Attitude pointing sun-avoiding maneuvering path predicting and planning method

Abstract

The invention relates to a method for predicting and planning an attitude pointing sun-avoiding maneuvering path, which is used for solving the problem of planning an attitude maneuvering path of a spacecraft with attitude pointing constraint. The method comprises two steps of maneuver prediction and attitude maneuver, if it is predicted that a conventional maneuver path can enable a pointing shaft to enter a prohibited area before attitude maneuver, a proper middle attitude is selected, a spacecraft is made to maneuver towards the middle attitude firstly, then the spacecraft is selected to maneuver towards a target attitude, and therefore a new maneuver path capable of avoiding the prohibited area is planned. Finally, the feasibility of the method is effectively proved through mathematical simulation analysis. According to the technical scheme, an effective and feasible method is provided for preventing the pointing shaft from entering the sun prohibited area in the attitude maneuver process.

Description

Technical Field

[0001] The present invention relates to a method for predicting and planning an attitude pointing anti-sun maneuver path, which is used to solve the problem of attitude maneuver path planning of a spacecraft with attitude pointing constraints, and belongs to the field of spacecraft control technology. This technical solution provides an effective and feasible method to prevent the pointing axis from entering the sun prohibited area during the attitude maneuver process. Background Art

[0002] When a spacecraft performs various space missions, it needs to reorient its attitude. The attitude maneuver process is usually physically constrained by payloads. For example, during the attitude maneuver process, to ensure the normal use of an optical sensor, strong light such as the sun should be avoided from entering the field of view. To ensure the normal use of the payload, attitude prohibition constraints must be established, but the attitude pointing constraints limit the attitude maneuver range of the spacecraft and increase the difficulty of spacecraft attitude maneuver. Therefore, it is particularly important to plan an attitude maneuver path of a spacecraft with attitude pointing constraints to avoid the payload from entering the prohibited area. Summary of the Invention

[0003] The technical problem solved by the present invention is: to ensure the normal use of an optical sensor during the attitude maneuver process and avoid the situation of sunlight entering the field of view, a method for predicting and planning an attitude pointing anti-sun maneuver path is proposed to solve the problem of attitude maneuver path planning of a spacecraft with attitude pointing constraints.

[0004] The technical solution adopted by the present invention is: a method for predicting and planning an attitude pointing anti-sun maneuver path, including:

[0005] Performing attitude maneuver prediction; the attitude maneuver prediction is to predict whether the maneuver path will cause the pointing axis to enter the prohibited area under the attitude prohibition constraint. If not, the path planning is completed using the initial attitude quaternion at the maneuver start time and the target attitude quaternion at the maneuver end time; otherwise, an intermediate attitude is selected, and the spacecraft is maneuvered towards the intermediate attitude first, and then opportunistically towards the target attitude to plan a new maneuver path that avoids the prohibited area.

[0006] After the attitude maneuver prediction is completed, attitude maneuver is performed according to the planned path, and no attitude maneuver is performed when the path planning fails during the attitude maneuver prediction process.

[0007] Preferably, the following method is used to predict whether the maneuver path will cause the pointing axis to enter the prohibited area:

[0008] Step S1, setting parameters: setting time parameters, spacecraft orbit parameters, the initial attitude quaternion at the maneuver start time, and the target attitude quaternion at the maneuver end time;

[0009] Step S2, according to the parameters set in Step S1, perform the operation from the initial attitude quaternion qbi_0 To the target attitude quaternion q xi Prediction of the maneuvering path. First, calculate the relative quaternion dq between the two _0 ;

[0010] Step S3: According to the relative quaternion dq _k-1 Calculate the predicted next attitude quaternion; from the current attitude quaternion q bi_k-1 And the limited relative quaternion dq lim_k-1 Calculate the next attitude quaternion q bi_k ;

[0011] Step S4: According to the attitude prediction result q in Step S3 bi_k Calculate the corresponding attitude transformation matrix A bi_k , and then according to A bi_k And the solar vector sun in the inertial system i_k Calculate the corresponding solar vector sun in the body system b_k , thereby calculating the angle θ between the +X axis in the body system and the solar vector _k ;

[0012] Step S5: If the angle between the +X axis and the solar vector is less than or equal to the constraint angle, i.e., θ _k ≤θ ys , it is predicted that the maneuvering path will cause the pointing axis to enter the prohibited area; otherwise, update the relative quaternion dq according to the current attitude quaternion and the target attitude quaternion _k , and sequentially loop through Step S3, Step S4, and Step S5 until the prediction of the entire attitude maneuvering process is completed.

