A combined navigation method based on pulsar / inter-satellite ranging / landmarks

Abstract

The present invention provides a combined navigation method based on pulsar / inter-satellite ranging / landmark, firstly, a state model and a measurement model of a pulsar / inter-satellite ranging / landmark combined navigation system are established, and the pulsar pulse arrival time TDOA, the inter-satellite relative position distance, and the landmark absolute position distance are obtained as the quantity measurement, and the pulsar / inter-satellite ranging / landmark combined navigation system is established by using the measurement information of the three navigation systems; the state model of the pulsar / inter-satellite ranging / landmark combined navigation system is established, the pulse arrival time TDOA is provided by pulsar navigation, the relative position information is provided by inter-satellite ranging navigation, and the absolute position information is provided by landmark navigation, and the measurement model is obtained by combining them, and the position and speed of the spacecraft are estimated by using the Unscented Kalman filter method. The present invention is aimed at the lunar exploration mission, and finally improves the accuracy of the navigation system.

Description

Technical Field

[0001] The invention belongs to the technical field of aerospace navigation, and in particular relates to a combined navigation method based on pulsar / inter-satellite ranging / landmarks. Background Art

[0002] Astronomical navigation is currently the most important autonomous navigation method for spacecraft. Astronomical navigation uses celestial bodies as navigation beacons, does not rely on other external information, and does not radiate energy to the outside. It obtains navigation information by passively receiving light radiated or reflected by celestial bodies. It is a completely autonomous navigation method.

[0003] According to the different measurements, the astronomical navigation system can be divided into astronomical angle measurement navigation, astronomical velocity measurement navigation and astronomical distance measurement navigation. Astronomical angle measurement navigation usually observes the angle between near celestial bodies and far celestial bodies. Since the motion law of celestial bodies in the universe is stable, the position and orientation of celestial bodies can be found through the ephemeris at a certain moment. By combining the detected observation quantity with the inherent motion law of celestial bodies, and performing geometric analysis and filtering on the obtained results, the position, speed and other information of the spacecraft can be obtained. The theoretical basis of astronomical velocity measurement navigation is the optical Doppler effect, that is, when the distance between the navigation target and the observed object changes, the frequency of light will also change accordingly. In deep space exploration, the spacecraft is the navigation target. In the process of observing the light emitted by natural celestial bodies, the frequency change of light is obtained, which reflects the relative motion between the spacecraft and the celestial body. Combined with the inherent motion law of celestial bodies, the obtained frequency change is processed accordingly, and the position, speed and other information of the spacecraft can be obtained from it. Astronomical ranging navigation mainly uses X-ray pulsars to obtain information, measures the arrival time of X-ray pulsars relative to the spacecraft, compares it with the pulse arrival time of the X-ray pulsar relative to the center of mass of the solar system, and obtains the projection of the distance between the spacecraft and the center of mass of the solar system in the direction of the X-ray pulse. Select a suitable filtering method to filter the obtained distance, and through geometry or equation solving, the position, speed and other information of the spacecraft can be obtained. The present invention adopts astronomical speed measurement navigation.

[0004] A pulsar is a rapidly rotating neutron star, the product of a supernova explosion caused by a massive star at the end of its life. The rotation period of a pulsar has excellent long-term stability and can be observed in different bands at the same time in many cases, such as radio, optical, and X-ray. Since X-ray radiation can be easily detected by small-area detectors, the X-ray signal emitted by a pulsar can be used for autonomous navigation of spacecraft. This navigation method is called X-ray pulsar navigation.

[0005] X-ray pulsar-based navigation (XPNAV) is an autonomous navigation method that obtains spacecraft status information by observing the pulse radiation signals of X-ray pulsars and combining relevant signal processing algorithms. Pulsars are natural celestial bodies in space. They have stable physical properties, are far away, and are not easily interfered by human factors. By establishing a phase time model of the pulsar at the solar system barycenter (SSB), the time when a certain pulse arrives at the solar system barycenter can be calculated. At the same time, by processing the photon measurement data on-orbit, the time when the pulse arrives at the spacecraft can be obtained. The difference between the two reflects the projection of the spacecraft's position relative to the SSB in the direction of the pulsar. By processing the measurement information in different directions, the position and time of the spacecraft can be estimated. As a type of astronomical navigation, X-ray pulsar navigation has the common characteristics of astronomical navigation: strong autonomy, strong anti-interference ability, high reliability, synchronous positioning and attitude determination, and navigation errors that do not accumulate over time.

