CN103197669B - Satellite multiple attitude control mode test system based on double gimbal control moment gyroscope (DGCMG) structure - Google Patents

Satellite multiple attitude control mode test system based on double gimbal control moment gyroscope (DGCMG) structure Download PDF

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Publication number
CN103197669B
CN103197669B CN201310125910.8A CN201310125910A CN103197669B CN 103197669 B CN103197669 B CN 103197669B CN 201310125910 A CN201310125910 A CN 201310125910A CN 103197669 B CN103197669 B CN 103197669B
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satellite
dgcmg
control
star
gyro
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CN201310125910.8A
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Chinese (zh)
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CN103197669A (en
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杨照华
余远金
王浩
郭雷
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北京航空航天大学
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Abstract

The invention relates to a satellite multiple attitude control mode test system based on a double gimbal control moment gyroscope (DGCMG) structure. Performance of three satellite attitude control systems which are based on a DGCMG, a single gimbal control moment gyroscope (CMG) and a flywheel is tested and verified through a DGCMG actuating mechanism. The satellite multiple attitude control mode test system based on the DGCMG structure comprises a platform system, the satellite attitude control system, a space environment simulation system and a ground station system. The platform system is composed of a tri-axial air bearing table, a satellite service comprehensive management system, a power source and a wireless bridge and used for simulating satellite dynamic characteristics and information management. The satellite attitude control system is composed of a jet propulsion system, the DGCMG, a fiber-optic gyroscope, a star sensor, a sun sensor and a global position system (GPS) receiver and used for determination of an attitude and an orbit and control of a satellite platform. The space environment simulation system is composed of a GPS simulator, a sliding block which is of a pyramidal structure, a sun simulator and a star simulator and used for simulating space interference torque, and part performance of a GPS satellite and a celestial body. The satellite multiple attitude control mode test system based on the DGCMG structure can provide ground testing and verification for multiple attitudes of a satellite.

Description

Based on a kind of satellite many attitude control model test macro of DGCMG configuration

Technical field

The present invention relates to based on the two framework control-moment gyro of a kind of DGCMG() the satellite many attitude control model test macro of configuration, be applicable to based on the satellite gravity anomaly conceptual design under two framework control-moment gyro, single-gimbal control momentum gyro and counteraction flyback three class attitude control actuator and different configuration thereof.

Background technology

21st century, space science technology is fast-developing, and subhost kinetic force significantly promotes, and the multiple satellite such as small-sized quick maneuvering satellite, over the ground high-accuracy stable satellite attracts Liao Duojia mechanism to carry out large quantity research.The attitude topworks of satellite mainly contains jet-propulsion, control-moment gyro and counteraction flyback etc.As small satellite attitude control ground simulating device and method in research in the past, open (bulletin) number CN1119310031A discloses a kind of method being applicable to the attitude control ground emulation of multiple different model satellite; The simulated test device of small satellite attitude control reliability demonstration and method of testing, then paid close attention to the research of gesture stability reliability aspect in open (bulletin) number CN111444899.The then research being absorbed in a kind of attitude topworks in the research of attitude of satellite topworks in addition more, as adopted the quick attitude of satellite/angular momentum of DGCMG to jointly control, the articles such as bias momentum wheel control satellite gravity anomaly are studied the attitude such as DGCMG and momenttum wheel topworks.Control-moment gyro has the application relative maturity of multiple classical configuration as attitude topworks, mostly is at present and adopts a certain configuration in the use of control-moment gyro, installs after configuration is determined and can not change again.As the control-moment gyro of other configurations need be adopted can only to be disassembled as attitude topworks, more again build the control-moment gyro group of required configuration, very loaded down with trivial details, waste a large amount of time costs and financial cost.

Summary of the invention

Technology of the present invention is dealt with problems and is: overcome the deficiencies in the prior art, provides the satellite many attitude control model test macro based on a kind of DGCMG configuration, utilizes this test platform, for the multiple appearance control of satellite provides ground test and checking.

