CN111457919A - Multipurpose constantan constellation system of sun static orbit - Google Patents

Abstract

The invention provides a multipurpose constellations system of solar static orbit, comprising: the satellites are arranged in a ecliptic plane and are uniformly deployed at equal radius by taking the sun as the center of a circle; each satellite includes: the energy transmission module is used for generating power by utilizing solar energy and transmitting energy to the detector through high-energy laser; the relay module is used for completing communication between the detector and the earth; the navigation module adds a modulation signal in the laser beam based on the position information of the satellite, so that high-energy-density energy charging is realized in the fixed star or satellite detection process, communication relay between the detector and the earth is realized, the detector can determine an accurate position through demodulation through the modulation signal, the detector can quickly and accurately complete the flight process, and a task target is realized.

Description

Multipurpose constantan constellation system of sun static orbit

Technical Field

The invention relates to the technical field of satellite navigation, in particular to a multipurpose constellations system of a sun stationary orbit and a state determination method of a detector.

Background

The current flying mission scheme of the near-earth planet detection adopts a gravitational orbit transfer mode, namely, a carrier rocket is used for directly launching a planet detector to an earth orbit, orbit change is carried out in an earth gravitational field, the planet detector is changed into an earth hyperbolic escape orbit, then the planet detector enters an earth fire cruising track mainly influenced by a solar gravitational field and flies to a mars without power, and an attitude orbit control engine carrying a propulsion system is used for track correction in the flying process. When approaching the influence range of the Mars gravitation, the speed is reduced, the orbit is changed, and the Mars orbit influenced by the Mars gravitational field is entered.

In the process that the detector flies from the earth orbit to the target planet, the electric energy used by an electrical system of the detector and the like is provided by the solar cell of the detector. As the distance between the detector and the sun increases, the illumination intensity decreases, the power generation power of the solar cell decreases, and the high-efficiency work of the detector is influenced. In addition, the detector is communicated with the ground by means of the satellite-borne equipment in the flying process, the satellite-borne equipment and the ground station are required to have higher power, and the size of an antenna of a detector measurement and control system and the ground station is required to be increased. The power consumption of the detector is increased, the size of the antenna is increased, the detector is inconvenient to maintain effective and reliable communication with the ground, and the envelope size of the detector is also large. Meanwhile, the detector runs in a transfer orbit, and information such as speed, position and the like of the detector is obtained mainly by measuring the star by the star sensor and comparing a star map, so that the method is single.

Disclosure of Invention

To address at least one of the above deficiencies, one aspect of the present invention provides a multipurpose constellations system for solar stationary orbits, comprising:

the satellites are arranged in a ecliptic plane and are uniformly deployed at equal radius by taking the sun as the center of a circle;

each satellite includes:

the energy transmission module is used for generating power by utilizing solar energy and transmitting energy to the detector through high-energy laser;

the relay module is used for completing communication between the detector and the earth;

and the navigation module is used for increasing a modulation signal in the laser beam based on the position information of the satellite.

In a preferred embodiment, the number of the satellites is 8-36, and an included angle between each two adjacent satellites and a connecting line of the sun is 10-45 degrees.

In a preferred embodiment, the energy transfer module comprises: a solar cell array, a laser and a reflector;

the solar cell array is used for generating electricity by utilizing solar energy;

the laser is used for converting electric energy into laser;

the mirror is used to track the motion of the detector.

In a preferred embodiment, the navigation module comprises: an inertia measurement system, a star sensor and an energy storage flywheel.

In a preferred embodiment, the relay module comprises: a measurement and control antenna and a transponder.

In a preferred embodiment, further comprising:

and the power module is used for adjusting the attitude and the orbit of the satellite.

In a preferred embodiment, the laser is a semiconductor laser.

In a preferred embodiment, the solar cell array is a thin film type solar cell array.

In a preferred embodiment, the modulation signal is a high-low logic signal formed by switching off the laser.

Another embodiment of the present invention provides a method for determining a state of a detector, where the state of the detector includes a speed and a position, and the method includes:

sending laser beams to a detector of which the position is to be determined through a plurality of satellites; wherein the content of the first and second substances,

in the process of sending the laser beam, high and low logic signals are formed by switching off the laser; and the detector receives the laser beams sent by the satellites, and the corresponding high-low logic signals are resolved to obtain the speed and the position of the detector.

The invention has the following beneficial effects:

the invention provides a multipurpose constellations system of a solar stationary orbit and a state determination method of a detector, which realize high-energy density energy charging in the detection process of a fixed star or a satellite, realize communication relay between the detector and the earth, and enable the detector to determine an accurate position through demodulation by modulating signals, so that the detector can quickly and accurately complete a flight process and realize a task target.

