1. A space attack and defense method based on a solar eclipse effect is used for shielding and controlling sunlight received by a target spacecraft, and comprises the following steps:

determining the fixed-point hovering orbit configuration of the task spacecraft and the target spacecraft;

based on the fixed-point hovering orbit configuration, calculating an ideal relative position of the task spacecraft and the target spacecraft according to the operation orbit and the sun position of the target spacecraft, and controlling the task spacecraft to maneuver to the ideal relative position so as to form a solar eclipse effect on the target spacecraft;

calculating and updating the ideal relative position of the task spacecraft and the target spacecraft at a preset time interval, and judging whether the actual relative position of the task spacecraft and the target spacecraft is within the range of a preset trajectory tracking error ball of the ideal relative position, wherein the preset trajectory tracking error ball represents a preset ball area with the ideal relative position as the center, and when the task spacecraft is located within the range of the preset ball area, the daily eating effect can be formed on the target spacecraft;

and if the actual relative position is not within the range of the preset trajectory tracking error ball, controlling the task spacecraft to maneuver to the ideal relative position at the corresponding moment.

2. The space defense and attack method based on the eclipse effect according to claim 1, characterized in that the fixed point hovering orbit configuration of the mission spacecraft and the target spacecraft is set to a circular formation configuration, and the configuration radius of the circular formation configuration is set to the nominal working distance of the mission spacecraft.

3. The solar eclipse effect-based space defense method according to claim 2, characterized in that: the O-XYZ coordinate system represents an orbit coordinate system of the target spacecraft, an origin O is the centroid of the target spacecraft, a Y axis points to the earth geocentric from the centroid of the target spacecraft, an X axis is positioned in the orbit plane of the target spacecraft and points to the motion direction of the target spacecraft, and a Z axis, the X axis and the Y axis form a right-hand rectangular coordinate system; unit vector of solar inertial system is k_iThe solar unit vector under the orbital coordinate system of the target spacecraft is k_o＝[m,n,p]^TM, n and p each represent a vector k_oComponents in the X, Y and Z directions of the orbital coordinate system of the target spacecraft;

calculating the ideal relative position of the mission spacecraft and the target spacecraft by using the following formula I;

wherein r is_t＝[x_t,y_t,z_t]^TRepresenting the ideal relative position, x, of the mission spacecraft and the target spacecraft_t、y_tAnd z_tRespectively representing the ideal relative position coordinates, D, of the mission spacecraft in the X, Y and Z directions of the orbital coordinate system of the target spacecraft_tRepresenting the nominal working distance of the mission spacecraft.

4. The space attacking and defending method based on the eclipse effect as claimed in claim 3, wherein if the actual relative position of the mission spacecraft and the target spacecraft satisfies the following formula three, the actual relative position is within the range of the preset trajectory tracking error ball of the ideal relative position;

wherein r is_f＝[x_f,y_f,z_f]^TRepresenting the actual relative position, x, of the mission spacecraft and the target spacecraft_f、y_fAnd z_fRespectively represents the actual relative position coordinates of the mission spacecraft in the X direction, the Y direction and the Z direction of the orbit coordinate system of the target spacecraft, eta represents the orbit control target area coefficient, 0 < eta < 1,representing the effective occlusion radius of the mission spacecraft.

5. The space attacking and defending method based on the eclipse effect as claimed in claim 4, wherein when controlling the task spacecraft to maneuver to an ideal relative position at a corresponding moment, the operation trajectory of the task spacecraft is calculated and determined by using a task spacecraft operation trajectory optimization design model shown in the following formula IV;

wherein the content of the first and second substances,representing the relative state of the mission spacecraft at the moment t in the orbital coordinate system of the target spacecraft, t₀Indicates the initial control time, t_fExpressing the corresponding time, J expressing the target function of the thrust of the orbit control engine of the corresponding task spacecraft, u (t) expressing the thrust of the orbit control engine of the task spacecraft at the time t, X (theta, h), Y (theta, h) and Z (theta, h) respectively expressing the X coordinate, the Y coordinate and the Z coordinate of a path constraint area of the task spacecraft in a coordinate system of a target spacecraft orbit, and r (theta, h)_min＝[x_min,y_min,z_min]^TRepresenting the minimum working distance, r, between the mission spacecraft and the target spacecraft_maxExpressing the maximum working distance between the task spacecraft and the target spacecraft, theta and h expressing equation parameters, u_minRepresents the minimum thrust of the mission spacecraft orbit control engine, u_maxRepresents the maximum thrust, x, of the mission spacecraft orbit control engine_f(t₀) Represents t₀Actual relative state, x, of the time mission spacecraft in the orbital coordinate system of the target spacecraft_t(t_f) Represents t_fAnd (3) the ideal relative state of the task spacecraft at the moment under the orbit coordinate system of the target spacecraft.

