Energy-conserving satellite step-by-step maneuvering method

Abstract

The application discloses a satellite stepping motor method for energy conservation, which regards the observation direction in inertial space in a task strategy as a satellite Z-axis vector, simultaneously calculates a target attitude matrix by using a sun vector; calculates the angles of three rotations around the body XYZ axes according to a current satellite attitude matrix; rotates the satellite body to a target attitude around the XYZ axes; when rotating around the Y axis, each time rotates by a unit step motor, and after the satellite body rotates by a step angle, offsets the sailboard by an opposite angle; after completing the work task in the target attitude, the satellite restores to the attitude before the task is issued in a ZYX sequence, and the sailboard is returned to the original position when rotating around the Y axis. The application can ensure that the sun is perpendicular to the sailboard as much as possible during the attitude motor process of the satellite pointing to the target, and the satellite obtains the most energy.

Description

Technical Field

[0001] This invention relates to a satellite stepping maneuvering method for conserving energy, belonging to the field of satellite maneuvering technology. Background Technology

[0002] With the development of spacecraft technology and the diversification of space missions, spacecraft sometimes need to maneuver satellites to point in specific directions according to mission requirements, thereby achieving specific observation tasks. These specific pointing maneuvers are spatially arbitrary. During these maneuvers, current attitude maneuvers often focus on achieving both rapid attitude pointing control and high-precision, high-stability control. Typically, the design of wide-bandwidth control parameters is considered to achieve rapid attitude pointing and steady-state control to achieve high-precision, high-stability control pointing, without regard to the satellite's solar panels' alignment with the sun; in some cases, the solar panels may even stop rotating during the maneuver.

[0003] However, for satellites with stringent energy requirements, failing to ensure their energy supply through necessary measures can easily compromise their operational safety. Therefore, during such satellite attitude maneuvers, it is necessary to consider maneuvering methods that guarantee satellite energy supply and the requirements for solar panel drive control. Summary of the Invention

[0004] The technical problem solved by this invention is to overcome the shortcomings of the prior art and propose an energy-saving satellite stepping maneuvering method that can ensure that the sun shines on the solar panels as much as possible during the satellite attitude maneuvering process, so that the satellite obtains the most energy.

[0005] The technical solution of this invention is:

[0006] A method for energy-conserving satellite stepping maneuvering includes:

[0007] Using the observation direction in inertial space as the satellite's Z-axis vector in the mission strategy, we calculate the angle between the solar vector and the satellite's XOZ plane, and the target's OY plane. t The target attitude matrix with the angle between the x-axis and the current +Y-axis vector not exceeding 90°;

[0008] Obtain the current satellite attitude matrix, combine it with the target attitude matrix, calculate the rotation attitude matrix, and obtain the angles of the three maneuvers around the XYZ axes of the satellite body. The satellite then maneuvers to the target attitude according to the angle values ​​corresponding to the X-axis, Y-axis, and Z-axis in sequence, and begins to execute the observation mission.

[0009] After the observation mission is completed, the satellite will be restored to its attitude before the mission was initiated, based on the angles of the three maneuvers along the XYZ axes of the satellite body, in the order of Z-axis, Y-axis, and X-axis.

[0010] Preferably, the satellite maneuvers sequentially according to the angle values ​​corresponding to the X-axis, Y-axis, and Z-axis. The rotation method is as follows:

[0011] The satellite first maneuvers according to the angle value corresponding to the X-axis;

[0012] The satellite then rotates step by step according to the angle value corresponding to the Y-axis, until the rotation reaches the angle value, and after each step is completed, the solar panel is negatively offset by the preset step angle value.

[0013] The satellite finally maneuvers according to the angle value corresponding to the Z-axis.

[0014] Preferably, when the satellite rotates step by step around the angle value corresponding to the Y-axis, the rotation angle of the last step is the remainder obtained by dividing the corresponding angle value by the preset step value.

[0015] Preferably, the feature is that, based on the rotation attitude matrix, the angles of the three maneuvers around the XYZ axes of the satellite body are obtained, specifically including:

[0016] The satellite maneuvers around its X-axis, with the following angle:

[0017]

[0018] The second maneuver around the satellite's Y-axis has the following angle:

[0019] θ = asin(-M) tb (3,1))

[0020] The third maneuver around the satellite's Z-axis has the following angle:

[0021]

[0022] Among them, M tb Here is the rotation attitude matrix. θ and ψ are the angles of the three maneuvers around the X-axis, Y-axis, and Z-axis.