Technical solution of the present invention: based on a kind of satellite many attitude control model test macro of DGCMG configuration, comprising: plateform system, satellite attitude control system and space environment simulation system; Described plateform system comprises three-axis air-bearing table 1, Star Service total management system 2, power-supply system 16 and the first wireless bridge 12; Satellite attitude control system comprises the two framework control-moment gyro DGCMG15 of jet propulsion system 4, four, optical fibre gyro 13, star sensor 10, sun sensor 8 and GPS 5; Described space environment simulation system comprises GPS simulator 14, metal block 6, pyramid configuration slide bar 7 solar simulator 9, star emulator 11; Earth station system comprises ground simulation computing machine 18 and the second wireless bridge 17;

The stage body of described three-axis air-bearing table 1 adopts pocketed hollow cylindrical structure, be divided into three cabin bodies, wherein Star Service total management system 2, power-supply system 16 is positioned at bottom cabin body, and the first wireless bridge 12 is positioned at the superiors' cabin body and is connected with Star Service total management system 2; Star Service total management system 2 receive optical fibre gyro 13 in satellite control system, star sensor 10, sun sensor 8 and GPS 5 information carry out the attitude track real-time simulated animation of satellite, Star Service total management system 2 is communicated with ground simulation computing machine 18 by the first wireless bridge 12, second wireless bridge 17 simultaneously, and send order according to the instruction of ground simulation computer 18 to DGCMG15, carry out the gesture stability of satellite platform;

Jet propulsion system 4 is positioned at cabin, middle layer body; The spaced 100 ° of bottoms being arranged on the bottom cabin body of three-axis air-bearing table 1 of four two framework control-moment gyros 15, single frame moment gyro and the counteraction flyback of multiple configuration is formed respectively, for realizing the attitude control system of three class variety classes topworkies by locking inside casing or housing, inside casing and housing; Optical fibre gyro 13, star sensor 10, sun sensor 8 and GPS 5 are all positioned at the superiors' cabin body, and above-mentioned parts are all connected with Star Service total management system 2; GPS simulator 14 is positioned at the superiors' cabin body; Metal block 6 is installed on the connected pyramid configuration slide bar 7 of the superiors' cabin body, and solar simulator 9, star emulator 11 are installed on the superiors' cabin body, and above-mentioned parts are all connected with Star Service total management system 2; Solar simulator 9, star emulator 11, GPS simulator 14 are connected with GPS 5 with sun sensor 8, star sensor 10 respectively.

Principle of the present invention is: utilize two framework control-moment gyro group to carry out the motor-driven of the attitude of satellite.Two framework control-moment gyro can be used as single-gimbal control momentum gyro by locking housing or inside casing, and change multiple configuration according to the variable-angle of locking housing or inside casing; Two framework control-moment gyro can also can be used as counteraction flyback by locking housing and inside casing; The satellite attitude control system checking based on two framework control-moment gyro, single-gimbal control momentum gyro and counteraction flyback three class attitude topworks can be completed.

The present invention's advantage is compared with prior art: the present invention adopts the two framework control-moment gyro of 4 parallel configuration as the topworks of test platform, two framework control-moment gyro can be used as single-gimbal control momentum gyro by locking housing or inside casing, and the variable-angle according to locking housing or inside casing changes multiple configuration, two framework control-moment gyro can be used as counteraction flyback by locking housing and inside casing, and the variable-angle according to locking housing and inside casing changes multiple configuration, can realize based on two framework control-moment gyro, the satellite attitude control system checking of single-gimbal control momentum gyro and counteraction flyback three class attitude topworks.Namely the checking of the satellite attitude control system that three class topworkies are formed can be completed by a kind of configuration of topworks.

Accompanying drawing explanation

Fig. 1 is overall layout chart of the present invention;

Fig. 2 is satellite spatial disturbance torque equivalent mechanism of the present invention;

Fig. 3 is measuring system of satellite attitude of the present invention;

Fig. 4 is jet propulsion system of the present invention;

Fig. 5 is total system on star of the present invention;

Fig. 6 is DGCMG group of the present invention;

Fig. 7 is signal flow diagram between present system.

Embodiment

As shown in Figure 1, the present invention includes: plateform system, satellite attitude control system and space environment simulation system; Plateform system comprises three-axis air-bearing table 1, Star Service total management system 2, power-supply system 16 and the first wireless bridge 12; Satellite attitude control system comprises the two framework control-moment gyro DGCMG15 of jet propulsion system 4, four, optical fibre gyro 13, star sensor 10, sun sensor 8 and GPS 5; Described space environment simulation system comprises GPS simulator 14, metal block 6, pyramid configuration slide bar (7), solar simulator 9, star emulator 11; Earth station system comprises ground simulation computing machine 18 and the second wireless bridge 17.

Three-axis air-bearing table 1 is for analog satellite platform, carry useful load, the stage body of three-axis air-bearing table 1 adopts pocketed hollow cylindrical structure, be divided into three cabin bodies, wherein Star Service total management system 2, power-supply system 16 is positioned at bottom cabin body, first wireless bridge 12 is positioned at the superiors' cabin body and is connected with House keeping system 2, power-supply system 16, for whole test macro is powered, Star Service total management system 2 receives the optical fibre gyro 13 in satellite control system, star sensor 10, the information of sun sensor 8 and GPS 5 carries out the attitude track real-time simulated animation of satellite, Star Service total management system 2 is by the first wireless bridge 12 simultaneously, second wireless bridge 17 communicates with ground simulation computing machine 18, and send order according to the instruction of ground simulation computer 18 to DGCMG15, carry out the gesture stability of satellite platform,

Jet propulsion system 4 is positioned at cabin, middle layer body, and jet propulsion system 4 is responsible for initial rate damping and angular momentum dumping; The topworks that four two framework control-moment gyros 15 adjust as the attitude of satellite, spaced 100 ° of bottoms being arranged on the bottom cabin body of three-axis air-bearing table 1, single frame moment gyro and the counteraction flyback of multiple configuration is formed respectively, for realizing the attitude control system of three class variety classes topworkies by locking inside casing or housing, inside casing and housing; Optical fibre gyro 13, star sensor 10, sun sensor 8 and GPS 5 are all positioned at the superiors' cabin body, and above-mentioned parts are all connected with Star Service total management system 2; GPS simulator 14 is positioned at the superiors' cabin body, for the signal of each gps satellite in simulation space, solar simulator 9 and star emulator 11, for simulated solar irradiation and starlight, carry out virtual space disturbance torque by the metal block 6 be arranged on three-axis air-bearing table 1 top gold tower configuration slide bar (7); Metal block 6 is installed on the connected pyramid configuration slide bar 7 of the superiors' cabin body, and solar simulator 9, star emulator 11 are installed on the superiors' cabin body, and above-mentioned parts are all connected with Star Service total management system 2; Solar simulator 9, star emulator 11, GPS simulator 14 are connected with GPS 5 with sun sensor 8, star sensor 10 respectively, adopt star sensor 10 and sun sensor 8 obtain the starlight of space environment simulation system culminant star emulator 11 and solar simulator 9 simulation and sunshine respectively and combine optical fibre gyro 13 to determine the attitude information of satellite; The orbit information that gps signal that GPS simulator 14 in space environment simulation system simulates determines satellite is received by GPS 5; Solar simulator 9 in space environment simulation system and star emulator 11 are respectively used to sunshine and starlight in virtual space, and in satellite control system, sun sensor 8 and star sensor 10 utilize the attitude of this information determination satellite.

If Fig. 2 is vertical view of the present invention, four slide bars 7 are fixed on the edge of top layer, form Pyramid, and a metal block 6 installed by each slide bar 7.Pyramid configuration slide bar 7 and metal block 6 form the slide block structure of gold tower configuration.Virtual space disturbance torque is carried out by the slip of metal block 6 on pyramid configuration slide bar 7.

Be illustrated in figure 3 three-axis air-bearing table 1 the superiors of the present invention cabin body distribution plan, sun sensor 8 is arranged on the second quadrant, solar simulator 9 is arranged on the photosensitive head of sun sensor 8, in simulation process, Star Service total management system controls solar simulator 9 and produces pumping signal, carry out platform stance angle after sun sensor 8 obtains this signal to determine, and pass attitude angle information back Star Service total management system 2.Star sensor 10 is arranged on first quartile, star emulator 11 is arranged on the photosensitive head of star sensor 10, in simulation process, Star Service total management system 2 controls star emulator 11 and produces pumping signal, star sensor 10 obtains measuring table attitude angle after this signal, and passes attitude angle information back Star Service total management system 2.Three orthogonal optical fibre gyros 13 are arranged on fourth quadrant, measure three-axis air-bearing table 1 attitude angular rate in three directions, and this information is passed to Star Service total management system 2 by three optical fibre gyros 13.The attitude information that Star Service total management system 2 overall treatment sun sensor 8, star sensor 10 and three optical fibre gyros 13 provide carries out the determination of final carriage.GPS 5, GPS simulator 14 are arranged on third quadrant, and GPS 5 receives the information of the GPS simulator 14 from space environment simulation system, and information are sent to Star Service total management system 2 and are used for determining orbital position.

Be illustrated in figure 4 the layout of cabin, the middle layer body of three-axis air-bearing table 1 of the present invention, this layer of cabin body installs jet propulsion system 4.At installation two jet propulsion systems 4 of this cabin body symmetry, these two jet propulsion systems 4 are all connected with Star Service total management system 2, control its jet size and Orientation carry out attitude maneuver by Star Service total management system 2.

Be illustrated in figure 5 the distribution plan of the bottom cabin body of three-axis air-bearing table 1 of the present invention, each quadrant all lays circuit board and the balancing weight 3 of Star Service total system 2, and the quality of balancing weight 3 can be regulated when the counterweight of the stage body of three-axis air-bearing table 1 is unbalanced to make the counterweight of stage body balanced.

Be illustrated in figure 6 three-axis air-bearing table 1 upward view of the present invention, power-supply system 16 is fixedly mounted on stage body bottom surface, four even fixing being arranged on bottom stage body of two framework control-moment gyro 15.Sun sensor 8, solar simulator 9, star sensor 10, star emulator 11, three optical fibre gyros 13, GPS, jet propulsion system 4, Star Service total management system 2 are all connected with power-supply system 16 by RS422.Above-mentioned all RS422 are guided to bottom stage body by the cylinder of the centre of three-axis air-bearing table 1, and are connected with power-supply system 16.

Shown in Fig. 6, four two framework control-moment gyros 15 of the present invention form the bottom that parallel configuration is installed on air supporting stage body 1.For each pair of framework control-moment gyro sets up coordinate system, X1, X2, X3, X4 all point to direction, local due east; Y1, Y2, Y3, Y4 all point to local direct north; Z1, Z2, Z3, Z4 all point to the earth's core.Each when inter and outer gimbal angle is zero, pin the outside framework of two framework control-moment gyro 15, then form the single-gimbal control momentum gyro of two parallel configuration.α angle (0< α <100) is turned over clockwise and locking housing around housing axle Y3 at the 3rd DGCMG15, 4th DGCMG15 turns over α angle and locking housing around between the housing axle Y4 inverse time, one GCMG turns over α angle and locking inside casing counterclockwise around housing axle Y1, 2nd DGCMG15 turns over α angle clockwise around housing axle Y2 and pins the single-gimbal control momentum gyro that inside casing then forms pyramid configuration, if at inside casing and the housing of the single-gimbal control momentum gyro finger lock all pairs of framework control-moment gyros 15 of formation pyramid configuration, then constitute the counteraction flyback of pyramid configuration.Therefore, the single frame utilizing two framework control-moment gyros 15 of configuration as shown in Figure 6 can form multiple typical configurations controls gyro group and counteraction flyback group.

Be illustrated in figure 7 signal flow diagram between present system.Star Service total management system 2 is information processing centres of this test macro.Control solar simulator 9 by Star Service total management system 2, star emulator 11, GPS simulator 14 export corresponding sunshine signal, starlight signal and gps satellite signal.Four slide blocks 6 controlled on pyramid configuration slide bar 7 by Star Service manager 2 do cyclical movement, virtual space disturbance torque, and three-axis air-bearing table 1 is trembled under this moment.Sun sensor 8, star sensor 10, GPS 5 receive the corresponding sunshine signal, starlight signal and the gps satellite signal that are exported by solar simulator 9, star emulator 11, GPS simulator 14 respectively and carry out the measurement of satellite platform attitude angle and the measurement of gps satellite information.Optical fibre gyro 13 is for measuring the attitude angular rate of three-axis air-bearing table 1.The attitude angle that sun sensor 8, star sensor 10, optical fibre gyro 13 are measured and the gps satellite information that GPS 5 is measured send Star Service total management system 2 to and preserve.Star Service total management system 2 calculates the track of satellite platform by gps satellite information, and carries out track and attitude maneuver according to the two framework control-moment gyro 15 of attitude information control of orbit information and satellite platform.Star Service total management system 2 carries out information interaction by the first wireless bridge 12 and the second wireless bridge 17 with ground simulation computing machine 18.

The content be not described in detail in instructions of the present invention belongs to the known prior art of professional and technical personnel in the field.

Claims (4)

1., based on a kind of satellite many attitude control model test macro of two framework control-moment gyro DGCMG (15) configuration, it is characterized in that comprising: plateform system, satellite attitude control system and space environment simulation system, described plateform system comprises three-axis air-bearing table (1), Star Service total management system (2), power-supply system (16) and the first wireless bridge (12), described satellite attitude control system comprises jet propulsion system (4), four two framework control-moment gyro DGCMG (15), optical fibre gyro (13), star sensor (10), sun sensor (8) and GPS (5), described space environment simulation system comprises GPS simulator (14), metal block (6), pyramid configuration slide bar (7) solar simulator (9), star emulator (11), earth station system comprises ground simulation computing machine (18) and the second wireless bridge (17), the stage body of described three-axis air-bearing table (1) adopts pocketed hollow cylindrical structure, be divided into three cabin bodies, wherein Star Service total management system (2), power-supply system (16) is positioned at bottom cabin body, and the first wireless bridge (12) is positioned at the superiors' cabin body and is connected with Star Service total management system (2), Star Service total management system (2) receives the optical fibre gyro (13) in satellite control system, star sensor (10), the information of sun sensor (8) and GPS (5) carries out the attitude track real-time simulated animation of satellite, Star Service total management system (2) is by the first wireless bridge (12) simultaneously, second wireless bridge (17) communicates with ground simulation computing machine 18, and send order according to the instruction of ground simulation computer (18) to two framework control-moment gyro DGCMG (15), carry out the gesture stability of satellite platform,
Jet propulsion system (4) is positioned at cabin, middle layer body; The spaced 100 ° of bottoms being arranged on the bottom cabin body of three-axis air-bearing table (1) of four two framework control-moment gyros (15), single frame moment gyro and the counteraction flyback of multiple configuration is formed respectively, for realizing the attitude control system of three class variety classes topworkies by locking inside casing or housing, inside casing and housing; Optical fibre gyro (13), star sensor (10), sun sensor (8) and GPS (5) are all positioned at the superiors' cabin body, and optical fibre gyro (13), star sensor (10), sun sensor (8) are all connected with Star Service total management system (2) with GPS (5); GPS simulator (14) is positioned at the superiors' cabin body; Metal block (6) is installed on the connected pyramid configuration slide bar (7) of the superiors' cabin body, solar simulator (9), star emulator (11) are installed on the superiors' cabin body, and above-mentioned parts are all connected with Star Service total management system (2); Solar simulator (9), star emulator (11), GPS simulator (14) are connected with GPS (5) with sun sensor (8), star sensor (10) respectively.
2. the satellite multiple appearance control pattern test macro based on a kind of two framework control-moment gyro DGCMG (15) configuration according to claim 1, it is characterized in that: described pair of framework control-moment gyro DGCMG (15) is as topworks, when its inter and outer gimbal angle is all zero, pin the outside framework of two framework control-moment gyro DGCMG (15), then form the single-gimbal control momentum gyro of two parallel configuration, for studying the performance of single-gimbal control momentum gyro as the Satellite Attitude Control System of topworks of two parallel configuration.
3. the satellite multiple appearance control pattern test macro based on a kind of two framework control-moment gyro DGCMG (15) configuration according to claim 1, it is characterized in that: described pair of framework control-moment gyro DGCMG (15) is as topworks, α angle is turned over clockwise around housing axle Y3 at the 3rd two framework control-moment gyro DGCMG (15), 0< α <100, and locking housing, 4th two framework control-moment gyro DGCMG (15) turns over α angle and locking housing around between the housing axle Y4 inverse time, first two framework control-moment gyro DGCMG (15) turns over α angle and locking inside casing counterclockwise around housing axle Y1, second two framework control-moment gyro DGCMG (15) turns over α angle clockwise around housing axle Y2 and pins the single-gimbal control momentum gyro that inside casing then forms pyramid configuration.
4. the satellite multiple appearance control pattern test macro based on a kind of two framework control-moment gyro DGCMG (15) configuration according to claim 1, it is characterized in that: described pair of framework control-moment gyro DGCMG (15) is as topworks, 3rd pair of framework control-moment gyro DGCMG (15) turns over α angle clockwise around housing axle Y3, 0< α <100, and locking housing, 4th pair of framework control-moment gyro DGCMG (15) turns over α angle and locking housing around between the housing axle Y4 inverse time, first pair of framework control-moment gyro DGCMG (15) turns over α angle and locking inside casing counterclockwise around housing axle Y1, second pair of framework control-moment gyro DGCMG (15) turns over α angle clockwise around housing axle Y2, and also pinning housing and inside casing then form the counteraction flyback of the pyramid configuration of pyramid configuration simultaneously.
CN201310125910.8A 2013-04-12 2013-04-12 Satellite multiple attitude control mode test system based on double gimbal control moment gyroscope (DGCMG) structure CN103197669B (en)