6. The eclipse effect-based space defense method according to claim 5, wherein an effective shielding radius of the mission spacecraft is adjusted by mounting a shielding device on the mission spacecraft.

7. The solar eclipse effect-based space defense method according to claim 6, wherein the shielding device is in a structure form capable of being unfolded and folded.

8. The solar eclipse effect-based space defense method according to any one of claims 1 to 7, further comprising:

if the actual relative position is within the range of the preset trajectory tracking error ball of the ideal relative position, repeating the process of calculating and updating the ideal relative position of the task spacecraft and the target spacecraft at a preset time interval, and judging whether the actual relative position of the task spacecraft and the target spacecraft is within the range of the preset trajectory tracking error ball of the ideal relative position.

9. The solar eclipse effect-based space defense method according to any one of claims 1 to 8, further comprising: and when the sunlight received by the target spacecraft does not need to be shielded, controlling the task spacecraft to leave the fixed-point hovering orbit.

Background

With the continuous development of the aerospace technology, the number of launched spacecrafts of various countries is increasing day by day, the space technology is taken as an important mark of the country in the field of aviation, and the battlefield of the game of the big country is gradually migrated to the space. Solar energy is widely used as a clean energy source in renewable energy sources, is widely used in various industries, is more and more emphasized in the field of aerospace, and is a main energy source of the current on-orbit operation spacecraft. The solar cell provided on the spacecraft converts solar energy into electric energy to provide electric energy for the system work of the spacecraft, and stores redundant electric energy in the storage battery to be used as a supplementary power supply in a shadow period or when the solar energy is insufficient. However, to ensure that the system power remains balanced when the spacecraft is operating in orbit, while at the same time saving design costs, the energy margin of the solar cells of the spacecraft is typically not more than 10%.

At present, the main attacking and defending technologies in space countermeasure comprise two means of hard killing and soft killing. The hardware system of the target spacecraft is irreversibly destroyed or destroyed by the hard killing means through adopting a kinetic energy weapon, a directional energy weapon, space maneuvering and operating equipment and the like, the attack mode is irreversible in effect, space fragments are easy to generate, and the space fragments are easy to discover by all parties of the target spacecraft. The soft killing means interferes the target spacecraft by adopting methods such as electromagnetic interference, battery array spraying, parasitic interference and the like, so that the spacecraft cannot normally work or the performance is reduced, the attack mode has controllable effect or recoverable damage, and the soft killing means is usually used for deterrence. However, the adoption of the soft killing means requires the task star and the target star to meet or accompany in a close range, requires maneuvering in advance and approach to the target star, has high requirements on a control system, is easy to be found by an enemy, and has collision risk. If evasive measures are taken in advance, task failure is easily caused.

Disclosure of Invention

In order to solve part or all of the technical problems in the prior art, the invention provides a space attack and defense method based on the eclipse effect.

The technical scheme of the invention is as follows:

a space attack and defense method based on the eclipse effect is provided, the method is used for shielding and controlling sunlight received by a target spacecraft, and comprises the following steps:

determining the fixed-point hovering orbit configuration of the task spacecraft and the target spacecraft;

based on the fixed-point hovering orbit configuration, calculating an ideal relative position of the task spacecraft and the target spacecraft according to the operation orbit and the sun position of the target spacecraft, and controlling the task spacecraft to maneuver to the ideal relative position so as to form a solar eclipse effect on the target spacecraft;

calculating and updating the ideal relative position of the task spacecraft and the target spacecraft at a preset time interval, and judging whether the actual relative position of the task spacecraft and the target spacecraft is within the range of a preset trajectory tracking error ball of the ideal relative position, wherein the preset trajectory tracking error ball represents a preset ball area with the ideal relative position as the center, and when the task spacecraft is located within the range of the preset ball area, the daily eating effect can be formed on the target spacecraft;

and if the actual relative position is not within the range of the preset trajectory tracking error ball, controlling the task spacecraft to maneuver to the ideal relative position at the corresponding moment.

