A navigation method and device for spacecraft in unknown gravity field

Abstract

The present invention provides a navigation method and device for a spacecraft under an unknown gravity field, which belongs to the field of space navigation technology and can navigate the spacecraft when the spacecraft has insufficient observation data on the detection target. The method determines the motion state estimation error of the spacecraft and the gravity field parameter estimation error of the central celestial body according to the motion state information of the spacecraft at the last moment and the gravity field parameters of the central celestial body predicted at the last moment; predicts the gravity field parameters of the central celestial body at the current moment according to the motion state estimation error, the gravity field parameter estimation error and the observation data of the motion state information of the spacecraft at the current moment; uses the gravity field parameters of the central celestial body predicted at the current moment and the motion state information of the spacecraft at the last moment to filter the observation data of the motion state information of the spacecraft at the current moment, and obtains the motion state information of the spacecraft at the current moment, so as to realize navigation of the spacecraft.

Description

Technical Field

[0001] The present invention relates to the field of space navigation technology, and in particular to a navigation method and device for a spacecraft in an unknown gravity field. Background Art

[0002] When a spacecraft is carrying out a detection mission, it is necessary to navigate the spacecraft based on the gravity field information (also called gravity field parameters) of the detection target (also called the central celestial body) of the spacecraft. The gravity field information of the detection target needs to be determined offline based on a large amount of observation data obtained during the previous spacecraft orbiting the detection target.

[0003] With the development of space exploration technology, exploration missions have gradually become smaller and longer-distance, making exploration targets more diverse. However, due to limited spacecraft resources, it is impossible to conduct a large number of observations on each exploration target. Therefore, in some cases, the spacecraft cannot navigate during the exploration mission because the spacecraft's observation data on the exploration target is not sufficient to determine the gravity field information of the exploration target. Summary of the invention

[0004] The present invention provides a method and device for navigating a spacecraft in an unknown gravity field, which can navigate the spacecraft when the spacecraft has insufficient observation data on a detection target.

[0005] In order to achieve the above object, the present invention adopts the following technical solution:

[0006] In a first aspect, the present invention provides a navigation method for a spacecraft in an unknown gravity field, comprising: determining the motion state estimation error of the spacecraft and the gravity field parameter estimation error of the central celestial body according to the motion state information of the spacecraft at the last moment and the gravity field parameters of the central celestial body predicted at the last moment. Predicting the gravity field parameters of the central celestial body at the current moment according to the motion state estimation error, the gravity field parameter estimation error and the observation data of the motion state information of the spacecraft at the current moment. Using the gravity field parameters of the central celestial body predicted at the current moment and the motion state information of the spacecraft at the last moment, filtering the observation data of the motion state information of the spacecraft at the current moment, obtaining the motion state information of the spacecraft at the current moment, so as to realize navigation of the spacecraft; wherein the motion state information of the spacecraft at the current moment is used to represent the real motion state of the spacecraft at the current moment.

[0007] In a navigation method for a spacecraft in an unknown gravity field provided by the present invention, the motion state information of the spacecraft at the last moment (i.e., the precise positioning of the spacecraft at the last moment) and the predicted gravity field parameters of the central celestial body are used to determine the motion state estimation error of the spacecraft and the gravity field parameter estimation error of the central celestial body; after obtaining the above two errors, the gravity field parameters of the central celestial body predicted at the current moment are predicted by the above two errors, and then the gravity field parameters of the central celestial body predicted at the current moment, the motion state information of the spacecraft at the last moment, and the observation data of the motion state information of the spacecraft at the current moment are used to jointly realize the navigation of the spacecraft in the unknown gravity field, and determine the motion state information of the spacecraft at the current moment (i.e., the precise positioning of the spacecraft at the current moment). In summary, the navigation method can continuously and in real time predict and update the gravity field parameters of the central celestial body according to the motion state information of the spacecraft and the observation data of the motion state information during the navigation of the spacecraft. Therefore, in this process, not only can the gravity field parameters of the central celestial body be continuously refined, but also the spacecraft can be navigated by the refined gravity field parameters when the observation data of the spacecraft on the detection target is insufficient.

