CN115027702A - Design method of cube star zero momentum attitude control system structure - Google Patents
Design method of cube star zero momentum attitude control system structure Download PDFInfo
- Publication number
- CN115027702A CN115027702A CN202210851880.8A CN202210851880A CN115027702A CN 115027702 A CN115027702 A CN 115027702A CN 202210851880 A CN202210851880 A CN 202210851880A CN 115027702 A CN115027702 A CN 115027702A
- Authority
- CN
- China
- Prior art keywords
- star
- frame
- momentum
- attitude control
- mounting bracket
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
Links
- 238000013461 design Methods 0.000 title claims abstract description 33
- 238000000034 method Methods 0.000 title claims abstract description 17
- 238000005259 measurement Methods 0.000 claims abstract description 9
- 238000009434 installation Methods 0.000 claims abstract description 6
- 238000005516 engineering process Methods 0.000 claims description 8
- 238000012360 testing method Methods 0.000 claims description 4
- 239000000654 additive Substances 0.000 claims description 3
- 230000000996 additive effect Effects 0.000 claims description 3
- 238000004891 communication Methods 0.000 claims description 3
- 238000004519 manufacturing process Methods 0.000 claims description 3
- 230000006698 induction Effects 0.000 claims description 2
- 238000004804 winding Methods 0.000 claims description 2
- 230000010354 integration Effects 0.000 abstract description 4
- 238000011161 development Methods 0.000 description 4
- 240000004282 Grewia occidentalis Species 0.000 description 3
- 238000011160 research Methods 0.000 description 3
- 238000010586 diagram Methods 0.000 description 2
- 244000205725 Boronia megastigma Species 0.000 description 1
- XEEYBQQBJWHFJM-UHFFFAOYSA-N Iron Chemical group [Fe] XEEYBQQBJWHFJM-UHFFFAOYSA-N 0.000 description 1
- 108091092878 Microsatellite Proteins 0.000 description 1
- 101100460147 Sarcophaga bullata NEMS gene Proteins 0.000 description 1
- 238000011949 advanced processing technology Methods 0.000 description 1
- 230000008094 contradictory effect Effects 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 238000004377 microelectronic Methods 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 150000003071 polychlorinated biphenyls Chemical class 0.000 description 1
- 238000005507 spraying Methods 0.000 description 1
- 230000005477 standard model Effects 0.000 description 1
- 238000012795 verification Methods 0.000 description 1
Images
Classifications
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/244—Spacecraft control systems
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02E—REDUCTION OF GREENHOUSE GAS [GHG] EMISSIONS, RELATED TO ENERGY GENERATION, TRANSMISSION OR DISTRIBUTION
- Y02E10/00—Energy generation through renewable energy sources
- Y02E10/50—Photovoltaic [PV] energy
Abstract
The invention discloses a cubic star zero momentum attitude control system structure design method, which comprises the following steps: firstly, determining a frame structure of a cubic star zero momentum attitude control system according to the size requirement of a cubic star, then determining the installation position of a measurement and control assembly, designing an intermediate bracket, and finally designing a control module according to the residual space of the system. The measurement and control assembly comprises a three-axis zero momentum wheel, an obliquely-installed zero momentum wheel, a star sensor, a sun sensor, a three-axis magnetic torquer, a gyroscope, a GPS receiver and two magnetometers, the three-axis zero momentum wheel is positively installed, the obliquely-installed zero momentum wheel is fixed in the main frame, the three sun sensors are respectively fixed on three outer side edges of the main frame, and the GPS receiver and the two magnetometers are integrated on the control module. The control modules are connected through the flexible flat cables, so that the reliability of the system is improved, the space is saved, and the integration and modularization of the attitude control system are facilitated.
Description
Technical Field
The invention relates to the technical field of cuboids, in particular to a design method of a cuboids zero-momentum attitude control system structure.