[0006] Traditional astronomical navigation methods achieve spacecraft positioning by measuring the spatial angle between a reference celestial body and the spacecraft, and the navigation accuracy depends on the distance from the spacecraft to the reference celestial body. For deep space probes in the cruise phase, traditional astronomical navigation methods can only obtain a positioning accuracy of several thousand kilometers. Pulsar signals have extremely high periodic stability characteristics, and their characteristic signals can be used to identify changes in the space position of the aircraft, thereby achieving high-precision navigation and positioning. At the same time, because pulsars are very far away from the solar system, observing pulsars within the solar system can assume that their directions are basically unchanged. Therefore, under the same conditions, the accuracy of X-ray pulsar navigation can be better than 10km.

[0007] X-ray pulsar navigation has unique advantages, which are mainly reflected in the following two aspects: 1. Providing a high-precision reference time base. The rotation period of X-ray pulsars is highly stable. Using the observation information of pulsars, on the one hand, a comprehensive pulsar time can be established to maintain the time of the spacecraft navigation system, and on the other hand, the clock error of the onboard atomic clock can be corrected while achieving spacecraft positioning. 2. High navigation accuracy.

[0008] Intersatellite ranging navigation refers to a navigation method that uses a spacecraft with known position, speed and other status information as a reference, and uses the detection equipment carried on the spacecraft to measure and solve the relative distance information between two spacecraft. At present, the mainstream technologies used to achieve intersatellite ranging include the following: First, with the help of global positioning navigation satellites, GNSS ranging technology is used for intersatellite ranging positioning and navigation. This method uses small equipment and high measurement accuracy. The relative position accuracy can reach 1.5m, and the time synchronization accuracy can reach 6~12ns. It is one of the current mainstream ranging and positioning methods, but it has limitations in high orbit and deep space exploration. The second method is satellite autonomous ranging and speed measurement. This method does not rely on external equipment at all. The on-board equipment independently completes the measurement of the relative position between satellites. Commonly used methods include visible light ranging and speed measurement, laser ranging and speed measurement, radio ranging and speed measurement technology, etc. The limitation of visible light ranging technology is that the ranging range is limited. Laser ranging technology can achieve very high accuracy, but the equipment is large and the processing technology is complex. It is not suitable for small-volume, short-development cycle and low-cost application scenarios. Intersatellite measurement and navigation has the following advantages: ① High navigation accuracy. ② Strong adaptability of the navigation system. ③ Strong fault tolerance. ④ High navigation system performance.

[0009] Landmarks are a general term for significant fixed objects with accurate positions that can be observed visually or by radar for navigation or positioning. The method and process of positioning by observing landmarks and calculating the relationship between the landmarks and the vehicle (such as azimuth, distance, and horizontal angle, etc.) is called landmark positioning. Landmark navigation in deep space exploration usually uses visual images as the measurement information source. Visual images mainly include star images, images of on-orbit targets, and images of landmarks on the surface of planets. The optical camera carried on the spacecraft uses image recognition and extraction technology to extract feature points from the optical images of the target and its background as observation information, analyze the changes of observation points in the image within a certain period of time, solve the position, speed and other states of the spacecraft, and realize the estimation of the operating state of the on-orbit spacecraft. The advantage of the landmark navigation system is that it can provide high-precision position information. Using visual images as the measurement information source and extracting navigation information through image recognition increases the compatibility of the navigation system and can be applied to different applications. Combining landmark navigation with other navigation systems can effectively improve the accuracy of the navigation system.

[0010] In deep space exploration, if only X-ray pulsar navigation is used, it may not meet the navigation accuracy requirements of spacecraft. At present, the navigation methods commonly used in combination with pulsar navigation are mainly inertial navigation and optical navigation. This combined navigation system can effectively improve the scope of application of navigation systems that rely solely on pulsars and reduce the problem of inertial navigation errors accumulating over time. In terms of pulsar / optical combined navigation, virtual observation values ​​generated by neural networks and starlight angular distance measurements are used for centralized filtering during the pulsar observation period, which improves the navigation accuracy to a certain extent. In addition, combined navigation of pulsar / astronomical Doppler difference, a combination of X-ray pulsar navigation, inertial navigation and starlight navigation, and a combined navigation method of X-ray pulsar / inertial / starlight have been proposed. However, X-ray pulsar navigation currently has problems such as difficulty in star selection and low signal acquisition accuracy in deep space exploration. Summary of the invention

