On-orbit task restorative design method for satellite control system

Abstract

The invention relates to a satellite control system-oriented on-orbit task restorative design method, which comprises the following steps of: forming a task sequence expression according to the current task planning time sequence characteristics of a satellite, and defining each characteristic moment and a corresponding state thereof in the task sequence expression as a characteristic subset; the state S0 at the moment t0 and the relative time tx-t0 are stored, the feature subsets completed before tx do not need to be stored, the feature subsets after the moment tx do not need to be stored if the feature subsets can be obtained through calculation according to the feature subsets at the initial moment and the current observed quantity of the satellite, and otherwise, the state Si and the relative time ti-t0 of the feature subsets are stored; storing the target state Sn to form a minimum data set; and after the computer is reset, obtaining an initial moment feature subset and a current moment tx according to the delta tx and the S0, and reasoning to obtain a feature subset after the moment tx to realize the recovery of the task sequence expression. The problem of satellite task scene recovery under the system reset working condition is effectively solved.

Description

TECHNICAL FIELD

[0001] The application belongs to the technical field of autonomous intelligent management of spacecraft, and relates to an on-orbit task recovery design method for a satellite control system. BACKGROUND

[0002] With the continuous improvement of the complexity and importance of space missions, the requirements for spacecraft software have not only been to ensure the safety of spacecraft, but also to ensure the successful completion of on-orbit business, and even to ensure business continuity under certain fault conditions.

[0003] For the control system of a satellite, its task is to ensure that the attitude of the satellite points to a predetermined target to ensure the energy of the satellite or complete certain specific pointing tasks, or even complete the current attitude pointing task without interruption under certain fault conditions, such as computer reset. This requires the software running in the control system to have the ability of autonomous recovery of task scenarios. However, the computing system of the satellite control system is an embedded system with limited computing power and storage capacity, and it is not possible to save all data during software operation. Therefore, it is necessary to design certain task scenario descriptions and recovery strategies to reduce the dimensionality of the expression semantics of the task scenario, minimize the process data, and meet the limited storage space restrictions. At the same time, a certain task scenario inversion strategy is also needed to be designed, so that once a reset occurs, the current time and future task scenarios can be inferred from the minimum data set and certain strategies to ensure that the task scenario is quickly restored to the state before the reset.

[0004] Taking the task of a certain satellite control system as an example, the satellite usually works in an attitude stabilization state of earth orientation or sun orientation, and also adjusts to a certain target pointing attitude according to business needs. According to the task requirements, if a reset occurs during the attitude stabilization state, it needs to continue to maintain the attitude stabilization state after the reset, and if a reset occurs during the attitude maneuver, it needs to quickly recover to the attitude maneuver process to continue to complete the attitude adjustment and finally point to the target attitude.

[0005] According to existing mission requirements and satellite attitude control design, the satellite attitude adjustment process depends not only on the relative relationship between the current attitude and the target attitude, but also on the time and angular velocity of the real-time planning process. Once a reset occurs, all relevant data, including the target attitude and the computational workload related to trajectory planning, will be lost, making it impossible to maintain subsequent maneuver control. To solve this problem, it is necessary to save a series of elements such as the target attitude, the angular velocity related to the current maneuver planning, and the time of each stage in real time during the attitude adjustment process. However, due to the limitations of the satellite's overall computing power, storage capacity, and bus transmission bandwidth, it is impossible to save all of the above data completely. How to utilize limited storage and bus bandwidth resources to achieve on-site recovery of attitude maneuvering functions, so as to meet the mission requirement of continuing attitude maneuvering after a reset during the maneuver, requires overall planning at the system level, removing redundancy, extracting key core elements, retaining the backbone logic, and fully analyzing the implicit relationships and conditions using associated data and newly observed data after the reset, and performing secondary calculations and processing. Summary of the Invention

[0006] The technical problem solved by this invention is to overcome the shortcomings of the prior art and propose an on-orbit mission recovery design method for satellite control systems. Addressing the limited computing power and storage resources of onboard embedded computing systems, this method effectively solves the problem of satellite mission scenario recovery under system reset conditions. It can flexibly adapt to diverse satellite mission scenarios, meeting both the continuity requirements of routine operations and the business continuity and system security requirements after system reset during special operations, thereby improving the security, robustness, and autonomous intelligence of system management.

[0007] The solution to the technical problem of this invention is: an on-orbit mission recoverability design method for satellite control systems, comprising the following steps:

[0008] Based on the current mission planning timing characteristics of the satellite, a mission description template is designed. The mission sequence expression is formed in the manner of {{initial time t0, initial state S0}, {time t1, state S1}...{target time tn, target state Sn}}. Each feature time and its corresponding state in the mission sequence expression is defined as a feature subset.

