An adaptive angular momentum clipping adjustment attitude maneuvering method

Abstract

The application discloses a kind of self-adapting angular momentum amplitude adjustment attitude maneuvering methods, comprising: obtaining the attitude quaternion deviation and angular velocity deviation of satellite body;According to the attitude quaternion deviation of satellite body, three-axis attitude maneuvering angular velocity amplitude value is obtained;According to the three-axis attitude maneuvering angular velocity amplitude value of satellite body, satellite body three-axis quaternion deviation amplitude value is obtained;According to satellite body three-axis quaternion deviation amplitude value, the attitude quaternion deviation of satellite body after vector amplitude limiting is calculated;According to satellite body three-axis quaternion deviation after vector amplitude limiting and angular velocity deviation, three-axis command control moment is obtained.The application can guarantee that the execution mechanism angular momentum does not exceed the premise of output capacity during maneuvering process, and simultaneously realize three-axis simultaneous maneuvering to place to the greatest extent using execution angular momentum.

Description

Technical Field

[0001] This invention belongs to the field of spacecraft attitude control technology, specifically relating to an attitude maneuvering method with adaptive angular momentum limiting adjustment. Background Technology

[0002] With the rapid development of aerospace-related technologies, satellite attitude control technology has matured. For satellites capable of attitude maneuvers, a fixed angular velocity maneuver is generally designed to achieve this.

[0003] Problems with fixed angular velocity maneuvering include: ① The angular momentum of the actuator must be strictly controlled before attitude maneuvering, because excessive angular momentum of the actuator before attitude maneuvering will lead to saturation of the actuator's output capacity and loss of control capacity during the maneuvering process; ② In order to ensure that the actuator is not saturated during attitude maneuvering, the attitude maneuvering angular velocity is often designed with a large margin, and the maneuvering capacity is not fully utilized. Summary of the Invention

[0004] The purpose of this invention is to overcome the aforementioned shortcomings and provide an attitude maneuvering method with adaptive angular momentum limiting adjustment. This method solves the technical problems of existing fixed angular velocity maneuvering methods, such as complex actuator angular momentum design methods, which easily lead to actuator output capacity saturation or insufficient maneuvering capacity during the maneuvering process. This invention can maximize the utilization of actuator angular momentum while ensuring that the actuator angular momentum does not exceed the output capacity during the maneuvering process, and simultaneously achieve three-axis simultaneous maneuvering to the correct position.

[0005] To achieve the above-mentioned objectives, the present invention provides the following technical solution:

[0006] An attitude maneuvering method with adaptive angular momentum limiting adjustment includes:

[0007] Obtain the attitude quaternion deviation and angular velocity deviation of the satellite body;

[0008] The three-axis attitude maneuver angular velocity limit of the satellite body is obtained based on the attitude quaternion deviation of the satellite body;

[0009] The three-axis quaternion deviation limit value of the satellite body is obtained based on the limit value of the three-axis attitude maneuver angular velocity of the satellite body;

[0010] The three-axis quaternion deviation of the satellite body is obtained after vector limiting based on the three-axis quaternion deviation limit value of the satellite body;

[0011] The three-axis command control torque of the satellite body is obtained based on the three-axis quaternion deviation and angular velocity deviation of the satellite body after vector limiting.

[0012] Furthermore, the attitude quaternion deviation of the satellite itself Calculate using the following formula:

[0013] ;

[0014] in, To represent quaternion multiplication, Let be the attitude quaternion of the satellite body relative to the orbital coordinate system. Let be the attitude quaternion of the attitude maneuvering target relative to the orbital coordinate system.

[0015] Furthermore, the angular velocity deviation of the satellite itself Calculate using the following formula:

[0016] ;

[0017] in, This represents the attitude angular velocity of the satellite body relative to the orbital coordinate system in the body coordinate system. Indicates quaternion deviation The corresponding rotation matrix, , for The vector part, Let ω represent the attitude angular velocity of the attitude maneuvering target relative to the orbital coordinate system.

[0018] Furthermore, .

[0019] Furthermore, the limit value of the satellite's three-axis attitude maneuver angular velocity. Calculate using the following formula:

[0020]

[0021] in, This is a column vector composed of the satellite's principal inertia. Dot product; The maximum output angular momentum modulus of the actuator. This represents the angular momentum modulus of the actuator at the start of the attitude maneuver.

[0022] Furthermore, the three-axis quaternion deviation limit value of the satellite body. Calculate using the following formula:

[0023]

[0024] in, This indicates taking the minimum value. This represents the three-axis moment limit value for the satellite body; . / represents point division. D These are the angular velocity deviation control parameters. K These are the attitude quaternion deviation vector control parameters.