[0023] Preferably, after the observation mission is completed, the satellite is restored to its pre-mission attitude according to the angles of the three maneuvers along the X, Y, and Z axes, in the order of Z-axis, Y-axis, and X-axis. Specifically, this includes:

[0024] First, maneuver around the satellite's Z-axis by -ψ, keeping the solar panel angle unchanged; then, maneuver around the satellite's Y-axis by a preset angle step value to -θ, while simultaneously offsetting the solar panel in the opposite direction by the same angle step value after each step; finally, maneuver around the satellite's X-axis by -ψ, keeping the solar panel angle unchanged.

[0025] Preferably, the satellite rotates in steps according to the angle value corresponding to the Y-axis, with a preset step angle value, while the solar panels are simultaneously offset in the opposite direction. After each step, a correctness judgment is made. If an error is found, the system immediately reverts to the correct attitude of the previous step to ensure the satellite's energy safety.

[0026] Preferably, the preset step angle value is 10°.

[0027] Preferably, the calculation minimizes the angle between the solar vector and the satellite's XOZ plane and the target's OY plane. t The target attitude matrix is ​​obtained by setting the angle between the x-axis and the current +Y-axis vector to no greater than 90°. The specific method is as follows:

[0028] Calculate the target OY t Axis vector:

[0029]

[0030] Where dir is the direction of maneuver, when OY b T ·(OZ t When ×Si)≥0, dir=1; otherwise, it is -1. b Let OY be the current satellite +Y axis vector in the inertial frame. b Axis and target OY t Axis ≤ 90°; Si is the solar vector in the inertial frame;

[0031] The target attitude matrix is ​​then:

[0032] M ti =[OY t ×OZ t OY t OZ t ] T

[0033] In the formula, M ti Let be the target attitude matrix.

[0034] Preferably, the method for obtaining the current satellite attitude matrix is ​​as follows:

[0035] Based on the measured value q of the star sensor si Star sensor mounting matrix q bs Obtain the current attitude of the satellite under its current operating conditions:

[0036]

[0037] Where, q si q is the attitude quaternion in the star-sensor coordinate system obtained from the star sensor measurement. bs For the installation quaternion of the star sensor relative to the satellite body, Multiply by quaternions;

[0038] The current satellite attitude matrix is:

[0039]

[0040] Preferably, the rotation matrix from the current satellite attitude to the target attitude is obtained based on the target attitude matrix and the current satellite attitude matrix:

[0041] M tb =M ti ×M bi T

[0042] Among them, M ti Let M be the target attitude matrix. bi This is the current satellite attitude matrix.

[0043] The advantages of this invention compared to the prior art are:

[0044] (1) This invention utilizes the pointing requirements of the mission target to design a rotation method for energy conservation. During the pointing process, based on the target attitude design, while ensuring that the +Z axis points to the observation target, the angle between the solar vector and the satellite XOZ plane is minimized. At the same time, the rotation axis of the solar panel is driven step by step to maximize the energy supply.

[0045] (2) The present invention uses a stepping maneuver to minimize the impact of sunlight on the deviation of the solar panel when the satellite is stepping, ensuring that the satellite has a power supply throughout the attitude maneuver and maximizing the safety of the satellite.

[0046] (3) The present invention uses a step-by-step driving method, which greatly improves the fault tolerance of maneuvering and the safety of satellite energy. The correctness can be judged step by step through step operation, and errors that occur during maneuvering can be immediately reversed to the previous correct attitude. Attached Figure Description

[0047] Various other advantages and benefits will become apparent to those skilled in the art upon reading the following detailed description of preferred embodiments. The accompanying drawings are for illustrative purposes only and are not intended to limit the invention. Furthermore, the same reference numerals denote the same parts throughout the drawings. In the drawings:

[0048] Figure 1 This is a flowchart of a satellite stepping maneuvering method for conserving energy, according to an embodiment of the present invention. Detailed Implementation

[0049] Exemplary embodiments of the present disclosure will now be described in more detail with reference to the accompanying drawings. While exemplary embodiments of the present disclosure are shown in the drawings, it should be understood that the present disclosure may be implemented in various forms and should not be limited to the embodiments set forth herein. Rather, these embodiments are provided so that this disclosure will be thorough and complete, and will fully convey the scope of the disclosure to those skilled in the art.