Background
The development requirements of the aerospace technology lead to the defects of more and more complex functions and larger volume of the satellite, high cost, long development period and the like of the satellite. With the continuous research of space, micro-satellites with light weight, low cost and good performance have gradually attracted more attention and rapidly developed in the nineties. With the progress of advanced technologies such as microelectronics, small high-performance electronic components are gradually applied to minisatellites, so that the minisatellites are further miniaturized. The concept of cubic satellites, which was proposed by the university of Stanford and specified to have a mass of 1 kg and a structural size of 10 cm, was first introduced as one of the scientific studies of California's college of science and Stanford university
10 cm
The 10 cm cube is a standard unit, and can be flexibly configured into a unit, a double unit, a triple unit and the like according to task requirements. Different from the design of a spacecraft taking tasks as guidance, the cubic satellite sets a series of standards including structure, electrical interface, test flow, working mode definition and the like, so that the satellite design flow is standard, and the repeated design cost among the tasks is reduced.
The attitude determination and control system is one of subsystems with the highest complexity in the satellite and is also a core system in the whole satellite system. The main task of the system is to realize the pointing of a satellite in a space specific coordinate system, and the system is generally composed of a sensor for measuring an attitude vector, an attitude control computer and an execution part. The principle of cubic satellite attitude determination and control system design is that on the basis of comprehensively considering cost, development period and system reliability, the system design can meet the requirements of the satellite on attitude control accuracy and the limitations of the quality, volume and power consumption of the attitude control system. In the aspect of attitude measurement and control devices, research on micro devices based on advanced processing technologies such as MEMS/NEMS/MOEMS and the like is widely carried out, such as micro gyros, micro accelerometers, sun sensors manufactured by the technology, three-axis micro magnetometers with high integration, micro momentum wheels and the like, and the micro gyros and the micro accelerometers have application examples.
Momentum wheels are divided into offset momentum wheels and zero momentum wheels, wherein a zero momentum control scheme is generally adopted in modern satellite engineering due to the advantages of good control precision, high reliability, strong maneuverability and the like. In practical engineering application, the main control mode of the zero momentum satellite is angular momentum exchange due to the limitation of satellite service life, energy supply, control stability and the like. The momentum wheel is the first choice for the control system actuator because of its small volume and weight, high control accuracy, stability and reliability. On the premise of considering both the reliability of a control system, the weight of a satellite and the cost, the three-front-mounted one-inclined-mounted momentum wheel system is most commonly applied in engineering, the control algorithm is simple, and the momentum wheel mounting requirement is low.
Canx-2 is a three-unit cubic satellite developed by space laboratories of Toronto university, Canada, and the attitude determination and control system hardware of the satellite comprises six coarse sun sensors, a magnetometer, three orthogonally arranged flywheels and a three-axis magnetic moment coil. The ALMASat-1 three-unit cubic satellite designed by the university of Boronia Italy realizes the attitude determination through a three-axis magnetic moment coil and a momentum wheel and is provided with a cold air micro-spraying system. A Nanjing Ringman second cube satellite attitude determination system developed by Nanjing Ringman university adopts a nano satellite sensor and an execution component which are independently developed and based on an MEMS technology, a novel control load and a high-precision GPS receiver are carried, and the attitude control adopts the design of a zero-momentum wheel and a magnetic moment coil.
Generally, the research aspect of the attitude control system of the pico-nano satellite conforming to the standard of the cubic satellite is in the basic technical verification stage. However, with the current development of MEMS technology and various micro-components, future cubic satellites are expected to reach the same attitude control level of hundreds of kilograms of minisatellites. The currently designed structure of the cubic star zero momentum attitude control system occupies too much space in a cubic star, which reaches 1.5U or even 2U, and is not suitable for the cubic star with the size of only 2U, and the system design needs to consider both the intra-star size constraint and the system circuit connection, so that the zero momentum attitude control system in the prior art is complex in design, cannot design the system independently, needs to redesign the system according to different satellites, and is low in system reliability.