Drawings

In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly described below, it is obvious that the drawings in the following description are only some embodiments of the present invention, and for those skilled in the art, other drawings can be obtained according to the drawings without creative efforts.

FIG. 1 is a schematic diagram of a constellation satellite deployment of the present invention;

FIG. 2 is a schematic diagram of energy transmission of a constellation satellite generator according to the present invention;

FIG. 3 is a schematic diagram of a relay function of a constellation satellite according to the present invention;

FIG. 4 is a schematic diagram of a constellation satellite assembly according to the present invention;

FIG. 5 is a schematic diagram of a Beidou satellite position correction of the constellation satellite of the present invention.

Detailed Description

The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.

The invention provides a multipurpose constellations system of solar stationary orbit, as shown in fig. 1, fig. 2 and fig. 4, comprising: the satellites are arranged in a ecliptic plane and are uniformly deployed at equal radius by taking the sun as the center of a circle; each satellite includes: the energy transmission module is used for generating power by utilizing solar energy and transmitting energy to the detector through high-energy laser; the relay module is used for completing communication between the detector and the earth; and the navigation module is used for increasing a modulation signal in the laser beam based on the position information of the satellite.

In a preferred embodiment, the number of the satellites is 8-36, and an included angle between each two adjacent satellites and a connecting line of the sun is 10-45 degrees.

In a preferred embodiment, the energy transfer module comprises: a solar cell array 3, a laser and a reflector 6; the solar cell array is used for generating electricity by utilizing solar energy; the laser is used for converting electric energy into laser; the mirror is used to track the motion of the detector.

In a preferred embodiment, the navigation module comprises: an inertia measurement system, a star sensor 2 and an energy storage flywheel.

In a preferred embodiment, the relay module comprises: a measurement and control antenna 4 and a transponder.

In a preferred embodiment, further comprising: and the power module 1 is used for adjusting the attitude and the orbit of the satellite.

In a preferred embodiment, the laser is a semiconductor laser.

In a preferred embodiment, the solar cell array is a thin film type solar cell array.

In a preferred embodiment, the modulation signal is a high-low logic signal formed by switching off the laser.

The following description is made with reference to specific scenarios:

in a specific implementation, a constellations system of a stationary orbit of the sun can be constructed in the interval of 1.2a.u. to 1.35a.u. from the sun in the ecliptic plane (the optimal distance is 1.25A.U.), wherein 1A.U. (astronomical unit) refers to the average distance between the earth and the sun, and the data is published according to the international union of astronomy, and 1A.U. (astronomical unit) 149,597,870km (kilometer). The satellites are uniformly deployed in the ecliptic plane at equal radius by taking the sun as the center of a circle, the number of the deployed satellites is between 8 and 36, namely the included angle between the connecting line of two adjacent satellites and the sun is 45-10 degrees; the optimal deployment quantity is 24, namely the included angle between the adjacent two satellites and the sun connecting line is 15 degrees. The deployed satellites remain in a heliocentric inertial frame with positions that remain constant, i.e., the relative positions of the satellites to the sun in the solar frame remain constant. The constellation deployment is schematically shown in fig. 1.

In addition, in the invention, the capacity transmission module has the following capacity transmission function: the device consists of a laser, a reflector and the like and is used for realizing energy transmission to a detector in a wireless mode. In addition, a tracking angle reflecting mirror is arranged on the satellite and used for tracking the motion of the detector, and the satellite generates electricity by converting solar energy into electric energy through a large-area thin-film solar cell array at the position of a constellation orbit. And energy transmission is performed by using high-energy laser, which is detailed in fig. 2.

The relay module has a relay function: the satellite earth observation and control system is composed of measurement and control antennas, transponders and other equipment, and besides the earth measurement and control of the satellites, a star chain is formed among the satellites, so that the relay function between the detector and the ground is completed. If the relay is not carried out through the satellite, the farthest distance between the detector and the earth is 2.5A.U. when the detector reaches a mars, namely about 3.73 hundred million kilometers, the minimum aperture of a measurement and control antenna on the detector is 1m, and the requirement on equipment on the detector is high due to the fact that X-frequency band radio frequency is used. The satellites in the orbit form a relay constellation, when the detector is far away from the earth, the detector is used as a relay through an inter-satellite link between the constellations, the communication function between the detector and the earth is completed, the technical requirements on a measurement and control system on the detector, particularly an antenna and the like are reduced, and detailed description is given in figure 3.

