CN111268173A - Satellite for constructing space telescope array and control method - Google Patents
Satellite for constructing space telescope array and control method Download PDFInfo
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- CN111268173A CN111268173A CN202010105591.4A CN202010105591A CN111268173A CN 111268173 A CN111268173 A CN 111268173A CN 202010105591 A CN202010105591 A CN 202010105591A CN 111268173 A CN111268173 A CN 111268173A
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- 238000005259 measurement Methods 0.000 claims description 18
- 239000000919 ceramic Substances 0.000 claims description 16
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- B64G1/10—Artificial satellites; Systems of such satellites; Interplanetary vehicles
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- B—PERFORMING OPERATIONS; TRANSPORTING
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- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
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- B—PERFORMING OPERATIONS; TRANSPORTING
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Abstract
The invention discloses a satellite for constructing a space telescope array and a control method, and relates to the field of spacecraft control. The satellite includes: the telescope device is arranged on the satellite platform and is connected with the satellite platform through the vibration isolation device. The method is suitable for constructing a space telescope array, realizes centimeter-level control precision of the length of the interference arm and high-precision cooperative control of the orbital attitude of a plurality of satellites, thereby having high-angular-resolution astronomical observation capability and meeting the requirement of high-precision astronomical observation angular resolution of milli-second to micro-second level.
Description
Technical Field
The invention relates to the field of spacecraft control, in particular to a satellite for constructing a space telescope array and a control method.
Background
The current technical field of astronomical observation is that the observation angle resolution of an astronomical telescope is generally improved to obtain better imaging effect. The observation angle resolution of the telescope is determined by the antenna aperture D, i.e. 1.22 λ/D, taking into account the celestial radiation at the observation wavelength λ. In the history of astronomical observation of human beings, a single large-aperture telescope has become an important facility for astronomical observation, however, the aperture of the single telescope cannot be infinitely increased, and therefore, the imaging definition of the astronomical telescope cannot be further improved.
Disclosure of Invention
The invention aims to solve the technical problem of the prior art, and provides a satellite for constructing a space telescope array and a control method thereof.
The technical scheme for solving the technical problems is as follows:
a satellite for constructing a spatial telescope array, comprising: satellite platform, telescope device and vibration isolation device, the telescope device sets up on the satellite platform, through the vibration isolation device with the satellite platform is connected, wherein:
a formation flight control device is arranged in the satellite platform and is used for controlling the movement of the satellite platform, so that the telescope devices and telescope devices on other satellite platforms form a space telescope array with a preset formation;
the vibration isolation device is used for isolating disturbance between the satellite platform and the telescope device;
and an arm length locking control device is arranged in the telescope device and is used for controlling the movement of the telescope device and locking the interference arm length of the space telescope array.
The invention has the beneficial effects that: the satellite provided by the invention is suitable for constructing a space telescope array, the movement of the satellite is controlled by the formation flight control device, the satellite and other satellites form a space telescope array with a preset formation, the centimeter-level control precision of the interference arm length is realized, the disturbance of a satellite platform on an interference light path is isolated by the active vibration isolation control device, the control precision of the interference arm length is further improved to the millimeter level, the interference arm length is locked to the nanometer level by the arm length locking control device, the high-precision cooperative control of the orbital attitudes of a plurality of satellites is realized, and the high-angle-resolution astronomical observation capability is realized, and the high-precision astronomical angle resolution requirement from milli-second to micro-angle-second is met.
Another technical solution of the present invention for solving the above technical problems is as follows:
a satellite control method for constructing a spatial telescope array, the satellite comprising: the telescope device is arranged on the satellite platform and is connected with the satellite platform through the vibration isolation device, and the control method comprises the following steps:
the formation flight control device arranged in the satellite platform controls the movement of the satellite platform, so that the telescope device and telescope devices on other satellite platforms form a space telescope array with a preset formation;
the vibration isolation device isolates disturbance between the satellite platform and the telescope device;
and an arm length locking control device arranged in the telescope device controls the movement of the telescope device to lock the interference arm length of the space telescope array.