In some optional implementations, the fixed-point hovering orbit configuration of the mission spacecraft and the target spacecraft is set to a circular formation configuration, and a configuration radius of the circular formation configuration is set to a nominal working distance of the mission spacecraft.

In some alternative implementations, it is set that: the O-XYZ coordinate system represents the orbit coordinate system of the target spacecraft, the origin O is the centroid of the target spacecraft, the Y axis points to the earth geocentric from the centroid of the target spacecraft, and the X axis is positioned on the orbit of the target spacecraftA Z axis, an X axis and a Y axis form a right-hand rectangular coordinate system; unit vector of solar inertial system is k_iThe solar unit vector under the orbital coordinate system of the target spacecraft is k_o＝[m,n,p]^TM, n and p each represent a vector k_oComponents in the X, Y and Z directions of the orbital coordinate system of the target spacecraft;

calculating the ideal relative position of the mission spacecraft and the target spacecraft by using the following formula I;

wherein r is_t＝[x_t,y_t,z_t]^TRepresenting the ideal relative position, x, of the mission spacecraft and the target spacecraft_t、y_tAnd z_tRespectively representing the ideal relative position coordinates, D, of the mission spacecraft in the X, Y and Z directions of the orbital coordinate system of the target spacecraft_tRepresenting the nominal working distance of the mission spacecraft.

In some optional implementation manners, if the actual relative position of the task spacecraft and the target spacecraft satisfies the following formula three, it indicates that the actual relative position is within the range of the preset trajectory tracking error ball of the ideal relative position;

wherein r is_f＝[x_f,y_f,z_f]^TRepresenting the actual relative position, x, of the mission spacecraft and the target spacecraft_f、y_fAnd z_fRespectively represents the actual relative position coordinates of the mission spacecraft in the X direction, the Y direction and the Z direction of the orbit coordinate system of the target spacecraft, eta represents the orbit control target area coefficient, 0 < eta < 1,representing the effective occlusion radius of the mission spacecraft.

In some optional implementation manners, when the task spacecraft is controlled to maneuver to an ideal relative position at a corresponding moment, calculating and determining the operation trajectory of the task spacecraft by using a task spacecraft operation trajectory optimization design model shown in the following formula four;

wherein the content of the first and second substances,representing the relative state of the mission spacecraft at the moment t in the orbital coordinate system of the target spacecraft, t₀Indicates the initial control time, t_fExpressing the corresponding time, J expressing the target function of the thrust of the orbit control engine of the corresponding task spacecraft, u (t) expressing the thrust of the orbit control engine of the task spacecraft at the time t, X (theta, h), Y (theta, h) and Z (theta, h) respectively expressing the X coordinate, the Y coordinate and the Z coordinate of a path constraint area of the task spacecraft in a coordinate system of a target spacecraft orbit, and r (theta, h)_min＝[x_min,y_min,z_min]^TRepresenting the minimum working distance, r, between the mission spacecraft and the target spacecraft_maxExpressing the maximum working distance between the task spacecraft and the target spacecraft, theta and h expressing equation parameters, u_minRepresents the minimum thrust of the mission spacecraft orbit control engine, u_maxRepresents the maximum thrust, x, of the mission spacecraft orbit control engine_f(t₀) Represents t₀Actual relative state, x, of the time mission spacecraft in the orbital coordinate system of the target spacecraft_t(t_f) Represents t_fAnd (3) the ideal relative state of the task spacecraft at the moment under the orbit coordinate system of the target spacecraft.

In some alternative implementations, the effective shielding radius of the mission spacecraft is adjusted by mounting a shielding device on the mission spacecraft.

In some alternative implementations, the shielding device takes the form of a structure that can be expanded and collapsed.

In some optional implementations, the method further includes:

if the actual relative position is within the range of the preset trajectory tracking error ball of the ideal relative position, repeating the process of calculating and updating the ideal relative position of the task spacecraft and the target spacecraft at a preset time interval, and judging whether the actual relative position of the task spacecraft and the target spacecraft is within the range of the preset trajectory tracking error ball of the ideal relative position.

In some optional implementations, the method further includes: and when the sunlight received by the target spacecraft does not need to be shielded, controlling the task spacecraft to leave the fixed-point hovering orbit.

The technical scheme of the invention has the following main advantages:

the space attack and defense method based on the eclipse effect performs orbit maneuver control on the task spacecraft according to the fixed-point hovering orbit configuration, the operation orbit of the target spacecraft and the relative position of the sun, so that the task spacecraft performs shielding control on the sunlight received by the target spacecraft, the normal work of the target spacecraft can be interfered, the target spacecraft cannot be irreversibly damaged, the concealment is strong, the collision risk does not exist, and the safety and the reliability are high.