Disclosure of Invention
The invention provides a design method of a structure of a cubic star zero momentum attitude control system, which solves the problems of complex installation, overlarge volume occupation and the like of the zero momentum attitude control system on the premise of realizing functions due to limited cubic star volume.
The technical solution for realizing the invention is as follows: a design method of a cubic star zero momentum attitude control system structure comprises the following steps:
the cubic star zero momentum attitude control system framework comprises a main framework, a middle support, a three-axis momentum wheel mounting support, an oblique momentum wheel mounting support, a star sensor mounting support and an attitude control computer mounting support.
The main frame comprises four layers of frames and four stand columns, the four layers of frames are arranged in parallel from top to bottom, each frame is fixedly connected with the four stand columns, and the four layers of frames are a first square frame, a second U-shaped frame, a third square frame and a fourth square frame from top to bottom. The middle support is positioned in the middle of the main frame and is vertically and fixedly connected with the first square frame and the third square frame. The three-axis momentum wheel mounting bracket is fixed between the first square frame and the second U-shaped frame. The oblique momentum wheel mounting bracket is arranged between the second U-shaped frame and the third square frame. One end of the star sensor mounting bracket is fixed on the first square frame and the third square frame, and the other end of the star sensor mounting bracket is fixedly connected with the middle bracket. The attitude control computer mounting bracket is arranged between the third frame and the fourth frame.
the measurement and control assembly comprises a three-axis zero momentum wheel, an obliquely-installed zero momentum wheel, a star sensor, a three-axis magnetic torquer, a gyroscope, a GPS receiver, three sun sensors and two-way magnetometers, the three-axis momentum wheel is fixed on the upper portion of the main frame, the obliquely-installed momentum wheel is fixed in the middle of the main frame, the three sun sensors are respectively fixed on three outer side edges of the main frame, the star sensors and the gyroscope are fixed on a star sensor mounting bracket, the three-axis magnetic torquer comprises a first magnetic bar, a second magnetic bar and a third magnetic bar which are arranged in a pairwise orthogonal mode, the first magnetic bar is respectively fixedly connected with a first square frame and a three-axis momentum wheel mounting bracket, the second magnetic bar is fixedly connected with a middle bracket, and the third magnetic bar is fixedly connected with a third square frame.
the middle support is of an H-shaped structure so as to be conveniently installed in the main frame, and two parallel edges of the middle support are fixedly connected with the first square frame and the third square frame of the main frame through bolts respectively.
the control module comprises 6 PCBs which are sequentially stacked, five control circuit boards and a wiring board are adopted from top to bottom, the GPS receiver and the two magnetometers are integrated on the control module, and the control circuit boards are connected through interfaces and are integrated on the whole circuit board. The wiring board and the magnetometer modules in the control circuit board are respectively and directly connected with the attitude control computer mounting bracket, and all the control circuit boards in the control modules are axially connected through stud supports.
Compared with the prior art, the invention has the remarkable advantages that:
(1) the invention has high function density, adopts the integrated design idea, and the system only occupies 0.8U volume in the star, completely meets the requirements of the cube star on miniaturization and light weight, reduces the limit to the mass and space of other systems, and improves the function density of the cube star.
(2) The invention has wide application range, is structurally reserved with four-corner through holes, is fixed in the cube star through the studs, is suitable for the design standard of the existing cube star, and can be used for the standard 2U, 3U, 6U and other cube star zero-momentum attitude control systems.
(3) The control modules are connected by adopting the flexible flat cables, so that the reliability of the system is improved, the space is saved, and the high integration and modularization of the attitude control system can be realized.
(4) The invention greatly simplifies the design of the attitude control system of the cube star, adopts the installation mode of the standard cube star and is connected with the outside through the reserved interface, thereby realizing the independent design of the attitude control system, reducing the design difficulty of the attitude control system and simplifying the design flow.
