CN114814968B - Space gravitational wave detection device based on single inspection quality - Google Patents

Abstract

The present invention discloses a space gravitational wave detection device based on a single test mass, comprising three spacecrafts, wherein the spacecraft platform in each spacecraft comprises an inertial sensing component, a driving component and two optical components, wherein the inertial sensing component comprises a polygonal columnar test mass, wherein the test mass comprises two sets of sides with a normal direction angle of 60° and four sides which are perpendicular or parallel to each other for detecting displacement sensing and electrostatic feedback control; the optical component comprises a local interferometric measurement unit and an intersatellite long-arm interferometric measurement unit, wherein the local interferometric measurement unit measures the relative displacement change of the spacecraft platform relative to the test mass, and controls the driving component therein to drive the spacecraft platform to follow the test mass movement in the orbital plane through a drag-free control algorithm; and the intersatellite long-arm interferometer measures the relative displacement change between the optical platforms in the two spacecrafts, thereby realizing the measurement of gravitational wave signals. The present invention can effectively simplify the drag-free control strategy of traditional space gravitational wave detection.

Description

A space gravitational wave detection device based on single test mass

Technical Field

The present invention belongs to the technical field of space gravitational wave detection, and more specifically, relates to a space gravitational wave detection device based on a single test mass.

Background technique

Gravitational wave detection is of great significance to fields such as general relativity, astrophysics and cosmology. Space gravitational wave detection refers to the method of using spacecraft formations or constellations to build large laser interferometers in space to detect gravitational waves. The spacecraft used for space gravitational wave detection is mainly composed of inertial sensors, optical measurement systems and spacecraft platforms. The basic principle of space gravitational wave detection is to use drag-free control technology to ensure that the test mass follows the geodesic motion in the direction of the laser interferometer, and to use laser interferometers to measure the optical path change between the two test masses in space caused by gravitational waves.

At present, the international space gravitational wave detection spacecraft formation is mainly an equilateral triangle configuration, such as LISA (Laser Interferometer Space Antenna), Tianqin Project and Taiji Project. In these plans, the spacecraft formation consists of three spacecraft, each of which has two test masses, and each test mass provides a surface as the end mirror of the laser interferometer. Drag-free control requires that the two test masses move along the geodesic line in the sensitive axis direction (interferometer measurement direction) with an angle of 60°. When a test mass has a relative displacement relative to the spacecraft in the sensitive axis direction, a micro-newton thruster is required to push the spacecraft in the non-sensitive axis direction of the other test mass to compensate for this relative displacement. At the same time, an electrostatic actuator is required in the non-sensitive axis direction to pull the two test masses back to the center position. Therefore, the drag-free control strategy of the traditional dual-test mass scheme is very complicated, and it requires the cooperation of micro-newton thrusters and electrostatic feedback actuators to achieve drag-free control.

Summary of the invention

To achieve the above-mentioned object, the present invention provides a space gravitational wave detection device based on a single test mass, comprising a main control system and a spacecraft formation in an equilateral triangle configuration, wherein the spacecraft formation comprises three spacecrafts with completely identical configurations and different working modes, and the three spacecrafts are divided into a master spacecraft and two slave spacecrafts according to the working modes, and the spacecraft platform in each spacecraft is equipped with an inertial sensor component, a control component, a drive component and two optical components, wherein:

The inertial sensor component includes a polygonal columnar test mass, the test mass includes two sets of first side surfaces with a normal direction angle of 60° and four second side surfaces that are perpendicular or parallel to each other, and the four second side surfaces are surrounded by plates for detecting displacement sensing and electrostatic feedback control;

The optical components include a laser, a local interferometry unit, an intersatellite long-arm interferometry unit and an optical phase-locking unit. In the same spacecraft, the laser is used to emit laser light; the local interferometry unit is used to emit local laser light to one of the first side faces of the inspection mass and the interferometry unit using the laser light, and obtain first optical path information of the spacecraft platform of the spacecraft relative to the inspection mass therein by using the interference measurement generated by the local laser light reflected back from the first side face and the emitted local laser light; the control component in each spacecraft is used to control the driving component therein to drive the spacecraft platform of the spacecraft to follow the movement of the inspection mass therein in the orbital plane according to the first optical path information measured by the local interferometry unit therein through a drag-free algorithm;

