CN115027702B - Design method of cube star zero momentum attitude control system structure - Google Patents

Abstract

The invention discloses a structure design method of a cube star zero momentum attitude control system, which comprises the following steps: firstly, determining a frame structure of a cube star zero momentum attitude control system according to the dimension requirement of the cube star, then determining the installation position of a measurement and control component, then designing a middle bracket, and finally designing a control module according to the residual space of the system. The measurement and control assembly comprises a triaxial zero momentum wheel, an oblique zero momentum wheel, a star sensor, a sun sensor, a triaxial magnetic torquer, a gyroscope, a GPS receiver and two paths of magnetometers, wherein the three-positive-installation oblique zero momentum wheel is fixed in the main frame, the three sun sensors are respectively fixed on three outer sides of the main frame, and the GPS receiver and the two paths of magnetometers are integrated on the control module. The control modules are connected through the soft flat cable, so that the system reliability is improved, the space is saved, and the attitude control system is convenient to integrate and modularize.

Description

A Design Method of CubeSat Zero Momentum Attitude Control System Structure

technical field

The invention relates to the technical field of cubesats, in particular to a design method for a cubesat zero-momentum attitude control system structure.

Background technique

Since the Soviet Union successfully launched the first satellite in 1957, the world's major powers and organizations have joined the space science and technology competition. The requirements of the development of aerospace technology have led to more and more complex functions and larger volumes of satellites, resulting in defects such as high cost and long development cycle of satellites. Since the 1990s, with the continuous deepening of space research, light weight, low cost, and good performance micro-satellites have gradually attracted more attention and achieved rapid development. With the advancement of advanced technologies such as microelectronics, small high-performance electronic components are gradually applied to small satellites, making small satellites further miniaturized. Then the concept of cube satellite was born. The cube satellite appeared for the first time as one of the scientific researches of Caltech and Stanford University. A cube of cm x 10 cm is a standard unit, which can be flexibly configured into one unit, two units, three units, etc. according to task requirements. Different from mission-oriented spacecraft design, CubeSat has formulated a series of standards including structure, electrical interface, test process, and working mode definition, which makes the satellite design process standardized and reduces the cost of repetitive design between tasks.

The attitude determination and control system is one of the most complex subsystems in the satellite, and it is also the core system in the entire satellite system. Its main task is to realize the pointing of the satellite in a specific space coordinate system, and it generally consists of a sensor for measuring the attitude vector, an attitude control computer and execution components. The principle of CubeSat attitude determination and control system design is based on comprehensive consideration of cost, development cycle, and system reliability, so that the system design can meet the overall requirements of the satellite for attitude control accuracy and the quality, volume, and power consumption of the attitude control system. limit. In terms of attitude measurement and control devices, the research on micro devices based on advanced processing technologies such as MEMS/NEMS/MOEMS has been extensively carried out, such as micro gyroscopes, micro accelerometers, solar sensors made by technology, and highly integrated three-axis micro magnetic Meters, micro-momentum wheels, etc., have application examples on pico and nano-satellites.

Momentum wheels are divided into offset momentum wheels and zero momentum wheels. The zero momentum control scheme is widely used in modern satellite engineering because of its good control accuracy, high reliability, and strong maneuverability. In practical engineering applications, limited to satellite life, energy supply, control stability and other reasons, the main control method of zero-momentum satellites is angular momentum exchange. Because of the small size and weight, high control accuracy, stability and reliability of the momentum wheel, the momentum wheel has become the first choice for the actuator of the control system. On the premise of taking into account the reliability of the control system and the weight and cost of the satellite, the three-mounted-one-slant-mounted momentum wheel system is the most commonly used in engineering. Its control algorithm is simple and the momentum wheel installation requirements are low.

Canx-2 is a three-unit cube satellite developed by the Space Laboratory of the University of Toronto in Canada. The hardware of the satellite attitude determination and control system includes six coarse sun sensors, one magnetometer, three orthogonally placed flywheels and three Shaft magnetic torque coil. The ALMASat-1 three-unit cube satellite designed by the University of Bologna in Italy uses three-axis magnetic torque coils and momentum wheels to achieve attitude determination, and is equipped with a cold air micro-injection system. The NTU-2 CubeSat attitude determination system developed by Nanjing University of Science and Technology adopts self-developed nano-satellite sensors and actuators based on MEMS technology, equipped with new control loads and high-precision GPS receivers, and adopts zero-momentum wheels plus magnetic force for attitude control Torque coil design.