(5) The invention does not need to change the fixed direction of the system in the cube star according to the direction of the system in the satellite, and realizes the momentum exchange between the cube star and the flywheel by adjusting the rotating speed direction of the momentum wheel; meanwhile, the direction of the obliquely installed momentum wheel and the inclination angles of the three star shafts are equal, and the obliquely installed momentum wheel and the three star shafts are used as backup in fault, so that the reliability of the system is improved, and the zero-crossing problem of the momentum wheel is avoided.
(6) The invention has high precision, integrates two paths of magnetometers in the control circuit board, and adopts the star sensor, thereby effectively improving the precision of the attitude control system.
Drawings
FIG. 1 is a flow chart of a design method of a cube-star zero-momentum attitude control system structure of the present invention.
FIG. 2 is a three-dimensional structure diagram of the whole cube-star zero-momentum attitude control system of the present invention.
FIG. 3 is an exploded view of the cube-star zero momentum attitude control system of the present invention.
FIG. 4 is a perspective view of a main frame of the cube-star zero momentum attitude control system of the present invention.
FIG. 5 is a perspective view of the intermediate support of the cube-star zero momentum attitude control system of the present invention.
FIG. 6 is a perspective view of a star sensor mount of the cubic star zero momentum attitude control system of the present invention.
FIG. 7 is a perspective view of a control module of the cube-star zero momentum attitude control system of the present invention.
Fig. 8 is an expanded schematic diagram of the control module of the cubic star zero momentum attitude control system of the invention.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without inventive step, are within the scope of the present invention.
It should be noted that all the directional indicators (such as up, down, left, right, front, and rear … …) in the embodiment of the present invention are only used to explain the relative position relationship between the components, the movement situation, etc. in a specific posture (as shown in the drawing), and if the specific posture is changed, the directional indicator is changed accordingly.
In addition, the descriptions related to "first", "second", etc. in the present invention are only for descriptive purposes and are not to be construed as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include at least one such feature. In the description of the present invention, "a plurality" means at least two, e.g., two, three, etc., unless explicitly specifically defined otherwise.
In the present invention, unless otherwise expressly specified or limited, the terms "connected," "secured," and the like are to be construed broadly, e.g., "secured" may be fixedly connected, releasably connected, or integral; "connected" may be mechanically or electrically connected. The specific meanings of the above terms in the present invention can be understood by those skilled in the art according to specific situations.
In addition, the technical solutions in the embodiments of the present invention may be combined with each other, but it must be based on the realization of those skilled in the art, and when the technical solutions are contradictory or cannot be realized, such a combination of technical solutions should not be considered to exist, and is not within the scope of the claimed invention.
The following further introduces specific embodiments, technical difficulties and inventions of the present invention with reference to the design examples.
With reference to fig. 1 to 8, a method for designing a structure of a cubic star zero momentum attitude control system includes the following steps:
the cubic star zero momentum attitude control system framework comprises a main framework 1, a middle support 2, a three-axis momentum wheel mounting support 9, an oblique momentum wheel mounting support 11, a star sensor mounting support 13 and an attitude control computer mounting support 23.
Further, with reference to fig. 2 to 4, the main frame 1 includes four layers of frames and four upright posts 4, the four layers of frames are arranged in parallel from top to bottom, each frame is fixedly connected with the four upright posts 4, and the four layers of frames are a first square frame 5, a second U-shaped frame 6, a third square frame 7 and a fourth square frame 8 from top to bottom; the middle bracket 2 is positioned in the middle of the main frame 1; the three-axis momentum wheel mounting bracket 9 is mounted between the first square frame 5 and the second U-shaped frame 6 of the four-layer frame; the oblique momentum wheel mounting bracket 11 is arranged between the second U-shaped frame 6 and the third square frame 7; one end of the star sensor mounting bracket 13 is fixed on the first square frame 5 and the second U-shaped frame 6, and the other end is fixedly connected with the middle bracket 2; the attitude control computer mounting bracket 23 is positioned between the third frame 7 and the fourth frame 8 of the four-layer frame; the volume of the main frame is 0.8U cube star standard model.