The navigation module has a constellation navigation function: (GNC System): the satellite velocity measurement system comprises an inertial measurement system, a star sensor, an energy storage flywheel and the like, and is used for measuring the motion states of the satellite such as the velocity, the position, the attitude and the like. The satellite uses energy-transmitting laser, and the position information and the like are modulated in the light beam based on the relative fixation of the position of the satellite. The detector is used for detecting the satellite laser energy, resolving the modulation information in the laser, and resolving to obtain the information such as the relatively accurate speed and position of the detector.

In some embodiments, the satellite further comprises:

a structural module: a lightweight structure is adopted, and platform support is provided for on-board equipment installation;

a power supply module: the solar photovoltaic power generation system comprises a solar cell array, a storage battery, a power supply and distribution management unit and the like. The solar cell generates electric energy which is used as a source for the self-operation of the satellite and the wireless transmission energy of the laser.

A power module: and an electric propulsion power system is adopted to complete the functions of orbit control and attitude control of the satellite.

The satellite is schematically illustrated in fig. 4, in which some of the components are identified, and in particular, the identified components include: the system comprises a power module 1, a star sensor 2, a solar cell array 3, a measurement and control antenna 4, a relay inter-satellite communication transceiving antenna 5, a reflector 6 and a structural system 7.

The following is a description with specific scenarios:

satellite deployment implementation

The satellite transmission process is as follows:

when the earth revolves to the position near the satellite deployment point, the satellite is launched to the earth orbit by using a carrier rocket, and the satellite moves on the earth orbit, changes the orbit, accelerates and enters a transfer orbit. The electric propulsion system provides thrust to accelerate the satellite to fly to a target orbit along the transfer orbit and reach a target parking point.

Satellite function and performance implementation

1. Realization of satellite constellation energy transmission function

(1) Satellite power supply system

Taking the three-junction gallium arsenide solar cell which is most widely used at present as an example for calculation. At earth orbit, i.e. 1a.u., the generated power of the currently mature triple junction gaas solar cell is about 300W/m 2. Since the solar illumination intensity is inversely proportional to the square of the distance from the light source, and the generated power of the solar cell is inversely proportional to the square of the distance from the light source, the generated power of the solar cell is about 192W/m2 at the satellite constellation orbit, i.e. 1.25a.u.

The area of the flexible thin film type solar cell is designed to be 20m 2-100 m2, the typical value can be 50m2, at the moment, the electric energy obtained by the solar cell is 192W/m2 × 30, 30m2 is 9600W, and the requirements of satellite work and energy transmission are met.

(2) Energy transmission module

The laser is used for converting electric energy into laser wireless energy, the laser can be a solid laser, a gas laser and a semiconductor laser, performance parameters such as conversion efficiency are considered, and the semiconductor laser can be selected in a typical design; the energy transmission equipment transmits high-energy laser to an energy-receiving target (detector), and the typical energy transmission equipment adopts a concave reflector controlled by a servo system, so that the laser transmission at different transmission angles can be met by rotating the angle according to the requirement. Meanwhile, the high-energy laser can be effectively received for tracking the movement of the detector.

2. Implementation of satellite relay function

The relay function of the satellite measurement and control system is divided into three steps

(1) Satellite and detector

The detector and the satellite closest to the detector are in measurement and control communication, can be realized by adopting a small-caliber antenna, have the function of communicating with two adjacent satellites simultaneously, and are convenient to switch.

(2) Inter-satellite communication

Fixed link communication is adopted among satellites, and a typical design can adopt an X-band communication design.

(3) Satellite-to-ground communication

The satellite-ground measurement and control communication adopts an X wave band to complete measurement and control communication with each measurement and control station on the ground.

3. Implementation of constellation navigation function

In the process of transmitting energy to the detector by using laser, the satellite can modulate information such as satellite position, time scale and the like in the energy transmission laser, and in view of the fact that the energy density of the energy transmission laser is high, a typical modulation mode is to form high-low logic signals by cutting off the laser.

And the detector end demodulates the signal transmitted by the cut-off laser after receiving the signal. The detector receives information of 4 or more satellites, and motion information such as relatively accurate speed and position of the detector is obtained through calculation, so that the function of detector navigation is realized.

Satellite performance

1. Attitude assurance

(1) GNC system: measuring the speed, position and attitude information of the satellite by using an inertial measurement system, a star sensor and the like, comparing a star map, and controlling the attitude of the satellite by using a flywheel;

(2) fall beidou system: because the positions of the satellites in the constellation are kept unchanged in the centroid inertial coordinate system and are uniformly distributed in the orbit, the period of traversing all the satellites is 1 year. When the satellite passes through a near place, the satellite and a ground measurement and control station built on the earth can form an inverted Beidou system, namely, the ground station (4 or more) with the accurately measured position is utilized to communicate with the satellite, the satellite carries out resolving on the satellite, and the space position information of the satellite is corrected, which is shown in detail in figure 5.