The two intersatellite long-arm interferometry units in each spacecraft are used to respectively transmit intersatellite long-arm lasers to one intersatellite long-arm interferometry unit in two adjacent spacecrafts; the optical phase-locking unit in the slave spacecraft is used to phase-lock the intersatellite long-arm laser it transmits with the phase of the intersatellite long-arm laser transmitted by the master spacecraft, and transmit the phase-locked intersatellite long-arm laser back to the intersatellite long-arm interferometry unit in the master spacecraft through the intersatellite long-arm interferometry unit in the spacecraft, and obtain the second optical path information of the spacecraft platform in the master spacecraft and the spacecraft platforms in the two slave spacecrafts through the interference measurement generated by the local laser emitted by the local interferometry unit in the master spacecraft;

The main control system is used to obtain a space gravitational wave signal through data processing on the ground according to the second optical path information and the first optical path information measured by the local interferometric measurement unit in each spacecraft.

Compared with the traditional gravitational wave detection device using dual test masses, the space gravitational wave detection device based on a single test mass provided by the present invention is equipped with a single easy-to-process polygonal cylindrical test mass in each spacecraft, and the polygonal cylindrical test mass has two sets of side faces with a normal direction angle of 60°, which can simplify the complex drag-free control strategy in the traditional orbital plane into a spacecraft platform that completely follows the movement of the test mass through driving components, and no longer requires the cooperation of electrostatic feedback actuators, which can effectively simplify the traditional drag-free control strategy for space gravitational wave detection; at the same time, the polygonal cylindrical test mass also has four side faces that are perpendicular or parallel to each other, and the pole plates are arranged around the four side faces. Compared with the spherical test mass, the accuracy of the pole plate displacement measurement can be effectively improved; and compared with the traditional use of dual test masses, the present embodiment uses a single test mass, which can also effectively reduce the overall mass of the spacecraft.

In one embodiment, the driving component adopts a micro-newton-level propeller.

In one of the embodiments, the local interferometry unit includes a first beam splitter, a second beam splitter, a third beam splitter, a fourth beam splitter, a fifth beam splitter, a first fast deflection mirror, a first reflector, a first four-quadrant photodetector, a first acousto-optic modulator, and a second acousto-optic modulator;

Among them, in the same spacecraft, the laser emitted by the laser is divided into two laser beams C by the first beam splitter, one laser beam C is divided into two laser beams by the second beam splitter, and the two laser beams are frequency-shifted by the first acousto-optic modulator and the second acousto-optic modulator respectively; the laser frequency-shifted by the first acousto-optic modulator is emitted to the first side surface of the inspection mass through the first reflecting mirror and the first fast deflection mirror in sequence, and the laser reflected by the first side surface of the inspection mass is irradiated on the first four-quadrant photodetector after passing through the first fast deflection mirror and the third beam splitter in sequence; the laser frequency-shifted by the second acousto-optic modulator is irradiated on the first four-quadrant photodetector after passing through the fifth beam splitter and the fourth beam splitter.

In one of the embodiments, the intersatellite long-arm interferometry unit includes a satellite-borne telescope and an intersatellite long-arm interferometer, and the intersatellite long-arm interferometer includes a sixth beam splitter, a seventh beam splitter, an eighth beam splitter, a second reflector, and a second four-quadrant photoelectric detector;

Among them, in an optical platform in a spacecraft, the laser frequency-shifted by the second acousto-optic modulator is divided into two laser beams D after passing through the sixth beam splitter, one of the laser beams D is emitted through the satellite telescope to emit intersatellite long-arm laser, and the other laser beam D is formed into interference laser after passing through the eighth beam splitter, and the interference laser is irradiated on the second four-quadrant photodetector; the intersatellite long-arm laser B emitted by the spacecraft adjacent to the spacecraft passes through the satellite telescope, the seventh beam splitter, the second reflector, and the eighth beam splitter in sequence, and then is irradiated on the second four-quadrant photodetector to form interference measurement with the interference laser to obtain the second optical path information of the spacecraft platform in the master spacecraft and the spacecraft platforms in the two slave spacecraft.