The satellite control method provided by the invention is suitable for constructing a space telescope array, the movement of the satellite is controlled by the formation flight control device, the satellite and other satellites form a space telescope array with a preset formation, the centimeter-level control precision of the interference arm length is realized, the disturbance of a satellite platform on an interference light path is isolated by the active vibration isolation control device, the control precision of the interference arm length is further improved to the millimeter level, the interference arm length is locked to the nanometer level by the arm length locking control device, the high-precision cooperative control of the orbital attitudes of a plurality of satellites is realized, and therefore, the astronomical observation capability with high angular resolution is realized, and the requirement of high-precision astronomical observation angular resolution from milli second to micro-angle second is met.
Advantages of additional aspects of the invention will be set forth in part in the description which follows, and in part will be obvious from the description, or may be learned by practice of the invention.
Drawings
FIG. 1 is a schematic diagram of a satellite used for constructing a spatial telescope array according to an embodiment of the present invention;
FIG. 2 is a schematic structural diagram of a formation flight control device provided in another embodiment of a satellite for constructing a spatial telescope array according to the present invention;
FIG. 3 is a schematic structural diagram of a vibration isolation device provided in another embodiment of the satellite for constructing a spatial telescope array according to the present invention;
FIG. 4 is a schematic structural diagram of an arm length locking control device provided in another embodiment of the satellite for constructing a space telescope array according to the present invention;
fig. 5 is a schematic flow chart provided by an embodiment of the satellite control method for constructing the spatial telescope array according to the present invention.
Detailed Description
The principles and features of this invention are described below in conjunction with the following drawings, which are set forth to illustrate, but are not to be construed to limit the scope of the invention.
As shown in fig. 1, a schematic structural diagram is provided for an embodiment of the satellite for constructing a spatial telescope array according to the present invention, and the satellite includes: satellite platform 1, telescope device 2 and vibration isolation mounting 3, telescope device 2 sets up on satellite platform 1, is connected with satellite platform 1 through vibration isolation mounting 3, wherein:
a formation flight control device is arranged in the satellite platform 1 and is used for controlling the movement of the satellite platform 1, so that the telescope device 2 and the telescope devices 2 on the rest satellite platforms 1 form a space telescope array with a preset formation;
the vibration isolation device 3 is used for isolating disturbance between the satellite platform 1 and the telescope device 2;
an arm length locking control device is arranged in the telescope device 2 and used for controlling the movement of the telescope device 2 and locking the interference arm length of the space telescope array.
It should be understood that the whole structure of the satellite is generally divided into two parts, namely a satellite platform 1 and a loading platform, and the satellite platform 1 is provided with sensors and actuators required by attitude and orbit control, and movable parts and flexible parts such as a solar sailboard. The load platform is provided with a sensitive load with high microgravity level and high precision and fast pointing requirement, and the telescope device 2 in the embodiment is arranged on the load platform. The satellite platform 1 can be connected with the telescope device 2 through an umbilical cord and is used for supplying power to the telescope device 2 and transmitting data.
The vibration isolation device 3 can adopt a magnetic suspension active vibration isolation control scheme to actively attenuate the vibration disturbance transmitted from the satellite platform 1 to the telescope device 2, and the vibration isolation device 3 can be provided with a sensor and an actuating mechanism required by relative displacement and attitude control, so that the mounting platform of the telescope device 2 moves along with the satellite platform 1.
In the structure shown in fig. 1, stator electrified coil and float permanent magnet can be installed on satellite platform 1 and the load platform respectively, and the two are arranged relatively, and the middle is an air gap, and the current through electrified coil is controlled, and then the magnetic force size and the direction that the control permanent magnet received to realize the motion control to telescope device 2, make it follow satellite platform 1 motion.
The formation flight control device needs to include a propulsion device, for example, a micro-thruster group, in order to realize a space telescope array forming a preset formation with the telescope devices 2 on the rest satellite platforms 1, and needs to include a measurement device, for example, a laser measurement group, to measure the relative displacement and the relative attitude of each satellite platform 1 compared with the other satellite platforms 1 through laser ranging, and then to realize corresponding motion control according to a preset control program in the controller.
The micro-thruster group comprises a plurality of micro-thrusters, and the micro-thrusters can be arranged on the surface of the satellite platform 1, and the specific positions can be arranged according to actual requirements without limitation.