[0011] The present invention provides a combined navigation method based on pulsar / inter-satellite ranging / landmarks, which overcomes the problems of difficulty in selecting stars and low signal acquisition accuracy in current X-ray pulsar navigation in deep space exploration. The present invention establishes a pulsar / inter-satellite ranging / landmark combined navigation system, uses inter-satellite ranging navigation to provide relative position information, and uses landmark navigation to provide absolute position information, thereby improving the accuracy of the navigation system. The present invention verifies the theoretical part using computer simulation and Kalman filtering, and analyzes the advantages of combined navigation compared with other navigation.

[0012] In order to achieve the above object, the present invention adopts the following technical scheme:

[0013] A combined navigation method based on pulsars / intersatellite ranging / landmarks integrates X-ray pulsar observation information, intersatellite relative measurement information, and landmark information, and uses intersatellite relative measurement information and landmark information to provide a benchmark for the accurate acquisition of pulsar measurement information. This method studies the X-ray pulsar / intersatellite ranging / landmark combined navigation method, thereby obtaining a more accurate spacecraft state estimate and improving the performance of the navigation system. Since selecting a suitable pulsar can effectively improve the positioning accuracy of X-ray pulsar navigation, it specifically includes the following steps:

[0014] Step 1: Establish a combined navigation state model of pulsar / inter-satellite ranging / landmarks;

[0015] Step 2: Establish a combined navigation measurement model of pulsar / inter-satellite ranging / landmarks;

[0016] Step 3: discretize the state model and measurement model in step 1 and step 2;

[0017] Step 4: Get the arrival time of pulsar pulses As a measurement for pulsar navigation, it can provide the absolute position information of the spacecraft;

[0018] Step 5: Obtain the relative position information between spacecraft through inter-satellite ranging navigation; obtain the distance information of the spacecraft relative to the lunar surface landmarks through the landmark observation system, which is the absolute position information;

[0019] Step 6: Use the Unscented Kalman filter to process the system measurement information to obtain the estimated vector of the spacecraft's position and velocity.

[0020] Furthermore, the step 1 includes: for the lunar probe, the coordinate system selects the lunar center inertial coordinate system of epoch (J2000.0), considering the gravity, radiation pressure, rocket thrust during orbital maneuvers and high-precision ephemeris of the sun, moon, earth and other celestial bodies on the lunar probe, and the state model of the spacecraft celestial navigation system is as follows:

[0021] (1)

[0022] The above formula can be simplified as:

[0023] +W(t) (2)

[0024] In the formula, is the state vector of the state model, for The differential of x, y, z, are the position and velocity of the lunar probe in the X, Y, and Z directions respectively; is the system nonlinear continuous state transfer function of the state model; , , are the gravitational constants of the Sun, Moon, and Earth respectively; is the vector from the heliocenter to the detector; is the vector from the center of the moon to the detector; is the vector from the center of the earth to the detector; is the vector from the center of the moon to the center of the sun; is the vector from the center of the moon to the center of the earth in the geocentric coordinate system; is the coordinate of the moon's position in the solar mass center coordinate system; is the coordinate of the Earth's position in the geocentric coordinate system, where the coordinates of the Moon and the Earth are functions of time and can be obtained from the ephemeris table; , , , , , are the system noise respectively.

[0025] Furthermore, the step 2 comprises:

[0026] The non-gravitational acceleration is measured by the accelerometer carried on the spacecraft, and the gravitational acceleration is calculated by the spacecraft orbital dynamics. The two are added together to obtain the acceleration information of the spacecraft. After integration, the velocity information is obtained to correct the pulse arrival time caused by the Doppler frequency shift. The influence of measurement accuracy is used to obtain high-precision pulsar measurement information; inter-satellite ranging navigation is used to provide relative position information between spacecraft; the landmark observation system is used to extract the distance information of the spacecraft relative to the lunar surface landmarks as absolute position information; X-ray pulsar navigation obtains the pulse arrival time by calculating the difference between the time when the pulsar signal arrives at the center of mass of the solar system and the time when it arrives at the spacecraft. , thereby obtaining the projection of the distance of the spacecraft relative to the center of mass of the solar system in the direction of the pulsar, and thus obtaining the position information of the spacecraft:

[0027]

[0028] in, is the position vector of the solar system mass center in the solar mass center coordinate system; is the position vector of the spacecraft relative to the lunar mass center; is the position vector of the pulsar in the solar mass center coordinate system; b and r are , size; is the constant measurement error of the pulse arrival time; is the random measurement error of the pulse arrival time; c is the speed of light; n is the direction vector of the pulsar relative to the center of mass of the solar system.