[0013] Preferably, the selection of the intermediate attitude, planning a new maneuvering path to avoid the prohibited area by maneuvering the spacecraft to the intermediate attitude first and then opportunistically to the target attitude, includes:

[0014] Step S6: According to the initial attitude quaternion q set in Step S1 bi_0 And the target attitude quaternion q xi , calculate the intermediate attitude quaternion q mi ; Then, perform the prediction of the maneuvering path from the initial attitude quaternion q bi_0 To the intermediate attitude quaternion q mi , first calculate the relative quaternion dq between the two _0 ; Sequentially loop through Step S3, Step S4, Step S7, and Step S8;

[0015] Step S7: Opportunistically select to switch the intermediate attitude quaternion q mi To the target attitude quaternion q xi , and calculate and update the relative quaternion dq _k ; Otherwise, do not switch;

[0016] Step S8. Determine the new maneuver path prediction result according to the included angle in Step S4; if the included angle between the +X axis and the solar vector is less than or equal to the constraint angle, i.e., θ _k ≤θ lim , the path planning fails; otherwise, continue to loop until the prediction of the entire attitude maneuver process is completed.

[0017] Preferably, the intermediate attitude is selected in the following manner, i.e., according to the initial attitude quaternion q bi_0 and the target attitude quaternion q xi , calculate the intermediate attitude quaternion q mi ;

[0018]

[0019] wherein, represents the quaternion multiplication symbol, and inv() represents the quaternion inverse operation.

[0020] Preferably,

[0021] Preferably, the timing condition is to start judging the difference in the included angle between the +X axis of the system and the solar vector at the previous and current moments after a preset time is reached during the prediction process from the initial attitude to the intermediate attitude. When there is an included angle θ _k <θ _k-1 , start switching the target attitude and switch the intermediate attitude quaternion to the target attitude quaternion; k represents the next prediction beat.

[0022] Preferably, the preset time is determined according to the spacecraft attitude maneuver mode, which is 10 s for jet attitude maneuver and 100 s for flywheel attitude maneuver.

[0023] Preferably, during the attitude maneuver prediction process, the attitude maneuver is guided by setting a prediction flag; when the pointing axis does not enter the prohibited area during the entire prediction process, the prediction flag is in the default state 0, i.e., flag _yc =0; when a new maneuver path that avoids the prohibited area is successfully planned by selecting the intermediate attitude during the prediction process, the prediction flag is set to 1, i.e., flag _yc =1; when the path planning fails during the attitude maneuver prediction process, the prediction flag is set to 2, i.e., flag _yc =2.

[0024] Preferably, guiding the attitude maneuver using the prediction flag includes:

[0025] If flag _yc =0, directly maneuver from the initial attitude quaternion q bi_0 to the target attitude quaternion q xi . If flag_yc = 1, first start from the initial attitude quaternion q bi_0 to the intermediate attitude quaternion q mi for maneuver; opportunistically switch q mi to q xi , and start maneuvering towards the target attitude quaternion q xi ; if flag _yc = 2, stop maneuvering.

[0026] The advantages of the present invention compared with the prior art are as follows:

[0027] By predicting before attitude maneuver that the conventional maneuver path will cause the pointing axis to enter the prohibited area, the present invention selects a suitable intermediate attitude, enabling the spacecraft to first maneuver towards the intermediate attitude and then opportunistically maneuver towards the target attitude, thereby planning a new maneuver path that can avoid the prohibited area. This technical solution provides an effective and feasible method to prevent the pointing axis from entering the solar prohibited area during the attitude maneuver process.