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Families Citing this family (11)

* Cited by examiner, † Cited by third party
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US10202208B1 (en) 2014-01-24 2019-02-12 Arrowhead Center, Inc. High control authority variable speed control moment gyroscopes
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CN105035370B (en) * 2015-07-31 2017-03-29 上海卫星工程研究所 Three-axis air-bearing table platform analog
CN105173129B (en) * 2015-09-18 2017-05-17 南京航空航天大学 Triaxial air bearing table leveling method
CN105865432B (en) * 2016-03-31 2017-07-18 北京航空航天大学 A kind of mixed filtering method and test platform for many source noises of gyroscope
CN105807780B (en) * 2016-05-30 2017-06-20 北京航空航天大学 A kind of anti-interference attitude control method and checking device based on flywheel output bias
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Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN101301934A (en) * 2008-04-22 2008-11-12 北京航空航天大学 Double-frame magnetic suspension control moment gyroscope control system
CN101599670A (en) * 2009-05-27 2009-12-09 北京航空航天大学 A kind of integrating double-framework magnetically suspended control moment gyroscope (MSCMG) magnetic bearing control system
CN101995824A (en) * 2010-10-26 2011-03-30 哈尔滨工业大学 Semi-physical simulation system for attitude control of star-arrow integrated spacecraft
CN102289211A (en) * 2011-06-24 2011-12-21 北京航空航天大学 Satellite attitude control semiphysical simulation system based on multi-target machine
CN102323825A (en) * 2011-07-18 2012-01-18 北京航空航天大学 Torque compensation control method of DGMSCMG (double-gimbal magnetically suspended control moment gyroscope) system for spacecraft maneuver
CN102354123A (en) * 2011-07-18 2012-02-15 北京航空航天大学 Cross-platform extendible satellite dynamic simulation test system

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN101301934A (en) * 2008-04-22 2008-11-12 北京航空航天大学 Double-frame magnetic suspension control moment gyroscope control system
CN101599670A (en) * 2009-05-27 2009-12-09 北京航空航天大学 A kind of integrating double-framework magnetically suspended control moment gyroscope (MSCMG) magnetic bearing control system
CN101995824A (en) * 2010-10-26 2011-03-30 哈尔滨工业大学 Semi-physical simulation system for attitude control of star-arrow integrated spacecraft
CN102289211A (en) * 2011-06-24 2011-12-21 北京航空航天大学 Satellite attitude control semiphysical simulation system based on multi-target machine
CN102323825A (en) * 2011-07-18 2012-01-18 北京航空航天大学 Torque compensation control method of DGMSCMG (double-gimbal magnetically suspended control moment gyroscope) system for spacecraft maneuver
CN102354123A (en) * 2011-07-18 2012-02-15 北京航空航天大学 Cross-platform extendible satellite dynamic simulation test system

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