In one embodiment, the spaceborne telescope includes a spaceborne telescope primary mirror and a second fast deflection reflector.

In one embodiment, the optical phase-locking unit includes a ninth beam splitter, a tenth beam splitter, a third four-quadrant photodetector, a first fiber coupler, a second fiber coupler, a phase meter, and an optical phase-locked loop;

Wherein, in an optical platform in the main spacecraft, another beam of the laser C is split into two beams of laser E by the ninth beam splitter, one beam of laser E is emitted to another optical platform of the main spacecraft through the first optical fiber coupler, and the other beam of laser E interferes with the laser emitted by the laser in another optical platform of the main spacecraft through the second optical fiber coupler on the third four-quadrant photodetector, and the detection signal of the third four-quadrant photodetector is passed through the phase meter to obtain the phase difference of the lasers emitted by the two lasers in the main spacecraft; the optical phase-locked loop is used to adjust the phase of the laser emitted by a laser in the main spacecraft according to the phase difference information, so that the phases of the lasers emitted by the two lasers in the main spacecraft remain consistent;

In the slave spacecraft, the second four-quadrant photoelectric detector is used to measure the second optical path information of the spacecraft platform in the master spacecraft and the spacecraft platforms in the two slave spacecrafts, and the phase meter is used to calculate the phase difference between the intersatellite long-arm laser received by the onboard telescope in the slave spacecraft and the intersatellite long-arm laser emitted by its intersatellite long-arm interferometry measurement unit; the optical phase-locked loop adjusts the laser phase emitted by the laser in the slave spacecraft according to the phase difference information to make it consistent with the phase of the intersatellite long-arm laser received by the onboard telescope.

In one embodiment, the proof mass is in the shape of a right octagonal prism.

In one embodiment, the inspection mass is in the shape of a cuboid with four arc-shaped corners.

In one embodiment, the inspection mass is in the shape of a cuboid with specific structures cut off at both ends of a pair of side surfaces, and the specific structure is a cuboid with a triangular prism cut off at a corner.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG1 is a structural block diagram of a space gravitational wave detection device based on a single test mass provided by an embodiment of the present invention;

FIG2 is a schematic diagram of the structure of a spacecraft formation provided by an embodiment of the present invention;

FIG3 is a schematic diagram of the structure of an optical component provided by an embodiment of the present invention;

FIG4 is a schematic diagram of a structure of a quality inspection system according to an embodiment of the present invention;

5 is a schematic diagram of a structure of a quality inspection provided by another embodiment of the present invention;

FIG6 is a schematic structural diagram of a quality inspection method provided by another embodiment of the present invention.

Detailed ways

In order to make the purpose, technical solution and advantages of the present invention more clearly understood, the present invention is further described in detail below in conjunction with the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present invention and are not intended to limit the present invention.

It should be noted that the basic principle of space gravitational wave detection in this field is to use the fact that gravitational waves will cause the optical path between two test masses moving along geodesics to change, and to extract gravitational wave signals by performing data processing on the optical path change.

As shown in Figures 1 and 2, the space gravitational wave detection device based on a single test mass provided by the present invention includes a main control system and a spacecraft formation in an equilateral triangle configuration, the spacecraft formation includes three spacecrafts with exactly the same configuration and different working modes, and can be specifically divided into a master spacecraft and two slave spacecrafts according to the working mode. The spacecraft platform 50 in each spacecraft is equipped with a control component, an inertial sensor component, a drive component and two optical components.

In this embodiment, the inertial sensing component includes a proof mass 10, which is a component used to provide an inertial reference moving along the ground for gravitational wave detection. The existence of space gravitational waves will cause the relative distance between the two proof masses in each spacecraft to change. According to the basic principles of space gravitational wave detection in this field, the space gravitational wave detection provided by the present invention is correspondingly achieved by respectively measuring the optical path change between the proof mass 10 in the master spacecraft and the proof masses 10 in the two slave spacecraft.