[0008] In an implementation of the first aspect, the gravity field parameters of the central celestial body predicted at the current moment satisfy:

[0009] ;

[0010] in, Indicates the current moment, Indicates the last moment, It represents the gravity field parameters of the central celestial body predicted at the current moment, represents the gravity field parameters of the central celestial body predicted at the last moment, Indicates the gain of gravity field parameters. Satisfy the following formula;

[0011] ;

[0012] represents the motion state estimation error, represents the estimation error of gravity field parameters;

[0013] Observation data representing the motion state information of the spacecraft at the current moment, , The first observation data representing the motion state information of the spacecraft at the current moment i Prediction values, It is predicted through the motion state information of the spacecraft at the previous moment.

[0014] In an implementation of the first aspect, the motion state estimation error Satisfy the following formula;

[0015] ;

[0016] in, Indicates the current moment, represents the transposed matrix; , Indicates the last moment, The first observation data representing the motion state information of the spacecraft at the current moment i Prediction values, It is obtained by predicting the motion state information of the spacecraft at the previous moment. represents the gravity field parameters of the central celestial body predicted at the last moment, and , and The field harmonic coefficients representing the gravity field of the central celestial body at the last moment; Represents the error in the predicted value of the spacecraft's observation data at the current moment.

[0017] In one implementation of the first aspect, the predicted value of the observation data of the spacecraft at the current moment is , satisfying the following formula;

[0018] ;

[0019] in, , represents the motion state information of the spacecraft at the last moment, and , represents the position of the spacecraft at the last moment, Indicates the speed of the spacecraft at the last moment; represents the covariance estimate of the gravity field parameters of the central body; represents the gradient operator, represents the gravitational potential energy of the central celestial body at the previous moment, represents the system error matrix of the spacecraft.

[0020] In an implementation of the first aspect, the gravity field parameter estimation error Satisfy the following formula;

[0021] ;

[0022] in, Indicates the current moment, represents the transposed matrix; Represents the error in the predicted value of the spacecraft's observation data at the current moment.

[0023] In a second aspect, the present invention provides a navigation device for a spacecraft in an unknown gravity field, which is used to navigate the spacecraft while the spacecraft is orbiting a central celestial body with unknown gravity field parameters, and the device includes an error determination module, a gravity field parameter determination module and a navigation module;

[0024] The error determination module is used to determine the estimated error of the motion state of the spacecraft and the estimated error of the gravity field parameters of the central celestial body according to the motion state information of the spacecraft at the last moment and the gravity field parameters of the central celestial body predicted at the last moment;

[0025] The gravity field parameter determination module is used to predict the gravity field parameters of the central celestial body at the current moment according to the motion state estimation error, the gravity field parameter estimation error and the observation data of the motion state information of the spacecraft at the current moment;

[0026] The navigation module is used to filter the observation data of the motion state information of the spacecraft at the current moment by using the gravity field parameters of the central celestial body predicted at the current moment and the motion state information of the spacecraft at the previous moment, so as to obtain the motion state information of the spacecraft at the current moment and realize navigation of the spacecraft.

[0027] In a third aspect, the present invention provides an electronic device, comprising a processor and a memory coupled to the processor; the memory is used to store computer instructions, and when the electronic device is running, the processor executes the computer instructions stored in the memory, so that the electronic device performs the method described in the first aspect or any one of its implementations.

[0028] In a fourth aspect, the present invention provides a computer-readable storage medium, comprising computer program instructions, which, when executed by a computer, enable the computer to execute the method as described in the first aspect or any one of its implementations.

[0029] In a fifth aspect, the present invention provides a computer program product, comprising computer program instructions, which, when executed on a computer, enable the computer to execute the method as described in the first aspect or any one of its implementations.

[0030] The technical effects corresponding to the above-mentioned second to fifth aspects and their possible implementations can refer to the above-mentioned description of the technical effects of the first aspect and its possible implementations, and will not be repeated here. BRIEF DESCRIPTION OF THE DRAWINGS

[0031] Figure 1 It is a schematic diagram of a navigation method for a spacecraft in an unknown gravity field provided by an embodiment of the present application;

[0032] Figure 2It is a structural schematic diagram of a navigation device for a spacecraft in an unknown gravity field provided in an embodiment of the present application. DETAILED DESCRIPTION

[0033] The terms "first" and "second" etc. in the description and claims of the present invention are used to distinguish different objects rather than to describe a specific order of the objects.

[0034] The “and / or” in the embodiments of the present application indicates the relationship between objects. For example, A and / or B may indicate the following three situations: A exists alone, B exists alone, and A and B exist at the same time.