It should be understood that those skilled in the art may select other propulsion devices and measurement devices capable of implementing the above functions according to actual needs, and detailed description thereof is omitted here.
The arm length locking control device also needs to comprise a propelling device and a measuring device in order to realize the locking of the interference arm length of the space telescope array, wherein the propelling device can select a piezoelectric ceramic actuator group to realize high-precision control, and the measuring device can select a laser interferometer to realize high-precision measurement of displacement variation between the two telescope devices 2.
The piezoelectric ceramic actuator group comprises a plurality of piezoelectric ceramic actuators, and the piezoelectric ceramic actuators can be arranged on the surface of the satellite platform 1, and specific positions can be set according to actual requirements, which is not limited herein.
It should be understood that those skilled in the art may select other propulsion devices and measurement devices capable of implementing the above functions according to actual needs, and detailed description thereof is omitted here.
The satellite that this embodiment provided, be applicable to and found the space telescope array, through the motion of formation flight control device control place satellite, make it form the space telescope array of predetermineeing the formation with other satellites, the centimeter level control accuracy of interference arm length has been realized, and keep apart the disturbance of satellite platform 1 to interfering the light path through initiative vibration isolation controlling means, the arm length control accuracy of will interfering further improves to millimeter magnitude, rethread arm length locking controlling means will interfere the arm length and lock to nanometer magnitude, the high accuracy cooperative control of many satellite orbit gesture has been realized, thereby possess high angular resolution ratio astronomical observation ability, satisfy the high accuracy astronomical angular resolution ratio demand of milli angle second to micro angle second level.
Optionally, in some embodiments, as shown in fig. 2, the formation flight control device comprises: laser survey group 21, first displacement controller 22, first attitude controller 23 and microthrusters 24, wherein:
the laser measurement group 21 is used for measuring a first relative displacement and a first relative attitude between the satellite platform 1 and the rest satellite platforms 1;
the first displacement controller 22 is used for calculating a first control force required to be applied to the satellite platform 1 according to the first relative displacement;
the first attitude controller 23 is configured to calculate a first control moment to be applied to the satellite platform 1 according to the first relative attitude;
the micro-thruster group 24 is used for applying corresponding thrust to the satellite platform 1 according to the first control force and applying corresponding torque to the satellite platform 1 according to the first control torque.
For example, assuming that the relative displacement or attitude between any two satellite platforms 1 in the spatial telescope array changes due to some reason during the operation of the satellite, after the laser measurement group 21 detects the relative displacement and the relative attitude, the first displacement controller 22 calculates the control force to be applied to keep the relative displacement constant, the first attitude controller 23 calculates the control torque to be applied to keep the relative attitude constant, and then sends the control torque to the micro-thruster group 24, and the micro-thruster group 24 applies the control force and the control torque to the satellite platforms 1, thereby keeping the formation configuration constant.
It should be understood that the process of calculating the control force and the control moment according to the relative displacement and the relative attitude by the first displacement controller 22 and the first attitude controller 23 belongs to the prior art, and can be implemented by programming, and will not be described in detail herein.
It should be noted that, in order to measure the relative displacement and the relative attitude angle between any two satellite platforms 1 in the telescope array, multiple sets of laser measurement sensors are required, and corresponding cooperative targets are deployed, and the laser measurement set 21 shown in fig. 2 is only a functional representation thereof, and does not represent the number thereof. By detecting the relative displacement and attitude angle between the satellite platform 1 and the satellite platform 1 through the laser measurement group 21, various layout schemes can be designed, for example, one laser measurement sensor is arranged on the outer surface of the satellite platform 1 at a certain distance, and the layout schemes can be realized by those skilled in the art without creative labor, and are not described herein again.
Similarly, the micro-thruster sets 24 shown in fig. 2 actually work together in multiple sets, and in practical applications, at least 12 sets can apply corresponding force and moment effects to the satellite platform 1.
It should be understood that, in the formation flight control process, the relative displacement reference input and the attitude angle reference input are set to be constant values according to task requirements, so that stable measurement base lines and cooperative pointing are achieved.