[0028] (1) Under the attitude prohibited pointing constraint conditions, this technology adopts a pre-prediction method to flexibly select the maneuver path as needed;

[0029] (2) Compared with the current control methods relying on complex controllers such as potential functions and backstepping methods, the present invention can achieve attitude maneuver control of a spacecraft with attitude pointing constraints by using a reliable and simple PD controller. BRIEF DESCRIPTION OF THE DRAWINGS

[0030] Figure 1 is a flowchart of the attitude pointing solar avoidance maneuver path prediction and planning scheme provided by the present invention;

[0031] Figure 2 is a diagram of the target quaternion and maneuver process quaternion before avoidance provided by the present invention;

[0032] Figure 3 is a diagram of the angle between the +X axis of the present system and the solar vector before avoidance provided by the present invention;

[0033] Figure 4 is a diagram of the attitude control attitude angle before avoidance provided by the present invention;

[0034] Figure 5 is a diagram of the attitude control angular velocity before avoidance provided by the present invention;

[0035] Figure 6 is a diagram of the target quaternion and maneuver process quaternion after avoidance provided by the present invention.

[0036] Figure 7 is a diagram of the angle between the +X axis of the present system and the solar vector after avoidance provided by the present invention.

[0037] Figure 8 It is the diagram of the attitude angle in the attitude control after avoidance provided by the present invention.

[0038] Figure 9 It is the diagram of the angular velocity of the attitude in the attitude control after avoidance provided by the present invention. Specific embodiments

[0039] The following further describes the present invention in detail in conjunction with the appended Figures 1-9 drawings and specific embodiments.

[0040] When a spacecraft performs various space missions, it needs to reorient its attitude. The attitude maneuvering process is usually physically restricted by payloads. For example, during the attitude maneuvering process, to ensure the normal use of optical sensors, strong light such as sunlight should be avoided from entering the field of view. To ensure the normal use of the payload, attitude prohibition constraints must be established, but the attitude pointing constraints limit the attitude maneuvering range of the spacecraft and increase the difficulty of spacecraft attitude maneuvering. Currently, the attitude maneuvering method under attitude prohibition constraints usually introduces a Lyapunov function to limit the angular velocity tracking error and uses the backstepping method to complete the design of the attitude control law, which is often complex. The method proposed by the present invention is divided into two steps: maneuver prediction and attitude maneuvering. If it is predicted before the attitude maneuvering that the conventional maneuvering path will cause the pointing axis to enter the prohibited area, a suitable intermediate attitude is selected, so that the spacecraft first maneuvers to the intermediate attitude and then maneuvers to the target attitude opportunistically, thereby planning a new maneuvering path that can avoid the prohibited area. Finally, the feasibility of the method is effectively proved through mathematical simulation analysis. This technical solution provides an effective and feasible method to prevent the pointing axis from entering the sun prohibited area during the attitude maneuvering process.

[0041] The present invention can be used in the attitude maneuvering scenario of a spacecraft with attitude pointing constraints and can solve the problem that the pointing axis enters the sun prohibited area during the attitude maneuvering process. A method for predicting and planning an attitude pointing sun-avoiding maneuvering path, as Figure 1 shown, includes the following steps:

[0042] (1) Step S1: Set parameters: Set time parameters, spacecraft operating orbit parameters, the initial attitude quaternion at the starting moment of the maneuver, and the target attitude quaternion at the ending moment of the maneuver.

[0043] Set the attitude maneuvering conditions:

[0044] ① The starting time of the attitude maneuver;

[0045] ② The starting orbit of the attitude maneuver;

[0046] ③ The initial attitude quaternion;

[0047] ④ The target attitude quaternion;

[0048] (2) Step S2: According to the parameters set in Step S1, perform the prediction of the maneuver path from the initial attitude quaternion q bi_0 to the target attitude quaternion q xi . First, calculate the relative quaternion dq _0 .

[0049] Calculate the initial relative quaternion dq _0 :

[0050]

[0051] (3) Step S3: According to the relative quaternion dq _k-1 calculate the predicted next attitude quaternion. From the current attitude quaternion q bi_k-1 and the limited relative quaternion dq lim_k-1 , calculate the next attitude quaternion q bi_k .

[0052] First, limit the relative quaternion dq _k-1 to obtain dq lim_k-1 :

[0053]

[0054] Among them, the value of w is related to the actuator:

[0055] If jet control is performed: w = 2° / s; Δt = 1; the total number of prediction steps is 100 steps.