Drawings

The accompanying drawings, which are included to provide a further understanding of embodiments of the invention and are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and together with the description serve to explain the invention and not to limit the invention. In the drawings:

fig. 1 is a flowchart of a space defense method based on the eclipse effect according to an embodiment of the present invention.

Detailed Description

In order to make the objects, technical solutions and advantages of the present invention more apparent, the technical solutions of the present invention will be clearly and completely described below with reference to the specific embodiments of the present invention and the accompanying drawings. It is to be understood that the described embodiments are merely a few embodiments of the invention, and not all embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments of the present invention without making any creative effort, shall fall within the protection scope of the present invention.

The technical scheme provided by the embodiment of the invention is described in detail below with reference to the accompanying drawings.

Referring to fig. 1, an embodiment of the present invention provides a space defense and attack method based on a solar eclipse effect, where the method is used for shielding and controlling sunlight received by a target spacecraft, and includes:

determining the fixed-point hovering orbit configuration of the task spacecraft and the target spacecraft;

based on the fixed-point hovering orbit configuration, calculating an ideal relative position of the task spacecraft and the target spacecraft according to the operation orbit and the sun position of the target spacecraft, and controlling the task spacecraft to maneuver to the ideal relative position so as to form a solar eclipse effect on the target spacecraft;

calculating and updating the ideal relative position of the task spacecraft and the target spacecraft at a preset time interval, and judging whether the actual relative position of the task spacecraft and the target spacecraft is within the range of a preset trajectory tracking error ball of the ideal relative position, wherein the preset trajectory tracking error ball represents a preset ball area with the ideal relative position as the center, and when the task spacecraft is located within the range of the preset ball area, the daily eating effect can be formed on the target spacecraft;

and if the actual relative position is not within the range of the preset trajectory tracking error ball, controlling the task spacecraft to maneuver to the ideal relative position at the corresponding moment.

In an embodiment of the invention, the solar eclipse effect on the target spacecraft means that sunlight received by the target spacecraft is shielded, so that the target spacecraft cannot absorb solar energy.

The space attack and defense method based on the eclipse effect provided by the embodiment of the invention performs orbit maneuver control on the task spacecraft according to the fixed-point hovering orbit configuration, the operation orbit of the target spacecraft and the relative position of the sun, so that the task spacecraft can shield and control the sunlight received by the target spacecraft, the normal work of the target spacecraft can be interfered, the target spacecraft cannot be irreversibly damaged, the concealment is strong, the collision risk does not exist, and the safety and the reliability are high.

The following is a detailed description of the steps of the space defense and attack method based on the eclipse effect according to an embodiment of the present invention.

In an embodiment of the invention, because the task to be executed by the task spacecraft is to shield sunlight received by the target spacecraft, based on the characteristics of the task to be executed by the task spacecraft, the fixed-point hovering orbit configuration of the task spacecraft and the target spacecraft can be set to be a circular formation configuration, and the configuration radius of the circular formation configuration is set to be the nominal working distance of the task spacecraft.

The nominal working distance of the task spacecraft is designed and determined according to the actual task condition, and the actual task condition comprises the following steps: the structural size and performance parameters of the mission spacecraft, the orbit and structural size of the target spacecraft, and the relative position of the target spacecraft to the sun.

Because the target spacecraft is in an on-orbit running state, the positions of the target spacecraft at different time points are different, and accordingly, the ideal relative position capable of forming the solar eclipse effect on the mission spacecraft can be changed. A specific example is given in an embodiment of the present invention as to how to calculate and determine the ideal relative position of the computing mission spacecraft and the target spacecraft.