[0035] In the embodiments of the present application, words such as "exemplary" or "for example" are used to indicate examples, illustrations or descriptions. Any embodiment or design described as "exemplary" or "for example" in the embodiments of the present application should not be interpreted as being more preferred or more advantageous than other embodiments or designs. Specifically, the use of words such as "exemplary" or "for example" is intended to present related concepts in a specific way.

[0036] The method and device provided in the embodiments of the present application relate to spacecraft navigation, and are used to navigate a spacecraft when the spacecraft performs a space exploration mission. Specifically, they are used to navigate a spacecraft while the spacecraft is orbiting a central celestial body (for example, a satellite of some planet) whose gravity field parameters are unknown.

[0037] It should be noted that in the embodiment of the present application, the spacecraft operates on a set orbit to perform a detection mission. In this process, navigating the spacecraft means that the spacecraft obtains the observation data of the ground observation station on the motion state information of the spacecraft in real time through the uplink, and compares the observation data with the set orbit. When the spacecraft deviates from the set orbit, the motion state of the spacecraft is adjusted to keep the spacecraft on the set orbit; at the same time, the spacecraft determines the motion state information of the spacecraft in real time through the observation data of the motion state information of the spacecraft and the gravity field parameters of the central celestial body that the spacecraft orbits. Among them, the motion state information of the spacecraft includes the position of the spacecraft and the speed of the spacecraft.

[0038] With the development of science and technology, detection targets are becoming more and more diverse. However, due to limited spacecraft resources, it is impossible to conduct a large number of observations on each detection target. Therefore, in some cases, the observation data of the spacecraft on the detection target is not sufficient to determine the gravity field parameters of the detection target, making it impossible for the spacecraft to navigate during the detection mission. In order to solve the above problem, the embodiment of the present application provides a navigation method and device for a spacecraft in an unknown gravity field, which can continuously and in real time predict and update the gravity field parameters of the central celestial body according to the motion state information of the spacecraft and the observation data of the motion state information during the navigation of the spacecraft. Therefore, in this process, not only can the gravity field parameters of the central celestial body be continuously refined, but also the spacecraft can be navigated by the refined gravity field parameters when the observation data of the spacecraft on the detection target is insufficient.

[0039] It should be understood that the above-mentioned unknown gravity field refers to the gravity field parameters of the gravity field of the central celestial body around which the spacecraft orbits are unknown.

[0040] Exemplarily, the navigation method of a spacecraft in an unknown gravity field provided by an embodiment of the present invention may be executed by an electronic device having a processing function, for example, the electronic device may be a computer, a server, etc. Taking the electronic device as a computer as an example, the hardware part of the computer may include: a processor, a memory, a network interface, a user interface, a communication bus, etc.

[0041] The processor is used to control the electronic device to perform related processing and computing tasks, for example, to determine the motion state estimation error and gravity field parameter estimation error, predict gravity field parameters, determine the motion state information of the spacecraft, etc. The processor may include a central processing unit (CPU) or other processors, and the processor may be single-core or multi-core, for example, the processor may include multiple CPUs.

[0042] The memory is used to store computer instructions and related data, for example, to store motion state estimation errors, gravity field parameter estimation errors, gravity field parameters, and motion state information. The memory can be a random access memory (RAM), a read only memory (ROM), an erasable programmable read-only memory (EPROM), a flash memory, or an optical memory, a disk storage medium or other magnetic storage device, or any other medium that can be used to store program codes or data that can be accessed by a computer. Optionally, the memory can be integrated into the processor, or the memory can be independent of the processor.

[0043] The network interface is used for the computer to communicate with other devices or communication networks. The network interface can be a transceiver with transceiver functions. Optionally, the network interface can include a standard wired interface, a wireless interface (such as a WI-FI interface, a Bluetooth interface, a 5G interface).

[0044] The communication bus is used to realize the connection and communication between various components. For example, the processor, memory, network interface and user interface mentioned above can be interconnected through the communication bus.

[0045] The user interface may include a display screen and an input unit (such as a keyboard). Optionally, the user interface may also include a standard wired interface and a wireless interface.

[0046] Those skilled in the art will appreciate that the above-mentioned computer may also include more or fewer components, or a combination of certain components, or different arrangements of components, which is not limited in the embodiments of the present application.