In the embodiment, the laser measurement group 21 is used for measuring the relative displacement and the relative attitude angle between any two satellite platforms 1 in the telescope array, the first displacement controller 22 and the first attitude controller 23 are used for calculating the corresponding control force and moment, and the micro-thruster group 24 is used for motion control, so that formation maintenance control between the satellite platforms 1 and the satellite platforms 1 is realized, and the centimeter-level control precision of the length of the interference arm is achieved.
Optionally, in some embodiments, as shown in fig. 3, the vibration isolation device 3 includes: a two-dimensional position sensor group 31, a second displacement controller 32, a second attitude controller 33, and a two-dimensional electromagnetic actuator group 34, wherein:
the two-dimensional position sensor group 31 is used for measuring a second relative displacement and a second relative attitude between the satellite platform 1 and the telescope device 2;
the second displacement controller 32 is configured to calculate a second control force to be applied to the float of the vibration isolating device 3 based on the second relative displacement;
the second attitude controller 33 is configured to calculate a second control torque to be applied to the float of the vibration isolating device 3, based on the second relative attitude;
the two-dimensional electromagnetic actuator group 34 is configured to apply a corresponding electromagnetic force to the float of the vibration isolation device 3 according to the second control force, and apply a corresponding electromagnetic torque to the float of the vibration isolation device 3 according to the second control torque.
The float of the vibration damping device 3 is a mounting platform of the telescope device 2.
For example, in the operation process of the satellite, for some reason, the satellite platform 1 generates a disturbance in the positive direction of the movement direction, at this time, the disturbance is transmitted to the telescope device 2 through an umbilical cord between the satellite platform 1 and the mounting platform of the telescope device 2, at this time, a small displacement is generated between the satellite platform 1 and the mounting platform of the telescope device 2, after the two-dimensional position sensor set 31 detects the relative displacement, a control force required to be applied to counteract the disturbance is calculated through the first displacement controller and then is sent to the two-dimensional electromagnetic actuator set 34, and the two-dimensional electromagnetic actuator set 34 applies the control force to the mounting platform of the telescope device 2, so that the disturbance is counteracted.
It will be appreciated that in order to detect the relative displacement and attitude angles of the satellite platform 1 and the telescopic arrangement 2, at least 3 sets of two-dimensional position sensors are required. Similarly, the two-dimensional electromagnetic actuator group 34 shown in fig. 3 actually works together for a plurality of groups, and applies corresponding electromagnetic force and electromagnetic torque to the mounting platform of the telescope device 2.
It will be appreciated that during active vibration isolation control, both the relative displacement reference input and the attitude angle reference input may be set to zero values to avoid collision of the telescopic device 2 with the satellite platform 1.
In this embodiment, through active vibration isolation control, on the one hand, vibration disturbance transmitted from the satellite platform 1 to the telescope device 2 can be effectively isolated, and on the other hand, millimeter-scale control accuracy of the interference arm length is realized.
Optionally, in some embodiments, as shown in fig. 4, the arm length lock control means comprises: a laser interferometer 41, a third displacement controller 42, a third attitude controller 43, and a piezoelectric ceramic actuator group 44, wherein:
the laser interferometer 41 is used for measuring the relative displacement variation between the telescope device 2 and the telescope devices 2 on the rest satellite platforms 1;
the third displacement controller 42 is used for calculating a third control force required to be applied to the telescope device 2 according to the relative displacement variation;
the third attitude controller 43 is configured to calculate a third control torque to be applied to the telescope device 2 according to the first relative attitude calculated by the laser measurement group 21 and the second relative attitude calculated by the two-dimensional position sensor group 31;
the piezo-ceramic actuator group 44 is configured to apply a corresponding force to the telescope device 2 according to the third control force, and apply a corresponding moment to the telescope device 2 according to the third control moment.
For example, when a spatial telescope array is observed astronomically, for some reason, a small displacement is generated between the telescope devices 2 of any two satellite platforms 1 in the spatial telescope array, the laser interferometer 41 detects a change in the relative displacement, a control force required to be applied to eliminate the change in the relative displacement is calculated by the third displacement controller 42 and is then transmitted to the piezoelectric ceramic actuator group 44, and the piezoelectric ceramic actuator group 44 applies the control force to the telescope devices 2, thereby achieving precise locking of the interference arm length.