Specifically, the following are set: the O-XYZ coordinate system represents an orbit coordinate system of the target spacecraft, an origin O is the centroid of the target spacecraft, a Y axis points to the earth geocentric from the centroid of the target spacecraft, an X axis is positioned in the orbit plane of the target spacecraft and points to the motion direction of the target spacecraft, and a Z axis, the X axis and the Y axis form a right-hand rectangular coordinate system; unit vector of solar inertial system is k_iThe solar unit vector under the orbital coordinate system of the target spacecraft is k_o＝[m,n,p]^TM, n and p each represent a vector k_oThe components in the X, Y and Z directions of the orbital coordinate system of the target spacecraft, the vector k_iSum vector k_oThe relationship between is A transformation matrix representing the solar inertial system to the orbital coordinate system of the target spacecraft;

calculating the ideal relative position of the mission spacecraft and the target spacecraft by using the following formula I;

in the formula, r_t＝[x_t,y_t,z_t]^TRepresenting the ideal relative position, x, of the mission spacecraft and the target spacecraft_t、y_tAnd z_tRespectively representing the ideal relative position coordinates, D, of the mission spacecraft in the X, Y and Z directions of the orbital coordinate system of the target spacecraft_tRepresenting the nominal working distance of the mission spacecraft.

Based on formula one, the ideal relative position represents the unit vector k with the sun_oCollinear co-directional and at a distance from the target spacecraft that is the nominal working distance of the mission spacecraft.

Because the target spacecraft is in an on-orbit running state, the position of the target spacecraft is a time variable, the solar unit vector under the orbit coordinate system of the target spacecraft is also a time variable, and correspondingly, the ideal relative position of the mission spacecraft and the target spacecraft, which is determined based on the first formula, is also a time variable.

Further, based on the set and calculated ideal relative position, the ideal relative speed of the mission spacecraft in the orbit coordinate system of the target spacecraft can be calculated and determined by the following formula two;

in the formula (I), the compound is shown in the specification,representing the ideal relative velocity, r, of the mission spacecraft in the orbital coordinate system of the target spacecraft_t(k) Indicating the ideal relative position of the current state update period, r_t(k-1) represents the ideal relative position of the previous state update period, and T represents the state update period, i.e. the above-mentioned preset time interval.

The preset time interval can be determined according to specific task requirements.

Accordingly, the ideal relative state of the mission spacecraft in the orbital coordinate system of the target spacecraft can be expressed asThe relative state includes a relative position and a relative velocity.

Further, in an embodiment of the present invention, the following formula three may be utilized to determine whether the actual relative position of the task spacecraft and the target spacecraft is within the range of the preset trajectory tracking error ball of the ideal relative position; specifically, if the actual relative position of the task spacecraft and the target spacecraft meets the following formula three, it indicates that the actual relative position of the task spacecraft and the target spacecraft is within the range of the preset trajectory tracking error ball of the ideal relative position;

in the formula, r_f＝[x_f,y_f,z_f]^TRepresenting the actual relative position, x, of the mission spacecraft and the target spacecraft_f、y_fAnd z_fRespectively represents the actual relative position coordinates of the mission spacecraft in the X direction, the Y direction and the Z direction of the orbit coordinate system of the target spacecraft, eta represents the orbit control target area coefficient, 0 < eta < 1,representing the effective occlusion radius of the mission spacecraft.

According to the formula III, idealThe predetermined tracking error sphere of the relative position is expressed as an ideal relative position r_t＝[x_t,y_t,z_t]^TAs an origin, inA sphere region formed by the radius.

The effective shielding radius of the task spacecraft depends on the structural size of the task spacecraft, and the orbit control target area coefficient is determined according to the effective shielding radius of the task spacecraft, the operation orbit and the structural size of the target spacecraft and the relative position of the target spacecraft and the sun, so that the solar eclipse effect on the target spacecraft can be formed when the actual relative position of the task spacecraft and the target spacecraft is within the range of a preset trajectory tracking error ball of an ideal relative position.

In one embodiment of the invention, the effective shielding radius of the task spacecraft can be adjusted by installing the shielding device on the task spacecraft.

Alternatively, the shielding device may be in the form of a deployable and collapsible structure, and may comprise, for example, an inflatable balloon or a telescoping umbrella.

By utilizing the shielding device which can be unfolded and folded, the effective shielding radius adjustment of the task spacecraft can be realized, the solar food effect on the target spacecraft is convenient to form, and the shielding device can be repeatedly used, and cannot influence the launching and running of the task spacecraft.