[0009] For the task sequence expression, the association between each feature subset is extracted, and the minimum data set is formed according to the following principles: firstly, the state S0 at t0 is saved, and secondly, the relative time Δtx of the satellite current time tx relative to t0 is saved, and for the feature subset completed before tx, it is not necessary to be saved; for the feature subset after tx, if any feature subset can be derived from the initial time feature subset and the satellite current observation through calculation or reasoning, the feature subset does not need to be saved, otherwise the state Si of the feature subset and the relative time Δti relative to t0 are stored in the minimum data set; finally, the target state Sn of the target feature subset is stored in the minimum data set to form the final minimum data set.

[0010] After the computer reset occurs, the minimum data set is obtained, the {initial time t0, initial state S0} feature subset and the current time tx are obtained according to the read Δtx and S0, the time sequence position of the satellite in the task sequence expression is determined, and then the feature subset after tx is derived from the initial time feature subset and the satellite current observation according to the previously prepared strategy, so as to realize the recovery of the task sequence expression.

[0011] Further, the state Sx of each feature subset includes the key feature data of the current task at the current time, the condition expression for entering the feature subset, and the condition expression for entering the next feature subset.

[0012] Further, for the satellite attitude stable state period, the key feature data includes: current time, attitude angle and angular velocity, and orbit elements.

[0013] The condition expression for entering the feature subset includes: first entering the attitude stable state, or reaching the state update period.

[0014] The condition expression for entering the next feature subset includes: turning out of the attitude stable state and entering the attitude maneuver state, or reaching the state update period.

[0015] Further, for the satellite attitude maneuver state period, the key feature data includes: current time, current attitude, target attitude, this time attitude maneuver start time, acceleration segment / constant speed segment / deceleration segment start time, attitude maneuver direction, attitude maneuver maximum angular velocity, and attitude maneuver angular acceleration.

[0016] The condition expression for entering the feature subset includes: first entering the attitude maneuver state, or reaching the state update period.

[0017] The condition expression for entering the next feature subset includes: the switching time of the acceleration segment / constant speed segment / deceleration segment, or turning out of the attitude maneuver state and entering the attitude stable state, or reaching the state update period.

[0018] Further, the target time tn of the target feature subset is calculated or inferred from the previous subset:

[0019] 1) For the attitude stable state, if a state transition occurs, tn is the attitude transition time, tn is reset to the current time t0; if no state transition occurs, the target time of the target feature subset is updated according to a fixed period T, that is, tn is the time that increases by the period from t0;

[0020] 2) For the attitude maneuvering state, the target feature subset corresponds to the acceleration segment, the constant speed segment, and the deceleration segment, so the target time tn of the feature subset corresponds to the start time t0 of the acceleration segment, the start time tm1 of the constant speed segment, the start time tm2 of the deceleration segment, and the end time tm3 of the attitude maneuvering state;

[0021] As long as t0 is determined, tm1, tm2, and tm3 can be calculated by linear equations according to the maximum angular velocity of the attitude maneuvering and the angular acceleration of the attitude maneuvering.

[0022] Further, the feature subset after the time tx is inferred from the initial time feature subset and the current observation of the satellite as follows:

[0023] For the attitude stable state, the attitude angle and angular velocity in the target feature subset are directly obtained from the current observation, and the current orbit element is extrapolated from the orbit element at t0 according to Δtx:

[0024] Orbitelement(tx)=f(Orbitelement(t0),Δtx);

[0025] f(Orbitelement(t0),Δtx) is obtained by extrapolation from the two-body orbit dynamics equation;

[0026] For the attitude maneuvering state, the current stage is determined according to tm1, tm2, and tm3 in the target feature subset, including the acceleration segment, the constant speed segment, and the deceleration segment; the attitude angle and angular velocity at the current time are directly obtained from the observation, and then the remaining attitude maneuvering task is completed according to the saved target attitude and maneuvering direction according to the following steps:

[0027] 1) According to the stage at the current time, the attitude angular velocity control quantity is determined;

[0028] 2) According to the attitude maneuvering direction, the attitude angular velocity control polarity is determined;

[0029] 3) According to the difference between the current attitude angle and the target attitude angle, it is determined whether to stop the attitude maneuvering process.

[0030] Further, the complete minimum data set is periodically saved in a predetermined location at a preset fixed time interval.

[0031] The on-orbit mission recovery design system for a satellite control system comprises a core backbone data and logic extraction module, a semantic dimension reduction design module, and a scene recovery strategy module.