Generally speaking, the research on the attitude control system of the Pina satellite conforming to the cube satellite standard is still in the basic technical verification stage. However, with the current MEMS technology and the development of various micro-components, the future cube satellite is expected to achieve the same attitude control level as the 100-kilogram small satellite. The current design of the CubeSat zero-momentum attitude control system structure occupies too much space in the CubeSat, reaching 1.5U or even 2U, which is not suitable for a CubeSat with a size of only 2U, and the system design must take into account the size constraints in the CubeSat , and the circuit connection of the system must be considered. Therefore, the design of the existing zero-momentum attitude control system is complicated and the system cannot be designed independently. The system needs to be redesigned according to different satellites, and the system reliability is not high.

Contents of the invention

The invention proposes a design method for the structure of the cubic star zero momentum attitude control system, which solves the problems of complicated installation and excessive volume occupation of the zero momentum attitude control system under the premise of realizing functions due to the limited volume of the cube star.

Realize the technical solution of the present invention as: a kind of design method of cubic star zero-momentum attitude control system structure, the steps are as follows:

Step 1. Determine the framework structure of the CubeSat zero-momentum attitude control system according to the size requirements of the CubeSat:

The cubic star zero momentum attitude control system framework includes the main frame, the middle bracket, the three-axis momentum wheel mounting bracket, the oblique momentum wheel mounting bracket, the star sensor mounting bracket, and the attitude control computer mounting bracket.

The main frame includes a four-layer frame and four uprights, the four-layer frame is arranged in parallel from top to bottom, each frame is fixedly connected with four uprights, and the four-layer frame is the first square frame and the second U frame from top to bottom. frame, the third frame, and the fourth frame. The middle bracket is located in the middle of the main frame, and is vertically fixed to the first frame and the third frame at the same time. The three-axis momentum wheel mounting bracket is fixed between the first square frame and the second U-shaped frame. The oblique momentum wheel installation bracket is installed between the second U-shaped frame and the third frame. One end of the star sensor mounting bracket is fixed on the first frame and the third frame, and the other end is fixedly connected with the middle bracket. The attitude control computer mounting bracket is arranged between the third frame and the fourth frame.

Step 2. Determine the installation location of the measurement and control components according to the main frame structure and the functional requirements of the CubeSat:

The measurement and control assembly includes a three-axis momentum wheel, a slanted momentum wheel, a star sensor, a three-axis magnetic torque device, a gyroscope, a GPS receiver, three sun sensors and two-way magnetometers, and the three-axis momentum wheel is fixed on On the upper part of the main frame, the oblique momentum wheel is fixed in the middle of the main frame, three sun sensors are respectively fixed on the three outer sides of the main frame, the star sensor and gyroscope are fixed on the star sensor mounting bracket, three-axis The magnetic torque device includes a first magnetic bar, a second magnetic bar, and a third magnetic bar that are arranged orthogonally in pairs. The first magnetic bar is fixedly connected to the first square frame and the three-axis momentum wheel mounting bracket respectively, and the second magnetic bar is connected to the three-axis momentum wheel mounting bracket. The middle bracket is fixedly connected, and the third magnetic bar is fixedly connected with the third frame.

Step 3. Design the middle bracket:

The middle bracket is an H-shaped structure so as to be easily installed in the main frame, and the two parallel sides on the middle bracket are respectively fixedly connected with the first square frame and the third frame of the main frame by bolts.

Step 4. Design the control module according to the remaining space of the system:

The control module includes 6 PCB boards stacked in sequence. Five control circuit boards and one wiring board are used from top to bottom. The GPS receiver and two-way magnetometer are integrated on the control module, and the control circuit boards are connected through interfaces. Thus integrated into a whole circuit board. The wiring board and the magnetometer module in the control circuit board are directly connected to the installation bracket of the attitude control computer, and the control circuit boards in the control module are connected axially through stud support.