It should be noted that, when measuring space gravitational waves, in order to ensure that the test mass 10 moves along the geodesic, it is necessary to ensure that the spacecraft platform 50 follows the movement of the test mass 10. Therefore, when using the test mass 10 in each spacecraft to measure space gravitational waves, it is also necessary to maintain the relative displacement between the spacecraft platform 50 of each spacecraft and the test mass 10 therein. When the spacecraft platform of a certain spacecraft and the test mass 10 therein change in displacement, it is necessary to adopt a suitable control strategy to control the spacecraft platform to follow the movement of the test mass 10 therein. Specifically, the motion state of the test mass 10 includes the motion state of 3 translational degrees of freedom (2 translational degrees of freedom in the orbital plane and 1 translational degree of freedom perpendicular to the orbital plane) and 3 rotational degrees of freedom.

In combination with the basic principle of space gravitational wave detection in the art, the structures of the inspection mass 10 and the optical components in each spacecraft provided by the present invention are as follows:

The inspection mass 10 provided in this embodiment adopts a polygonal columnar structure, which includes four side surfaces 10b that are perpendicular or parallel to each other. A pole plate 11 for detecting displacement sensing and electrostatic feedback control is arranged around the four side surfaces 10b. The pole plate 11 is used to realize the displacement sensing measurement of the inspection mass 10 in one translational degree of freedom and three rotational degrees of freedom perpendicular to the orbital plane using the pole plate 11, and the electrostatic feedback control is controlled by the control component 60 to make the inspection mass 10 move according to the spacecraft platform on which it is located. In this embodiment, the pole plates 11 are arranged on four side surfaces that are perpendicular or parallel to each other in the polygonal columnar inspection mass 10. Compared with the spherical inspection mass, the linearity of the pole plate displacement detection can be effectively improved, and the cross-coupling coefficient can be reduced to ensure the measurement accuracy.

For the displacement detection of the two translational degrees of freedom of the inspection mass 10 in the orbital plane, this embodiment uses another set of side surfaces of the inspection mass 10 in conjunction with the optical platform to achieve it. Specifically, the inspection mass provided in this embodiment also includes two sets of side surfaces 10a (only one of the side surfaces is marked in the figure) with a normal direction angle of 60°, and one set of side surfaces 10a is used to act as an end mirror of the local laser A emitted in the front-end optical component.

Specifically, the optical components provided in this embodiment include a laser 51, a local interferometer measurement unit 20, an intersatellite long-arm interferometer measurement unit 21, and an optical phase-locking unit. In the same spacecraft, the laser 51 is used to emit laser light; the local interferometer measurement unit 20 is used to use the laser light to emit local laser light A to one of the side surfaces 10a of the inspection mass 10 and the interferometer measurement unit 21, respectively, and to obtain the first optical path information of the spacecraft platform 50 of the spacecraft relative to the inspection mass 10 therein by using the interference measurement generated by the local laser light reflected back from the side surface 10a and the emitted local laser light.

The control component 60 in each spacecraft is used to obtain the first optical path information according to the local interferometric measurement unit 20 therein, and calculate the relative displacement change of the spacecraft platform 50 of the spacecraft relative to the inspection mass 10 therein through the optical path-displacement conversion formula, and then the control component 60 controls the driving component therein to drive the spacecraft platform 50 of the spacecraft to follow the inspection mass 10 therein in the orbital plane according to the drag-free control algorithm, without the need to cooperate with the electrostatic feedback actuator, which can effectively simplify the drag-free control strategy of the traditional gravitational wave detection device in the orbital plane. Specifically, the driving component can use a micro-newton-level thruster.

The two intersatellite long-arm interferometry units 21 in each spacecraft provided in this embodiment are used to respectively transmit the intersatellite long-arm laser B to one intersatellite long-arm interferometry unit 21 in two adjacent spacecrafts. The optical phase-locking unit in the slave spacecraft is used to phase-lock the intersatellite long-arm laser it transmits with the phase of the intersatellite long-arm laser it receives from the master spacecraft, and transmit the phase-locked intersatellite long-arm laser back to the intersatellite long-arm interferometry unit in the master spacecraft through the intersatellite long-arm interferometry unit in the spacecraft, and obtain the second optical path information of the spacecraft platform 50 in the master spacecraft and the spacecraft platforms 50 in the two slave spacecrafts through the interference measurement generated by the local laser emitted by the local interferometry unit in the master spacecraft.