[0032] The core backbone data and logic extraction module is configured to design a mission description template according to a current mission planning time sequence characteristic of the satellite, form a mission sequence expression in the form of {{initial time t0, initial state S0}, {time t1, state S1},..., {target time tn, target state Sn}}, and define each characteristic time and corresponding state in the mission sequence expression as a characteristic subset.

[0033] The semantic dimension reduction design module is configured to extract the association between the characteristic subsets in the mission sequence expression, and form a minimum data set according to the following principles: first, save the state S0 at the time t0; second, save the relative time Δtx of the current time tx of the satellite relative to the time t0, and for the characteristic subsets completed before tx, there is no need to save them; for the characteristic subsets after the time tx, if any characteristic subset can be derived from the initial time characteristic subset and the current observation of the satellite through calculation or reasoning, the characteristic subset does not need to be saved, otherwise, the state Si of the characteristic subset and the relative time Δti of the characteristic subset relative to the time t0 are stored in the minimum data set; and finally, the target state Sn of the target characteristic subset is stored in the minimum data set to form the final minimum data set.

[0034] The scene recovery strategy module is configured to, after a computer reset, acquire the minimum data set, acquire the initial time characteristic subset {initial time t0, initial state S0} and the current time tx according to the read Δtx and S0, determine the time sequence position of the satellite in the mission sequence expression, and according to the previously prepared strategy, derive the characteristic subsets after the time tx through the initial time characteristic subset and the current observation of the satellite to realize the recovery of the mission sequence expression.

[0035] Further, the core backbone data and logic extraction module, the state Sx of each characteristic subset comprises key feature data of the current time of the current mission, condition expression for entering the characteristic subset, and condition expression for entering the next characteristic subset.

[0036] For the satellite attitude stable state, the key feature data comprises the current time, attitude angle and angular velocity, and orbit elements; the condition expression for entering the characteristic subset comprises first entering the attitude stable state or reaching the state update period; and the condition expression for entering the next characteristic subset comprises exiting the attitude stable state and entering the attitude maneuver state or reaching the state update period.

[0037] For the satellite attitude maneuver state, the key feature data includes: current time, current attitude, target attitude, this time attitude maneuver starting time, acceleration segment / constant speed segment / deceleration segment starting time, attitude maneuver direction, attitude maneuver maximum angular velocity, attitude maneuver angular acceleration; the condition expression for entering the current feature subset includes: first entering the attitude maneuver state, or reaching the state update period; the condition expression for turning into the next feature subset includes: acceleration segment / constant speed segment / deceleration segment switching time, or turning out of the attitude maneuver state into the attitude stable state, or reaching the state update period.

[0038] Further, the scene recovery strategy module infers the feature subset after the tx time through the initial time feature subset and the current observation of the satellite, and the method is as follows:

[0039] For the attitude stable state, the attitude angle and angular velocity in the target feature subset are directly obtained through the current observation, and the current orbit element is obtained through extrapolation based on the orbit element at the t0 time according to the Δtx:

[0040] Orbitelement(tx)=f(Orbitelement(t0),Δtx);

[0041] f(Orbitelement(t0),Δtx) is obtained through extrapolation of the two-body orbit dynamics equation;

[0042] For the attitude maneuver state, the current stage is determined according to the constant speed segment starting time tm1, the deceleration segment starting time tm2 and the attitude maneuver state ending time tm3 in the target feature subset, including the acceleration segment, the constant speed segment and the deceleration segment; the attitude angle and angular velocity at the current time are directly obtained through the observation, and then the remaining attitude maneuver task is completed according to the saved target attitude and maneuver direction according to the following steps:

[0043] 1) According to the stage at the current time, the attitude angular velocity control amount is determined;

[0044] 2) According to the attitude maneuver direction, the attitude angular velocity control polarity is determined;

[0045] 3) According to the difference between the current attitude angle and the target attitude angle, it is determined whether to stop the attitude maneuver process.

[0046] The beneficial effects of the present application compared with the prior art are:

[0047] (1) The on-orbit task recovery design method for a satellite control system proposed in the application can meet the business continuity after reset during the satellite attitude stable state and the business continuity after reset during the satellite attitude maneuvering state on the basis of ensuring the safety of the on-orbit operation of the satellite, and improves the robustness and on-orbit autonomous stable operation capability of the satellite.