It should be understood that the third relative attitude is calculated indirectly from the first relative attitude measured by the laser measurement set and the second relative attitude measured by the two-dimensional position sensor set.
It should be understood that the process of calculating the control force and the control moment according to the relative displacement and the relative attitude by the third displacement controller 42 and the third attitude controller 43 belongs to the prior art, and can be implemented by programming, and will not be described in detail herein.
It should be noted that, in order to achieve precise locking of the interference arm length, multiple sets of piezo ceramic actuators are required to cooperate, and as shown in fig. 4, the piezo ceramic actuator set 44 actually works together to apply corresponding forces and moments to the telescope device 2, and fig. 4 is only a functional representation and does not represent the number thereof.
It should be appreciated that during arm length lock control, both the relative displacement reference input and the attitude angle reference input are set to zero values. And the interference arm length nanometer-level control precision is realized through arm length locking control.
In the present embodiment, the interference arm length can be locked to the order of nanometers by accurately measuring the amount of relative displacement change between the two telescope devices 2 using the laser interferometer 41 and performing high-precision control by the piezoelectric ceramic actuator group 44.
Optionally, in some embodiments, the vibration isolation device 3 further comprises: and when the satellite is in a launching process, the satellite platform 1 and the telescope device 2 are locked through the locking mechanism.
It is understood that some or all of the alternative embodiments described above may be included in some embodiments.
As shown in fig. 5, a schematic flow chart is provided for an embodiment of a satellite control method for constructing a spatial telescope array according to the present invention, where the satellite includes: the control method comprises the following steps of:
s1, controlling the movement of the satellite platform by a formation flight control device arranged in the satellite platform, so that the telescope device and telescope devices on the other satellite platforms form a space telescope array with a preset formation;
s2, isolating disturbance between the satellite platform and the telescope device by the vibration isolation device;
and S3, controlling the movement of the telescope device by an arm length locking control device arranged in the telescope device, and locking the interference arm length of the space telescope array.
The satellite control method provided by the embodiment is suitable for constructing a space telescope array, the movement of a satellite is controlled by a formation flight control device, the satellite and other satellites form a space telescope array with a preset formation, the centimeter-level control precision of the interference arm length is realized, the disturbance of a satellite platform on an interference light path is isolated by an active vibration isolation control device, the millimeter-level control precision of the interference arm length is further improved, the nanometer-level interference arm length is locked by an arm length locking control device, the high-precision cooperative control of the orbital attitudes of a plurality of satellites is realized, and therefore the astronomical observation capability with high angular resolution is achieved, and the requirement of high-precision astronomical observation angular resolution from milli second to micro-angle second is met.
Optionally, in some embodiments, the formation flight control device comprises: laser survey group, first displacement controller, first attitude control ware and microthruster, formation flight control device control satellite platform's motion specifically includes:
the laser measurement group measures a first relative displacement and a first relative attitude between the satellite platform and the rest satellite platforms;
the first displacement controller calculates a first control force required to be applied to the satellite platform according to the first relative displacement;
the first attitude controller calculates a first control moment required to be applied to the satellite platform according to the first relative attitude;
the micro-thruster group applies corresponding thrust to the satellite platform according to the first control force and applies corresponding torque to the satellite platform according to the first control torque.
Optionally, in some embodiments, the vibration isolation device comprises: two-dimensional position sensor group, second displacement controller, second attitude control ware and two-dimensional electromagnetic actuator group, the vibration isolation device keeps apart the disturbance between satellite platform and the telescope device, specifically includes:
the two-dimensional position sensor group measures a second relative displacement and a second relative attitude between the satellite platform and the telescope device;
the second displacement controller calculates a second control force to be applied to the float of the vibration isolation device according to the second relative displacement;
the second attitude controller calculates a second control torque to be applied to the floater of the vibration isolation device according to the second relative attitude;
and the two-dimensional electromagnetic actuator group applies corresponding electromagnetic force to the floater of the vibration isolation device according to the second control force and applies corresponding electromagnetic torque to the floater of the vibration isolation device according to the second control torque.