[0047] like Figure 1 As shown, the navigation method of a spacecraft in an unknown gravity field provided in an embodiment of the present application is used to navigate the spacecraft during the process of the spacecraft orbiting a central celestial body with unknown gravity field parameters. The following describes the navigation method of the spacecraft in an unknown gravity field by taking a moment of the spacecraft in the above process as an example. Exemplarily, the navigation method of the spacecraft in an unknown gravity field includes S101-S103:

[0048] S101, determining the estimated error of the motion state of the spacecraft and the estimated error of the gravity field parameters of the central celestial body according to the motion state information of the spacecraft at the last moment and the gravity field parameters of the central celestial body predicted at the last moment;

[0049] In the embodiment of the present application, the motion state information of the above-mentioned spacecraft includes the position and speed of the spacecraft; the gravity field parameters of the central celestial body include the field harmonic coefficients of the gravity field of the central celestial body; it can be understood that the field harmonic coefficients refer to the coefficients of different orders and degrees in the spherical harmonic function expansion of the gravitational potential of the central celestial body, which are used to describe the degree to which the geometric shape of the central celestial body deviates from the standard sphere;

[0050] It should be noted that the above-mentioned spacecraft motion state estimation error refers to the error between the motion state information of the spacecraft at the current moment estimated by the motion state information of the spacecraft at the previous moment and the gravity field parameters of the central celestial body predicted at the previous moment, and the motion state information of the spacecraft at the current moment under actual circumstances;

[0051] The above-mentioned central celestial body gravity field parameter estimation error refers to the error between the gravity field parameters of the central celestial body predicted at the last moment and the gravity field parameters of the central celestial body predicted at the last moment, which are estimated from the motion state information of the spacecraft at the last moment and the motion state information of the spacecraft at the current moment;

[0052] The detailed process of determining the estimation error of the motion state of the spacecraft and the estimation error of the gravity field parameters of the central celestial body is introduced below;

[0053] Step 1: Use the covariance of the gravity field parameters of the central celestial body at the previous moment and forgetting factor , calculate the system estimation process error of the spacecraft ;

[0054] The system estimation process error of the above spacecraft The calculation formula is: ;

[0055] in, Indicates the current moment, Indicates the last moment, forgetfulness The value range of the sub is 0-1; the covariance of the gravity field parameters of the central celestial body at the last moment The gravitational field parameters of the central celestial body can be predicted based on the last moment Calculated, The calculation formula of is not described here;

[0056] It can be understood that the value of the covariance of the gravity field parameters of the central celestial body predicted at the initial moment can be set based on experience. For example, the value of the covariance of the gravity field parameters of the central celestial body predicted at the initial moment can be set to diag(1,1,…,1);

[0057] Step 2: Use the system to estimate process error and the covariance of the gravitational field parameters of the central celestial body at the previous moment , calculate the covariance estimate of the gravity field parameters of the central body ;

[0058] The estimated covariance of the gravitational field parameters of the above central celestial body The calculation formula is ,in ; , Represents a vector collection E The i column vector elements;

[0059] Optionally, a vector collection ;

[0060] Step 3: Estimation of the covariance of the gravity field parameters of the central celestial body and the spacecraft's motion status information at the previous moment , calculate the observation data of the spacecraft's motion state information at the current moment i Predicted values ;

[0061] The observation data of the spacecraft's motion state information at the current moment i Predicted values Satisfies the following formula:

[0062] ;

[0063] in, , , represents the position of the spacecraft at the last moment, represents the speed of the spacecraft at the last moment, represents the gradient operator, It represents the gravitational potential energy of the central celestial body at the last moment. The gravitational field parameters of the central celestial body predicted at the last moment include the field harmonic coefficients of the gravitational field of the central celestial body at the last moment. represents the system error matrix of the spacecraft, which is a known quantity. Represents the predicted value ; Optionally, =12;

[0064] It should be noted that the above It is obtained by integrating and discretizing the dynamic model of the spacecraft. The above integration and discretization process will not be repeated here; the dynamic model of the spacecraft satisfies the following formula:

[0065] ;

[0066] in, represents the state function, represents the observation error matrix of the spacecraft (known quantity), represents the position of the spacecraft, is the speed of the spacecraft, and is a vector;

[0067] In one implementation, the gravitational potential energy of the central celestial body satisfies the following formula:

[0068] ;