[0048] (2) The on-orbit task recovery design method for a satellite control system proposed in the application effectively solves the problem of limited computing power, storage resources and bus bandwidth resources of the satellite-borne embedded computing system, and makes overall planning from the system level, removes redundancy, refines key core elements, retains backbone logic, analyzes the implicit relationship and conditions in the associated data and the newly observed data after reset, and performs secondary calculation and processing instead of original code saving, so that the maximum information mining and on-site recovery are realized with limited resources.

[0049] (3) The method and the software design thereof can be applied to autonomous intelligent management tasks in space tasks, and are flexible, easy to use, and strong in universality and portability. BRIEF DESCRIPTION OF DRAWINGS

[0050] Figure 1 A flowchart of the on-orbit task recovery design method for a satellite control system is shown in the application.

[0051] Figure 2 A satellite attitude maneuvering trajectory planning diagram of an embodiment of the application is shown in the application. DETAILED DESCRIPTION

[0052] The application will be further described below in combination with the drawings and embodiments.

[0053] The application proposes an on-orbit task recovery design method for a satellite control system, as shown in the accompanying drawings, which comprises the following steps: Figure 1

[0054] According to the current task planning time sequence characteristics of the satellite, a task description template is designed, a task sequence expression is formed in the manner of {{initial time t0, initial state S0}, {time t1, state S1}…{target time tn, target state Sn}}, and each characteristic time and the corresponding state in the task sequence expression are defined as a characteristic subset;

[0055] ​For the task sequence expression, the association between each feature subset is extracted, and the minimum data set is formed according to the following principles: firstly, the state S0 at t0 is saved, and secondly, the relative time Δtx (tx-t0) of the satellite current time tx relative to t0 is saved, and for the feature subset completed before tx, it is not necessary to be saved; for the feature subset after tx, if any feature subset can be derived from the initial time feature subset and the satellite current observation through calculation or reasoning, the feature subset does not need to be saved, otherwise the state Si of the feature subset and the relative time Δti (ti-t0) relative to t0 are stored in the minimum data set; finally, the target state Sn of the target feature subset is stored in the minimum data set, thereby forming the final minimum data set; the complete minimum data set is periodically saved at the agreed location according to the preset fixed time interval, so as to ensure that the data is not lost after the computer is reset.

[0056] After the computer is reset, the minimum data set is obtained, the initial time t0, the initial state S0 feature subset and the current time tx are obtained according to the read Δtx and S0, the time sequence position of the satellite in the task sequence expression is determined, and the feature subset after tx is derived through the initial time feature subset and the satellite current observation according to the previously formulated strategy, thereby realizing the recovery of the task sequence expression and ensuring the continuous business after the reset.

[0057] Wherein, the state Sx of each feature subset includes the key feature data of the current task at the current time, the condition expression for entering the feature subset, the condition expression for entering the next feature subset, etc.

[0058] Specifically, for the satellite attitude stable state period, the key feature data includes: current time, attitude angle and angular velocity, and orbit elements; the condition expression for entering the feature subset includes: first entering the attitude stable state, or reaching the state update period; and the condition expression for entering the next feature subset includes: exiting the attitude stable state and entering the attitude maneuver state, or reaching the state update period.

[0059] For the satellite attitude maneuver state period, the key feature data includes: current time, current attitude, target attitude, this time attitude maneuver start time, acceleration segment / constant speed segment / deceleration segment start time, attitude maneuver direction, attitude maneuver maximum angular velocity, and attitude maneuver angular acceleration; the condition expression for entering the feature subset includes: first entering the attitude maneuver state, or reaching the state update period; and the condition expression for entering the next feature subset includes: switching time of the acceleration segment / constant speed segment / deceleration segment, or exiting the attitude maneuver state and entering the attitude stable state, or reaching the state update period.

[0060] Wherein, the target time tn of the target feature subset can be obtained by calculation or reasoning according to the preceding subset:

[0061] 1) For the attitude stable state, if a state transition occurs (such as from the attitude stable state to the attitude maneuvering state), tn is the attitude transition time, tn is reset to the current time t0; if no state transition occurs, the target time of the target feature subset is updated according to a fixed period T, that is, tn is the time that increases by the period from t0, tn = t0 + T*n, n is a known quantity.

[0062] 2) For the attitude maneuvering state, the target feature subset corresponds to the acceleration segment, the constant speed segment, the deceleration segment, so the target time tn of the feature subset corresponds to the start time t0 of the acceleration segment, the start time tm1 of the constant speed segment, the start time tm2 of the deceleration segment, and the end time tm3 of the attitude maneuvering state.

[0063] As long as t0 is determined, tm1, tm2 and tm3 can be calculated by linear equations according to the maximum angular velocity of the attitude maneuvering and the angular acceleration of the attitude maneuvering.