Optionally, in some embodiments, the arm length locking control means comprises: the laser interferometer, the third displacement controller, the third attitude controller and the piezoelectric ceramic actuator group, the arm length locking control device controls the movement of the telescope device, and the device specifically comprises:
the laser interferometer measures the relative displacement variation between the telescope device and the telescope devices on the rest satellite platforms;
the third displacement controller calculates a third control force required to be applied to the telescope device according to the relative displacement variation;
the third attitude controller calculates a third control torque to be applied to the telescope device according to the first relative attitude and the second relative attitude;
the piezoelectric ceramic actuator group applies corresponding force to the telescope device according to the third control force and applies corresponding torque to the telescope device according to the third control torque.
Optionally, in some embodiments, the vibration isolation device further comprises: the locking mechanism and the satellite control method further comprise the following steps:
when the satellite is in the launching process, the satellite platform and the telescope device are locked by the locking mechanism.
It is understood that some or all of the alternative embodiments described above may be included in some embodiments.
It should be noted that the above embodiments are product embodiments corresponding to the previous method embodiments, and for the description of each optional implementation in the product embodiments, reference may be made to corresponding descriptions in the above method embodiments, and details are not described here again.
The reader should understand that in the description of this specification, reference to the description of the terms "one embodiment," "some embodiments," "an example," "a specific example," or "some examples," etc., means that a particular feature, structure, material, or characteristic described in connection with the embodiment or example is included in at least one embodiment or example of the invention. In this specification, the schematic representations of the terms used above are not necessarily intended to refer to the same embodiment or example. Furthermore, the particular features, structures, materials, or characteristics described may be combined in any suitable manner in any one or more embodiments or examples. Furthermore, various embodiments or examples and features of different embodiments or examples described in this specification can be combined and combined by one skilled in the art without contradiction.
In the several embodiments provided in the present application, it should be understood that the disclosed apparatus and method may be implemented in other ways. For example, the above-described method embodiments are merely illustrative, and for example, the division of steps into only one logical functional division may be implemented in practice in another way, for example, multiple steps may be combined or integrated into another step, or some features may be omitted, or not implemented.
While the invention has been described with reference to specific embodiments, it will be understood by those skilled in the art that various changes in form and details may be made therein without departing from the spirit and scope of the invention as defined by the appended claims. Therefore, the protection scope of the present invention shall be subject to the protection scope of the claims.
Claims (10)
1. A satellite for constructing a spatial telescope array, comprising: satellite platform, telescope device and vibration isolation device, the telescope device sets up on the satellite platform, through the vibration isolation device with the satellite platform is connected, wherein:
a formation flight control device is arranged in the satellite platform and is used for controlling the movement of the satellite platform, so that the telescope devices and telescope devices on other satellite platforms form a space telescope array with a preset formation;
the vibration isolation device is used for isolating disturbance between the satellite platform and the telescope device;
and an arm length locking control device is arranged in the telescope device and is used for controlling the movement of the telescope device and locking the interference arm length of the space telescope array.
2. The satellite for constructing a spatial telescope array according to claim 1, wherein the formation flight control device comprises: laser survey group, first displacement controller, first attitude control ware and microthruster, wherein:
the laser measurement group is used for measuring a first relative displacement and a first relative attitude between the satellite platform and the rest satellite platforms;
the first displacement controller is used for calculating a first control force required to be applied to the satellite platform according to the first relative displacement;
the first attitude controller is used for calculating a first control moment required to be applied to the satellite platform according to the first relative attitude;
the micro-thruster group is used for applying corresponding thrust to the satellite platform according to the first control force and applying corresponding moment to the satellite platform according to the first control moment.
3. The satellite for constructing a spatial telescope array according to claim 2, wherein the vibration isolation means comprises: two-dimensional position sensor group, second displacement controller, second gesture controller and two-dimensional electromagnetic actuator group, wherein:
the two-dimensional position sensor group is used for measuring a second relative displacement and a second relative attitude between the satellite platform and the telescope device;
the second displacement controller is used for calculating a second control force which needs to be applied to the floater of the vibration isolation device according to the second relative displacement;
the second attitude controller is used for calculating a second control torque which needs to be applied to the floater of the vibration isolation device according to the second relative attitude;
and the two-dimensional electromagnetic actuator group is used for applying corresponding electromagnetic force to the floater of the vibration isolation device according to the second control force and applying corresponding electromagnetic torque to the floater of the vibration isolation device according to the second control torque.