[0025] Furthermore, the three-axis quaternion deviation of the satellite body after vector limiting. Calculate using the following formula:

[0026] ;

[0027] .

[0028] Furthermore, the three-axis command control torque of the satellite body Calculate using the following formula:

[0029]

[0030] in, for according to The value of the vector-limited output. This indicates that the amplitude is limited according to Tmax_int, where Tmax_int = 0.005~0.02 Nm; K int The quaternion deviation vector integral control parameters for the satellite body are... t Indicates time on the satellite. Indicates to Integrate over time.

[0031] Compared with the prior art, the present invention has at least one of the following advantages:

[0032] (1) The present invention utilizes the magnitude of the angular momentum of the actuator before maneuvering to dynamically adjust the attitude quaternion deviation limit during the maneuvering process, thereby maximizing the utilization of the actuator's angular momentum while ensuring that the angular momentum of the actuator does not exceed the output capacity during the maneuvering process, thus improving the maneuvering speed.

[0033] (2) The present invention adopts a quaternion deviation vector partial vector limiter in the maneuvering process. The limiter value is designed to take into account the principal inertia of each axis and the quaternion deviation magnitude to distribute the angular momentum, so as to realize the simultaneous maneuvering of the three axes and their simultaneous arrival at the destination. Attached Figure Description

[0034] Figure 1 This is a schematic diagram of the angular velocity during the three-axis attitude maneuver of the satellite body according to the present invention;

[0035] Figure 2 This is a schematic diagram of the angular momentum modulus curve of the actuator during the attitude maneuvering process of the present invention. Detailed Implementation

[0036] The features and advantages of the present invention will become clearer and more apparent from the following detailed description.

[0037] The term “exemplary” as used herein means “serving as an example, embodiment, or illustration.” Any embodiment illustrated herein as “exemplary” is not necessarily to be construed as superior to or better than other embodiments. Although various aspects of embodiments are shown in the accompanying drawings, the drawings are not necessarily drawn to scale unless specifically indicated otherwise.

[0038] This invention provides an attitude maneuvering method with adaptive angular momentum limiting adjustment. The method adaptively adjusts the angular momentum limiting during the maneuvering process based on the angular momentum of the actuator before the maneuver, thereby maximizing the utilization of the actuator's capabilities for rapid attitude maneuvering.

[0039] Satellite attitude maneuver control typically employs a pre-designed angular velocity. This angular velocity must be designed with a certain margin, based on the angular momentum output capability of the actuators, to avoid saturation during maneuvering. This invention provides a three-axis shortest path attitude maneuvering method that adaptively adjusts the angular velocity based on the pre-maneuver angular momentum and angular momentum output capability. Specifically designed for satellites requiring rapid attitude maneuvers, this invention presents an attitude maneuver angular momentum limiting method and achieves simultaneous three-axis maneuvers and simultaneous arrival at the destination.

[0040] This invention discloses an attitude maneuvering method with adaptive angular momentum limiting adjustment, comprising the following steps:

[0041] Step 1: Calculate attitude quaternion deviation and angular velocity deviation

[0042] Attitude quaternion deviation is

[0043]

[0044] Angular velocity deviation is

[0045]

[0046] in, To represent quaternion multiplication, Let be the attitude quaternion of the satellite body relative to the orbital coordinate system. This represents the satellite's attitude angular velocity relative to the orbital coordinate system in its body coordinate system. Let be the attitude quaternion of the attitude maneuvering target relative to the orbital coordinate system. The target's attitude angular velocity relative to the orbital coordinate system; This represents the inversion of a quaternion, i.e., the inversion of the vector part; Indicates quaternion deviation The corresponding rotation matrix is ​​calculated as follows:

[0047]

[0048] Step 2: Calculate the angular velocity limit for three-axis attitude maneuvering.

[0049]

[0050] in, The maximum output angular momentum modulus of the actuator. This represents the angular momentum modulus of the actuator during attitude maneuver initiation. attitude quaternion deviation The vector part; This is a column vector composed of the satellite's principal inertia. The symbols . and . / represent dot product and dot division, respectively, which means multiplying or dividing corresponding values ​​of two vectors to output a new vector.

[0051] Step 3: Calculate the triaxial quaternion deviation limit value

[0052]

[0053] in, This refers to the triaxial torque limit value; D These are the angular velocity deviation control parameters. K These are attitude quaternion deviation vector control parameters. This indicates taking the minimum value.