[0069] Among them, the gravity field parameters of the central celestial body are , The field harmonic coefficients representing the gravity field of the central celestial body at the last moment; represents the gravitational constant of the central celestial body, represents the distance from the spacecraft to the center of mass of the central celestial body, and , represents the equatorial radius of the central celestial body, and They represent the geographical longitude and latitude of the spacecraft relative to the central celestial body, Indicates the last moment n Step m Since Legendre function belongs to the common knowledge in the technical field, the expansion form of Legendre function will not be described here;

[0070] Step 4: Pass The first observation data of the spacecraft's motion state information at the current moment i Predicted values Calculate the gravity field parameter estimation error ; Through the observation data of the spacecraft's motion state information at the current moment i Predicted values And the gravity field parameters of the central celestial body predicted at the last moment , calculate the motion state estimation error ;

[0071] Motion state estimation error Satisfies the following formula:

[0072] ;

[0073] in, , , represents the transposed matrix, ;

[0074] Gravity field parameter estimation error Satisfies the following formula:

[0075] ;

[0076] S102, predicting the gravity field parameters of the central celestial body at the current moment according to the motion state estimation error, the gravity field parameter estimation error and the observation data of the motion state information of the spacecraft at the current moment;

[0077] The following is a detailed description of the process of predicting the gravity field parameters of the central celestial body at the current moment;

[0078] Step 1: Calculate the gain of the gravity field parameter at the current moment through the observation error and the estimation error; the gain of the gravity field parameter at the current moment satisfies the following formula:

[0079] ;

[0080] in, Indicates the gain of the gravity field parameters of the central celestial body at the current moment;

[0081] Step 2: The gravity field parameters of the central celestial body predicted at the last moment , the gain of the gravity field parameters of the central celestial body at the current moment , observation data of the spacecraft's motion state information at the current moment As well as the above , calculate the gravity field parameters of the central celestial body predicted at the current moment ;

[0082] The gravity field parameters of the central celestial body predicted at the current moment satisfy: ;

[0083] From this, we can know that the gravity field parameters of the central celestial body predicted by S101-S102 at the previous moment The precision and update of the gravity field parameters of the central celestial body predicted at the current moment are obtained. ; The gravity field parameters of the central celestial body predicted at the current moment obtained through S101-S102 The gravitational field parameters of the central celestial body predicted at the previous moment Compared with the previous prediction, it is closer to the actual operation of the spacecraft, so it can be inferred that the gravity field parameters of the central celestial body predicted at the current moment The information on the spacecraft's motion state at the current moment is also more accurate;

[0084] S103, using the gravity field parameters of the central celestial body predicted at the current moment and the motion state information of the spacecraft at the previous moment, filtering the observation data of the motion state information of the spacecraft at the current moment, obtaining the motion state information of the spacecraft at the current moment, so as to realize navigation of the spacecraft;

[0085] It can be understood that the motion state information of the spacecraft at the current moment is used to represent the real motion state of the spacecraft at the current moment; the observation data of the motion state information of the spacecraft at the current moment is obtained by observing the motion state of the spacecraft at the ground observation station, and there is a certain error between the real motion state of the spacecraft; therefore, by using the gravity field parameters of the central celestial body predicted at the current moment and the motion state information of the spacecraft at the previous moment, the observation data of the motion state information of the spacecraft at the current moment is filtered, and the error between the observation data of the motion state information of the spacecraft at the current moment and the real motion state of the spacecraft can be removed, so as to obtain more accurate motion state information of the spacecraft at the current moment;

[0086] The detailed process of S103 is described below;

[0087] Step 1: Use the motion state information of the spacecraft at the last moment , the covariance of the spacecraft's motion state information at the previous moment , calculate the estimated value of the spacecraft's motion state information at the previous moment;

[0088] The estimated value of the motion state information of the spacecraft at the last moment satisfies the following formula:

[0089] ;

[0090] in, The first one represents the motion state information of the spacecraft at the last moment. i An estimated value, represents the number of estimates, and optionally, =12; For the definition and value of , refer to the relevant description of step 2 in S102;

[0091] Step 2: Based on the estimated value of the spacecraft's motion state information at the previous moment and the gravity field parameters of the central celestial body predicted at the current moment , calculate the predicted value of the motion state information of the spacecraft at the current moment;

[0092] The predicted value of the spacecraft's motion state information at the current moment satisfies the following formula:

[0093] ;

[0094] in, The first one represents the motion state information of the spacecraft at the current moment. i Prediction values, , this formula is obtained by integrating and discretizing the dynamic model of the spacecraft. The above integration and discretization process will not be described in detail here;

[0095] Step 3: Calculate the estimated value of the motion state information of the spacecraft at the current moment and the square root of the state variance matrix through the predicted value of the motion state information of the spacecraft at the current moment;

[0096] The estimated value of the spacecraft's motion state information at the current moment and the square root of the state variance matrix Satisfy the following two formulas respectively:

[0097] ;

[0098] in, is the spacecraft system error matrix The square root of By calculating the spacecraft's system error matrix Perform Cholesky decomposition to obtain; , It represents the triangular matrix obtained by performing QR decomposition on the matrix in the brackets;

[0099] Step 4: Use the estimated value of the spacecraft's motion state information at the current moment and the state variance matrix The square root of the current spacecraft motion state information is calculated to obtain the estimated value of the observation data and the square root of the innovation variance matrix ;

[0100] The estimated value of the observation data of the spacecraft's motion state information at the current moment Satisfies the following formula:

[0101] ;

[0102] in, , , It is obtained by integrating and discretizing the dynamic model of the spacecraft. The above integration and discretization process will not be described in detail here; ;

[0103] The square root of the above innovation variance matrix is Satisfies the following formula:

[0104] ;

[0105] ; is the spacecraft observation error matrix The square root of By using the spacecraft observation error matrix Perform Cholesky decomposition to obtain;

[0106] Step 5: Estimated value of the spacecraft's motion state information at the current moment , the predicted value of the spacecraft's motion state information at the current moment, and , calculate the covariance of the spacecraft's motion state information and the observed value of the motion state information ;

[0107] The covariance between the motion state information of the above spacecraft and the observed value of the motion state information satisfy: ,in, ;

[0108] Step 6: The covariance between the spacecraft's motion state information and the observed value of the motion state information as well as , calculate the gain of the spacecraft's motion state information at the current moment ;

[0109] The gain of the above-mentioned spacecraft's motion state information at the current moment Satisfies the following formula:

[0110] ;

[0111] Step 7: Gain of the spacecraft's motion state information at the current moment , the spacecraft observation error matrix The square root of , as well as , calculate the covariance of the motion state information of the spacecraft at the current moment;

[0112] The covariance of the motion state information of the spacecraft at the current moment satisfies the following formula:

[0113] ;

[0114] in, ;

[0115] Step 8: Estimated value of the spacecraft's motion state information at the current moment , the gain of the spacecraft's motion state information at the current moment , observation data of the spacecraft's motion state information at the current moment and the estimated value of the observed data of the spacecraft's motion state information at the current moment , calculate the motion state information of the spacecraft at the current moment;

[0116] The motion state information of the spacecraft at the current moment satisfies the following formula:

[0117] ;

[0118] In summary, in the navigation method of a spacecraft under an unknown gravity field provided in the embodiment of the present application, the motion state information of the spacecraft at the last moment (i.e., the precise positioning of the spacecraft at the last moment) and the predicted gravity field parameters of the central celestial body are used to determine the motion state estimation error of the spacecraft and the gravity field parameter estimation error of the central celestial body; after obtaining the above two errors, the gravity field parameters of the central celestial body predicted at the current moment are predicted by the above two errors, and then the gravity field parameters of the central celestial body predicted at the current moment, the motion state information of the spacecraft at the last moment, and the observation data of the motion state information of the spacecraft at the current moment are used to jointly realize the navigation of the spacecraft under the unknown gravity field, and determine the motion state information of the spacecraft at the current moment (i.e., the precise positioning of the spacecraft at the current moment). In summary, the navigation method can continuously and in real time predict and update the gravity field parameters of the central celestial body according to the motion state information of the spacecraft and the observation data of the motion state information during the navigation of the spacecraft. Therefore, in this process, not only can the gravity field parameters of the central celestial body be continuously refined, but also the spacecraft can be navigated by the refined gravity field parameters when the observation data of the spacecraft on the detection target is insufficient.

[0119] Accordingly, an embodiment of the present application provides a navigation device for a spacecraft in an unknown gravity field, which is used to navigate the spacecraft while the spacecraft is orbiting a central celestial body with unknown gravity field parameters, such as Figure 2 As shown, the navigation device includes an error determination module 501 , a gravity field parameter determination module 502 and a navigation module 503 .