On the ground, the main control system measures the intersatellite long-arm interferometry measurement unit 21 in the master spacecraft to obtain the second optical path information of the spacecraft platform 50 in the master spacecraft and the spacecraft platforms 50 in the two slave spacecrafts, and calculates the relative displacement change of the spacecraft platform 50 of the master spacecraft and the spacecraft platform 50 of the slave spacecraft through the optical path-displacement conversion formula, and then combines the two sections of relative displacement changes of the spacecraft platform 50 of each spacecraft relative to the inspection mass 10 therein obtained by the local interferometry measurement units 20 of the master spacecraft and the slave spacecraft, and adds these three sections of measurement data to extract the gravitational wave signal through data processing. Specifically, the data processing method can adopt the processing method commonly used in the field, such as the matched filtering processing method, which is not limited in this embodiment.

Compared with the traditional gravitational wave detection device using dual test masses, the space gravitational wave detection device based on a single test mass provided in the present embodiment is equipped with a single easily processed polygonal cylindrical test mass 10 in each spacecraft, and the polygonal cylindrical test mass 10 has two sets of side faces 10a with a normal direction angle of 60°, which can simplify the complex drag-free control strategy in the traditional orbital plane to the spacecraft platform completely following the movement of the test mass through driving components, and no longer requires the cooperation of electrostatic feedback actuators, which can effectively simplify the traditional drag-free control strategy for space gravitational wave detection; at the same time, the polygonal cylindrical test mass 10 also has four side faces 10b that are perpendicular or parallel to each other, and the pole plates 11 are arranged around the four side faces, which can effectively improve the accuracy of the pole plate displacement measurement; and compared with the traditional use of dual test masses, the present embodiment uses a single test mass, which can also effectively reduce the overall mass of the spacecraft.

In one embodiment, as shown in Figure 3, the local interference measurement unit 20 includes a laser 51, a first beam splitter 531, a second beam splitter 532, a third beam splitter 533, a fourth beam splitter 534, a fifth beam splitter 535, a first fast deflection mirror 55, a first reflector 542, a first four-quadrant photodetector 522, a first acousto-optic modulator 561 and a first acousto-optic modulator 562.

The working principle of the local interferometric measurement unit 20 provided in this embodiment is as follows: in an optical platform 20 in a spacecraft, the laser emitted by the laser 51 is divided into two laser beams C by the first beam splitter 531, one of which is divided into two beams by the second beam splitter 532, and the two beams are frequency-shifted by the first acousto-optic modulator 561 and the second acousto-optic modulator 562, respectively. Among them, the laser frequency-shifted by the first acousto-optic modulator 561 is the measurement laser, and the laser frequency-shifted by the second acousto-optic modulator 562 is the reference laser. The measurement laser passes through the first reflector 542 and the first fast deflection mirror 55 and is emitted to the inspection mass 10. The laser reflected by a group of side surfaces 10a of the inspection mass 10 passes through the first fast deflection mirror 55 and the third beam splitter 533 and is irradiated on the first four-quadrant photodetector 522. The reference laser passes through the fifth beam splitter 535 and the fourth beam splitter 534 and is irradiated on the first four-quadrant photodetector 522. Finally, the measuring laser and the reference laser form interference at the first four-quadrant photodetector 522, and the first optical path information of the spacecraft platform relative to the inspection mass therein is measured.

In one embodiment, as shown in FIG3 , the intersatellite long-arm interferometry unit 21 includes a satellite telescope and an intersatellite long-arm interferometer, wherein the satellite telescope includes a second fast deflection mirror 42 and a primary mirror 41. The intersatellite long-arm interferometer includes a sixth beam splitter 536, a seventh beam splitter 537, an eighth beam splitter 538, a second reflector 543, and a second four-quadrant photodetector 521.