[0056] If flywheel control is performed: w = 0.2° / s; Δt = 1; the total number of prediction steps is 1000 steps.

[0057] Then, calculate the next attitude quaternion:

[0058]

[0059] (4) Step S4: According to the attitude prediction result q bi_k in Step S3, calculate the corresponding attitude transformation matrix A bi_k , and then according to A bi_k and the solar vector sun i_k in the inertial system, calculate the corresponding solar vector sun b_k in the body system, so as to calculate the angle θ _k between the +X axis of the body system and the solar vector.

[0060] First, convert the unit solar vector in the inertial system to the body system:

[0061] sun b_k = A bi_k · sun i_k , k = 1, 2...

[0062] Then, calculate the included angle between the system + X - axis x b =

[100] and the solar vector:

[0063] θ _k =acos<sun b_k ·x b >>, k = 1, 2...

[0064] (5) Step S5: Determine whether to set the intermediate attitude quaternion based on the included angle in Step S4. If the included angle between the + X - axis and the solar vector is less than or equal to the constraint angle, i.e., θ _k ≤θ ys , then proceed to Step S6; otherwise, calculate and update the relative quaternion dq _k according to the current attitude quaternion and the target attitude quaternion, and sequentially loop through Step S3, Step S4, and Step S5. Wait until dq _k ≤1e - 3 and jump to Step S9. At this time, the prediction flag is the default value 0, i.e., flag _yc =0.

[0065] If there is θ _k ≤θ ys during the prediction process of maneuvering along the current path, then jump to Step S6 and continue the prediction, that is, calculate the intermediate attitude quaternion and re - predict according to the new path.

[0066] If there is always θ _k >θ ys during the prediction process of maneuvering along the current path, then the prediction flag is the default value 0, i.e., flag _yc =0, and jump to Step S9 to end the prediction.

[0067] (6) Step S6: Calculate the intermediate attitude quaternion q bi_0 based on the initial attitude quaternion q xi and the target attitude quaternion q mi set in Step S1, and perform the prediction of the maneuvering path from the initial attitude quaternion q bi_0 to the intermediate attitude quaternion q mi . First, calculate the relative quaternion dq bi_0 between q mi and q _0 , and then sequentially loop through Step S3, Step S4, Step S7, and Step S8.

[0068] First, calculate the intermediate attitude quaternion q mi :

[0069]

[0070] Then, update the relative quaternion dq_0 。

[0071]

[0072] (7) Step S7: Determine whether to switch the target attitude quaternion based on the angle in Step S4. Start the determination after 10 s (jet) / 100 s (flywheel) of attitude maneuver prediction. If the angle between the +X axis of this system and the solar vector is θ _k <θ _k-1 , switch the intermediate attitude quaternion q mi to the target attitude quaternion q xi , and calculate the relative quaternion dq. Otherwise, do not switch.

[0073] Do not switch, and update the relative quaternion dq _k :

[0074]

[0075] After switching, update the relative quaternion dq _k :

[0076]

[0077] (8) Step S8: Determine the new maneuver path prediction result based on the angle in Step S4. If the angle between the +X axis and the solar vector is less than or equal to the constraint angle, i.e., θ _k ≤θ ys , set the prediction flag to 2 and jump to Step S9; otherwise, wait until dq _k ≤1e-3, set the prediction flag to 1 and jump to Step S9.

[0078] If there is θ _k ≤θ ys during the maneuver prediction process according to the current path, then set the prediction flag to 2, i.e., flag _yc = 2, and jump to Step S9 to end the prediction.

[0079] If there is always θ _k >θ ys during the maneuver prediction process according to the current path, then set the prediction flag to 1, i.e., flag _yc = 1, and jump to Step S9 to end the prediction.

[0080] (9) Step S9: Perform attitude maneuver according to the maneuver prediction flag flag _yc in Steps S8 and S5. If flag _yc = 0, maneuver directly from the initial attitude quaternion q bi_0 to the target attitude quaternion q xi . If flag _yc = 1, first from the initial attitude quaternion qbi_0 To the intermediate attitude quaternion q mi Maneuver; After predicting the attitude maneuver for 10 s (jet) / 100 s (flywheel), start to judge the difference in the angle between the +X axis of this system and the solar vector at the previous and current moments. When there is an angle θ between the +X axis of this system and the solar vector _k < θ _k-1 , start to switch the target attitude, switch the intermediate attitude quaternion to the target attitude quaternion, that is, switch q mi to q xi , and start to maneuver towards the target attitude quaternion q xi . If flag _yc = 2, stop maneuvering.