[0050] This invention proposes a method for energy-conserving satellite stepping maneuvers, such as... Figure 1 As shown, it specifically includes 5 steps:

[0051] S1) Target attitude design

[0052] Define the satellite's +Z axis as pointing towards the observation target, with the solar panels mounted along the satellite's ±Y direction. The observation direction in inertial space, as defined in the mission strategy, is taken as the satellite's Z-axis vector OZ. t The vector obtained by cross-product of the solar vector at the moment of maneuvering and Z is the direction of the rotation axis of the solar panel (i.e., the Y-axis):

[0053]

[0054] Where dir is the direction of maneuver, when OY b T ·(OZ t When ×Si)≥0, dir=1; otherwise, it is -1. b The current satellite's +Y axis vector in the inertial frame, ensuring OY b Axis and target OY t Axis ≤ 90°.

[0055] The vector obtained by cross-product of Y and Z is the satellite's X-direction vector. Based on the orientation of the XYZ vectors in inertial space, the inertial target attitude matrix M of the satellite at the moment of maneuvering is determined. ti :

[0056] M ti =[OY t ×OZ t OY t OZ t ] T

[0057] S2) Obtain the current satellite attitude

[0058] Based on the measured value q of the star sensor si And the installation matrix q of the star sensor bs Obtain the attitude q of the satellite in its current operating state. bi :

[0059]

[0060] Where q si q is the attitude quaternion in the star-sensor coordinate system obtained from the star sensor measurement. bs For the installation quaternion of the star sensor relative to the satellite body, This is a quaternion round multiplication. Therefore, the current satellite attitude matrix M... bi for:

[0061]

[0062] S3) Calculate the angles of the three rotations.

[0063] Using the target attitude M pointed to by the observation ti and the current posture M bi Calculate the rotation matrix M between the current attitude and the target attitude being observed. tb =M ti ×M bi T .

[0064] The satellite reached its target attitude through three step maneuvers, ensuring that its solar panels were illuminated throughout the process. The angles of the three rotations around the X, Y, and Z axes were calculated based on the rotation matrix. θ and ψ, specifically:

[0065] The first maneuver around the X-axis of the body has the following angle:

[0066]

[0067] The second maneuver around the Y-axis of the body has the following angle:

[0068] θ = asin(-M) tb (3,1))

[0069] The third maneuver around the Z-axis of the main body has the following angle:

[0070]

[0071] S4) Energy-saving stepping motor

[0072] Based on the three calculated angles θ and ψ rotate sequentially around the satellite's body along the X, Y, and Z axes. During Y-axis rotation, the rotation increments are 10°. Simultaneously, after each rotation, the solar panels are offset by 10° in the opposite direction. The final rotation angle is the remainder of θ divided by 10 degrees, ensuring that the solar panels' deviation from sunlight coverage is minimized during satellite rotation. The solar panels do not require offset drive during X and Z axis rotations.

[0073] S5) Return to initial attitude after observation ends.

[0074] After observation, the system returns along the original path, first maneuvering -ψ around the Z-axis while maintaining the sail angle. Then, it maneuvers around the Y-axis in 10° increments to -θ, simultaneously offsetting the sail by θ at the same angle. Finally, it maneuvers -ψ around the X-axis while maintaining the sail angle. The system recovers its original attitude through these ZYX sequential maneuvers, with the sail returning to its original position during the Y-axis maneuver.

[0075] The calculation can be performed using general numerical calculation software. The calculation formula has been given. Through a simple calculation method, it is possible to ensure that the sun shines directly on the solar panels as much as possible during the satellite's attitude maneuvering towards the target, so that the satellite can obtain the most energy.

[0076] The embodiments described above are merely preferred embodiments of the present invention. Ordinary variations and substitutions made by those skilled in the art within the scope of the technical solution of the present invention should be included within the protection scope of the present invention.

