft relative to the predicted position of the spacecraft in the Earth's inertial coordinate system based on GNSS:

[0013] Through the pseudo-range between the spacecraft and each satellite Calculating the Earth's inertial coordinate system Next Observation position of the spacecraft at this moment and through Calculating the Earth's inertial coordinate system Next Spacecraft predicted position at this moment The observed position error ;

[0014] (4) Calculate the observed position error of the spacecraft relative to the predicted position of the spacecraft in the Earth's inertial coordinate system based on XPNAV:

[0015] The distance between the spacecraft and the spacecraft through the unit vector in the direction of each pulsar Calculating the Earth's inertial coordinate system Next Observation position of the spacecraft at this moment and through Calculate the Earth's inertial coordinate system Next Spacecraft predicted position at this moment The observed position error :

[0016] (5) Obtain the current position and velocity of the spacecraft in the Earth's inertial coordinate system:

[0017] Fusion of Earth Inertial Coordinate System via Federated Filter Structure Next The predicted position error of the spacecraft at time , spacecraft predicted position calculated based on GNSS and XPNAV The observed position error 、 Calculate the spacecraft's position in the Earth's inertial coordinate system Next The a posteriori estimated position error of the spacecraft at time and speed error , and then through and Calculating the Earth's inertial coordinate system Next Spacecraft position at time and speed .

[0018] Compared with the prior art, the present invention has the following advantages:

[0019] 1. The integrated navigation system designed in this invention for deep space scenarios uses an inertial navigation system. GNSS navigation is added to the INS / XPNAV integrated navigation. Compared with the former, it can integrate GNSS navigation information and calculate the navigation accuracy of the spacecraft with higher accuracy.

[0020] 2. The present invention uses a double-layer filter structure in the federated filter structure instead of the standard Kalman filter, which can improve the overall navigation accuracy and the accuracy of the navigation results by an order of magnitude, making the integrated navigation system more robust. BRIEF DESCRIPTION OF THE DRAWINGS

[0021] Figure 1 This is a flowchart for implementing the navigation method of the present invention.

[0022] Figure 2 The figure is a comparison chart of the simulation results of the position error of the present invention and the prior art. DETAILED DESCRIPTION

[0023] The present invention is described in further detail below with reference to the accompanying drawings and specific embodiments.

[0024] The navigation system of the present invention comprises an inertial navigation system INS carried on a spacecraft on the Halo orbit of the Earth-Moon libration point and an inertial navigation system INS The Global Navigation Satellite System (GNSS) is composed of 100 satellites. , also includes the X-ray pulsar navigation system XPNAV. In this example, navigation is performed in the Earth-Moon L1 point scenario, and the Halo orbit is mathematically modeled. At the same time, GNSS satellite data is obtained from the IGS Data Center of Wuhan University. At the same time, when fusing the information of the three navigation systems, a federal filtering structure based on double-layer filtering is used, and a fault detection algorithm based on anomaly ratio is also added. The X-ray pulsar navigation system XPNAV includes pulsars, .

[0025] Reference Figure 1 The navigation method of the present invention comprises the following steps:

[0026] Step 1) Initialize parameters:

[0027] Initialize the spacecraft in the body coordinate system measured by the inertial navigation system INS Next The angular increment of the spacecraft at time , Included in the body coordinate system The angle increments of the pitch, roll and yaw angles of the spacecraft 、 and ,in , represents the time when the spacecraft navigation solution is performed; the non-gravitational acceleration is , the length of the adjacent time interval is , the Earth's inertial coordinate system measured by the Global Navigation Satellite System GNSS Next The pseudo-range between the spacecraft and each satellite at time , Earth's inertial coordinate system measured by the X-ray pulsar navigation system XPNAV Next The distance between the spacecraft and the spacecraft in the direction of each pulsar unit vector at the moment is , , represents the number of pulsars, 3:

[0028] Step 2) Construct the direction cosine matrix at each moment in the spacecraft body coordinate system relative to the Earth's inertial coordinate system. The expression is as follows:

[0029]

[0030]

[0031]

[0032]

[0033] in, 、 Represent the Earth's inertial coordinate system Next Direction cosine matrix of the spacecraft at this moment, body coordinate system Next Time relative to the The direction cosine matrix of the spacecraft at time t, represents the 3rd-order identity matrix, Indicates angle increment The equivalent rotation vector of express The antisymmetric matrix of .