Compared with the prior art, the present invention has significant advantages in that:

(1) The present invention has high functional density and adopts an integrated design concept. The system only occupies 0.8U volume in the star, which fully meets the requirements of miniaturization and light weight of the cube star, reduces the restrictions on the quality and space of other systems, and improves the capacity of the cube star. functional density.

(2) The invention has a wide range of applications, and the structure reserves four-corner through holes, which are fixed in the cube star by studs. It is suitable for the existing design standards of the cube star and can be used for zero momentum attitude control of standard 2U, 3U, 6U cube stars system.

(3) The control modules of the present invention are connected by flexible cables, which improves system reliability and saves space, and can realize high integration and modularization of the attitude control system.

(4) The present invention greatly simplifies the design of the attitude control system of the cube star, adopts the installation method of the standard cube star, connects with the outside through the reserved interface, realizes the independent design of the attitude control system, reduces the design difficulty of the attitude control system and simplifies Design Flow.

(5) The present invention does not need to change its fixed direction in the cube star according to the direction of the satellite itself, and realizes the momentum exchange between the cube star and the flywheel by adjusting the rotational speed direction of the momentum wheel; The shafts are equi-inclined, as a backup in case of failure, improving system reliability and avoiding the zero-crossing problem of the momentum wheel.

(6) The invention has high precision, integrates two-way magnetometers in the control circuit board, and uses a star sensor, which can effectively improve the precision of the attitude control system.

Description of drawings

Fig. 1 is the flow chart of the design method of the CubeSat zero-momentum attitude control system structure of the present invention.

Fig. 2 is an overall three-dimensional structural diagram of the CubeSat zero-momentum attitude control system of the present invention.

Fig. 3 is an exploded view of the CubeSat zero-momentum attitude control system of the present invention.

Fig. 4 is a perspective view of the main frame of the cubesat zero-momentum attitude control system of the present invention.

Fig. 5 is a perspective view of the middle bracket of the cubesat zero-momentum attitude control system of the present invention.

Fig. 6 is a perspective view of the star sensor bracket of the cubic star zero-momentum attitude control system of the present invention.

Fig. 7 is a perspective view of the control module of the CubeSat zero-momentum attitude control system of the present invention.

Fig. 8 is an expanded schematic view of the control module of the cubesat zero-momentum attitude control system of the present invention.

Implementation

The following will clearly and completely describe the technical solutions in the embodiments of the present invention with reference to the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of them. Based on the embodiments of the present invention, all other embodiments obtained by persons of ordinary skill in the art without creative efforts fall within the protection scope of the present invention.

It should be noted that all directional indications (such as up, down, left, right, front, back...) in the embodiments of the present invention are only used to explain the relationship between the components in a certain posture (as shown in the figure). Relative positional relationship, movement conditions, etc., if the specific posture changes, the directional indication will also change accordingly.

In addition, in the present invention, descriptions such as "first", "second" and so on are used for description purposes only, and should not be understood as indicating or implying their relative importance or implicitly indicating the quantity of indicated technical features. Thus, the features defined as "first" and "second" may explicitly or implicitly include at least one of these features. In the description of the present invention, "plurality" means at least two, such as two, three, etc., unless otherwise specifically defined.

In the present invention, unless otherwise specified and limited, the terms "connection" and "fixation" should be interpreted in a broad sense, for example, "fixation" can be a fixed connection, a detachable connection, or an integral body; " "Connection" may be a mechanical connection or an electrical connection. Those of ordinary skill in the art can understand the specific meanings of the above terms in the present invention according to specific situations.

In addition, the technical solutions of the various embodiments of the present invention can be combined with each other, but it must be based on the realization of those skilled in the art. When the combination of technical solutions is contradictory or cannot be realized, it should be considered as a combination of technical solutions. Does not exist, nor is it within the protection scope of the present invention.

The specific implementation, technical difficulties and invention points of this invention will be further introduced below in conjunction with this design example.

Combining Figures 1 to 8, a structural design method for a cube star zero-momentum attitude control system, the steps are as follows:

Step 1. Determine the framework structure of the CubeSat zero-momentum attitude control system according to the size requirements of the CubeSat:

Cube star zero momentum attitude control system framework includes main frame 1, intermediate support 2, three-axis momentum wheel mounting bracket 9, oblique momentum wheel mounting bracket 11, star sensor mounting bracket 13, and attitude control computer mounting bracket 23.