Claims

1. A method for energy-conserving satellite stepping maneuvering, characterized in that, include: Using the observation direction in inertial space as the satellite's Z-axis vector in the mission strategy, we calculate the angle between the solar vector and the satellite's XOZ plane, and the target's OY plane. t The target attitude matrix with the angle between the x-axis and the current +Y-axis vector not exceeding 90°; Obtain the current satellite attitude matrix, combine it with the target attitude matrix, calculate the rotation attitude matrix, and obtain the angles of the three maneuvers around the XYZ axes of the satellite body. The satellite then maneuvers to the target attitude according to the angle values ​​corresponding to the X-axis, Y-axis, and Z-axis in sequence, and begins to execute the observation mission. After the observation mission is completed, the satellite body will be restored to its attitude before the mission was initiated, based on the three maneuver angles of the XYZ axes mentioned above, in the order of Z-axis, Y-axis, and X-axis. The satellite maneuvers sequentially according to the angle values ​​corresponding to the X-axis, Y-axis, and Z-axis. The rotation method is as follows: the satellite first maneuvers according to the angle value corresponding to the X-axis; then the satellite rotates step by step according to the angle value corresponding to the Y-axis with a preset step angle value until the rotation reaches the angle value, and after each step is completed, the solar panels are negatively offset by the preset step angle value; finally, the satellite maneuvers according to the angle value corresponding to the Z-axis. Based on the rotation attitude matrix, the angles of the three maneuvers around the XYZ axes of the satellite body are obtained, specifically including: The satellite maneuvers around its X-axis, with the following angle: The second maneuver around the satellite's Y-axis has the following angle: The third maneuver around the satellite's Z-axis has the following angle: in, Here is the rotational attitude matrix. , , The angles of the three maneuvers around the X-axis, Y-axis, and Z-axis.

2. The energy-conserving satellite stepping maneuvering method according to claim 1, characterized in that, When the satellite rotates in steps around the Y-axis, the rotation angle of the last step is the remainder obtained by dividing the corresponding angle value by the preset step value.

3. A method for energy-conserving satellite stepping maneuvering according to claim 1 or 2, characterized in that, After the observation mission is completed, based on the angles of the satellite's three maneuvers along the X, Y, and Z axes, it will be restored to its attitude before the mission was initiated, in the order of Z-axis, Y-axis, and X-axis. Specifically, this includes: First, maneuver around the satellite's Z-axis. The angle of the solar panel remains unchanged; then, it moves around the satellite's Y-axis in a step-by-step manner according to a preset angle step value until it reaches its destination. Simultaneously, the solar panels are reversed by the same angular step value after each step; finally, they maneuver around the satellite's X-axis. The angle of the windsurfing board remains unchanged.

4. The energy-conserving satellite stepping maneuvering method according to claim 1, characterized in that, The satellite rotates step by step according to the angle value corresponding to the Y-axis, with the solar panels simultaneously offset in the opposite direction. After each step, the correctness is judged. If an error is found, the system immediately reverts to the correct attitude of the previous step.

5. The energy-conserving satellite stepping maneuvering method according to claim 1, characterized in that, The preset step angle value is 10°.

6. The energy-conserving satellite stepping maneuvering method according to claim 1, characterized in that, Calculate to minimize the angle between the solar vector and the satellite's XOZ plane and the target's OY plane. t The target attitude matrix is ​​obtained by setting the angle between the x-axis and the current +Y-axis vector to no greater than 90°. The specific method is as follows: Calculate the target OY t Axis vector: in, For the direction of maneuver, when hour =1, otherwise -1. OY b Let the current satellite's +Y axis vector be in the inertial frame, so that... OY b Axis and Target OY t Axis ≤ 90°; The solar vector in the inertial frame; The target attitude matrix is ​​then: In the formula, Let be the target attitude matrix.

7. The energy-conserving satellite stepping maneuvering method according to claim 1, characterized in that, The method for obtaining the current satellite attitude matrix is ​​as follows: Based on the measurements from the star sensor Star sensor mounting matrix Obtain the current attitude of the satellite under its current operating conditions: in, The attitude quaternion in the star-sensor coordinate system is obtained from the measurements of the star sensor. For the installation quaternion of the star sensor relative to the satellite body, Multiply by quaternions; The current satellite attitude matrix is: 。 8. A satellite stepping maneuvering method for conserving energy according to claim 1, characterized in that, Based on the target attitude matrix and the current satellite attitude matrix, the rotation matrix from the current satellite attitude to the target attitude is obtained: in, Let be the target attitude matrix. This is the current satellite attitude matrix.

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