[0034] Step 3) Calculate the predicted position error of the spacecraft at each moment in the Earth's inertial coordinate system:

[0035] By angle increment Constructing the Earth's inertial coordinate system Next The direction cosine matrix of the spacecraft at time , and through non-gravitational acceleration and Calculating the Earth's inertial coordinate system Next Real-time acceleration of the spacecraft at this moment , then by and Calculating the Earth's inertial coordinate system Next The predicted position error of the spacecraft at time , the calculation formula is as follows

[0036]

[0037]

[0038]

[0039]

[0040] in, 、 Represent the Earth's inertial coordinate system Next The gravitational acceleration and non-gravitational acceleration of the spacecraft at this moment, Represents the Earth's inertial coordinate system Next The predicted velocity error of the spacecraft at time and denote the gravitational perturbations of the Earth and the Moon on the spacecraft, It represents the attitude error of the pitch angle, deflection angle and yaw angle of the spacecraft in the body coordinate system, express The initial bias of:

[0041] Step 4) Calculate the observed position error of the spacecraft relative to the predicted position of the spacecraft in the Earth inertial coordinate system based on GNSS:

[0042] Through the pseudo-range between the spacecraft and each satellite Calculating the Earth's inertial coordinate system Next Observation position of the spacecraft at this moment and through Calculating the Earth's inertial coordinate system Next Spacecraft predicted position at this moment The observed position error , the calculation process is as follows;

[0043]

[0044]

[0045]

[0046]

[0047] in, Indicates the Earth's inertial coordinate system The predicted spacecraft speed at time, Indicates the Earth's inertial coordinate system The predicted position of the spacecraft at time is the speed of light, Indicates that the GNSS signal is connected to the The deviation of the signal received by the receiver at that moment, express The measurement noise, Expressed in the Earth's inertial coordinate system Here are the positions of each satellite:

[0048] Step 5) Calculate the observed position error of the spacecraft relative to the predicted position of the spacecraft in the Earth inertial coordinate system based on XPNAV:

[0049] The distance between the spacecraft and the spacecraft through the unit vector in the direction of each pulsar Calculating the Earth's inertial coordinate system Next Observation position of the spacecraft at this moment and through Calculate the Earth's inertial coordinate system Next Spacecraft predicted position at this moment The observed position error , the calculation formula is as follows:

[0050]

[0051] =

[0052] in, Indicates the Earth's inertial coordinate system The unit vector of the spacecraft in the direction of each pulsar at time:

[0053] Step 6) Construct a federated filtering structure based on double-layer filtering:

[0054] First, the fusion is performed through federated filtering, which contains two sub-filters. The sub-filters are double-layer filtering algorithms that combine adaptive anti-error filtering and robust Kalman filtering. The two filters are used in different situations. When the observation information is valid, the adaptive anti-error filtering is used, and when the prediction information is valid, the robust Kalman filtering is used, making the combined navigation more robust.

[0055] 6a) The adaptive robust filtering algorithm process is as follows:

[0056] Predicted state vector:

[0057]

[0058] Covariance matrix of the one-step-forecast state:

[0059]

[0060] The filter gain is:

[0061]

[0062] Robustness Factor The construction method is as follows:

[0063]

[0064]

[0065]

[0066] in Represents the prediction residual:

[0067] The state estimate is:

[0068]

[0069] The mean square error is estimated as:

[0070]

[0071] in, is the system equation matrix, For the The result at this moment is the observation equation matrix ,

[0072] 6b) The robust Kalman filter algorithm process is:

[0073] The robust Kalman filter algorithm differs from the adaptive robustness algorithm in two ways:

[0074]

[0075] Robust construction factor The construction is as follows:

[0076]

[0077]

[0078] 6c) The results of the double-layer filtering algorithm structure are as follows:

[0079]

[0080]

[0081]

[0082]

[0083] in Indicates Always move forward A moment, and Take 30, in this example Taking 0.95, the double-layer filtering structure is:

[0084]

[0085] is the threshold value, The judgment factor is used to determine whether to use robust Kalman filtering or adaptive anti-error filtering:

[0086] Step 7) Fusion of the Earth Inertial Coordinate System via a Federated Filtering Structure Based on Double-Layer Filtering Next The predicted position error of the spacecraft at time , spacecraft predicted position calculated based on GNSS and XPNAV The observed position error 、 ,Will As the predicted value, 、 As the observation value of the two sub-filters, calculate the spacecraft in the Earth's inertial coordinate system Next Spacecraft position error at time and speed error ,

[0087] Step 8) Get the current position and velocity of the spacecraft in the Earth's inertial coordinate system:

[0088] pass and Calculating the Earth's inertial coordinate system Next Spacecraft position at time ,speed :

[0089]

[0090]

[0091] in, 、 Represents the spacecraft in the Earth's inertial coordinate system Next The position error and velocity error of the spacecraft at this moment:

[0092] The following is an explanation of the technical effects of the present invention in conjunction with simulation experiments:

[0093] 1Simulation conditions and contents:

[0094] The simulation hardware is a microcomputer with the following parameters: CPU: Intel(R) Core(TM) i5-8265U CPU @1.60GHz ~1.80GHz; RAM: 8.00GB; operating system: Windows 11. The simulation software is: Computer software MATALB2021b.

[0095] In this embodiment, the relevant parameters of the inertial navigation system used are shown in Table 1:

[0096] Table 1

[0097] Gyro bias (deg / h) Accelerometer bias (ug) Gyroscope measurement noise (deg / sqrt(h)) Accelerometer measurement noise (ug / sqrt(Hz) data 0.03 100 0.001 1

[0098] In this embodiment, the data of the GNSS navigation system used is from the IGS Data Center of Wuhan University: the pulsar photon arrival accuracy is 5 .

[0099] The position error of the present invention is compared with that of the existing spacecraft navigation method based on XPNAV and INS combination. The results are as follows: Figure 2 shown.

[0100] 2 Simulation results analysis:

[0101] Reference Figure 2 , Figure 2 (a) is the result of the comprehensive position error of the existing technology. The horizontal axis is the total simulation time of 20000s, and the vertical axis is the comprehensive position error in meters. The convergence result is maintained at m to m, and it can be seen that the convergence time is about 4000s. Figure 2 (b) shows the comprehensive position error results of the present invention, with convergence maintained at approximately 179 meters. Convergence time is within 1000 seconds. Comparing the two results, it is clear that the simulated navigation accuracy of the present invention is higher. The curve trend of the simulated curve of the present invention is more stable, which also indirectly demonstrates the greater stability of the proposed method.

[0102] The above description is only a specific example of the present invention and does not constitute any limitation to the present invention. Obviously, for professionals in this field, after understanding the content and principles of the present invention, it is possible to make various modifications and changes in any form and details without departing from the principles and structure of the present invention. However, these modifications and changes based on the ideas of the present invention are still within the scope of protection of the claims of the present invention.

Claims

1. A spacecraft integrated navigation system based on INS / GNSS / XPNAV, comprising an inertial navigation system INS and an X-ray pulsar navigation system XPNAV carried on a spacecraft in a Halo orbit of the Earth-Moon libration point, characterized in that: Also included are The Global Navigation Satellite System (GNSS) is composed of 100 satellites. .