The working principle of the intersatellite long-arm interferometric measurement unit 21 provided in this embodiment is as follows: in an optical platform 20 in a spacecraft, after the reference laser (the laser frequency-shifted by the second acousto-optic modulator 562) is split into two laser beams D by the sixth beam splitter 536, one of the laser beams D emits the intersatellite long-arm laser B through the onboard telescope (the second fast deflection mirror 42 and the primary mirror 41), and the other laser beam D forms an interference laser through the eighth beam splitter 538 and irradiates the second four-quadrant photodetector 521. The intersatellite long-arm laser B emitted by the spacecraft adjacent to the spacecraft is converged by the onboard telescope primary mirror 41 in an optical platform 20 in the spacecraft, and then sequentially passes through the second fast deflection mirror 42, the seventh beam splitter 537, the second reflector 543, and the eighth beam splitter 538, and then irradiates the second four-quadrant photodetector 521 to form interference with the aforementioned interference laser, and the second optical path information of the spacecraft platform of the spacecraft relative to the spacecraft platform of the remote adjacent spacecraft is obtained by combining with the optical phase-locked unit measurement.

In this embodiment, the onboard telescope provided in the present invention is fixed on the spacecraft platform of each spacecraft. Since the change of the spacecraft formation configuration is easy to generate a breathing angle, in order to achieve compensation for the breathing angle, this embodiment can be roughly adjusted by rotating the onboard telescope, and then fine-tuned by the second fast deflection mirror 42 inside the onboard telescope or the first fast deflection mirror 55 in the local interferometric measurement unit 20. When the intersatellite long arm laser B deviates from the onboard telescope, the change can be measured and monitored by the second four-quadrant photodetector 521, and compensated by rotating the onboard telescope and the second fast deflection mirror 42 inside it. The optical platform is fixedly connected to the onboard telescope. When the onboard telescope rotates, the local laser A will deviate from the normal direction of the inspection mass 10. The change is measured and detected by the first four-quadrant photodetector 522, and compensated by the first fast deflection mirror 55 in the local interferometric measurement unit 20.

In one embodiment, as shown in FIG. 3 , the optical phase-locking unit includes a ninth beam splitter 539 , a tenth beam splitter 530 , a third four-quadrant photodetector 523 , a first fiber coupler 571 , a second fiber coupler 572 , a phase meter 58 and an optical phase-locked loop 59 .

The working principle of the optical phase-locking unit provided in this embodiment is as follows: in an optical platform 20 of the main spacecraft, the laser emitted by the laser 51 is divided into two laser beams C by the first beam splitter 531, and the other laser beam C is divided into two laser beams E by the ninth beam splitter 539, and one laser beam E is emitted to another optical platform 20 of the main spacecraft through the first fiber coupler 571, and the other laser beam E interferes with the laser emitted by the laser in another optical platform of the main spacecraft through the second fiber coupler 572 on the third four-quadrant photodetector 523. The detection signal of the third four-quadrant photodetector 523 is passed through the phase meter 58 to obtain the phase difference of the two laser beams. The optical phase-locked loop 59 adjusts the phase of the laser emitted by the laser in one of the optical platforms according to the phase difference information, so that the phases of the lasers emitted by the two lasers in the main spacecraft remain consistent.

In the slave spacecraft, the optical path information of the spacecraft platform of the slave spacecraft relative to the spacecraft platform of the remote master spacecraft is measured by the second four-quadrant photodetector 521, and the phase difference between the intersatellite long-arm laser received by the onboard telescope in the slave spacecraft and the intersatellite long-arm laser emitted by its intersatellite long-arm interferometry unit is calculated by the phase meter 58. According to the phase difference information, the optical phase-locked loop 59 adjusts the laser phase emitted by the laser 51 in the slave spacecraft to keep it consistent with the intersatellite long-arm laser phase received by the onboard telescope.

In one embodiment, the polygonal columnar inspection mass provided by the present invention can be in the shape of a right octagonal prism, as shown in Figure 4. It can also be in the shape of a cuboid with four arc-shaped corners, as shown in Figure 5, wherein the four arc-shaped sides can act as reflectors for the local laser, and polar plates can be arranged around the remaining four sides.