[0120] The error determination module 501 is used to determine the estimated error of the motion state of the spacecraft and the estimated error of the gravity field parameters of the central celestial body according to the motion state information of the spacecraft at the last moment and the gravity field parameters of the central celestial body predicted at the last moment. For example, the error determination module 501 is used to implement S101 of the above navigation method.

[0121] The gravity field parameter determination module 502 is used to predict the gravity field parameters of the central celestial body at the current moment according to the motion state estimation error, the gravity field parameter estimation error and the observation data of the motion state information of the spacecraft at the current moment. For example, the gravity field parameter determination module 502 is used to implement S102 of the above navigation method.

[0122] The navigation module 503 is used to filter the observation data of the motion state information of the spacecraft at the current moment by using the gravity field parameters of the central celestial body at the current moment and the motion state information of the spacecraft at the previous moment, so as to obtain the motion state information of the spacecraft at the current moment, so as to realize navigation of the spacecraft. For example, the navigation module 503 is used to realize S103 of the above navigation method.

[0123] Each module of the above navigation device can also be used to execute other steps in the above method embodiment. All relevant contents involved in the above method embodiment can be referred to the functional description of the corresponding functional module, which will not be repeated here.

[0124] The embodiment of the present application also provides an electronic device, including: a processor and a memory coupled to the processor; the memory is used to store computer instructions, and when the electronic device is running, the processor executes the computer instructions stored in the memory, so that the electronic device executes the method in the above embodiment. Among them, the processor can implement the above error determination module 501, gravity field parameter determination module 502 and navigation module 503; the above memory can also be used to store motion state estimation error, gravity field parameter estimation error, gravity field parameter and motion state information.

[0125] An embodiment of the present application further provides a computer-readable storage medium, which includes a computer program. When the computer program runs on a computer, the method described in the above embodiment is executed.

[0126] An embodiment of the present application further provides a computer program product, which includes computer program instructions. When the computer program instructions are run on a computer, the method described in the above embodiment is executed.

[0127] The various embodiments in this specification are described in a progressive manner, and the same or similar parts between the various embodiments can be referenced to each other, and each embodiment focuses on the differences from other embodiments.

[0128] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, rather than to limit it. Although the present invention has been described in detail with reference to the aforementioned embodiments, those skilled in the art should understand that they can still modify the technical solutions described in the aforementioned embodiments, or make equivalent replacements for some of the technical features therein. However, these modifications or replacements do not deviate the essence of the corresponding technical solutions from the spirit and scope of the technical solutions of the embodiments of the present application.

Claims

1. A navigation method for a spacecraft in an unknown gravity field, characterized in that: For navigating the spacecraft during the process of the spacecraft orbiting a central celestial body with unknown gravitational field parameters, the method comprises: Determining an estimation error of the motion state of the spacecraft and an estimation error of the gravity field parameters of the central celestial body according to the motion state information of the spacecraft at the last moment and the gravity field parameters of the central celestial body predicted at the last moment; According to the motion state estimation error, the gravity field parameter estimation error and the observation data of the motion state information of the spacecraft at the current moment, the gravity field parameters of the central celestial body at the current moment are predicted; the gravity field parameters of the central celestial body predicted at the current moment satisfy: Among them, k represents the current moment, k-1 represents the previous moment, ω k represents the gravity field parameters of the central celestial body predicted at the current moment, ω k-1 represents the gravity field parameters of the central celestial body predicted at the previous moment, represents the gain of the gravity field parameter, the gain of the gravity field parameter Satisfy the following formula; represents the motion state estimation error, represents the estimation error of the gravity field parameters; z k Observation data representing the motion state information of the spacecraft at the current moment, Z k,i represents the ith predicted value of the observed data of the motion state information of the spacecraft at the current moment, Z k,i It is obtained by predicting the motion state information of the spacecraft at the previous moment; n represents the number of predicted values ​​of the observed data of the motion state information of the spacecraft at the current moment; Using the gravity field parameters of the central celestial body predicted at the current moment and the motion state information of the spacecraft at the previous moment, the observation data of the motion state information of the spacecraft at the current moment is filtered to obtain the motion state information of the spacecraft at the current moment, so as to realize navigation of the spacecraft; wherein the motion state information of the spacecraft at the current moment is used to represent the real motion state of the spacecraft at the current moment.