[0081] In addition, if during the actual maneuvering process, the angle between the +X axis of this system and the solar vector is less than the constraint angle redundancy by 2°, that is, θ _k < (θ ys + 2 ° ), then stop maneuvering.

[0082] According to the attitude pointing sun - avoidance maneuver path prediction and planning method described in steps S1 to S9, simulate and verify the feasibility and effectiveness of this method.

[0083] Taking jet control and the constraint angle θ ys = 60 ° as an example, conduct simulation verification.

[0084] Before avoidance, there is no intermediate attitude planning path, and the attitude maneuver is as follows:

[0085] The attitude maneuvers directly from the initial attitude quaternion q bi_0 to the target attitude quaternion q xi , as shown in Figure 2 .

[0086] During the direct maneuver, the angle between the +X axis of this system and the solar vector is as shown in Figure 3 . Before the maneuver, the angle between the +X axis of this system and the solar vector at the initial attitude is 78°, and after maneuvering to the target attitude, the angle is 65°. When directly performing the attitude maneuver, the angle is less than or equal to the constraint angle θ ys = 60° during the process, and at this time, the pointing axis +X will enter the sunlight prohibited area.

[0087] During the attitude maneuver, the control attitude angle and the control attitude angular velocity are as shown in Figure 4 and Figure 5 . The maximum control attitude angular velocity is 2° / s.

[0088] After avoidance, through the intermediate attitude planning path, the attitude maneuver is as follows:

[0089] The attitude first maneuvers towards the intermediate attitude quaternion, switches to the target attitude quaternion opportunistically, and finally maneuvers the attitude to the target quaternion, as Figure 6 shown. By adjusting the maneuvering path, the prohibited area is avoided, that is, the area where the angle between the +X axis of this system and the solar vector is less than or equal to 60° is avoided.

[0090] During the maneuvering according to the newly planned path, the angle between the +X axis of this system and the solar vector is as Figure 7 shown. Before the maneuver, the angle between the +X axis of this system and the solar vector in the initial attitude is 78°, and after maneuvering to the target attitude, the angle between the +X axis of this system and the solar vector is 65°. When maneuvering the attitude according to the newly planned path, the angle is always greater than the constraint angle θ ys = 60°, which can ensure that the +X axis of the pointing axis does not enter the sunlight prohibited area during the maneuver from the initial attitude to the target attitude.

[0091] During the attitude maneuvering, the control attitude angle and the control attitude angular velocity are as Figure 8 and Figure 9 shown. The maximum control attitude angular velocity is 2° / s.

[0092] The parts not detailed in the present invention belong to the common general knowledge of those skilled in the art.

Claims

1. A method for predicting and planning an attitude-pointing sun-avoidance maneuver path, characterized in that Including: Performing attitude maneuver prediction; The attitude maneuver prediction is to predict whether the maneuver path will cause the pointing axis to enter the prohibited area under the attitude prohibition constraint. If not, the path planning is completed using the initial attitude quaternion at the start time of the maneuver and the target attitude quaternion at the end time of the maneuver; otherwise, an intermediate attitude is selected, and the spacecraft is maneuvered towards the intermediate attitude first and then opportunistically towards the target attitude to plan a new maneuver path that avoids the prohibited area; After the attitude maneuver prediction is completed, attitude maneuvers are performed according to the planned path, and no attitude maneuvers are performed when the path planning fails during the attitude maneuver prediction process.