the measurement and control assembly comprises a three-axis zero momentum wheel 10, an obliquely-installed zero momentum wheel 12, a star sensor 14, a three-axis magnetic torquer, a gyroscope 15, a GPS receiver, three sun sensors 16 and two magnetometers, wherein the three-axis momentum wheel 10 is fixed on a three-axis momentum wheel mounting bracket 9, the obliquely-installed momentum wheel 12 is fixed on an obliquely-installed momentum wheel mounting bracket 11, the sun sensors 16 in three directions are fixed on three sides of a second U-shaped frame 5 and a third square frame 7, the star sensor 14 and the gyroscope 15 are fixed on the star sensor mounting bracket 13, the three-axis magnetic torquer comprises a first magnetic bar 17, a second magnetic bar 18 and a third magnetic bar 19 which are arranged in an orthogonal mode in pairs, the magnetic bar is formed by winding an enameled wire outside the magnetic induction iron core, and two ends of the first magnetic bar 17 are respectively 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; the second magnetic bars 18 are fixed on the inner walls of the brackets at the two sides of the middle bracket 2; the third magnetic bar 19 is fixed on the inner wall of the third frame 7.
the middle support 2 is of an H-shaped structure so as to be installed in the main frame 1, and two parallel edges of the middle support 2 are fixedly connected with the first square frame 5 and the third square frame 7 of the main frame 1 through bolts respectively.
Further, referring to fig. 5 to 6, the star sensor mounting bracket 13 is fixed to the inner opposite side of the middle bracket 2 and the first and second U-shaped frames 4 and 6 of the main frame 1, and the star sensor 14 and the gyroscope 15 are fixedly connected to the star sensor mounting bracket 13 by bolts. Two ends of the first magnetic bar 17 are respectively and 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; the second magnetic bars 18 are fixed on the inner walls of the brackets at the two sides of the middle bracket 2; the third magnetic bar 19 is fixed on the inner wall of the third frame 7.
further, with reference to fig. 7 to 8, the attitude control computer mounting bracket 23 is fixed on a set of opposite sides between the third frame 7 and the fourth frame 8 of the main frame 1; the control module 3 comprises a wiring board 21 and five control circuit boards 20 which are sequentially arranged in a stacked manner, the control circuit boards 20 are fixedly connected with an attitude control computer mounting bracket 23 through bolts, threaded sleeves 26 are axially arranged between the circuit boards of the control module 3 and used for determining the distance between the circuit boards, and the control module 3 is positioned in the mounting frame 1.
Furthermore, the control module comprises 6 PCB boards, and the circuit boards can be connected through flexible flat cables and interfaces by adopting FPC additive manufacturing technology so as to be integrated on a whole circuit board. The first circuit board 101 is a CPU board, on which 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 are provided; the second circuit board 102 is a wiring board, and has 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; an interface corresponding to the first circuit board 101 is arranged on the first circuit board; the fourth circuit board 104 is a magnetic bar-module board, and has interfaces corresponding to the magnetic bar, the second circuit board 102 and the fifth circuit board 105; the fifth circuit board 105 is a magnetic bar two-module board, and is provided with interfaces corresponding to the magnetic bar and the fourth circuit board 104; the sixth circuit board 106 is a momentum wheel module board, and has interfaces corresponding to the momentum wheel and the second circuit board 102.
Further, the bottom surface and the side surface of the control module board 3 are respectively provided with a test connector 24 and a standard J70 communication connector 25 corresponding to the positions of the rectangular windows, and the test connector and the standard J70 communication connector can be connected with the outside, so that the zero momentum attitude control system can be independently designed, and the design difficulty of the system is greatly reduced as long as the accuracy of the zero momentum attitude control system and an external interface is ensured.