Further, with reference to Figures 2 to 4, the main frame 1 includes a four-layer frame and four columns 4, the four-layer frames are arranged in parallel from top to bottom, each frame is fixedly connected to the four columns 4, and the four-layer frame From top to bottom are the first frame 5, the second U-shaped frame 6, the third frame 7, and the fourth frame 8; the middle bracket 2 is located in the middle of the main frame 1; the three-axis momentum wheel mounting bracket 9 is installed on four Between the first square frame 5 and the second U-shaped frame 6 of the layer frame; the oblique momentum wheel mounting bracket 11 is installed between the second U-shaped frame 6 and the third frame 7; one end of the star sensor mounting bracket 13 is fixed on On the first square frame 5 and the second U-shaped frame 6, the other end is fixedly connected with the middle bracket 2; the posture control computer installation bracket 23 is located between the third frame 7 and the fourth square frame 8 of the four-layer frame; the main frame The volume is 0.8U cube star standard model.

Step 2. According to the external dimensions of the main frame structure and the functional requirements of the CubeSat, determine the installation location of the measurement and control components:

The measurement and control assembly includes a three-axis momentum wheel 10, an oblique momentum wheel 12, a star sensor 14, a three-axis magnetic torque device, a gyroscope 15, a GPS receiver, three sun sensors 16 and two Magnetometer, the triaxial momentum wheel 10 is fixed on the triaxial momentum wheel mounting bracket 9, the obliquely mounted momentum wheel 12 is fixed on the obliquely mounted momentum wheel mounting bracket 11, and the sun sensors 16 in three directions are fixed on the second U On the three sides of the rectangular frame 6 and the third frame 7, the star sensor 14 and the gyroscope 15 are fixed on the star sensor mounting bracket 13, and the three-axis magnetic torque device includes two first magnetic torque devices arranged orthogonally. Rod 17, the second magnetic rod 18 and the 3rd magnetic rod 19, described magnetic rod is wound on the outside of the magnetic induction core by enameled wire and forms, and the two ends of described first magnetic rod 17 are respectively connected with the inner wall of the first square frame 5 and three The outer wall of the shaft momentum wheel mounting bracket 9 is fixedly connected; the second magnetic bar 18 is fixed on the inner walls of the two sides of the middle bracket 2; the third magnetic bar 19 is fixed on the inner wall of the third frame 7 .

Step 3. Design the middle bracket 2:

The middle bracket 2 is an H-shaped structure for easy installation in the main frame 1, and the two parallel sides on the middle bracket 2 are respectively fixedly connected with the first frame 5 and the third frame 7 of the main frame 1 by bolts.

Further, with reference to Fig. 5 ~ Fig. 6, the star sensor mounting bracket 13 is fixed on the inner opposite side of the middle bracket 2 and the first square frame 5 and the second U-shaped frame 6 of the main frame 1, the star sensor 14 and the gyro Instrument 15 is fixedly connected with star sensor mounting bracket 13 by bolts. The two ends of the first magnetic bar 17 are fixedly connected with the inner wall of the first square frame 5 and the outer wall of the three-axis momentum wheel mounting bracket 9 respectively; the second magnetic bar 18 is fixed on the inner walls of the two sides of the middle bracket 2 ; The third magnetic bar 19 is fixed on the inner wall of the third frame 7 .

Step 4. Design the control module 3 according to the remaining space of the system:

Further, referring to Fig. 7 to Fig. 8, the attitude control computer mounting bracket 23 is fixed on a group of opposite sides between the third frame 7 and the fourth frame 8 of the main frame 1; the control module 3 includes sequentially stacked A wiring board 21 and five control circuit boards 20 are arranged, the control circuit boards 20 are fixedly connected to the attitude control computer mounting bracket 23 through bolts, and a screw sleeve is arranged between the circuit boards of the control module 3 in the axial direction 26, used to determine the distance between boards, the control module 3 is located inside the main frame 1.