[0054] Step 4: Calculate the quaternion deviation value after vector limiting.

[0055]

[0056] Where max represents taking the maximum value; .

[0057] Step 5: Calculate the three-axis command control torque

[0058]

[0059] in K int These are quaternion deviation vector integral control parameters. This indicates that the amplitude is limited according to Tmax_int, where Tmax_int can be 0.01Nm.

[0060] Figure 1 This is a schematic diagram of the angular velocity during the three-axis attitude maneuver of the satellite body obtained using the method of this invention. Figure 1 It can be seen that the attitude maneuvering process of the present invention adopts the quaternion deviation vector partial vector limiting. The limiting value is designed to consider the principal inertia of each axis and the angular momentum distribution based on the magnitude of the quaternion deviation, so as to achieve simultaneous maneuvering of the three axes and simultaneous arrival at the destination. Figure 2 This is a schematic diagram of the angular momentum modulus curve of the actuator during the attitude maneuvering process obtained by the method of the present invention. Figure 2 It can be seen that the method of the present invention utilizes the magnitude of the angular momentum of the actuator before maneuvering to dynamically adjust the attitude quaternion deviation limit during the maneuvering process, thereby maximizing the utilization of the actuator's angular momentum while ensuring that the angular momentum of the actuator does not exceed the output capacity during the maneuvering process, thus improving the maneuvering speed.

[0061] The present invention has been described in detail above with reference to specific embodiments and exemplary examples; however, these descriptions should not be construed as limiting the present invention. Those skilled in the art will understand that various equivalent substitutions, modifications, or improvements can be made to the technical solutions and embodiments of the present invention without departing from the spirit and scope of the invention, and all such modifications and improvements fall within the scope of the present invention. The scope of protection of the present invention is defined by the appended claims.

[0062] The contents not described in detail in this specification are common knowledge to those skilled in the art.

Claims

1. A method for attitude maneuvering with adaptive angular momentum limiting adjustment, characterized in that, include: Obtain the attitude quaternion deviation and angular velocity deviation of the satellite body; The three-axis attitude maneuver angular velocity limit of the satellite body is obtained based on the attitude quaternion deviation of the satellite body; The three-axis quaternion deviation limit value of the satellite body is obtained based on the limit value of the three-axis attitude maneuver angular velocity of the satellite body; The three-axis quaternion deviation of the satellite body is obtained after vector limiting based on the three-axis quaternion deviation limit value of the satellite body; The three-axis command control torque of the satellite body is obtained based on the three-axis quaternion deviation and angular velocity deviation of the satellite body after vector limiting; Three-axis attitude maneuver angular velocity limit of the satellite body Calculate using the following formula: in, This is a column vector composed of the satellite's principal inertia. Dot product; The maximum output angular momentum modulus of the actuator. The angular momentum modulus of the actuator at the start of the attitude maneuver; Triaxial quaternion deviation limit of the satellite body Calculate using the following formula: in, This indicates taking the minimum value. This represents the three-axis moment limit value for the satellite body; . / represents point division. D These are the angular velocity deviation control parameters. K These are the attitude quaternion deviation vector control parameters; Satellite body three-axis quaternion deviation after vector limiting Calculate using the following formula: ； ； Three-axis command control torque of the satellite body Calculate using the following formula: in, for according to The value of the vector-limited output. This indicates that the amplitude is limited according to Tmax_int; K int The quaternion deviation vector integral control parameters for the satellite body are... t Indicates time on the satellite. Indicates to Integrating over time; Tmax_int = 0.005 ~ 0.02 Nm.

2. The attitude maneuvering method for adaptive angular momentum limiting adjustment according to claim 1, characterized in that, Attitude quaternion deviation of the satellite body Calculate using the following formula: ； in, To represent quaternion multiplication, Let be the attitude quaternion of the satellite body relative to the orbital coordinate system. Let be the attitude quaternion of the attitude maneuvering target relative to the orbital coordinate system.

3. The attitude maneuvering method for adaptive angular momentum limiting adjustment according to claim 2, characterized in that, Angular velocity deviation of the satellite body Calculate using the following formula: ； in, This represents the attitude angular velocity of the satellite body relative to the orbital coordinate system in the body coordinate system. Indicates quaternion deviation The corresponding rotation matrix, , for The vector part, Let ω represent the attitude angular velocity of the attitude maneuvering target relative to the orbital coordinate system.

4. The attitude maneuvering method for adaptive angular momentum limiting adjustment according to claim 3, characterized in that, 。

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