2. The method according to claim 1, characterized in that: Predict whether the maneuver path will cause the pointing axis to enter the prohibited area through the following method: Step S1, setting parameters: setting time parameters, spacecraft orbit parameters, the initial attitude quaternion at the start time of the maneuver, and the target attitude quaternion at the end time of the maneuver; Step S2: According to the parameters set in Step S1, perform the prediction of the maneuvering path from the initial attitude quaternion q bi_0 to the target attitude quaternion q xi . First, calculate the relative quaternion dq between the two _0 ; Step S3: Calculate the predicted next attitude quaternion based on the relative quaternion dq _k-1 Calculate the predicted next attitude quaternion; from the current attitude quaternion q bi_k-1 and the limited relative quaternion dq lim_k-1 , calculate the next attitude quaternion q bi_k ; Step S4. According to the attitude prediction result q in step S3 bi_k calculate the corresponding attitude transformation matrix A bi_k , and then according to A bi_k and the solar vector sun in the inertial system i_k calculate the corresponding solar vector sun in the body system b_k , thereby calculating the angle θ between the +X axis in the body system and the solar vector _k ; Step S5. If the angle between the +X axis and the solar vector is less than or equal to the constraint angle, i.e., θ _k ≤θ ys , it is predicted that the maneuver path will cause the pointing axis to enter the prohibited area; otherwise, calculate and update the relative quaternion dq according to the current attitude quaternion and the target attitude quaternion _k , and sequentially and cyclically execute Step S3, Step S4, and Step S5 until the prediction of the entire attitude maneuver process is completed.

3. The method according to claim 2, wherein: The selection of the intermediate attitude and the planning of a new maneuver path that avoids the prohibited area by maneuvering the spacecraft towards the intermediate attitude first and then opportunistically towards the target attitude include: Step S6: Set the initial attitude quaternion q according to Step S1 bi_0 and the target attitude quaternion q xi , and calculate the intermediate attitude quaternion q mi ; then, perform the prediction of the maneuver path from the initial attitude quaternion q bi_0 to the intermediate attitude quaternion q mi . First, calculate the relative quaternion dq between the two _0 ; sequentially loop through and execute Step S3, Step S4, Step S7, and Step S8; Step S7: Selectively choose to switch the intermediate attitude quaternion q mi to the target attitude quaternion q xi , and calculate and update the relative quaternion dq _k ; otherwise, do not switch; Step S8. Determine the new maneuver path prediction result according to the included angle in Step S4; if the included angle between the +X axis and the solar vector is less than or equal to the constraint angle, i.e., θ _k ≤θ lim , the path planning fails; otherwise, continue to loop until the prediction of the entire attitude maneuver process is completed.

4. The method according to claim 1 or 3, characterized in that: The intermediate attitude is selected in the following way, that is, according to the initial attitude quaternion q bi_0 and the target attitude quaternion q xi , calculate the intermediate attitude quaternion q mi ; Among them, represents the quaternion multiplication symbol, and inv() represents the quaternion inverse operation.

5. The method according to claim 3, characterized in that:

6. The method according to claim 1 or 3, characterized in that: The selected condition is to start judging the difference in the angle between the +X axis of the system and the solar vector at the previous and subsequent moments after a preset time is reached during the prediction process from the initial attitude to the intermediate attitude. When the angle θ _k < θ _k-1 , start switching the target attitude and switch the intermediate attitude quaternion to the target attitude quaternion; k represents the next prediction beat.

7. The method according to claim 6, wherein: The preset time is determined according to the spacecraft attitude maneuver mode, which is 10 s for jet attitude maneuvers and 100 s for flywheel attitude maneuvers.

8. The method according to claim 1, characterized in that: During the attitude maneuver prediction process, the prediction flag is set to guide the attitude maneuver. When the pointing axis does not enter the prohibited area during the entire prediction process, the prediction flag is in the default state 0, that is, flag _yc = 0; When a new maneuver path that avoids the prohibited area is successfully planned by selecting an intermediate attitude during the prediction process, the prediction flag is set to 1, that is, flag _yc = 1; When the path planning fails during the attitude maneuver prediction process, the prediction flag is set to 2, that is, flag _yc = 2.

9. The method according to claim 8, wherein: Using the prediction flag to guide attitude maneuvers includes: If flag _yc = 0, directly maneuver from the initial attitude quaternion q bi_0 to the target attitude quaternion q xi If flag _yc = 1, first maneuver from the initial attitude quaternion q bi_0 to the intermediate attitude quaternion q mi ; opportunistically switch q mi to q xi and start maneuvering towards the target attitude quaternion q xi If flag _yc = 2, stop maneuvering.

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