[0064] Wherein, for the feature subset after tx time, if any feature subset can be obtained by calculation or reasoning from the initial time feature subset and the current observation of the satellite, the calculation method is as follows:

[0065] For the attitude stable state, the attitude angle and angular velocity in the target feature subset can be directly obtained from the current observation, and the current orbit element can be extrapolated from the orbit element at t0 time according to Δtx (tx-t0):

[0066] Orbitelement(tx) = f(Orbitelement(t0), Δtx);

[0067] f(Orbitelement(t0), Δtx) can be extrapolated using the conventional two-body orbit dynamics equation, which is well known in the field.

[0068] For the attitude maneuvering state, tm1, tm2 and tm3 in the target feature subset can be calculated by the foregoing method, so as to determine the current stage (acceleration segment, constant speed segment, deceleration segment); the attitude at the current time (including attitude angle and angular velocity) can be directly obtained from the observation, and then the remaining attitude maneuvering task can be completed according to the saved target attitude and maneuvering direction according to the following steps:

[0069] 1) According to the stage (acceleration segment, constant speed segment, deceleration segment) at the current time, the attitude angular velocity control quantity is determined;

[0070] 2) According to the attitude maneuver direction, the attitude angular velocity control polarity is determined;

[0071] 3) According to the difference between the current attitude angle and the target attitude angle, it is determined whether to stop the attitude maneuver process.

[0072] The application also provides an on-orbit task recovery design system for a satellite control system, which comprises a core backbone data and logic extraction module, a semantic dimension reduction design module, and a scene recovery strategy module.

[0073] The core backbone data and logic extraction module is used for designing a task description template according to the current task planning time sequence characteristics of the satellite, forming a task sequence expression in the manner of {{initial time t0, initial state S0}, {time t1, state S1}…{target time tn, target state Sn}}, and defining each characteristic time and its corresponding state as a characteristic subset in the task sequence expression.

[0074] The semantic dimension reduction design module is used for extracting the correlation between the characteristic subsets in the task sequence expression, and forming a minimum data set according to the following principles: firstly, saving the state S0 at the time t0, secondly, saving the relative time Δtx (tx-t0) of the current time tx of the satellite relative to the time t0, and for the characteristic subsets completed before tx, there is no need to save; for the characteristic subsets after the time tx, if any characteristic subset can be obtained by calculation or reasoning from the initial time characteristic subset and the current observation of the satellite, the characteristic subset does not need to be saved, otherwise, the state Si of the characteristic subset and the relative time Δti (ti-t0) of the characteristic subset relative to the time t0 are stored in the minimum data set; finally, the target state Sn of the target characteristic subset is stored in the minimum data set, thereby forming the final minimum data set.

[0075] The scene recovery strategy module is used for obtaining the minimum data set after the computer is reset, obtaining the characteristic subset of the initial time t0 and the initial state S0 and the current time tx according to the read Δtx and S0, determining the time sequence position of the satellite in the task sequence expression, and reasoning the characteristic subsets after the time tx through the initial time characteristic subset and the current observation of the satellite according to the previously prepared strategy, thereby realizing the recovery of the task sequence expression.

[0076] Embodiment 1

[0077] During the satellite on-orbit operation, the satellite is required to point to a fixed target according to the mission requirements, and thus a pointing maneuver adjustment task is started to gradually adjust the current earth-pointing attitude to the target pointing attitude. The mainstream design of the current satellite attitude control is a maneuver trajectory planning algorithm based on trajectory planning. The algorithm sets a maximum maneuver angular velocity for the satellite attitude maneuver process, and divides the maneuver process into three stages of acceleration, constant speed and deceleration. The algorithm feature is that the angular acceleration during the maneuver is constant. The angular velocity and angular acceleration curve shape obtained according to the planning algorithm is as shown in Figure 2

[0078] In Figure 2 , t0 is the start time of the maneuver, [t0, tm1) is the acceleration segment, [tm1, tm2) is the sliding segment, and [tm1, tm3) is the deceleration segment. During the maneuver, the angular acceleration, angular velocity and angle at each time are planned in real time according to the current time interval, which are used as the input for closed-loop control.

[0079] This embodiment carries out mission recovery design for the above-mentioned mission according to the on-orbit mission recovery design method for satellite control system.