The cube star zero momentum attitude control system has small volume and high function density, completely meets the requirements of the cube star on miniaturization and light weight of the attitude control system, has the characteristics of integrated and independent system design, greatly simplifies the design of the cube star attitude control system, reduces the design difficulty of the attitude control system and simplifies the design flow. In addition, the invention has wide application range, the structure is reserved with four-corner through holes, the four-corner through holes are fixed in the cube star through studs, the invention is suitable for the design standard of the existing cube star, and the invention can be used for the bias momentum attitude control system of the cube star such as standard 2U, 3U, 6U and the like. The control module in the invention combines the FPC additive manufacturing technology, adopts the flexible flat cable and the interface for connection, improves the reliability of the system, saves space and is convenient for integration and modularization of the attitude control system. Meanwhile, the fixed direction of the system in the cube star does not need to be changed according to the direction of the system in the satellite, and momentum exchange between the cube star and the flywheel is realized by adjusting the rotating speed direction of the momentum wheel; meanwhile, the direction of the obliquely-installed momentum wheel and the inclination angles of the three shafts of the star body are equal, the direction is used as a backup in the fault, the magnetometer is integrated in the control circuit board, and the dual-machine cold backup mode is adopted, so that the reliability of the attitude control system can be effectively improved.
The foregoing illustrates and describes the principles, general features, and advantages of the present invention. It will be understood by those skilled in the art that the present invention is not limited to the embodiments described above, which are described in the specification and illustrated only to illustrate the principle of the present invention, but that various changes and modifications may be made therein without departing from the spirit and scope of the present invention, which fall within the scope of the invention as claimed. The scope of the invention is defined by the appended claims and equivalents thereof.
Claims (8)
1. A design method of a cubic star zero momentum attitude control system structure is characterized by comprising the following steps: the method comprises the following steps:
step 1, determining a cubic star zero momentum attitude control system framework structure according to the size requirement of a cubic star:
the cubic star zero momentum attitude control system framework comprises a main framework (1), a middle bracket (2), a three-axis momentum wheel mounting bracket (9), an obliquely-installed momentum wheel mounting bracket (11), a star sensor mounting bracket (13) and an attitude control computer mounting bracket (23);
the main frame (1) comprises four layers of frames and four upright posts (4), wherein the four layers of frames are arranged in parallel from top to bottom, each frame is fixedly connected with the four upright posts (4), and the four layers of frames are a first square frame (5), a second U-shaped frame (6), a third square frame (7) and a fourth square frame (8) from top to bottom; the middle bracket (2) is positioned in the middle of the main frame (1) and is vertically and fixedly connected with the first square frame (5) and the third square frame (7); the three-axis momentum wheel mounting bracket (9) is fixed between the first square frame (5) and the second U-shaped frame (6); the oblique momentum wheel mounting bracket (11) is mounted between the second U-shaped frame (6) and the third square frame (7); one end of the star sensor mounting bracket (13) is fixed on the first square frame (5) and the third square frame (7), and the other end of the star sensor mounting bracket is fixedly connected with the middle bracket (2); the attitude control computer mounting bracket (23) is arranged between the third box (7) and the fourth box (8);
step 2, determining the installation position of the measurement and control assembly according to the structure of the main frame (1) and the design of the function requirement of the cube star:
the measurement and control assembly comprises a three-axis zero momentum wheel (10), an obliquely-installed zero momentum wheel (12), a star sensor (14), a three-axis magnetic torquer, a gyroscope (15), a GPS receiver, three sun sensors (16) and two magnetometers, wherein the three-axis momentum wheel (10) is fixed on the upper part of the main frame (1), the obliquely-installed momentum wheel (12) is fixed in the middle of the main frame (1), the three sun sensors (16) are respectively fixed on three outer side edges of the main frame (1), the star sensor (14) and the gyroscope (15) are fixed on a star sensor mounting bracket (13), the three-axis magnetic torquer comprises a first magnetic bar (17), a second magnetic bar (18) and a third magnetic bar (19) which are orthogonally arranged in pairs, the first magnetic bar (17) is respectively fixedly connected with the first square frame (5) and the three-axis momentum wheel mounting bracket (9), and the second magnetic bar (18) is fixedly connected with the middle bracket (2), the third magnetic bar (19) is fixedly connected with the third frame (7);
step 3, designing an intermediate bracket (2):
the middle support (2) is of an H-shaped structure so as to be conveniently installed in the main frame (1), and two parallel edges on the middle support (2) are fixedly connected with a first square frame (5) and a third square frame (7) of the main frame (1) through bolts respectively;
step 4, designing a control module (3) according to the system residual space:
the control module (3) comprises 6 PCB boards which are sequentially stacked, five control circuit boards (20) and a wiring board (21) are adopted from top to bottom, the GPS receiver and the two magnetometers are integrated on the control module (3), and the control circuit boards (20) are connected through interfaces so as to be integrated on the whole circuit board; the wiring board (21) and a magnetometer module (22) in the control circuit board (20) are respectively and directly connected with an attitude control computer mounting bracket (23), and the control circuit boards (20) in the control module (3) are in supporting connection through studs (26) along the axial direction.