2. A method for navigating a spacecraft in an unknown gravity field as claimed in claim 1, characterized in that: The motion state estimation error satisfies the following formula: Where k represents the current time, (*) T represents the transposed matrix; k-1 represents the previous moment, Z k,i represents the ith predicted value of the observed data of the motion state information of the spacecraft at the current moment, Z k,i It is obtained by predicting the motion state information of the spacecraft at the previous moment, ω k-1 represents the gravity field parameters of the central celestial body predicted at the previous moment, and ω k-1 =(C nm,k-1 ,S nm,k-1 ) T , C nm,k-1 and S nm,k-1 The field harmonic coefficients representing the gravitational field of the central celestial body at the previous moment; Represents the error in the predicted value of the observation data of the spacecraft at the current moment.

3. A method for navigating a spacecraft in an unknown gravity field as claimed in claim 2, characterized in that: The predicted value Z of the observation data of the spacecraft at the current moment k,i , satisfying the following formula; Z k,i =f(x k-1 ,Ω k,i ),i=1,...,n in, x k-1 represents the motion state information of the spacecraft at the last moment, and r k-1 represents the position of the spacecraft at the last moment, v k-1 represents the speed of the spacecraft at the last moment; Ω k,i represents the covariance estimate of the gravity field parameters of the central body; represents the gradient operator, U(ω k-1 ) represents the gravitational potential energy of the central celestial body at the previous moment, and Q represents the system error matrix of the spacecraft.

4. The navigation method of a spacecraft in an unknown gravity field as claimed in claim 1, characterized in that: The gravity field parameter estimation error satisfies the following formula: Where k represents the current time, (*) T represents the transposed matrix; Represents the error in the predicted value of the observation data of the spacecraft at the current moment.

5. A navigation device for a spacecraft in an unknown gravity field, based on the navigation method for a spacecraft in an unknown gravity field according to claim 1, characterized in that: used for navigating the spacecraft during the process of the spacecraft orbiting a central celestial body whose gravity field parameters are unknown, the device comprising an error determination module, a gravity field parameter determination module and a navigation module; The error determination module is used to determine the estimated error of the motion state of the spacecraft and the estimated error of the gravity field parameters of the central celestial body according to the motion state information of the spacecraft at the last moment and the gravity field parameters of the central celestial body predicted at the last moment; The gravity field parameter determination module is used to predict the gravity field parameters of the central celestial body at the current moment according to the motion state estimation error, the gravity field parameter estimation error and the observation data of the motion state information of the spacecraft at the current moment; The gravity field parameters of the central celestial body predicted at the current moment satisfy: Among them, k represents the current moment, k-1 represents the previous moment, ω k represents the gravity field parameters of the central celestial body predicted at the current moment, ω k-1 represents the gravity field parameters of the central celestial body predicted at the previous moment, represents the gain of the gravity field parameter, the gain of the gravity field parameter Satisfy the following formula; represents the motion state estimation error, represents the estimation error of the gravity field parameters; z k Observation data representing the motion state information of the spacecraft at the current moment, Z k,i represents the ith predicted value of the observed data of the motion state information of the spacecraft at the current moment, Z k,i It is obtained by predicting the motion state information of the spacecraft at the previous moment; n represents the number of predicted values ​​of the observed data of the motion state information of the spacecraft at the current moment; The navigation module is used to filter the observation data of the motion state information of the spacecraft at the current moment by using the gravity field parameters of the central celestial body predicted at the current moment and the motion state information of the spacecraft at the previous moment, so as to obtain the motion state information of the spacecraft at the current moment, so as to realize navigation of the spacecraft.

6. An electronic device, characterized in that: It includes a processor and a memory coupled to the processor; the memory is used to store computer instructions. When the electronic device is running, the processor executes the computer instructions stored in the memory, so that the electronic device executes the navigation method of a spacecraft in an unknown gravity field as described in claim 1.

7. A computer-readable storage medium, characterized in that: It includes computer program instructions, which, when executed by a computer, enable the computer to execute the navigation method for a spacecraft in an unknown gravity field as claimed in claim 1.

8. A computer program product, characterized in that It includes computer program instructions, and when the computer program instructions are executed on a computer, the computer is caused to execute the navigation method of a spacecraft in an unknown gravity field as claimed in claim 1.

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