Furthermore, the control module includes 6 PCB boards, using FPC additive manufacturing technology, the circuit boards can be connected through flexible cables and interfaces to integrate into a complete circuit board. The first circuit board 101 is a CPU board, on which there are interfaces corresponding to the sun sensor 16, the star sensor 14, the gyroscope 15, the second circuit board 102, and the third circuit board 103; board, there are interfaces corresponding to the first circuit board 101, the fourth circuit board 104, and the sixth circuit board 106; the third circuit board 103 is a magnetometer module board; there are interfaces corresponding to the first circuit board 101 on it. Interface; the fourth circuit board 104 is a magnetic rod one module board, on which there are corresponding interfaces with the magnetic rod, the second circuit board 102, and the fifth circuit board 105; the fifth circuit board 105 is a magnetic rod two module board, on which There are interfaces corresponding to the magnetic bar and the fourth circuit board 104; the sixth circuit board 106 is a momentum wheel module board, on which there are interfaces corresponding to the momentum wheel and the second circuit board 102.

Further, the bottom surface and the side of the control module 3 are respectively provided with a test connector 24 corresponding to the position of the rectangular window and a standard J70 communication connector 25, which can be connected to the outside, so that the zero momentum attitude control system can be independently designed, as long as Ensure that the interface with the outside is correct, greatly reducing the difficulty of system design.

The cubic star zero-momentum attitude control system of the present invention is small in size and high in functional density, and fully meets the cube star attitude control system miniaturization and light weight requirements, and its integrated and independent system design features greatly simplify the cube star attitude control The design of the system reduces the design difficulty of the attitude control system and simplifies the design process. In addition, the invention has a wide range of applications, and the structure reserves four-corner through holes, which are fixed in the cube star by studs. It is suitable for the existing design standards of the cube star, and can be used for standard 2U, 3U, 6U cube star bias momentum attitude control system. The control module in the present invention is combined with FPC additive manufacturing technology, and is connected by a flexible cable and an interface, which improves system reliability and saves space, and facilitates the integration and modularization of the attitude control system. Simultaneously, the present invention does not need to change its fixed direction in the cube star according to the direction of the satellite's own system, and realizes momentum exchange between the cube star and the flywheel by adjusting the rotational speed direction of the momentum wheel; The equal tilt angle is used as a backup in case of failure, and the magnetometer is integrated in the control circuit board, which adopts the dual-machine cold backup method, which can effectively improve the reliability of the attitude control system.

The basic principles, main features and advantages of the present invention have been shown and described above. Those skilled in the industry should understand that the present invention is not limited by the above-mentioned embodiments. What are described in the above-mentioned embodiments and the description only illustrate the principle of the present invention. Without departing from the spirit and scope of the present invention, the present invention will also have Variations and improvements are possible, which fall within the scope of the claimed invention. The protection scope of the present invention is defined by the appended claims and their equivalents.

Claims (8)

