CN103197669A - 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 PDFInfo
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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
Technical field
The present invention relates to two framework control-moment gyros based on a kind of DGCMG() the multiple attitude control mode test macro of satellite of configuration, be applicable to based on two framework control-moment gyros, single frame control-moment gyro and counteraction flyback three class attitude control actuators and the attitude of satellite control conceptual design under the isomorphism type not thereof.
Background technology
21st century, the fast development of space science technology, the subhost kinetic force significantly promotes, and small-sized quick maneuvering satellite, multiple satellite such as high precision stabilized satellite has attracted how tame mechanism to carry out big quantity research over the ground.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, open (bulletin) number CN1119310031A has announced a kind of attitude control ground method of emulation that is applicable to multiple different model satellite in research in the past; Simulated test device and the method for testing of small satellite attitude control reliability demonstration have then been paid close attention to the research of attitude control reliability aspect among open (bulletin) number CN111444899.In addition in the then researchs of being absorbed in a kind of attitude topworks aspect the research of attitude of satellite topworks more, as adopting the quick attitude of satellite/angular momentum of DGCMG to jointly control, articles such as bias momentum wheel guard star attitude control are studied attitude topworkies such as DGCMG and momenttum wheels.Control-moment gyro has the application of multiple classical configuration ripe relatively as attitude topworks, and the use at control-moment gyro at present mostly is a certain configuration of employing, and the installation configuration can not change after determining again.Adopt the control-moment gyro of other configurations it can only be disassembled as attitude topworks as need, build the control-moment gyro group of required configuration more again, very loaded down with trivial details, waste a large amount of time cost 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, provide based on the multiple attitude control mode test macro of a kind of satellite of DGCMG configuration, utilize this test platform, for the multiple appearance control of satellite provides ground test and checking.
Technical solution of the present invention: based on the multiple attitude control mode test macro of a kind of satellite 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 first wireless bridge 12; Satellite attitude control system comprises jet propulsion system 4, four two framework control-moment gyro DGCMG15, optical fibre gyro 13, star sensor 10, sun sensor 8 and GPS receivers 5; Described space environment simulation system comprises GPS simulator 14, metal slide block 6, pyramid configuration slide bar 7 solar simulators 9, star emulator 11; Earth station system comprises ground simulation computing machine 18 and second wireless bridge 17;
The stage body of described three-axis air-bearing table 1 adopts pocketed open circles column type structure, be divided into three cabin bodies, wherein the Star Service total management system 2, and power-supply system 16 is positioned at bottom cabin body, and first wireless bridge 12 is positioned at the superiors' cabin body and links to each other with Star Service total management system 2; The information that Star Service total management system 2 receives optical fibre gyro 13, star sensor 10, sun sensor 8 and GPS receiver 5 in the satellite control system is carried out the attitude track real-time simulation of satellite and is calculated, Star Service total management system 2 is communicated by letter with ground simulation computing machine 18 by 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 attitude control of satellite platform;
Jet propulsion system 4 is positioned at cabin, middle layer body; Four two 100 ° of bottoms that are installed in the bottom cabin body of three-axis air-bearing table 1, framework control-moment gyro 15 spaces, constitute single frame moment gyro and the counteraction flyback of multiple configuration respectively by locking inside casing or housing, inside casing and housing, be used for realizing the attitude control system of three class variety classes topworkies; Optical fibre gyro 13, star sensor 10, sun sensor 8 and GPS receiver 5 all are positioned at the superiors' cabin body, and above-mentioned parts all link to each other with Star Service total management system 2; GPS simulator 14 is positioned at the superiors' cabin body; Metal slide block 6 is installed on the continuous 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 all link to each other with Star Service total management system 2; Solar simulator 9, star emulator 11, GPS simulator 14 link to each other with sun sensor 8, star sensor 10 and GPS receiver 5 respectively.
Principle of the present invention is: utilize two framework control-moment gyro groups to carry out the motor-driven of the attitude of satellite.Two framework control-moment gyros can be used as the single frame control-moment 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 gyros can also can be used as counteraction flyback by locking housing and inside casing; Can finish the satellite attitude control system checking based on two framework control-moment gyros, single frame control-moment gyro and counteraction flyback three class attitude topworkies.
The present invention's advantage compared with prior art is: the present invention adopts the two framework control-moment gyros of 4 parallel configuration as the topworks of test platform, two framework control-moment gyros can be used as the single frame control-moment 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 gyros 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 gyros, the satellite attitude control system checking of single frame control-moment gyro and counteraction flyback three class attitude topworkies.Namely can finish the checking of the satellite attitude control system of three class topworkies formation by a kind of configuration of topworks.
Description of drawings
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 the star of the present invention;
Fig. 6 is DGCMG group of the present invention;
Fig. 7 is signal flow diagram between system of the present invention.
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 first wireless bridge 12; Satellite attitude control system comprises jet propulsion system 4, four two framework control-moment gyro DGCMG15, optical fibre gyro 13, star sensor 10, sun sensor 8 and GPS receivers 5; Described space environment simulation system comprises GPS simulator 14, metal slide block 6, pyramid configuration slide bar (7), solar simulator 9, star emulator 11; Earth station system comprises ground simulation computing machine 18 and second wireless bridge 17.
Three-axis air-bearing table 1 is used for the analog satellite platform, carry useful load, the stage body of three-axis air-bearing table 1 adopts pocketed open circles column type structure, be divided into three cabin bodies, wherein the 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 links to each other with Star Service management system 2, power-supply system 16, be the whole test system power supply, the optical fibre gyro 13 that Star Service total management system 2 receives in the satellite control system, star sensor 10, the information of sun sensor 8 and GPS receiver 5 is carried out the attitude track real-time simulation of satellite and is calculated, Star Service total management system 2 is by first wireless bridge 12 simultaneously, second wireless bridge 17 is communicated by letter with ground simulation computing machine 18, and send order according to the instruction of ground simulation computer 18 to DGCMG15, carry out the attitude control 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; Four topworkies that two framework control-moment gyros 15 are adjusted as the attitude of satellite, 100 ° of bottoms that are installed in the bottom cabin body of three-axis air-bearing table 1, space, constitute single frame moment gyro and the counteraction flyback of multiple configuration respectively by locking inside casing or housing, inside casing and housing, be used for realizing the attitude control system of three class variety classes topworkies; Optical fibre gyro 13, star sensor 10, sun sensor 8 and GPS receiver 5 all are positioned at the superiors' cabin body, and above-mentioned parts all link to each other with Star Service total management system 2; GPS simulator 14 is positioned at the superiors' cabin body, the signal that is used for each gps satellite of simulation space, solar simulator 9 and star emulator 11 are used for simulated solar irradiation and starlight, come the virtual space disturbance torque by the metal slide block 6 that is installed on the three-axis air-bearing table 1 top gold tower configuration slide bar (7); Metal slide block 6 is installed on the continuous 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 all link to each other with Star Service total management system 2; Solar simulator 9, star emulator 11, GPS simulator 14 link to each other with sun sensor 8, star sensor 10 and GPS receiver 5 respectively, and employing star sensor 10 and sun sensor 8 obtain starlight and the sunshine of space environment simulation system culminant star emulator 11 and solar simulator 9 simulations respectively and unite the attitude information that optical fibre gyro 13 is determined satellite; Determine the orbit information of satellite with the gps signal of GPS simulator 14 simulations in the GPS receiver 5 reception space environment simulation systems; Solar simulator 9 in the space environment simulation system and star emulator 11 are respectively applied to sunshine and starlight in the virtual space, and sun sensor 8 and star sensor 10 utilize this information to determine the attitude of satellite in the satellite control system.
Be vertical view of the present invention as Fig. 2, four slide bars 7 are fixed on the edge of top layer, form Pyramid, and a metal slide block 6 is installed on each slide bar 7.Pyramid configuration slide bar 7 and metal slide block 6 constitute the slide block structure of gold tower configuration.Come the virtual space disturbance torque by the slip of metal slide block 6 on pyramid configuration slide bar 7.
Be illustrated in figure 3 as three-axis air-bearing table 1 the superiors' cabin body distribution plan of the present invention, sun sensor 8 is installed in second quadrant, solar simulator 9 is installed on the photosensitive head of sun sensor 8, Star Service total management system control solar simulator 9 produces pumping signal in simulation process, sun sensor 8 obtains carrying out behind this signal the platform attitude angle to be determined, and passes attitude angle information back Star Service total management system 2.Star sensor 10 is installed in first quartile, star emulator 11 is installed on the photosensitive head of star sensor 10, Star Service total management system 2 control star emulators 11 produce pumping signal in simulation process, star sensor 10 obtains measuring table attitude angle behind this signal, and passes attitude angle information back Star Service total management system 2.The optical fibre gyro 13 of three quadratures is installed in fourth quadrant, measures the attitude angle speed of three-axis air-bearing table 1 on three directions by three optical fibre gyros 13, and this information is passed to Star Service total management system 2.The attitude information that Star Service total management system 2 overall treatment sun sensors 8, star sensor 10 and three optical fibre gyros 13 provide carries out determining of final attitude.GPS receiver 5, GPS simulator 14 are installed in third quadrant, and GPS receiver 5 receives the information from the GPS simulator 14 of space environment simulation system, and information is sent to Star Service total management system 2 is used for determining orbital position.
Be illustrated in figure 4 as the layout of cabin, the middle layer body of three-axis air-bearing table 1 of the present invention, this layer cabin body is installed jet propulsion system 4.At two jet propulsion systems 4 of installation of this cabin body symmetry, these two jet propulsion systems 4 all link to each other with Star Service total management system 2, carry out attitude maneuver by Star Service total management system 2 its jet size and Orientations of control.
Be illustrated in figure 5 as the distribution plan of the bottom cabin body of three-axis air-bearing table 1 of the present invention, each quadrant is all laid circuit board and the balancing weight 3 of Star Service total system 2, and the quality that can regulate balancing weight 3 when the counterweight of the stage body of three-axis air-bearing table 1 is unbalanced makes the counterweight equilibrium of stage body.
Be illustrated in figure 6 as three-axis air-bearing table 1 upward view of the present invention, power-supply system 16 is fixedly mounted on the stage body bottom surface, four two framework control-moment gyros 15 even fixing being installed in bottom the stage body.Sun sensor 8, solar simulator 9, star sensor 10, star emulator 11, three optical fibre gyros 13, GPS receiver, jet propulsion system 4, Star Service total management system 2 all link to each other with power-supply system 16 by RS422.Above-mentioned all RS422 guide to the stage body bottom by the cylinder of the centre of three-axis air-bearing table 1, and link to each other with power-supply system 16.
Shown in Figure 6, four two framework control-moment gyros 15 of the present invention constitute the bottom that parallel configuration is installed on air supporting stage body 1.For each two framework control-moment gyro is set up coordinate system, X1, X2, X3, X4 all point to local due east direction; Y1, Y2, Y3, Y4 all point to local direct north; Z1, Z2, Z3, Z4 all point to the earth's core.When each at inside and outside framework angle is zero, pin the outside framework of two framework control-moment gyros 15, then constitute the single frame control-moment gyro of two parallel configuration.Turn over α angle (0＜α＜100) and locking housing at the 3rd DGCMG15 clockwise around housing axle Y3, the 4th DGCMG15 turns over α angle and locking housing around housing axle Y4 between the inverse time, the one GCMG turns over α angle and locking inside casing counterclockwise around housing axle Y1, the 2nd DGCMG15 turns over the α angle clockwise and pins the single frame control-moment gyro that inside casing then constitutes the pyramid configuration around housing axle Y2, if at inside casing and the housing of all two framework control-moment gyros 15 of the single frame control-moment gyro finger lock that constitutes the pyramid configuration, then constituted the counteraction flyback of pyramid configuration.Therefore, utilize two framework control-moment gyros 15 of configuration as shown in Figure 6 can constitute single frame control gyro group and the counteraction flyback group of multiple typical configurations.
Be illustrated in figure 7 as signal flow diagram between system of the present invention.Star Service total management system 2 is information processing centres of this test macro.By Star Service total management system 2 control solar simulators 9, star emulator 11, GPS simulator 14 output corresponding sunshine signal, starlight signal and gps satellite signals.Do cyclical movement by four slide blocks 6 on the Star Service manager 2 control pyramid configuration slide bars 7, the virtual space disturbance torque trembles three-axis air-bearing table 1 under this moment.Corresponding sunshine signal, starlight signal and gps satellite signal that sun sensor 8, star sensor 10, GPS receiver 5 receive respectively by solar simulator 9, star emulator 11,14 outputs of GPS simulator carry out the measurement of satellite platform attitude angle and the measurement of gps satellite information.Optical fibre gyro 13 is used for measuring the attitude angle speed of three-axis air-bearing table 1.The gps satellite information that the attitude angle that sun sensor 8, star sensor 10, optical fibre gyro 13 are measured and GPS receiver 5 are measured sends Star Service total management system 2 to and preserves.Star Service total management system 2 is passed through the track of gps satellite information calculations satellite platform, and carries out track and attitude maneuver according to the two framework control-moment gyros 15 of attitude information control of orbit information and satellite platform.Star Service total management system 2 is carried out information interaction by first wireless bridge 12 and second wireless bridge 17 with ground simulation computing machine 18.
The content that is not described in detail in the instructions of the present invention belongs to this area professional and technical personnel's known prior art.
Claims (4)
1. based on a kind of DGCMG(15) the multiple attitude control mode test macro of satellite of 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 first wireless bridge (12); Described satellite attitude control system comprises the two framework control-moment gyro DGCMG(15 of jet propulsion system (4), four), optical fibre gyro (13), star sensor (10), sun sensor (8) and GPS receiver (5); Described space environment simulation system comprises GPS simulator (14), metal slide block (6), pyramid configuration slide bar (7) solar simulator (9), star emulator (11); Earth station system comprises ground simulation computing machine (18) and second wireless bridge (17); The stage body of described three-axis air-bearing table (1) adopts pocketed open circles column type structure, be divided into three cabin bodies, Star Service total management system (2) wherein, power-supply system (16) is positioned at bottom cabin body, and first wireless bridge (12) is positioned at the superiors' cabin body and links to each other with Star Service total management system (2); Star Service total management system (2) receives the information of optical fibre gyro (13), star sensor (10), sun sensor (8) and GPS receiver (5) in the satellite control system and carries out the attitude track real-time simulation calculating of satellite, Star Service total management system (2) is communicated by letter with ground simulation computing machine 18 by first wireless bridge (12), second wireless bridge (17) simultaneously, and according to the instruction of ground simulation computer (18) to DGCMG(15) send order, carry out the attitude control of satellite platform;
Jet propulsion system (4) is positioned at cabin, middle layer body; Four two 100 ° of bottoms that are installed in the bottom cabin body of three-axis air-bearing table (1), framework control-moment gyro (15) space, constitute single frame moment gyro and the counteraction flyback of multiple configuration respectively by locking inside casing or housing, inside casing and housing, be used for realizing the attitude control system of three class variety classes topworkies; Optical fibre gyro (13), star sensor (10), sun sensor (8) and GPS receiver (5) all are positioned at the superiors' cabin body, and above-mentioned parts all link to each other with Star Service total management system (2); GPS simulator (14) is positioned at the superiors' cabin body; Metal slide block (6) is installed on the continuous 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 all link to each other with Star Service total management system (2); Solar simulator (9), star emulator (11), GPS simulator (14) link to each other with sun sensor (8), star sensor (10) and GPS receiver (5) respectively.
2. according to claim 1 based on a kind of DGCMG(15) the multiple appearance control of the satellite pattern test macro of configuration, it is characterized in that: described DGCMG(15) as topworks, when its inside and outside framework angle all is zero, pinning DGCMG(15) outside framework, then constitute the single frame control-moment gyro of two parallel configuration, be used for the single frame control-moment gyro of the two parallel configuration of research as the performance of the satellite posture control system of topworks.
3. according to claim 1 based on a kind of DGCMG(15) the multiple appearance control of the satellite pattern test macro of configuration, it is characterized in that: described DGCMG(15) as topworks, at the 3rd DGCMG(15) turn over the α angle clockwise around housing axle Y3,0＜α＜100, and locking housing, the 4th DGCMG(15) turn over α angle and locking housing between the inverse time around housing axle Y4, first DGCMG(15) turn over α angle and locking inside casing counterclockwise around housing axle Y1, second DGCMG(15) turn over the α angle clockwise and pin the single frame control-moment gyro that inside casing then constitutes the pyramid configuration around housing axle Y2.
4. according to claim 1 based on a kind of DGCMG(15) the multiple appearance control of the satellite pattern test macro of configuration, it is characterized in that: described DGCMG(15) as topworks, the 3rd DGCMG(15) turns over the α angle clockwise around housing axle Y3,0＜α＜100, and locking housing, the 4th DGCMG(15) turns over α angle and locking housing between the inverse time around housing axle Y4, the one DGCMG(15) turn over α angle and locking inside casing, the 2nd DGCMG(15 counterclockwise around housing axle Y1) turn over the α angle clockwise and pin the counteraction flyback that housing and inside casing then constitute the pyramid configuration of pyramid configuration simultaneously around housing axle Y2.
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