[0080] The first step of the process is to perform core backbone data and logic extraction

[0081] According to the current mission planning time sequence characteristics of the satellite, a mission description template is designed, and a mission sequence expression is formed in the form of {{initial time t0, initial state S0}, {time t1, state S1}…{target time tn, target state Sn}}. In this embodiment, the satellite attitude maneuver process can be divided into two steps of initialization setting and real-time planning:

[0082] (1) Initialization setting

[0083] According to the current attitude and target attitude, the attitude difference is obtained, and then the maximum maneuver angular velocity and maximum maneuver angular acceleration are set to obtain the longest time of the acceleration and deceleration segments, and three turning points tm1, tm2 and tm3 are obtained.

[0084] (2) Real-time planning

[0085] According to tm1, tm2 and tm3 obtained by the initialization setting, the current angular acceleration ar is calculated in real time according to the trapezoidal trajectory in the above figure, and the current angular velocity and angle are integrated.

[0086] Based on the above, the core backbone data of the satellite attitude trajectory planning algorithm can be extracted as follows:

[0087] ​Initialization setting of the core backbone data "six elements" of the task sequence list expression: start time t0, start time attitude qB, target attitude qT, maximum maneuver angular velocity wmax, maximum maneuver angular acceleration amax and maneuver direction e.

[0088] Real-time planning of the core backbone data "four elements" of the task sequence list expression: turning points tm1, tm2, tm3 and current time t.

[0089] The core backbone data are fitted into the template of claim 2, and are defined as follows:

[0090] {t0 = maneuver start time, S0 = {qB at t0, maximum maneuver angular velocity wmax, maximum maneuver angular acceleration amax, maneuver direction e}}

[0091] {t1 = uniform speed segment start time tm1, S1 = uniform speed segment angular velocity wmax}

[0092] {t2 = deceleration segment start time tm2, S2 = uniform speed segment initial angular velocity wmax}

[0093] {t3 = maneuver to position time tm3, S3 = target attitude qT}

[0094] In summary, the task sequence list expression of the embodiment is formed, containing four characteristic subsets.

[0095] The second step of the flow: performing semantic dimension reduction design to form the minimum data set

[0096] Initial state saving: saving the state S0 at t0, and then saving the relative time Δtx (tx-t0) of the satellite current time tx relative to t0.

[0097] Process state semantic dimension reduction: for the characteristic subsets completed before tx, there is no need to save them; for the characteristic subsets after tx, if the characteristic subsets can be calculated or inferred from the initial time characteristic subsets and the satellite current observations, the characteristic subsets do not need to be saved, otherwise the state Si of the characteristic subsets and the relative time Δt(ti-t0) of the state Si relative to t0 are stored in the minimum data set.

[0098] Target state saving: storing the target state Sn of the target characteristic subset in the minimum data set, thereby forming the final minimum data set.

[0099] Finally, the complete minimum data set is periodically saved at a predetermined location at a certain fixed time interval, so as to ensure that the data is not lost after the computer is reset.

[0100] Fitted into the embodiment, the semantic dimension reduction design of the core backbone data of the task sequence list expression of the current attitude maneuvering task is shown in the following table:

[0101]

[0102] Based on the above semantic dimension reduction design, a large number of floating point numbers are compressed, only less than 7 bytes of minimum data set is reserved, other information can be obtained by resetting the minimum data set, and other related information is calculated again, which greatly improves the storage efficiency.

[0103] The third step of the process is to execute the scene recovery strategy

[0104] Minimum data set acquisition: after the computer is reset, the minimum data set is obtained according to the agreed address and protocol format.

[0105] Task scene inversion: according to the obtained Δtx and S0, the feature subset of {initial time t0, initial state S0} and the current time tx are obtained, the time sequence position of the satellite in the task sequence expression is determined, and then according to the previously prepared strategy, the subsequent feature subset after tx is inferred through the initial time feature subset and the satellite current observation, so as to realize the recovery of the task sequence expression, and ensure the continuous business after the reset.

[0106] In this embodiment, the reset scene recovery strategy of the attitude maneuver task is designed as follows:

[0107] After resetting, the minimum data set is obtained, the initial time t0 of the attitude maneuver task is obtained, the attitude qB, target attitude qT and maneuver direction e at t0 are obtained, and then the maximum maneuver angular velocity wmax and maximum maneuver angular acceleration amax are obtained according to the existing binding value, so as to recover the feature subset {t0, S0};

[0108] According to the existing trajectory planning strategy, the feature subsets {t1, S1}, {t2, S2} and {t3, S3} are further generated;

[0109] According to the current time, the position of the planning trajectory before the system reset is determined, and the subsequent feature subset obtained by the foregoing inference is used to finally realize the recovery of the attitude maneuver task sequence expression, and ensure the continuity of the maneuver task and the safety of the satellite.