2. The method of designing a cube-star zero-momentum attitude control system structure according to claim 1, wherein: a three-axis momentum wheel mounting bracket (9) is fixed on the first square frame (5) and the upright post (4) of the second U-shaped frame (6) positioned in the middle; the three-axis momentum wheel (10) is fixedly connected with the three-axis momentum mounting bracket (9) through bolts; the oblique momentum wheel bracket (11) is fixed on the second U-shaped frame (6) and the third square frame (7); the oblique momentum wheel (12) is fixedly connected with the oblique momentum wheel bracket (11) through a bolt; sun sensors (16) in three directions are fixed on each side of the second U-shaped frame (6) and the third frame (7); the star sensor mounting bracket (13) is fixed on the middle bracket (2) and the first square frame (5) and the third square frame (7) of the main frame (1), and the star sensor (14) and the gyroscope (15) are fixedly connected with the star sensor mounting bracket (13) through bolts.
3. The method of designing a cube-star zero-momentum attitude control system structure according to claim 1, wherein: two ends of the first magnetic bar (17) are respectively 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); the second magnetic bars (18) are fixed on the inner walls of the brackets at the two sides of the middle bracket (2); the third magnetic bar (19) is fixed on the inner wall of the third frame (7).
4. The method of designing a cube-star zero-momentum attitude control system structure according to claim 1, wherein: the attitude control computer mounting bracket (23) is arranged between the third frame (7) and the fourth frame (8) and is respectively and fixedly connected with the four upright posts (4); the control module (3) is fixed on the attitude control computer mounting bracket (23) through a bolt.
5. The method of designing a cubic star zero momentum attitude control system of claim 1, wherein: and a testing and communication connector (24) corresponding to the position of the rectangular window is arranged on the bottom surface of the control module (3).
6. The method of designing a cubic star zero momentum attitude control system of claim 1, wherein: the cuboidal zero-momentum attitude control system is 0.8U in volume.
7. The method of designing a cubic star zero momentum attitude control system according to claim 1, wherein: the first magnetic bar (17), the second magnetic bar (18) and the third magnetic bar (19) are formed by winding enameled wires outside the magnetic induction core.