[0029] Three pulsars are selected as navigation stars to provide pulsar measurement information; inter-satellite ranging navigation is used to provide relative position information between spacecraft; the landmark observation system is used to extract the distance information of the spacecraft relative to the lunar surface landmarks as absolute position information; the measurement information of the three navigation systems is used to establish a pulsar / inter-satellite ranging / landmark combined navigation system, and the results are expressed as follows:

[0030]

[0031] in, To establish the quantity measurement of the integrated navigation system, , , These are the measurement information differences of the pulsar signals of three different pulsars selected arriving at the spacecraft and the auxiliary satellite. , , . is the difference between the time when the pulsar signal of the first selected pulsar reaches the center of mass of the solar system and the time when it reaches the spacecraft, is the difference between the time when the pulsar signal of the first selected pulsar arrives at the center of mass of the solar system and the time when it arrives at the auxiliary satellite; is the difference between the time when the pulsar signal of the second selected pulsar reaches the center of mass of the solar system and the time when it reaches the spacecraft, is the difference between the time when the pulsar signal of the second selected pulsar arrives at the center of mass of the solar system and the time when it arrives at the auxiliary satellite; The difference between the time when the pulsar signal of the selected third pulsar arrives at the center of mass of the solar system and the time when it arrives at the spacecraft, It is the difference between the time when the pulsar signal of the selected third pulsar arrives at the center of mass of the solar system and the time when it arrives at the auxiliary satellite. is the relative distance between stars, , is the distance of the landmark relative to the spacecraft. is the error term corresponding to the TDOA obtained from three pulsars, is the error term corresponding to the relative position information of inter-satellite ranging, is the error term corresponding to the landmark distance information.

[0032] Its abbreviation is:

[0033]

[0034] in, To establish a combined navigation system for measuring the quantity at a certain moment, is the measurement information vector of the integrated navigation system at the corresponding moment; represents the measurement noise of the integrated navigation system at the corresponding moment; It is the nonlinear continuous measurement function of the integrated navigation system.

[0035] Furthermore, the step five includes: combining pulsar navigation with inter-satellite ranging navigation, using inter-satellite ranging information as part of the observation quantity to replace part of the pulsar observation quantity; introducing landmark navigation to provide absolute position information, reducing the impact of pulsar system errors, and effectively improving the accuracy of the navigation system.

[0036] Beneficial effects:

[0037] X-ray pulsar navigation has several advantages. However, due to the long distance of X-ray pulsars, the received X-ray signals are relatively weak. In addition, in X-ray pulsar navigation measurements, the navigation system time is the coordinate time. When considering the effect of gravity, the coordinate time does not satisfy the invariance of the speed of light. Therefore, when the time when the pulsar radiation signal arrives at the SSB and the signal arrival time recorded by the spacecraft's onboard atomic clock are converted into SSB coordinates, the pulsar timing observation method will produce errors, affecting navigation accuracy. In deep space exploration, if only X-ray pulsar navigation is used, it may not meet the spacecraft navigation accuracy requirements.

[0038] Although the relative position of satellites can be located by using only intersatellite relative measurement information, the lack of inertial reference information during the extrapolation process will cause the filtering results to diverge after a long period of iteration. If long-term autonomous navigation is to be achieved, an absolute observation reference must be introduced. If pulsar navigation is combined with intersatellite ranging navigation, the intersatellite ranging information can be used as part of the observation quantity to replace part of the pulsar observation quantity, thereby giving full play to the outstanding advantages of pulsar navigation in deep space exploration, and reducing the number of pulsars required, alleviating the difficulty of pulsar selection. On this basis, the landmark navigation system can provide high-precision position information, thereby further improving the navigation accuracy of the system. BRIEF DESCRIPTION OF THE DRAWINGS

[0039] Figure 1 This is a diagram of a system measurement model involved in the combined navigation method based on pulsar / inter-satellite ranging / landmarks of the present invention.

[0040] Figure 2 The present invention is a flow chart of the combined navigation method based on pulsar / inter-satellite ranging / landmark. DETAILED DESCRIPTION

[0041] In order to make the purpose, technical solutions and advantages of the present invention more clearly understood, the present invention is further described in detail below in conjunction with the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present invention and are not intended to limit the present invention. In addition, the technical features involved in the various embodiments of the present invention described below can be combined with each other as long as they do not conflict with each other.

[0042] The present invention provides a combined navigation method based on pulsar / inter-satellite ranging / landmarks. In order to correct the system error of pulsar navigation, the measurement of pulsar navigation is used to provide the position information of the spacecraft; the relative position information between spacecrafts is provided by inter-satellite ranging navigation; and the distance information of the spacecraft relative to the lunar surface landmarks is extracted by the landmark observation system as the absolute position information. The measurement information of the three navigation systems is used to establish a pulsar / inter-satellite ranging / landmark combined navigation system, and the system measurement information is processed by Unscented Kalman filtering to obtain the estimated results and estimation accuracy of the position and speed of the spacecraft. The obtained results are used to feedback and correct the measurement information of the pulsar navigation, offset the random errors in the estimated pulse profile extraction, and improve the system estimation accuracy.

[0043] The present invention is described in detail below with reference to specific embodiments.

[0044] The present invention is applied to the lunar probe, using pulsar navigation based on Doppler frequency shift correction, using the accelerometer carried on the spacecraft to measure the non-gravitational acceleration, using the spacecraft orbital dynamics to calculate the gravitational acceleration, the two are added to obtain the acceleration information of the spacecraft, and the velocity information is obtained after integration to correct the Doppler frequency shift. (pulse arrival time) measurement accuracy, and obtain high-precision pulsar measurement information. Figure 2 As shown, the present invention specifically includes the following steps:

[0045] Step (1) Establish the combined navigation state model of pulsar / inter-satellite ranging / landmark:

[0046] For the lunar probe, the coordinate system is the lunar center inertial coordinate system of epoch (J2000.0), taking into account the gravitational force of the sun, moon, earth and other celestial bodies on the lunar probe, radiation pressure, rocket thrust during orbital maneuvers, and high-precision ephemeris, etc. The state model of the spacecraft celestial navigation system is shown in formula (1):

[0047]

[0048] The above formula can be simplified as:

[0049] +W(t) (2)

[0050] In the formula, is the state vector of the state model, for The differential of x, y, z, are the position and velocity of the lunar probe in the X, Y, and Z directions respectively; is the system nonlinear continuous state transfer function of the state model; , , are the gravitational constants of the Sun, Moon, and Earth respectively; is the vector from the heliocenter to the detector; is the vector from the center of the moon to the center of the earth in the geocentric coordinate system; is the vector from the center of the moon to the center of the sun; is the coordinate of the moon's position in the solar mass center coordinate system; is the coordinate of the Earth's position in the geocentric coordinate system, where the coordinates of the Moon and the Earth are functions of time and can be obtained from the ephemeris table; , , , , , are the system noise respectively.

[0051] Step (2) Establish a combined navigation measurement model of pulsar / inter-satellite ranging / landmarks:

[0052] The measurement model of pulsar navigation is as follows Figure 1 As shown. By establishing a phase-time model of the pulsar at the solar system barycenter (SSB), the time it takes for a certain pulse to reach the solar system barycenter can be calculated. At the same time, by processing the photon measurement data on-orbit, the time it takes for the pulse to reach the spacecraft can be obtained. The difference between the two reflects the projection of the spacecraft's position relative to the SSB in the direction of the pulsar. By processing measurement information in different directions, the position and time of the spacecraft can be estimated. The non-gravitational acceleration is measured using the accelerometer carried on the spacecraft, and the gravitational acceleration is calculated using the spacecraft's orbital dynamics. The two are added together to obtain the acceleration information of the spacecraft. The velocity information is obtained after integration to correct the Doppler shift caused by the accelerometer. The influence of measurement accuracy is used to obtain high-precision pulsar measurement information. Intersatellite ranging navigation is used to provide relative position information between spacecraft; the landmark observation system is used to extract the distance information of the spacecraft relative to the lunar surface landmarks as absolute position information. X-ray pulsar navigation mainly obtains the pulse arrival time by calculating the difference between the time when the pulsar signal arrives at the center of mass of the solar system and the time when it arrives at the spacecraft. , thereby obtaining the projection of the distance of the spacecraft relative to the center of mass of the solar system in the direction of the pulsar, and thus obtaining the position information of the spacecraft.