Of course, a rectangular parallelepiped with a pair of side ends cut off with specific structures can also be used, and the specific structure is a rectangular parallelepiped with a triangular prism cut off at a corner, as shown in Figure 6. Specifically, the actual shape of the polygonal columnar inspection mass 10 provided by the present invention can be designed accordingly according to the actual space gravitational wave detection requirements, and it only needs to satisfy that the polygonal columnar inspection mass 10 has two sets of side surfaces with a normal direction angle of 120° and four side surfaces that are perpendicular or parallel to each other, and this embodiment is not limited.

In one embodiment, the present invention also adjusts the center of mass of the spacecraft platform 50 to coincide with the center of mass of the inspection mass 10 through the center of mass adjustment technology, such as the least squares estimation of the center of mass position of the electrostatic suspension accelerometer on orbit, to reduce the influence of inertial force caused by the misalignment of the center of mass, thereby effectively improving the accuracy of space gravitational wave detection.

It will be easily understood by those skilled in the art that the above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions and improvements made within the spirit and principles of the present invention should be included in the protection scope of the present invention.

Claims (9)

1. A space gravitational wave detection device based on a single test mass, characterized in that it includes a main control system and a spacecraft formation in an equilateral triangle configuration, wherein the spacecraft formation includes three spacecrafts with completely identical configurations and different working modes, and the three spacecrafts are divided into a master spacecraft and two slave spacecrafts according to the working mode, and the spacecraft platform in each spacecraft is equipped with an inertial sensor component, a control component, a drive component and two optical components, wherein: The inertial sensor component includes a polygonal columnar test mass, the test mass includes two sets of first side surfaces with a normal direction angle of 60° and four second side surfaces that are perpendicular or parallel to each other, and the four second side surfaces are surrounded by plates for detecting displacement sensing and electrostatic feedback control; The optical components include a laser, a local interferometry unit, an intersatellite long-arm interferometry unit and an optical phase-locking unit. In the same spacecraft, the laser is used to emit laser light; the local interferometry unit is used to emit local laser light to one of the first side faces of the inspection mass and the interferometry unit using the laser light, and obtain first optical path information of the spacecraft platform of the spacecraft relative to the inspection mass therein by using the interference measurement generated by the local laser light reflected back from the first side face and the emitted local laser light; the control component in each spacecraft is used to control the driving component therein to drive the spacecraft platform of the spacecraft to follow the movement of the inspection mass therein in the orbital plane according to the first optical path information measured by the local interferometry unit therein through a drag-free algorithm; The two intersatellite long-arm interferometry units in each spacecraft are used to respectively transmit intersatellite long-arm lasers to one intersatellite long-arm interferometry unit in two adjacent spacecrafts; the optical phase-locking unit in the slave spacecraft is used to phase-lock the intersatellite long-arm laser it transmits with the phase of the intersatellite long-arm laser transmitted by the master spacecraft, and transmit the phase-locked intersatellite long-arm laser back to the intersatellite long-arm interferometry unit in the master spacecraft through the intersatellite long-arm interferometry unit in the spacecraft, and obtain the second optical path information of the spacecraft platform in the master spacecraft and the spacecraft platforms in the two slave spacecrafts through the interference measurement generated by the local laser emitted by the local interferometry unit in the master spacecraft; The main control system is used to obtain a space gravitational wave signal through data processing on the ground according to the second optical path information and the first optical path information measured by the local interferometric measurement unit in each spacecraft. 2. According to the space gravitational wave detection device based on single test mass according to claim 1, it is characterized in that the driving component adopts a micronewton-level thruster. 