4. The satellite for constructing a spatial telescope array according to claim 3, wherein the arm length lock control device comprises: laser interferometer, third displacement controller, third gesture controller and piezoceramics actuator group, wherein:
the laser interferometer is used for measuring the relative displacement variation between the telescope device and the telescope devices on the rest satellite platforms;
the third displacement controller is used for calculating a third control force required to be applied to the telescope device according to the relative displacement variable quantity;
the third attitude controller is used for calculating a third control moment required to be applied to the telescope device according to the first relative attitude and the second relative attitude;
the piezoelectric ceramic actuator group is used for applying corresponding force to the telescope device according to the third control force and applying corresponding torque to the telescope device according to the third control torque.
5. The satellite for constructing a spatial telescope array according to any one of claims 1 to 4, wherein the vibration isolation device further comprises: and when the satellite is in a launching process, the satellite platform and the telescope device are locked through the locking mechanism.
6. A satellite control method for constructing a spatial telescope array, the satellite comprising: the telescope device is arranged on the satellite platform and is connected with the satellite platform through the vibration isolation device, and the control method comprises the following steps:
the formation flight control device arranged in the satellite platform controls the movement of the satellite platform, so that the telescope device and telescope devices on other satellite platforms form a space telescope array with a preset formation;
the vibration isolation device isolates disturbance between the satellite platform and the telescope device;
and an arm length locking control device arranged in the telescope device controls the movement of the telescope device to lock the interference arm length of the space telescope array.
7. The satellite control method for constructing a spatial telescope array according to claim 6, wherein the formation flight control device comprises: laser survey group, first displacement controller, first attitude control ware and microthruster, formation flight control device control satellite platform's motion specifically includes:
the laser measurement group measures a first relative displacement and a first relative attitude between the satellite platform and the rest satellite platforms;
the first displacement controller calculates a first control force required to be applied to the satellite platform according to the first relative displacement;
the first attitude controller calculates a first control moment required to be applied to the satellite platform according to the first relative attitude;
the micro-thruster group applies corresponding thrust to the satellite platform according to the first control force and applies corresponding torque to the satellite platform according to the first control torque.
8. The satellite control method for constructing a spatial telescope array according to claim 7, wherein the vibration isolation device comprises: two-dimensional position sensor group, second displacement controller, second attitude control ware and two-dimensional electromagnetic actuator group, the vibration isolation mounting keeps apart satellite platform with the disturbance between the telescope device specifically includes:
the two-dimensional position sensor group measures a second relative displacement and a second relative attitude between the satellite platform and the telescope device;
the second displacement controller calculates a second control force required to be applied to the float of the vibration isolation device according to the second relative displacement;
the second attitude controller calculates a second control torque to be applied to the floater of the vibration isolation device according to the second relative attitude;
and the two-dimensional electromagnetic actuator group applies corresponding electromagnetic force to the floater of the vibration isolation device according to the second control force and applies corresponding electromagnetic torque to the floater of the vibration isolation device according to the second control torque.
9. The satellite control method for constructing a spatial telescope array according to claim 8, wherein the arm length lock control means comprises: the laser interferometer, the third displacement controller, the third attitude controller and the piezoelectric ceramic actuator group, the arm length locking control device controls the movement of the telescope device, and the device specifically comprises:
the laser interferometer measures the relative displacement variation between the telescope device and the telescope devices on the rest satellite platforms;
the third displacement controller calculates a third control force required to be applied to the telescope device according to the relative displacement variation;
the third attitude controller calculates a third control torque to be applied to the telescope device according to the first relative attitude and the second relative attitude;
and the piezoelectric ceramic actuator group applies corresponding force to the telescope device according to the third control force and applies corresponding torque to the telescope device according to the third control torque.
10. The satellite control method for constructing a spatial telescope array according to claims 6 to 9, wherein the vibration isolation device further comprises: the locking mechanism, the satellite control method still includes:
when the satellite is in the launching process, the satellite platform and the telescope device are locked by the locking mechanism.
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