[0110] Although the present application has been disclosed with the above preferred embodiments, it is not intended to limit the present application, and any person skilled in the art can make possible changes and modifications to the technical solutions of the present application by using the disclosed methods and technical contents without departing from the spirit and scope of the present application, therefore, any simple modification, equivalent change and modification made to the above embodiments according to the technical essence of the present application, which does not deviate from the technical solutions of the present application, all belong to the protection scope of the technical solutions of the present application.

[0111] Those aspects of the application which are not specifically described in the specification are well known to those skilled in the art.

Claims

1. A method for on-orbit mission recoverability design for satellite control systems, characterized in that, The method comprises the following steps: According to the current task planning timing characteristics of the satellite, a task description template is designed, a task sequence list expression is formed in the manner of {{initial time t0, initial state S0}, {time t1, state S1}…{target time tn, target state Sn}}, and each characteristic time and corresponding state in the task sequence list expression are defined as a characteristic subset; For the task sequence list expression, the correlation between each characteristic subset is extracted, and a minimum data set is formed according to the following principles: firstly, the state S0 at the time t0 is saved, and secondly, the relative time Δtx of the current time tx of the satellite relative to the time t0 is saved, and for the characteristic subsets completed before tx, it is not necessary to be saved again; for the characteristic subsets after the time tx, if any characteristic subset can be derived from the initial time characteristic subset and the current observation of the satellite through calculation or reasoning, the characteristic subset does not need to be saved, otherwise the state Si of the characteristic subset and the relative time Δti of the state Si relative to the time t0 are stored in the minimum data set; finally, the target state Sn of the target characteristic subset is stored in the minimum data set to form the final minimum data set; After the computer is reset, the minimum data set is acquired, the initial time characteristic subset and the current time tx are acquired according to the read Δtx and S0, the timing position of the satellite in the task sequence list expression is determined, and the characteristic subsets after the time tx are derived through the initial time characteristic subset and the current observation of the satellite according to the previously prepared strategy to realize the recovery of the task sequence list expression.

2. The method of claim 1, wherein, The state Sx of each characteristic subset comprises the key characteristic data of the current task at the current time, the condition expression for entering the characteristic subset and the condition expression for transferring to the next characteristic subset.

3. The method of claim 2, wherein, For the satellite attitude stable state, the key characteristic data comprises the current time, the attitude angle and angular velocity and the orbit element; The condition expression for entering the characteristic subset comprises the first entering of the attitude stable state or the reaching of the state update period; The condition expression for transferring to the next characteristic subset comprises the transferring out of the attitude stable state and entering the attitude maneuver state or the reaching of the state update period.

4. The on-orbit mission recovery design method for satellite control system according to claim 3, wherein, For the satellite attitude maneuver state, the key characteristic data comprises the current time, the current attitude, the target attitude, the starting time of the current attitude maneuver, the starting time of the acceleration / constant speed / deceleration section, the attitude maneuver direction, the maximum angular velocity of the attitude maneuver and the angular acceleration of the attitude maneuver; The condition expression for entering the characteristic subset comprises the first entering of the attitude maneuver state or the reaching of the state update period; The condition expression for transferring to the next characteristic subset comprises the switching time of the acceleration / constant speed / deceleration section, the transferring out of the attitude maneuver state and entering the attitude stable state or the reaching of the state update period.

5. The on-orbit mission recovery design method for satellite control system according to claim 4, wherein, The target time tn of the target characteristic subset is derived from the previous subset through calculation or reasoning: 1) for the attitude stable state, if the state conversion occurs, tn is the attitude conversion time, and tn is reset as the current time t0; if the state conversion does not occur, the target time of the target characteristic subset is updated according to the fixed period T, that is, tn is the time which is increased by the period from t0; 2) For the attitude maneuver state, the target feature subset corresponds to the acceleration segment, the uniform speed segment, and the deceleration segment, so the feature subset target time tn corresponds to the acceleration segment starting time t0, the uniform speed segment starting time tm1, the deceleration segment starting time tm2, and the attitude maneuver state ending time tm3 respectively; As long as t0 is determined, tm1, tm2 and tm3 can be calculated through linear equations according to the maximum angular velocity of the attitude maneuver and the angular acceleration of the attitude maneuver.