[0053]

[0054] in, is the position vector of the solar system mass center in the solar mass center coordinate system, is the position vector of the spacecraft relative to the lunar mass center; is the position vector of the pulsar in the solar mass center coordinate system; b and r are , size; is the constant measurement error of the pulse arrival time; is the random measurement error of the pulse arrival time. c is the speed of light; n is the direction vector of the pulsar relative to the center of mass of the solar system.

[0055] Three pulsars are selected as navigation stars to provide pulsar measurement information. Pulsar detectors are installed on both the spacecraft and the auxiliary satellite, and the arrival time difference of the pulses at two different satellites is used. As a measurement of pulsar navigation, it can provide the absolute position information of the spacecraft; using intersatellite ranging navigation to provide relative position information between spacecraft; using the landmark observation system, the distance information of the spacecraft relative to the lunar surface landmarks is extracted as absolute position information. Using the measurement information of the three navigation systems, a pulsar / intersatellite ranging / landmark combined navigation system is established, and the results are expressed as:

[0056]

[0057] in, To establish the quantity measurement of the integrated navigation system, , , These are the measurement information differences of the pulsar signals of three different pulsars selected arriving at the spacecraft and the auxiliary satellite. , , . is the difference between the time when the pulsar signal of the first selected pulsar reaches the center of mass of the solar system and the time when it reaches the spacecraft, is the difference between the time when the pulsar signal of the first selected pulsar arrives at the center of mass of the solar system and the time when it arrives at the auxiliary satellite; is the difference between the time when the pulsar signal of the second selected pulsar reaches the center of mass of the solar system and the time when it reaches the spacecraft, is the difference between the time when the pulsar signal of the second selected pulsar arrives at the center of mass of the solar system and the time when it arrives at the auxiliary satellite; The difference between the time when the pulsar signal of the selected third pulsar arrives at the center of mass of the solar system and the time when it arrives at the spacecraft, It is the difference between the time when the pulsar signal of the selected third pulsar arrives at the center of mass of the solar system and the time when it arrives at the auxiliary satellite. is the relative distance between stars, , is the distance of the landmark relative to the spacecraft. is the error term corresponding to the TDOA obtained from three pulsars, is the error term corresponding to the relative position information of inter-satellite ranging, is the error term corresponding to the landmark distance information.

[0058] Its abbreviation is:

[0059]

[0060] in, To establish a combined navigation system for measuring the quantity at a certain moment, is the measurement information vector of the integrated navigation system at the corresponding moment; represents the measurement noise of the integrated navigation system at the corresponding moment; It is the nonlinear continuous measurement function of the integrated navigation system.

[0061] Figure 1 In the figure, the first landmark 1 and the second landmark 2 are landmark sensors set on the moon, and r1 and r2 are their corresponding landmark distances. , ; is the difference in measured information between the pulsar signal of one of the selected pulsars reaching the spacecraft and the auxiliary satellite, , The pulsar signal of the selected pulsar reaches the spacecraft measurement information, The measurement information of the pulsar signal of the selected pulsar reaching the auxiliary satellite.

[0062] Step (3) discretizes the state model and measurement model in step (1) and step (2):

[0063] The results after discretization are as follows:

[0064]

[0065]

[0066] In the formula, , and They are and The discretized result is for After discrete Time to The nonlinear state transfer function at time , for After discrete The nonlinear measurement function of time, and for and After discrete The equivalent noise at time , and and Not related to each other.

[0067] Step (4) Obtain the arrival time of the pulsar pulse As a measurement for pulsar navigation, it can provide the absolute position information of the spacecraft;

[0068] Step (5) obtains the relative position information between spacecraft through inter-satellite ranging navigation; obtains the distance information of the spacecraft relative to the lunar surface landmarks through the landmark observation system, which is the absolute position information.

[0069] Step (6) uses the Unscented Kalman filter to process the system measurement information and obtain the estimated vector of the spacecraft's position and velocity. The obtained results are used to feed back and correct the measurement information of the pulsar navigation to offset the random error in the estimated pulse profile extraction and improve the system estimation accuracy.