Further, in an embodiment of the present invention, when controlling the task spacecraft to maneuver to an ideal relative position at a corresponding moment, the operation trajectory of the task spacecraft can be determined by calculating using a task spacecraft operation trajectory optimization design model shown in the following formula four;

in the formula (I), the compound is shown in the specification,representing the relative state of the mission spacecraft at the moment t in the orbital coordinate system of the target spacecraft, t₀Indicates the initial control time, t_fExpressing the corresponding time, J expressing the target function of the thrust of the orbit control engine of the corresponding task spacecraft, u (t) expressing the thrust of the orbit control engine of the task spacecraft at the time t, X (theta, h), Y (theta, h) and Z (theta, h) respectively expressing the X coordinate, the Y coordinate and the Z coordinate of a path constraint area of the task spacecraft in a coordinate system of a target spacecraft orbit, and r (theta, h)_min＝[x_min,y_min,z_min]^TRepresenting the minimum working distance, r, between the mission spacecraft and the target spacecraft_maxExpressing the maximum working distance between the task spacecraft and the target spacecraft, theta and h expressing equation parameters, u_minRepresents the minimum thrust of the mission spacecraft orbit control engine, u_maxRepresents the maximum thrust, x, of the mission spacecraft orbit control engine_f(t₀) Represents t₀Actual relative state, x, of the time mission spacecraft in the orbital coordinate system of the target spacecraft_t(t_f) Represents t_fAnd (3) the ideal relative state of the task spacecraft at the moment under the orbit coordinate system of the target spacecraft.

In the optimization design model of the running track of the task spacecraft, the objective function is as follows:since the total thrust output by the spacecraft is proportional to the fuel consumption, the objective function is to minimize the fuel consumption of the mission spacecraft.

In the optimization design model of the running track of the task spacecraft, a function corresponding to the following formula five forms a task spacecraft path region constraint, and the task spacecraft path region constraint is used for ensuring that the task spacecraft can always form a solar eclipse effect on a target spacecraft in the orbital maneuver control process of the task spacecraft.

The minimum working distance and the maximum working distance between the task spacecraft and the target spacecraft can be designed and determined according to the actual task situation.

In the optimization design model of the running track of the task spacecraft, the initial state constraint and the terminal state constraint of the task spacecraft are formed by the functions corresponding to the following formula six and formula seven.

x(t₀)＝x_f(t₀) Formula six

x(t_f)＝x_t(t_f) Formula seven

In the optimization design model of the running track of the task spacecraft, a function corresponding to the following formula eight forms a track tracking error ball constraint of the task spacecraft, and the track tracking error ball constraint is used for ensuring that the task spacecraft can always form a solar eclipse effect on the target spacecraft in the process of controlling the orbital maneuver of the task spacecraft.

In an embodiment of the invention, an optimization algorithm such as a pseudo-spectrum method can be adopted to perform optimization solution on the task spacecraft operation trajectory optimization design model, trajectories meeting the objective function and the constraint conditions are obtained, and the obtained trajectories are used as input of a formation configuration maintaining controller to realize the orbital maneuver control of the task spacecraft.

The orbit maneuver control of the task spacecraft is carried out according to the operation locus of the task spacecraft calculated and determined by the optimization design model of the operation locus of the task spacecraft, and the following conditions can be simultaneously met: the fuel consumption of the mission spacecraft is minimum, and the mission spacecraft always forms the solar diet effect on the target spacecraft.

Further, in an embodiment of the present invention, the method further includes:

if the actual relative position is within the range of the preset trajectory tracking error ball of the ideal relative position, repeating the process of calculating and updating the ideal relative position of the task spacecraft and the target spacecraft at a preset time interval, and judging whether the actual relative position of the task spacecraft and the target spacecraft is within the range of the preset trajectory tracking error ball of the ideal relative position.

Further, in an embodiment of the present invention, the method further includes:

and when the sunlight received by the target spacecraft does not need to be shielded, controlling the task spacecraft to leave the fixed-point hovering orbit.

It is noted that, in this document, relational terms such as "first" and "second," and the like, may be used solely to distinguish one entity or action from another entity or action without necessarily requiring or implying any actual such relationship or order between such entities or actions. Also, the terms "comprises," "comprising," or any other variation thereof, are intended to cover a non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements does not include only those elements but may include other elements not expressly listed or inherent to such process, method, article, or apparatus. In addition, "front", "rear", "left", "right", "upper" and "lower" in the present document are all referred to as a state of being placed (if any) in the drawings.

Finally, it should be noted that: the above examples are only for illustrating the technical solutions of the present invention, and not for limiting the same; although the present invention has been described in detail with reference to the foregoing embodiments, it will be understood by those of ordinary skill in the art that: the technical solutions described in the foregoing embodiments may still be modified, or some technical features may be equivalently replaced; and such modifications or substitutions do not depart from the spirit and scope of the corresponding technical solutions of the embodiments of the present invention.