6. The method of claim 5, wherein, The method for reasoning the feature subset after tx time through the initial time feature subset and the current observation of the satellite is as follows: For the attitude stable state, the attitude angle and angular velocity in the target feature subset are directly obtained through the current observation, and the current orbit element is extrapolated from the orbit element at t0 according to Δtx: Orbitelement(tx) = f(Orbitelement(t0), Δtx); f(Orbitelement(t0), Δtx) is obtained through extrapolation of the two-body orbit dynamics equation; For the attitude maneuver state, the current stage is determined according to tm1, tm2 and tm3 in the target feature subset, including the acceleration segment, the uniform speed segment, and the deceleration segment; the attitude angle and angular velocity at the current time are directly obtained through the observation, and then the remaining attitude maneuver task is completed according to the saved target attitude and maneuver direction according to the following steps: 1) According to the stage at the current time, the attitude angular velocity control quantity is determined; 2) According to the attitude maneuver direction, the attitude angular velocity control polarity is determined; 3) According to the difference between the current attitude angle and the target attitude angle, it is determined whether to stop the attitude maneuver process.

7. The method of claim 1, wherein, According to the preset fixed time interval, the complete minimum data set is periodically saved in the agreed position.

8. An on-orbit mission recoverability design system for satellite control systems, characterized in that, It includes a core backbone data and logic extraction module, a semantic dimension reduction design module, and a scene recovery strategy module. The core backbone data and logic extraction module is used to design a task description template according to the current task planning time sequence characteristics of the satellite, form a task sequence expression in the form of {{initial time t0, initial state S0}, {time t1, state S1}…{target time tn, target state Sn}}, and define each feature time and its corresponding state in the task sequence expression as a feature subset; The semantic dimension reduction design module is used to extract the correlation between each feature subset in the task sequence expression, and form the minimum data set according to the following principles: first, save the t0 time state S0, second, save the relative time Δtx of the satellite current time tx relative to the t0 time, for the feature subsets completed before tx, there is no need to save; for the feature subsets after tx time, if any feature subset can be obtained through calculation or reasoning from the initial time feature subset and the current observation of the satellite, the feature subset does not need to be saved, otherwise the state Si of the feature subset and the relative time Δti relative to t0 are stored in the minimum data set; finally, the target state Sn of the target feature subset is stored in the minimum data set to form the final minimum data set; The scene recovery strategy module is configured to acquire a minimum data set after a computer reset, acquire a {initial time t0, initial state S0} feature subset and a current time tx according to the read Δtx and S0, determine a timing position of the satellite in a task sequence expression, and infer a feature subset after the time tx according to the initial time feature subset and satellite current observation to achieve recovery of the task sequence expression.

9. The on-orbit mission recovery design system for satellite control system according to claim 8, wherein, The core backbone data and logic extraction module, the state Sx of each feature subset includes key feature data of a current task at a current time, a condition expression for entering the feature subset, and a condition expression for switching to a next feature subset; For a satellite attitude stable state, the key feature data includes a current time, an attitude angle and angular velocity, and orbit elements; the condition expression for entering the feature subset includes a first entry into the attitude stable state or reaching a state update period; and the condition expression for switching to the next feature subset includes switching out of the attitude stable state into an attitude maneuver state or reaching the state update period. For a satellite attitude maneuver state, the key feature data includes a current time, a current attitude, a target attitude, a current attitude maneuver start time, an acceleration segment / constant speed segment / deceleration segment start time, an attitude maneuver direction, an attitude maneuver maximum angular velocity, and an attitude maneuver angular acceleration; the condition expression for entering the feature subset includes a first entry into the attitude maneuver state or reaching a state update period; and the condition expression for switching to the next feature subset includes a switching time of the acceleration segment / constant speed segment / deceleration segment, switching out of the attitude maneuver state into the attitude stable state, or reaching the state update period.

10. The on-orbit mission recovery design system for satellite control system according to claim 9, wherein, The scene recovery strategy module infers the feature subset after the time tx through the initial time feature subset and satellite current observation, and the method is as follows: For the attitude stable state, the attitude angle and angular velocity in the target feature subset are directly obtained through current observation, and current orbit elements are extrapolated based on orbit elements at the time t0 according to Δtx: Orbitelement(tx) = f(Orbitelement(t0), Δtx); f(Orbitelement(t0), Δtx) is obtained through a two-body orbit dynamics equation; For the attitude maneuver state, a current phase is determined according to the constant speed segment start time tm1, the deceleration segment start time tm2, and the attitude maneuver state end time tm3, including an acceleration segment, a constant speed segment, and a deceleration segment; the attitude angle and angular velocity at the current time are directly obtained through observation, and the remaining attitude maneuver task is continued according to the saved target attitude and maneuver direction according to the following steps: 1) determining an attitude angular velocity control amount according to a current phase; 2) determining an attitude angular velocity control polarity according to an attitude maneuver direction; 3) determining whether to stop the attitude maneuver process according to a difference between a current attitude angle and a target attitude angle.

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