[0070] Based on the state model of the astronomical navigation system, the measurement model of the astronomical navigation system, pulsar navigation, intersatellite ranging technology, and the measurement obtained by landmark navigation, the astronomical navigation system Unscented Kalman filter is performed to obtain the position and velocity of the spacecraft in the inertial coordinate system. A series of sample points are selected near, and the means and covariances of these sample points are and . Assume that the state variable is n×1 dimensional.

[0071] A. Initialization

[0072]

[0073]

[0074] B. In the formula, are the estimated values ​​of the three-axis position and velocity of the spacecraft in the inertial coordinate system at time 0 (initial time), is the true value of the three-axis position and velocity of the spacecraft in the inertial coordinate system at the 0th moment (initial moment), is the initial mean square error matrix of the state vector.

[0075] Calculate the sampling points:

[0076] In the celestial navigation system State vector at the moment A series of sample points are selected near the vicinity, and the mean and mean square error matrix of these sample points are respectively and . Assuming the state vector is n×1 dimensional, the 2n+1 sample points and their weights are:

[0077]

[0078] In the formula are sampling points, and their distribution is The Gaussian distribution of is approximated; is the scale adjustment parameter, and its size change will affect the filtering effect; A is the system state transfer matrix. hour, Take the i-th row of A; when hour, Take the i-th column of A. n is the dimension of the state vector; is the initial value of the weight; For the The weight of each sampling point; For the The weight of the sampling points;

[0079] Time update:

[0080] One-step prediction of the state vector of celestial navigation system for:

[0081]

[0082] in, () is the system nonlinear continuous state transfer function of the state model.

[0083] One-step prediction weighted result of the state vector of all sampling points of the celestial navigation system for:

[0084]

[0085] In the formula, For the The weight of the sampling points;

[0086] One-step prediction of the estimated mean square error matrix of the state vector of the celestial navigation system for:

[0087]

[0088] In the formula, is the state model error covariance matrix of the celestial navigation system at time k;

[0089] The measurement estimation vector corresponding to the sampling point of the celestial navigation system :

[0090]

[0091] in, It is the nonlinear continuous measurement function of the integrated navigation system.

[0092] The weighted vector of all sampling points of the celestial navigation system :

[0093]

[0094] The measurements are updated as follows:

[0095] Celestial Navigation System Measurement Mean Square Error Matrix for:

[0096]

[0097] In the formula, is the measurement noise covariance matrix of the celestial navigation system at time k;

[0098] The mean square error matrix of state vector measurement of celestial navigation system :

[0099]

[0100] Celestial navigation system filter gain for:

[0101]

[0102] Estimated state vector of celestial navigation system and the estimated mean square error matrix for:

[0103]

[0104]

[0105] After the Unscented Kalman filter of the astronomical navigation system, the position and velocity of the spacecraft are obtained in the inertial coordinate system.

[0106] The contents not described in detail in the specification of the present invention belong to the prior art known to the professional and technical personnel in the field. It is easy for the technical personnel in the field to understand that the above description is only the preferred embodiment of the present invention and is not intended to limit the present invention. Any modification, equivalent replacement and improvement made within the spirit and principle of the present invention shall be included in the protection scope of the present invention.

Claims

1. A combined navigation method based on pulsar / inter-satellite ranging / landmarks, characterized in that: The following steps are involved: Step 1: Establish a combined navigation state model of pulsar / intersatellite ranging / landmarks, including: for the lunar probe, the coordinate system selects the lunar center inertial coordinate system of epoch (J2000.0), considering the gravitational force of the sun, moon, earth and other celestial bodies on the lunar probe, radiation pressure, rocket thrust during orbital maneuvers, and high-precision ephemeris. The state model of the spacecraft celestial navigation system is as follows: The above formula can be simplified as: +W(t) (2) In the formula, is the state vector of the state model, for The differential of x, y, z, are the position and velocity of the lunar probe in the X, Y, and Z directions respectively; is the system nonlinear continuous state transfer function of the state model; , , are the gravitational constants of the Sun, Moon, and Earth respectively; is the vector from the heliocenter to the detector; is the vector from the center of the moon to the detector; is the vector from the center of the earth to the detector; is the vector from the center of the moon to the center of the sun; is the vector from the center of the moon to the center of the earth in the geocentric coordinate system; is the coordinate of the moon's position in the solar mass center coordinate system; is the coordinate of the Earth's position in the geocentric coordinate system, where the coordinates of the Moon and the Earth are functions of time and can be obtained from the ephemeris table; , , , , , are system noise respectively; Step 2: Establish a combined navigation measurement model of pulsar / inter-satellite ranging / landmarks; Step 3: discretize the state model and measurement model in step 1 and step 2; Step 4: Get the arrival time of pulsar pulses As a measurement for pulsar navigation, it can provide the absolute position information of the spacecraft; Step 5: Obtain the relative position information between spacecraft through inter-satellite ranging navigation; obtain the distance information of the spacecraft relative to the lunar surface landmarks through the landmark observation system, which is the absolute position information; Step 6: Use the Unscented Kalman filter to process the system measurement information to obtain the estimated vector of the spacecraft's position and velocity.