2. The navigation method of the system according to claim 1, characterized in that The steps include: (1) Initialization parameters: Initialize the spacecraft in the body coordinate system measured by the inertial navigation system INS Next The angular increment of the spacecraft at time , Included in the body coordinate system The angle increments of the pitch, roll and yaw angles of the spacecraft 、 and ,in , represents the time when the spacecraft navigation solution is performed; the non-gravitational acceleration is , the length of the adjacent time interval is , the Earth's inertial coordinate system measured by the Global Navigation Satellite System GNSS Next The pseudo-range between the spacecraft and each satellite at time , Earth's inertial coordinate system measured by the X-ray pulsar navigation system XPNAV Next The distance between the spacecraft and the spacecraft in the direction of each pulsar unit vector at the moment is , , represents the number of pulsars, 3: (2) Calculate the predicted position error of the spacecraft at each moment in the Earth's inertial coordinate system: By angle increment Constructing the Earth's inertial coordinate system Next The direction cosine matrix of the spacecraft at time , and through non-gravitational acceleration and Calculating the Earth's inertial coordinate system Next Real-time acceleration of the spacecraft at this moment , then by and Calculating the Earth's inertial coordinate system Next The predicted position error of the spacecraft at time ; (3) Calculate the observed position error of the spacecraft relative to the predicted position of the spacecraft in the Earth's inertial coordinate system based on GNSS: Through the pseudo-range between the spacecraft and each satellite Calculating the Earth's inertial coordinate system Next Observation position of the spacecraft at this moment and through Calculating the Earth's inertial coordinate system Next Spacecraft predicted position at this moment The observed position error ; (4) Calculate the observed position error of the spacecraft relative to the predicted position of the spacecraft in the Earth's inertial coordinate system based on XPNAV: The distance between the spacecraft and the spacecraft through the unit vector in the direction of each pulsar Calculating the Earth's inertial coordinate system Next Observation position of the spacecraft at this moment and through Calculate the Earth's inertial coordinate system Next Spacecraft predicted position at this moment The observed position error ; (5) Obtain the current position and velocity of the spacecraft in the Earth's inertial coordinate system: Fusion of Earth Inertial Coordinate System via Federated Filter Structure Next The predicted position error of the spacecraft at time , spacecraft predicted position calculated based on GNSS and XPNAV The observed position error 、 Calculate the spacecraft's position in the Earth's inertial coordinate system Next The a posteriori estimated position error of the spacecraft at time and speed error , and then through and Calculating the Earth's inertial coordinate system Next Spacecraft position at time and speed .

3. The method according to claim 2, characterized in that The Earth's inertial coordinate system described in step (2) Next The direction cosine matrix of the spacecraft at time , its expression is; ； ； ； ； in, 、 Represent the Earth's inertial coordinate system Next Direction cosine matrix of the spacecraft at this moment, body coordinate system Next Time relative to the The direction cosine matrix of the spacecraft at time t, represents the 3rd-order identity matrix, Indicates angle increment The equivalent rotation vector of express The antisymmetric matrix of .

4. The method according to claim 2, characterized in that The Earth's inertial coordinate system described in step (2) Next Real-time acceleration of the spacecraft at this moment , predicted position error , the calculation formulas are: ； ； ； ； in, 、 Represent the Earth's inertial coordinate system Next The gravitational acceleration and non-gravitational acceleration of the spacecraft at this moment, Represents the Earth's inertial coordinate system Next The predicted velocity error of the spacecraft at time and denote the gravitational perturbations of the Earth and the Moon on the spacecraft, It represents the attitude error of the pitch angle, deflection angle and yaw angle of the spacecraft in the body coordinate system, express The initial bias of .

5. The method according to claim 2, characterized in that The observation location described in step (3) , Observation position error , the calculation formulas are: ； ； ； ； in, Indicates the Earth's inertial coordinate system The predicted spacecraft speed at time, Indicates the Earth's inertial coordinate system The predicted position of the spacecraft at time is the speed of light, Indicates that the GNSS signal is connected to the The deviation of the signal received by the receiver at that moment, express The measurement noise, Expressed in the Earth's inertial coordinate system The position of each satellite.

6. The method according to claim 4, characterized in that The observation position described in step (4) , Observation position error , the calculation formulas are: ； = ； in, Indicates the Earth's inertial coordinate system The unit vector of the spacecraft in the direction of each pulsar at that moment.

7. The method according to claim 4, characterized in that The position of the spacecraft in the Earth's inertial coordinate system described in step (5) ,speed , the calculation formula is; ； 。

Images