2. Satellite attitude and orbit control implementation

During the in-orbit operation of the satellite, the relative position of the satellite in the centroid inertial coordinate system is required to be constant. The satellite GNC system measures the speed and position attitude of the satellite and performs orbit maintenance on the satellite by using an electric propulsion system, such as a high-power ion propulsion system, a Hall propulsion system and the like; the attitude of the detector is controlled by using an energy storage flywheel and the like.

From the above analysis, it can be understood that the invention can realize the functions of high-density energy charging, communication relaying and navigation of the detector in the solar system during the flight process by using the satellite constellation on the solar stationary orbit in the solar system ecliptic plane, the laser energy transmission module and the inter-satellite communication link. The invention reduces the technical requirements and technical difficulty for the detector and correspondingly reduces the cost for research, development and production and manufacture of the detector. Meanwhile, the invention reduces the technical complexity of the detector, improves the task reliability, has higher practical value and provides important guarantee for ensuring the safe and high-reliability operation of the near-earth planet detection spacecrafts such as the sparks in the solar system.

Further, the present invention also provides a method for determining a state of a probe, where the state of the probe includes a speed and a position, the method comprising:

sending laser beams to a detector of which the position is to be determined through a plurality of satellites; wherein the content of the first and second substances,

in the process of sending the laser beam, high and low logic signals are formed by switching off the laser; and the detector receives the laser beams sent by the satellites, and the corresponding high-low logic signals are resolved to obtain the speed and the position of the detector.

According to the state determining method of the detector, the detector can determine an accurate position through demodulation by modulating the signal, so that the detector can quickly and accurately complete a flight process, and a task target is achieved.

The embodiments in the present specification are described in a progressive manner, and the same and similar parts among the embodiments are referred to each other, and each embodiment focuses on the differences from the other embodiments. In particular, for the system embodiment, since it is substantially similar to the method embodiment, the description is simple, and for the relevant points, reference may be made to the partial description of the method embodiment. In the description herein, references to the description of the term "one embodiment," "some embodiments," "an example," "a specific example," or "some examples," etc., mean 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 an embodiment of the specification. In this specification, the schematic representations of the terms used above are not necessarily intended to refer to the same embodiment or example. 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. The above description is only an example of the embodiments of the present disclosure, and is not intended to limit the embodiments of the present disclosure. Various modifications and variations to the embodiments described herein will be apparent to those skilled in the art. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the embodiments of the present specification should be included in the scope of the claims of the embodiments of the present specification.

Claims (10)

1. A multi-purpose constellations system for stationary orbits of the sun, comprising:

the satellites are arranged in a ecliptic plane and are uniformly deployed at equal radius by taking the sun as the center of a circle;

each satellite includes:

the energy transmission module is used for generating power by utilizing solar energy and transmitting energy to the detector through high-energy laser;

the relay module is used for completing communication between the detector and the earth;

and the navigation module is used for increasing a modulation signal in the laser beam based on the position information of the satellite.

2. The multi-purpose constantan constellation system of claim 1, wherein the number of said satellites is 8-36, and the angle between two adjacent satellites and the line connecting the sun is 10-45 °.

3. The multipurpose constellations system of claim 1, wherein the energy transfer module comprises: a solar cell array, a laser and a reflector;

the solar cell array is used for generating electricity by utilizing solar energy;

the laser is used for converting electric energy into laser;

the mirror is used to track the motion of the detector.

4. The multipurpose constellations system of claim 1, wherein the navigation module comprises: an inertia measurement system, a star sensor and an energy storage flywheel.

5. The multipurpose constellations system of claim 1, wherein the relay module comprises: a measurement and control antenna and a transponder.

6. The multipurpose constellations system of claim 1, further comprising:

and the power module is used for adjusting the attitude and the orbit of the satellite.

7. The multi-purpose constellations system of claim 3, wherein the laser is a semiconductor laser.

8. The multi-purpose constellations system of claim 3, wherein the array of solar cells is a thin-film solar array.

9. The multi-purpose constellations system of claim 3, wherein the modulated signal is a high-low logic signal formed by switching off a laser.

10. A method of determining a condition of a probe, the condition of the probe including a velocity and a position, the method comprising:

sending laser beams to a detector of which the position is to be determined through a plurality of satellites; wherein the content of the first and second substances,

in the process of sending the laser beam, high and low logic signals are formed by switching off the laser; and the detector receives the laser beams sent by the satellites, and the corresponding high-low logic signals are resolved to obtain the speed and the position of the detector.

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