1. a design method of cube star zero-momentum attitude control system structure is characterized in that: the steps are as follows: Step 1. Determine the framework structure of the CubeSat zero-momentum attitude control system according to the size requirements of the CubeSat: The framework of the cubic star zero-momentum attitude control system includes the main frame (1), the middle bracket (2), the three-axis momentum wheel mounting bracket (9), the oblique momentum wheel mounting bracket (11), the star sensor mounting bracket (13), Attitude control computer mounting bracket (23); The main frame (1) includes four-layer frames and four columns (4), the four-layer frames are arranged in parallel from top to bottom, each frame is fixedly connected with four columns (4), and the four-layer frames are arranged from top to bottom It is the first frame (5), the second U-shaped frame (6), the third frame (7), and the fourth frame (8); the middle bracket (2) is located in the middle of the main frame (1), and is vertical It is fixedly connected with the first frame (5) and the third frame (7); the three-axis momentum wheel mounting bracket (9) is fixed between the first frame (5) and the second U-shaped frame (6); The momentum wheel mounting bracket (11) is installed between the second U-shaped frame (6) and the third frame (7); one end of the star sensor mounting bracket (13) is fixed on the first frame (5) and the third frame ( 7), the other end is fixedly connected to the middle bracket (2); the attitude control computer installation bracket (23) is set between the third frame (7) and the fourth frame (8); Step 2. According to the structure of the main frame (1) and the functional requirements of the CubeSat, determine the installation position of the measurement and control components: The measurement and control assembly includes a three-axis momentum wheel (10), a slanted momentum wheel (12), a star sensor (14), a three-axis magnetic torque device, a gyroscope (15), a GPS receiver, three sun sensors ( 16) and two-way magnetometers, the three-axis momentum wheel (10) is fixed on the upper part of the main frame (1), the inclined momentum wheel (12) is fixed on the middle part of the main frame (1), and the three solar sensors (16 ) are respectively fixed on the three outer edges of the main frame (1), the star sensor (14) and the gyroscope (15) are fixed on the star sensor mounting bracket (13), and the three-axis magnetic torque device includes two orthogonal The first magnetic bar (17), the second magnetic bar (18) and the third magnetic bar (19) are set, and the first magnetic bar (17) is respectively connected with the first square frame (5) and the three-axis momentum wheel mounting bracket ( 9) Fixed connection, the second magnetic rod (18) is fixedly connected to the middle bracket (2), and the third magnetic rod (19) is fixedly connected to the third frame (7); Step 3. Design the middle bracket (2): The middle bracket (2) is an H-shaped structure for easy installation in the main frame (1), and the two parallel sides on the middle bracket (2) are respectively connected with the first square frame (5) and the second Three square frames (7) are fixedly connected; Step 4. Design the control module according to the remaining space of the system (3): The control module (3) includes 6 PCB boards stacked in sequence, five control circuit boards (20) and one wiring board (21) are used from top to bottom, and the GPS receiver and two-way magnetometer are integrated in the control module (3), the control circuit board (20) is connected through an interface to be integrated into a complete circuit board; the magnetometer module (22) in the wiring board (21) and the control circuit board (20) are respectively connected to the The control computer mounting bracket (23) is directly connected, and the control circuit boards (20) in the control module (3) are supported and connected by studs (26) in the axial direction. 2. The design method of the cubic star zero-momentum attitude control system structure according to claim 1, characterized in that: the three-axis momentum wheel mounting bracket (9) is fixed on the first frame (5) and the second U in the middle on the column (4) of the shaped frame (6); the three-axis momentum wheel (10) is fixedly connected with the three-axis momentum wheel mounting bracket (9) by bolts; 6) and the third frame (7); the oblique momentum wheel (12) is fixedly connected with the oblique momentum wheel bracket (11) by bolts; the sun sensors (16) in three directions are fixed on the second U-shaped frame (6) and each side of the third frame (7); the star sensor mounting bracket (13) is fixed on the first frame (5) and the third frame (7) of the middle bracket (2) and the main frame (1) ), the star sensor (14) and gyroscope (15) are fixedly connected with the star sensor mounting bracket (13) by bolts. 3. The design method of the cubic star zero-momentum attitude control system structure according to claim 1, characterized in that: the two ends of the first magnetic bar (17) are respectively connected to the inner wall of the first box (5) and three The outer wall of the shaft momentum wheel mounting bracket (9) is fixedly connected; the second magnetic rod (18) is fixed on the inner walls of the two side brackets of the middle bracket (2); the third magnetic rod (19) is fixed on the third frame (7) on the inner wall. 4. The design method of the cubic star zero momentum attitude control system structure according to claim 1, characterized in that: the attitude control computer mounting bracket (23) is arranged between the third frame (7) and the fourth frame (8) and are respectively fixedly connected with four columns (4); the control module (3) is fixed on the attitude control computer mounting bracket (23) by bolts. 5. The design method of the cubic star zero momentum attitude control system structure according to claim 1, characterized in that: the bottom surface of the control module (3) is provided with a test and communication connector (24) corresponding to the position of the rectangular window ). 6. The design method of the cubic star zero-momentum attitude control system structure according to claim 1, characterized in that: the volume of the cubic star zero-momentum attitude control system is 0.8U. 7. The design method of the cubic star zero-momentum attitude control system structure according to claim 1, characterized in that: the first magnetic rod (17), the second magnetic rod (18) and the third magnetic rod (19) all pass through The enameled wire is wound around the outside of the magnetic sensing core. 8. The design method of the cubic star zero momentum attitude control system structure according to claim 1, characterized in that: the control circuit board (20) is made into a flexible circuit board by using FPC combined with additive manufacturing technology.

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