8. The method of designing a cubic star zero momentum attitude control system of claim 1, wherein: the control circuit board (20) is manufactured into a flexible circuit board by adopting FPC combined additive manufacturing technology.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210851880.8A CN115027702B (en) | 2022-07-20 | 2022-07-20 | Design method of cube star zero momentum attitude control system structure |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210851880.8A CN115027702B (en) | 2022-07-20 | 2022-07-20 | Design method of cube star zero momentum attitude control system structure |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN115027702A true CN115027702A (en) | 2022-09-09 |
| CN115027702B CN115027702B (en) | 2023-05-05 |
Family
ID=83128094
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202210851880.8A Active CN115027702B (en) | 2022-07-20 | 2022-07-20 | Design method of cube star zero momentum attitude control system structure |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN115027702B (en) |
Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5279483A (en) * | 1990-12-21 | 1994-01-18 | Aerospatiale Societe Nationale Industrielle | Attitude control system for a three-axis stabilized satellite especially a remote sensing satellite |
| CN107368092A (en) * | 2017-08-20 | 2017-11-21 | 南京理工大学 | A kind of integrated cube of star posture control system platform |
| CN110562495A (en) * | 2019-08-15 | 2019-12-13 | 南京理工大学 | Cube star offset momentum attitude control system |
| CN112173171A (en) * | 2019-07-04 | 2021-01-05 | 南京理工大学 | High-integration-level three-unit cube satellite capable of being mechanically transferred into orbit |
| CN114056602A (en) * | 2020-08-04 | 2022-02-18 | 南京理工大学 | Variable-configuration cube star for simulating space debris and working method thereof |
-
2022
- 2022-07-20 CN CN202210851880.8A patent/CN115027702B/en active Active
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5279483A (en) * | 1990-12-21 | 1994-01-18 | Aerospatiale Societe Nationale Industrielle | Attitude control system for a three-axis stabilized satellite especially a remote sensing satellite |
| CN107368092A (en) * | 2017-08-20 | 2017-11-21 | 南京理工大学 | A kind of integrated cube of star posture control system platform |
| CN112173171A (en) * | 2019-07-04 | 2021-01-05 | 南京理工大学 | High-integration-level three-unit cube satellite capable of being mechanically transferred into orbit |
| CN110562495A (en) * | 2019-08-15 | 2019-12-13 | 南京理工大学 | Cube star offset momentum attitude control system |
| CN114056602A (en) * | 2020-08-04 | 2022-02-18 | 南京理工大学 | Variable-configuration cube star for simulating space debris and working method thereof |
Also Published As
| Publication number | Publication date |
|---|---|
| CN115027702B (en) | 2023-05-05 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN110562495B (en) | Cube star offset momentum attitude control system | |
| CN109335023B (en) | Cable-free high-density cube star and assembling method thereof | |
| JP5977780B2 (en) | Vertical die / chip on board | |
| US20100058860A1 (en) | Inertial measurement unit and method of constructing the same using two orthogonal surfaces | |
| CN102313548B (en) | Micro attitude and heading reference system based on 3D stereoscopic packaging technology | |
| US10746758B2 (en) | MEMS-based sensor suite | |
| CN103776448A (en) | Posture course reference system | |
| CN110440798B (en) | High-precision integrated micro-inertia measurement unit and inertial navigation system | |
| CN115027702B (en) | Design method of cube star zero momentum attitude control system structure | |
| CN108957513B (en) | Microminiature position and attitude measurement system | |
| CN112229400B (en) | Small micro-electromechanical gyro inertia/satellite integrated navigation system | |
| CN110440783B (en) | Split type umbrella-shaped mechanical shaking device of laser gyroscope | |
| CN111595338A (en) | High-reliability single-shaft redundant fiber-optic gyroscope inertia measuring device | |
| CN114910068A (en) | Hybrid inertial navigation and satellite combined navigation system | |
| CN113790721A (en) | Planar micro inertial navigation system | |
| CN104931050A (en) | Integrative structure of MEMS (micro electro mechanical system) inertial posture sensor | |
| CN108770183B (en) | Three-axis vibration reduction structure based on FR4 printed circuit board | |
| CN113218381A (en) | Three-axis gyroscope | |
| CN100365383C (en) | Dimesize attitude measurement system in magnetic infrared ray | |
| KR102371693B1 (en) | Gravity gradiometer and method for manufacturing the same | |
| CN218865117U (en) | Arrangement design structure of inertia devices of micro-electromechanical inertial navigation system | |
| CN213594559U (en) | Triaxial integration mechatronic magnetic torquer | |
| CN217585827U (en) | Integrated device of multisource sensor navigation | |
| CN219141850U (en) | Six-axis inertial measurement unit | |
| CN117168496A (en) | Three-dimensional combined navigation system and navigation method |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |