CN111077800B - Double-star formation semi-physical test system and method - Google Patents
Double-star formation semi-physical test system and method Download PDFInfo
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Abstract
The invention discloses a double-star formation semi-physical test system and method, which adopt a mode of combining a real satellite and a digital satellite, give consideration to semi-physical simulation test and double-star formation control test, adopt a closed-loop iterative correction mode, greatly improve the semi-physical simulation test precision of double-star formation, reduce the hardware resource investment of double-star formation test, and obviously improve the economic benefit.
Description
Technical Field
The invention relates to the technical field of satellite orbit navigation and formation control testing, in particular to a double-satellite formation semi-physical testing system and method based on a digital orbit model.
Background
The formation flight of the spacecraft is a new spacecraft space operation mode which appears in the late 80 th century and is accompanied with the development of the microsatellite. The flight of the satellite formation has very outstanding advantages compared with a single spacecraft, and is favored by all aerospace big countries in the world since the birth of the concept.
In order to ensure the quality of formation configuration control, detailed function and performance tests are required on the ground. As the satellite formation flight technology in China starts late, no mature and reliable formation flight test scheme is provided for reference in the practice of satellite formation control engineering. In the practice of satellite engineering, due to the importance of product system interfaces, semi-physical simulation tests are an essential satellite engineering development process. If a plurality of distributed semi-physical simulation test systems are adopted, the investment of manpower and equipment in the development process is increased undoubtedly, so that the factors of manpower management, economy, test coverage, test system expansibility and the like are considered comprehensively and balancedly in the satellite formation flight test so as to adapt to the actual requirements of engineering development.
Disclosure of Invention
The invention provides a double-star formation semi-physical test system and method, which are suitable for giving consideration to semi-physical simulation test and formation flight control technology test, can be conveniently expanded into a multi-star distributed semi-physical simulation test system, greatly improve the precision of double-star formation semi-physical simulation test by adopting a closed loop iterative correction mode, are easy to realize, and obviously improve the economic benefit.
In order to achieve the aim, the invention provides a double-satellite formation semi-physical test system which comprises a set of single-satellite attitude and orbit control semi-physical simulation closed loop test system, a set of digital satellite simulation system and a set of control strategy support system;
the single-satellite attitude and orbit semi-physical simulation system comprises semi-physical simulation test equipment and a GNC computer, wherein the semi-physical simulation test equipment is in communication connection with the GNC computer, and the GNC computer is used for state acquisition and control of a real satellite; the semi-physical simulation test equipment is used for providing an excitation source for the star sensor, the earth sensor, the sun sensor, the accelerometer combination and the gyroscope combination according to the orbit and attitude information of a real satellite, acquiring the working states of a reaction flywheel, a moment gyroscope and a thruster through a GNC computer, and applying the acquired working state data to an attitude orbit kinematics and dynamics model of the semi-physical simulation test equipment so as to form a closed-loop single-star attitude orbit semi-physical simulation system; the single-satellite attitude and orbit semi-physical simulation system is used for outputting the absolute orbit position of a real satellite;
the digital satellite simulation system is a set of high-precision orbit dynamics modules, is connected with the semi-physical simulation test equipment and performs information interaction with the semi-physical simulation test equipment, outputs the absolute orbit position of a virtual satellite and calculates the simulation data of the formation flight of the double satellites;
the control strategy support system is respectively connected with the single-satellite attitude and orbit semi-physical simulation system and the digital satellite simulation system and is used for receiving the real satellite absolute orbit position output by the single-satellite attitude and orbit semi-physical simulation system and the virtual satellite absolute orbit position output by the digital satellite simulation system and generating a control strategy of double-satellite formation; and the control strategy support system sends the generated control strategy to the single-satellite attitude and orbit semi-physical simulation system, and corrects the control strategy according to a comparison result of a control result output by the single-satellite attitude and orbit semi-physical simulation system and simulation data output by the digital satellite simulation system for double-satellite formation flying.
The invention also discloses a method for carrying out the double-star formation semi-physical test by utilizing the double-star formation semi-physical test system, which comprises the following steps:
s1, a single-satellite attitude and orbit semi-physical simulation system calculates the absolute orbit position of an auxiliary satellite in a double-satellite formation, and transmits the orbit information of a main satellite in the double-satellite formation to a digital satellite simulation system;
s2, the digital satellite simulation system calculates the absolute orbit position of a main satellite in the double-satellite formation, and obtains the simulation data of the double-satellite formation flight according to the orbit information of the main satellite and an auxiliary satellite in the double-satellite formation;
s3, the control strategy support system calculates relative orbit vectors of the primary satellite and the secondary satellite according to the absolute orbit positions of the primary satellite and the secondary satellite and generates configuration parameters of the double-satellite formation;
s4, the control strategy support system generates a double-star formation control strategy according to configuration parameters of the double-star formation;
and S5, the single-satellite attitude and orbit semi-physical simulation system operates the double-satellite formation control strategy, and the control strategy support system corrects the double-satellite formation control strategy according to the control result of the single-satellite attitude and orbit semi-physical simulation system until the control result of the single-satellite attitude and orbit semi-physical simulation system is consistent with the simulation data of the digital satellite simulation system.
Preferably, in step S3, the relative orbit vector ρ of the primary satellite and the secondary satellite is:
ρ＝r c1 -r c2 ，
namely:
in the formula, r c2 Is the absolute orbital position of the satellite, r c1 Is the absolute orbital position of the dominant star, mu is the gravitational constant, F c1 For control forces acting on the primary satellite,. DELTA.f d Is the relative perturbation acceleration, m c1 Mass of the dominant star;
the relative position and the relative velocity corresponding to the calculation time of the relative orbit vector ρ are respectively:
[D x D y D z ] T and [ D ] vx D vy D vz ] T ；
The double-star formation configuration parameters of the relative orbits of the main star and the auxiliary star are as follows:
θ FF ＝u main flat -arctan2(D vx /n Main flat ,-(D x -Δa p ))
ψ FF ＝u Main flat -arctan2(D z ,D vz /n Main flat )
l＝D y -2D vx /n Main flat
Wherein p represents a two-star formation surfaceInner dimension, s denotes the outer dimension of the two-star formation, θ FF Representing the phase, psi, in dual-star formation FF Expressing the out-of-plane phase of the double-star formation, wherein the in-plane refers to the tangential direction of the double-star formation coordinate system relative to the plane of the orbit, the out-of-plane refers to the normal direction of the double-star formation coordinate system relative to the plane of the orbit, l is the tangential drift distance of the major axis of the orbit of the main star, and a Main flat ,i Main flat ,u Main flat And calculating the absolute average number of the main stars corresponding to the time for the relative orbit vector.
Preferably, the step S4 includes the steps of:
s4.1, calculating a formation air injection control strategy in a double-star plane;
s4.2, calculating a formation air injection control strategy outside the plane of the double stars;
the step S4.1 includes the following steps:
s4.1.1, calculating the vector variation of the relative eccentricity according to the major and minor axes of the main satellite orbit and the configuration parameters of the double-satellite formation:
wherein k is a bias value in degrees (°);
s4.1.2, calculating actual latitude argument control quantity between two satellites in a plane;
Δu c ＝1.5(u 1 -u 0 )Δa/a 1 -δΔu
in the formula u 1 ＝arctan(δΔe y /δΔe x ) Delta. DELTA.u is u 0 Rotate counterclockwise to u 1 Delta u threshold range of [0,2 pi ]]；
S4.1.3, calculating the adjustment quantity of the relative semimajor axis between the two stars in the plane:
Δa c ＝-Δa；
s4.1.4, calculating a formation control strategy in a double-star plane:
in the formula,. DELTA.v i For the ith control speed increment, i =1,2,3, ·,
the 1 st air injection in the double-star plane is at the latitude argument u 1 ＝arctan(δΔe y /δΔe x ) The time of day; the 2 nd jet is at latitude argument u 2 ＝arctan(δΔe y /δΔe x ) A time of + π; the 3 rd time of air injection is carried out at the latitude argument u 3 ＝arctan(δΔe y /δΔe x ) A time of +2 π;
the step S4.2 includes the following steps:
s4.2.1, calculating the vector variation of the relative inclination angle according to the major axis of the main satellite orbit and the configuration parameters of the double-satellite formation:
s4.2.2, calculating out-of-plane theoretical control latitude argument:
u＝arctanδΔi y /δΔi x ；
s4.2.3, calculating a formation control strategy outside the double-star plane:
wherein, the delta v is the control speed increment of formation outside the double-star plane, and the air injection is at the latitude argument
Time of day (c).
Preferably, the roles of the primary and secondary stars in the two-star formation are interchangeable.
The invention has the following advantages:
according to the double-star formation semi-physical test system and method, a mode of combining a real satellite and a digital satellite is adopted, semi-physical simulation test and double-star formation control test are considered, hardware investment is saved, on the premise that double-star formation simulation test precision is met, hardware resource investment of double-star formation test is reduced, implementation is easy, and economic benefit is improved.
Drawings
Fig. 1 is a schematic diagram of a double star formation semi-physical testing system according to an embodiment of the present invention;
fig. 2 is a flowchart of a double star formation semi-physical testing method according to an embodiment of the present invention.
Detailed Description
The following describes a double-star formation semi-physical testing system and method provided by the present invention in further detail with reference to the accompanying drawings and specific embodiments. Advantages and features of the present invention will become apparent from the following description and from the claims. It is to be noted that the drawings are in a very simplified form and are all used in a non-precise ratio for the purpose of facilitating and distinctly aiding in the description of the embodiments of the invention.
As shown in fig. 1, the double-satellite formation semi-physical test system provided by the present invention includes a set of single-satellite attitude and orbit control semi-physical simulation closed-loop test system, a set of digital satellite simulation system, and a set of control strategy support system.
The single-satellite attitude and orbit semi-physical simulation system comprises semi-physical simulation test equipment and a GNC computer (a satellite-borne attitude orbit control computer), wherein the semi-physical simulation test equipment is in communication connection with the GNC computer, and the GNC computer is used for state acquisition and control of a real satellite; the semi-physical simulation test equipment is used for providing excitation sources for the star sensor, the earth sensor, the sun sensor, the accelerometer combination, the gyroscope combination and the like according to the orbit and attitude information of a real satellite, acquiring the working states of a reaction flywheel, a moment gyroscope, a thruster and the like through a GNC computer, and applying the acquired working state data to an attitude orbit kinematics and dynamics model of the semi-physical simulation test equipment so as to form a closed-loop single-star attitude orbit semi-physical simulation system; the single-satellite attitude and orbit semi-physical simulation system is used for outputting the absolute orbit position of a real satellite.
The digital satellite simulation system is a set of high-precision orbit dynamics modules and is provided with a network/serial port and other universal interfaces, the digital satellite simulation system is connected with the semi-physical simulation test equipment through the interfaces, and is used for simulating the orbit change of a virtual satellite in an air injection state, performing orbit information interaction with the semi-physical simulation test equipment, calculating the formation information of double stars and obtaining the simulation data of double-star formation flight; and the digital satellite simulation system outputs the absolute orbit position of the virtual satellite.
The control strategy support system is respectively connected with the single-satellite attitude and orbit semi-physical simulation system and the digital satellite simulation system, and is used for receiving the true satellite absolute orbit position output by the single-satellite attitude and orbit semi-physical simulation system and the virtual satellite absolute orbit position output by the digital satellite simulation system, calculating relative orbit data between two satellites, generating a control strategy of double-satellite formation, sending the generated control strategy to the GNC computer, and correcting the control strategy according to a comparison result of a control result output by the single-satellite attitude and orbit semi-physical simulation system and simulation data output by the digital satellite simulation system, wherein the control strategy is formed by combining the true satellite absolute orbit position output by the single-satellite attitude and orbit semi-physical simulation system and the virtual satellite absolute orbit position output by the digital satellite simulation system.
The single-satellite attitude and orbit semi-physical simulation system can be used for testing a main satellite or an auxiliary satellite in a double-satellite formation, and correspondingly, the digital satellite simulation system is used for simulating the auxiliary satellite or the main satellite in the double-satellite formation. Specifically, in the embodiment, a real satellite accessed to the single-satellite attitude and orbit semi-physical simulation system is used as a secondary satellite, and a virtual satellite accessed to the digital satellite simulation system is used as a primary satellite.
As shown in fig. 2, the method for testing the semi-physical object of the double-star formation provided by the invention comprises the following steps:
s1, a single-satellite attitude and orbit semi-physical simulation system calculates the absolute orbit position of an auxiliary satellite in a double-satellite formation, and transmits the orbit information of a main satellite in the double-satellite formation to a digital satellite simulation system;
establishing a single-satellite attitude and orbit semi-physical simulation system by adopting a general design method, accessing satellites in a double-satellite formation after the test system operates normally, and obtaining the absolute track position r of the satellites by adopting a GNSS receiver c2 And transmitting the orbit information of the main satellite to the digital satellite simulation system.
S2, the digital satellite simulation system calculates the absolute orbit position of a main satellite in the double-satellite formation, and obtains the simulation data of the double-satellite formation flight according to the orbit information of the main satellite and an auxiliary satellite in the double-satellite formation;
a digital satellite simulation system is established by adopting an algebraic method, input information of the simulation system comprises the initial position, the initial speed, the simulated satellite air injection time and the air injection length of a satellite under a selected inertial system, and output data of the simulation system comprises the absolute orbit position r of a main satellite c1 And speed information.
And calculating the absolute position orbit information of the main satellite by adopting an orbit dynamics equation of the satellite. In the selected inertial frame, r is recorded c For an absolute orbital position vector of a satellite, the orbital dynamics equation for the satellite is:
wherein μ is an earth gravity constant; f. of dc Representing all perturbation accelerations acting on the satellite; f c For control forces acting on the satellite, Δ f d Is the relative perturbation acceleration, m c Is the satellite mass.
The digital satellite simulation system adopts a dynamics simulation method, integrates simulation data of the flight of the double-star formation according to the orbit information of the main star and the auxiliary star in the double-star formation by using an orbit dynamics equation, and the simulation data is an ideal state of the flight of the main star and the auxiliary star in the double-star formation.
S3, the control strategy support system calculates relative orbit vectors of the primary satellite and the secondary satellite according to the absolute orbit position information of the primary satellite and the secondary satellite and generates configuration parameters of the double-satellite formation;
the relative orbit vectors ρ of the primary and secondary stars are:
ρ＝r c1 -r c2 ，
namely:
the relative position and the relative velocity corresponding to the calculation time of the relative orbit vector ρ are respectively:
[D x D y D z ] T and [ D vx D vy D vz ] T 。
Obtaining the double-satellite formation configuration parameters of the relative orbits of the main satellite and the auxiliary satellite after the absolute orbit vectors of the main satellite and the auxiliary satellite are differentiated:
θ FF ＝u main flat -arctan2(D vx /n Main flat ,-(D x -Δa p ))
ψ FF ＝u Main flat -arctan2(D z ,D vz /n Main flat )
l＝D y -2D vx /n Main flat
Wherein p represents the in-plane dimension of the two-star formation, s represents the out-of-plane dimension of the two-star formation, wherein the in-plane dimension refers to the tangential direction of the two-star formation relative to the orbital plane under the coordinate system, the out-of-plane dimension refers to the normal direction of the two-star formation relative to the orbital plane under the coordinate system, and theta FF Representing the two-star formation in-plane phase, # FF Expressing the out-of-plane phase of the dual-star formation, i is the tangential drift distance of the major axis of the orbit of the main star, a Main flat ,i Main flat ,u Main flat And calculating the absolute average number of the main stars corresponding to the time for the relative orbit vector.
S4, the control strategy support system generates a double-star formation control strategy according to configuration parameters of the double-star formation;
the step S4 comprises the following steps:
s4.1, calculating a formation air injection control strategy in a double-star plane;
s4.2, calculating a formation air injection control strategy outside the plane of the double stars;
the step S4.1 includes the following steps:
s4.1.1, according to the major star orbit semimajor axis a 1 、u 0 、n 1 And the two-star formation configuration parameters p and theta FF Delta a, l, calculating the vector variation delta e of the relative eccentricity x And delta e y ：
Wherein k is a bias value in degrees (°);
s4.1.2, calculating actual latitude argument control quantity delta u between two satellites in plane c ；
Δu c ＝1.5(u 1 -u 0 )Δa/a 1 -δΔu
In the formula u 1 ＝arctan(δΔe y /δΔe x ) δ Δ u is u 0 Rotate counterclockwise to u 1 Angle of [0,2 π ] in the threshold range]；
S4.1.3, calculating the adjustment quantity delta a of the relative semi-major axis between the two satellites in the plane c ：
Δa c ＝-Δa；
S4.1.4, calculating a formation control strategy in a double-star plane:
in the formula,. DELTA.v i For the ith control speed increment, i =1,2,3.
As can be seen from the above formula, the 1 st air injection in the double star plane is at the latitude argument u 1 ＝arctan(δΔe y /δΔe x ) The time of day; the 2 nd jet is at latitude argument u 2 ＝arctan(δΔe y /δΔe x ) A time of + π; the 3 rd time of air injection is at the latitude argument u 3 ＝arctan(δΔe y /δΔe x ) A time of +2 pi.
The step S4.2 includes the following steps:
s4.2.1, according to the semi-major axis a of the main satellite orbit 1 、u 0 、n 1 And a two-star formation configuration parameter s 1 、ψ FF1 Calculating the vector variation delta i of the relative inclination angle x And delta i y ：
S4.2.2, calculating an out-of-plane theoretical control latitude argument u:
u＝arctan(δΔi y /δΔi x )；
s4.2.3, calculating a formation control strategy outside the double-star plane:
in the formula: Δ v is the dual star out-of-plane formation control velocity increment, and the jet is the latitude argument
The time of day.
And S5, the single-satellite attitude and orbit semi-physical simulation system operates the double-satellite formation control strategy, and the control strategy support system corrects the double-satellite formation control strategy according to the control result of the single-satellite attitude and orbit semi-physical simulation system until the control result of the single-satellite attitude and orbit semi-physical simulation system is consistent with the simulation data of the digital satellite simulation system.
Inputting a double-star formation control strategy generated by a control strategy support system into a GNC computer, and operating the control strategy by a single-star attitude and orbit semi-physical simulation system; the control strategy support system receives a control result actually output by the single-satellite attitude and orbit semi-physical simulation system, and compares the control result with simulation data of the digital satellite simulation system; and the control strategy support system corrects the control strategies inside and outside the plane of the double-star formation according to the comparison result, and controls the out-of-tolerance parameters once or for multiple times until the control result is consistent with the theoretical calculation result or meets the design requirement.
According to the double-star formation semi-physical test system and method, a mode of combining a real satellite and a digital satellite is adopted, semi-physical simulation test and double-star formation control test are considered, hardware investment is saved, and on the premise that double-star formation simulation test precision is met, hardware resource investment of double-star formation test is reduced.
While the present invention has been described in detail with reference to the preferred embodiments, it should be understood that the above description should not be taken as limiting the invention. Various modifications and alterations to this invention will become apparent to those skilled in the art upon reading the foregoing description. Accordingly, the scope of the invention should be limited only by the attached claims.
Claims (4)
1. A double-star formation semi-physical testing method is characterized by comprising the following steps:
s1, a single-satellite attitude and orbit semi-physical simulation system calculates the absolute orbit position of an auxiliary satellite in a double-satellite formation, and transmits the orbit information of a main satellite in the double-satellite formation to a digital satellite simulation system;
s2, the digital satellite simulation system calculates the absolute orbit position of a main satellite in the double-satellite formation, and obtains the simulation data of the double-satellite formation flight according to the orbit information of the main satellite and an auxiliary satellite in the double-satellite formation;
s3, the control strategy support system calculates relative orbit vectors of the primary satellite and the secondary satellite according to the absolute orbit positions of the primary satellite and the secondary satellite and generates configuration parameters of the double-satellite formation;
s4, the control strategy support system generates a double-satellite formation control strategy according to configuration parameters of the double-satellite formation;
s5, the single-satellite attitude and orbit semi-physical simulation system operates a double-satellite formation control strategy, and the control strategy support system corrects the double-satellite formation control strategy according to the control result of the single-satellite attitude and orbit semi-physical simulation system until the control result of the single-satellite attitude and orbit semi-physical simulation system is consistent with the simulation data of the digital satellite simulation system;
in step S3, the relative orbit vector ρ of the primary satellite and the secondary satellite is:
ρ＝r c1 -r c2 ，
namely:
in the formula, r c2 As absolute orbital position of satellite, r c1 Is the absolute orbital position of the main star, mu is the gravitational constant of the earth, F c1 For control forces acting on the primary satellite,. DELTA.f d Is a relative perturbation acceleration m c1 Mass of the dominant star;
the relative position and the relative velocity corresponding to the calculation time of the relative orbit vector ρ are respectively:
[D x D y D z ] T and [ D vx D vy D vz ] T ；
The double-star formation configuration parameters of the relative orbit of the main star and the auxiliary star are as follows:
θ FF ＝u main flat -arctan2(D vx /n Main flat ,-(D x -Δa p ))
ψ FF ＝u Main flat -arctan2(D z ,D vz /n Main flat )
l＝D y -2D vx /n Main flat
Wherein p represents the inner dimension of the double-star formation, s represents the outer dimension of the double-star formation, and theta FF Representing the phase, psi, in dual-star formation FF The phase of the double-star formation is shown, wherein the in-plane refers to the tangential direction of the double-star formation coordinate system relative to the orbital plane, and the out-of-plane refers to the phase of the double-star formationIs the normal direction of the plane of the opposite orbit under the coordinate system of the two-star formation, i is the tangential drift distance of the semi-major axis of the orbit of the main star, a Main flat ,i Main flat ,u Main flat Calculating the absolute average root number of the main star corresponding to the moment for the relative orbit vector;
the step S4 comprises the following steps:
s4.1, calculating a formation air injection control strategy in a double-star plane;
s4.2, calculating a formation air injection control strategy outside the plane of the double stars;
the step S4.1 includes the following steps:
s4.1.1, calculating the vector variation of the relative eccentricity according to the major axis of the main satellite orbit and the configuration parameters of the double-satellite formation:
wherein k is a bias value in degrees (°);
s4.1.2, calculating actual inter-satellite latitude argument control quantity in a plane;
Δu c ＝1.5(u 1 -u 0 )Δa/a 1 -δΔu
in the formula u 1 ＝arctan(δΔe y /δΔe x ) δ Δ u is u 0 Rotate counterclockwise to u 1 Delta u threshold range of [0,2 pi ]]；
S4.1.3, calculating the adjustment quantity of the relative semimajor axis between the two stars in the plane:
Δa c ＝-Δa；
s4.1.4, calculating a formation control strategy in a double-star plane:
in the formula,. DELTA.v i For the ith control speed increment, i =1,2,3, ·,
the 1 st air injection in the double star plane is at the latitude argument u 1 ＝arctan(δΔe y /δΔe x ) Time of day (c); the 2 nd air injection is carried out at an amplitude of u 2 ＝arctan(δΔe y /δΔe x ) A time of + π; the 3 rd time of air injection is at the latitude argument u 3 ＝arctan(δΔe y /δΔe x ) A time of +2 π;
the step S4.2 includes the following steps:
s4.2.1, calculating the vector variation of the relative inclination angle according to the major axis of the main satellite orbit and the configuration parameters of the double-satellite formation:
s4.2.2, calculating out-of-plane theoretical control latitude argument:
u＝arctan(δΔi y /δΔi x )；
s4.2.3, calculating a formation control strategy outside the double-star plane:
in the formula, deltav is the control speed increment of the double-star out-of-plane formation, and the air injection is carried out at the latitude argument u =
The time of day.
2. The semi-physical test method for the two-star formation of claim 1, wherein the roles of the main star and the auxiliary star in the two-star formation can be interchanged.