2. The combined navigation method based on pulsar / inter-satellite ranging / landmark according to claim 1, characterized in that: The second step comprises: The non-gravitational acceleration is measured by the accelerometer carried on the spacecraft, and the gravitational acceleration is calculated by the spacecraft orbital dynamics. The two are added together to obtain the acceleration information of the spacecraft. After integration, the velocity information is obtained to correct the pulse arrival time caused by the Doppler frequency shift. The influence of measurement accuracy is used to obtain high-precision pulsar measurement information; inter-satellite ranging navigation is used to provide relative position information between spacecraft; the landmark observation system is used to extract the distance information of the spacecraft relative to the lunar surface landmarks as absolute position information; X-ray pulsar navigation obtains the pulse arrival time by calculating the difference between the time when the pulsar signal arrives at the center of mass of the solar system and the time when it arrives at the spacecraft. , thereby obtaining the projection of the distance of the spacecraft relative to the center of mass of the solar system in the direction of the pulsar, and thus obtaining the position information of the spacecraft: in, is the position vector of the solar system mass center in the solar mass center coordinate system, is the position vector of the spacecraft relative to the lunar mass center; is the position vector of the pulsar in the solar mass center coordinate system; b and r are , size; is the constant measurement error of the pulse arrival time; is the random measurement error of the pulse arrival time; c is the speed of light; n is the direction vector of the pulsar relative to the center of mass of the solar system; Three pulsars are selected as navigation stars to provide pulsar measurement information; inter-satellite ranging navigation is used to provide relative position information between spacecraft; the landmark observation system is used to extract the distance information of the spacecraft relative to the lunar surface landmarks as absolute position information; the measurement information of the three navigation systems is used to establish a pulsar / inter-satellite ranging / landmark combined navigation system, and the results are expressed as follows: in, To establish the quantity measurement of the integrated navigation system, , , are the measured information differences of the pulsar signals of three different pulsars selected reaching the spacecraft and the auxiliary satellite; , , ; is the difference between the time when the pulsar signal of the first selected pulsar reaches the center of mass of the solar system and the time when it reaches the spacecraft, is the difference between the time when the pulsar signal of the first selected pulsar arrives at the center of mass of the solar system and the time when it arrives at the auxiliary satellite; is the difference between the time when the pulsar signal of the second selected pulsar reaches the center of mass of the solar system and the time when it reaches the spacecraft, is the difference between the time when the pulsar signal of the second selected pulsar arrives at the center of mass of the solar system and the time when it arrives at the auxiliary satellite; is the difference between the time when the pulsar signal of the selected third pulsar arrives at the center of mass of the solar system and the time when it arrives at the spacecraft, is the difference between the time when the pulsar signal of the selected third pulsar arrives at the center of mass of the solar system and the time when it arrives at the auxiliary satellite, is the relative distance between stars, , is the distance of the landmark relative to the spacecraft, is the error term corresponding to the TDOA obtained from three pulsars, is the error term corresponding to the relative position information of inter-satellite ranging, is the error term corresponding to the landmark distance information; Its abbreviation is: in, To establish a combined navigation system for measuring the quantity at a certain moment, is the measurement information vector of the integrated navigation system at the corresponding moment; represents the measurement noise of the integrated navigation system at the corresponding moment; It is the nonlinear continuous measurement function of the integrated navigation system.

3. The combined navigation method based on pulsar / inter-satellite ranging / landmark according to claim 2, characterized in that: The step five comprises: Combining pulsar navigation with intersatellite ranging navigation, using intersatellite ranging information as part of the observation quantity to replace part of the pulsar observation quantity; introducing landmark navigation to provide absolute position information and reduce the impact of pulsar system errors can effectively improve the accuracy of the navigation system.

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