3. The space gravitational wave detection device based on a single test mass according to claim 1, characterized in that the local interferometry unit comprises a first beam splitter, a second beam splitter, a third beam splitter, a fourth beam splitter, a fifth beam splitter, a first fast deflection mirror, a first reflecting mirror, a first four-quadrant photodetector, a first acousto-optic modulator and a second acousto-optic modulator; Among them, in the same spacecraft, the laser emitted by the laser is divided into two laser beams C by the first beam splitter, one laser beam C is divided into two laser beams by the second beam splitter, and the two laser beams are frequency-shifted by the first acousto-optic modulator and the second acousto-optic modulator respectively; the laser frequency-shifted by the first acousto-optic modulator is emitted to the first side surface of the inspection mass through the first reflecting mirror and the first fast deflection mirror in sequence, and the laser reflected by the first side surface of the inspection mass is irradiated on the first four-quadrant photodetector after passing through the first fast deflection mirror and the third beam splitter in sequence; the laser frequency-shifted by the second acousto-optic modulator is irradiated on the first four-quadrant photodetector after passing through the fifth beam splitter and the fourth beam splitter. 4. The space gravitational wave detection device based on a single test mass according to claim 3, characterized in that the intersatellite long-arm interferometry measurement unit comprises a satellite-borne telescope and an intersatellite long-arm interferometer, and the intersatellite long-arm interferometer comprises a sixth beam splitter, a seventh beam splitter, an eighth beam splitter, a second reflector, and a second four-quadrant photoelectric detector; Among them, in an optical platform in a spacecraft, the laser frequency-shifted by the second acousto-optic modulator is divided into two laser beams D after passing through the sixth beam splitter, one of the laser beams D is emitted through the satellite telescope to emit intersatellite long-arm laser, and the other laser beam D is formed into interference laser after passing through the eighth beam splitter, and the interference laser is irradiated on the second four-quadrant photodetector; the intersatellite long-arm laser B emitted by the spacecraft adjacent to the spacecraft passes through the satellite telescope, the seventh beam splitter, the second reflector, and the eighth beam splitter in sequence, and then is irradiated on the second four-quadrant photodetector to form interference measurement with the interference laser to obtain the second optical path information of the spacecraft platform in the master spacecraft and the spacecraft platforms in the two slave spacecraft. 5. The space gravitational wave detection device based on a single test mass according to claim 4 is characterized in that the satellite-borne telescope includes a primary mirror and a second fast deflection mirror. 6. The space gravitational wave detection device based on a single test mass according to claim 4, characterized in that the optical phase-locked unit comprises a ninth beam splitter, a tenth beam splitter, a third four-quadrant photodetector, a first fiber coupler, a second fiber coupler, a phase meter and an optical phase-locked loop; Wherein, in an optical platform in the main spacecraft, another beam of the laser C is split into two beams of laser E by the ninth beam splitter, one beam of laser E is emitted to another optical platform of the main spacecraft through the first optical fiber coupler, and the other beam of laser E interferes with the laser emitted by the laser in another optical platform of the main spacecraft through the second optical fiber coupler on the third four-quadrant photodetector, and the detection signal of the third four-quadrant photodetector is passed through the phase meter to obtain the phase difference of the lasers emitted by the two lasers in the main spacecraft; the optical phase-locked loop is used to adjust the phase of the laser emitted by a laser in the main spacecraft according to the phase difference information, so that the phases of the lasers emitted by the two lasers in the main spacecraft remain consistent; In the slave spacecraft, the second four-quadrant photoelectric detector is used to measure the second optical path information of the spacecraft platform in the master spacecraft and the spacecraft platforms in the two slave spacecrafts, and the phase meter is used to calculate the phase difference between the intersatellite long-arm laser received by the onboard telescope in the slave spacecraft and the intersatellite long-arm laser emitted by its intersatellite long-arm interferometry measurement unit; the optical phase-locked loop adjusts the laser phase emitted by the laser in the slave spacecraft according to the phase difference information to make it consistent with the phase of the intersatellite long-arm laser received by the onboard telescope. 7. The space gravitational wave detection device based on a single test mass according to claim 1 is characterized in that the test mass is in the shape of a right octagonal prism. 8. The space gravitational wave detection device based on a single test mass according to claim 1 is characterized in that the test mass is in the shape of a rectangular parallelepiped with four arc-shaped corners. 9. The space gravitational wave detection device based on a single test mass according to claim 1 is characterized in that the test mass is in the shape of a rectangular parallelepiped with specific structures cut off at both ends of a pair of side surfaces, and the specific structure is in the shape of a rectangular parallelepiped with a triangular prism cut off at a corner.