3. The method for testing the semi-physical objects in the double-star formation according to claim 1, wherein the method for testing the semi-physical objects in the double-star formation is implemented based on a semi-physical object testing system in the double-star formation, and the semi-physical object testing system in the double-star formation comprises: the system comprises a single-satellite attitude and orbit control semi-physical simulation system, a digital satellite simulation system and a control strategy support system;
the single-satellite attitude and orbit semi-physical simulation system is used for outputting the absolute orbit position of a real satellite and operating a control strategy generated by the control strategy support system;
the digital satellite simulation system is connected with the semi-physical simulation test equipment and performs orbit information interaction with the single-satellite attitude and orbit semi-physical simulation system, outputs the absolute orbit position of the virtual satellite and calculates the simulation data of the double-satellite formation flight;
the control strategy support system is respectively connected with the single-satellite attitude and orbit semi-physical simulation system and the digital satellite simulation system and is used for receiving the real satellite absolute orbit position output by the single-satellite attitude and orbit semi-physical simulation system and the virtual satellite absolute orbit position output by the digital satellite simulation system and generating a control strategy of double-satellite formation; and the control strategy support system sends the generated control strategy to the single-satellite attitude and orbit semi-physical simulation system, and corrects the control strategy according to a comparison result of a control result output by the single-satellite attitude and orbit semi-physical simulation system and simulation data output by the digital satellite simulation system for double-satellite formation flying.
4. The method as claimed in claim 3, wherein the single-star formation semi-physical simulation system comprises a semi-physical simulation test device and a GNC computer, the semi-physical simulation test device is connected with the GNC computer in communication, and the GNC computer is used for state acquisition and control of real satellites; the semi-physical simulation test equipment is used for providing an excitation source for the star sensor, the earth sensor, the sun sensor, the accelerometer combination and the gyroscope combination according to the orbit and attitude information of a real satellite, acquiring the working states of a reaction flywheel, a moment gyroscope and a thruster through a GNC computer, and applying the acquired working state data to an attitude orbit kinematics and dynamics model of the semi-physical simulation test equipment to form a closed-loop single-star attitude orbit semi-physical simulation system.
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Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN106767768A (en) * | 2016-12-06 | 2017-05-31 | 东南大学 | A kind of autonomous navigation method of Double Satellite |
| CN107544467A (en) * | 2017-09-22 | 2018-01-05 | 上海卫星工程研究所 | Double Satellite's control closed loop test system and method under the conditions of whole star |
| CN108438255A (en) * | 2018-03-14 | 2018-08-24 | 上海航天控制技术研究所 | Satellite is diversion Formation Configuration initial method under a kind of engineering constraints |
| CN110377047A (en) * | 2019-06-03 | 2019-10-25 | 上海航天控制技术研究所 | A kind of double star satellites formation collision-proof method |
Family Cites Families (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104077490A (en) * | 2014-07-03 | 2014-10-01 | 哈尔滨工业大学 | Aircraft navigation guidance and control ground simulation system performance evaluating method |
| CN105973237B (en) * | 2016-05-04 | 2018-09-07 | 中国人民解放军国防科学技术大学 | Emulation dynamic trajectory based on practical flight data interpolating parses generation method |
| CN110060552B (en) * | 2019-03-26 | 2021-07-20 | 南京航空航天大学 | Satellite constellation orbit physical simulation experiment platform |
-
2019
- 2019-12-12 CN CN201911274105.5A patent/CN111077800B/en active Active
Patent Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN106767768A (en) * | 2016-12-06 | 2017-05-31 | 东南大学 | A kind of autonomous navigation method of Double Satellite |
| CN107544467A (en) * | 2017-09-22 | 2018-01-05 | 上海卫星工程研究所 | Double Satellite's control closed loop test system and method under the conditions of whole star |
| CN108438255A (en) * | 2018-03-14 | 2018-08-24 | 上海航天控制技术研究所 | Satellite is diversion Formation Configuration initial method under a kind of engineering constraints |
| CN110377047A (en) * | 2019-06-03 | 2019-10-25 | 上海航天控制技术研究所 | A kind of double star satellites formation collision-proof method |
Non-Patent Citations (3)
| Title |
|---|
| Hardware-in-the-loop simulations of GPS-based navigation and control for satellite formation flying;Jae-Ik Park 等;《Advances in Space Research》;20101201;全文 * |
| 基于双星编队的空间非合作目标联合定轨方法;刘光明 等;《宇航学报》;20100930;全文 * |
| 无阻力双星串行编队相对位置有限时间控制;张永合 等;《宇航学报》;20150831;全文 * |
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