CN110510154B - Off-orbit delivery rail attitude coupling adjustment method for geomagnetic energy storage low-orbit space debris - Google Patents

Off-orbit delivery rail attitude coupling adjustment method for geomagnetic energy storage low-orbit space debris Download PDF

Info

Publication number
CN110510154B
CN110510154B CN201910774236.3A CN201910774236A CN110510154B CN 110510154 B CN110510154 B CN 110510154B CN 201910774236 A CN201910774236 A CN 201910774236A CN 110510154 B CN110510154 B CN 110510154B
Authority
CN
China
Prior art keywords
orbit
energy
spacecraft
energy storage
geomagnetic
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Active
Application number
CN201910774236.3A
Other languages
Chinese (zh)
Other versions
CN110510154A (en
Inventor
李文皓
冯冠华
张珩
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Institute of Mechanics of CAS
Original Assignee
Institute of Mechanics of CAS
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Institute of Mechanics of CAS filed Critical Institute of Mechanics of CAS
Priority to CN201910774236.3A priority Critical patent/CN110510154B/en
Publication of CN110510154A publication Critical patent/CN110510154A/en
Application granted granted Critical
Publication of CN110510154B publication Critical patent/CN110510154B/en
Active legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Images

Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/242—Orbits and trajectories
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/244—Spacecraft control systems
    • B64G1/245—Attitude control algorithms for spacecraft attitude control
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/66—Arrangements or adaptations of apparatus or instruments, not otherwise provided for

Landscapes

  • Engineering & Computer Science (AREA)
  • Remote Sensing (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Radar, Positioning & Navigation (AREA)
  • Automation & Control Theory (AREA)
  • Measuring Magnetic Variables (AREA)
  • Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)

Abstract

The embodiment of the invention discloses a method for coupling and adjusting the off-orbit delivery track attitude of space debris of a geomagnetic energy storage low track, which comprises the following steps: step 100, decomposing the whole geomagnetic energy storage low orbit period into an energy accumulation arc section and an energy release arc section, and configuring a momentum wheel energy absorption device on three axes of the spacecraft; 200, respectively determining an energy accumulation arc section and an energy release arc section when the spacecraft accumulates energy around an orbital plane y axis and an orbital plane z axis; 300, after the energy-accumulating arc section is determined, for energy-accumulating targets of Y axes or Z axes of different orbit coordinate systems, finishing off-orbit operation of space debris based on a geomagnetic energy-accumulating low-orbit space debris off-orbit control method; and 400, after the energy release arc section is determined, unloading the momentum wheel energy absorption device on the three shafts of the spacecraft through the generated magnetic moment.

Description

Off-orbit delivery rail attitude coupling adjustment method for geomagnetic energy storage low-orbit space debris
Technical Field
The embodiment of the invention relates to the technical field of space fragment off-orbit, in particular to a method for adjusting the off-orbit delivery attitude coupling of space fragments of a geomagnetic energy storage low orbit.
Background
Space garbage with the diameter of 10cm can bring serious threat to spacecrafts and astronauts, and a Hubby telescope, a space shuttle and an international space station are impacted by space debris; with the increasing activity of human space activities and the rapidly increasing chance of threat from the Kessler effect of space debris (Kessler effect: the generation of more space debris due to the impact of space debris).
The Kessler effect further aggravates the threat of space debris, such as 2009, where russian obsolete Comos satellites collide with U.S. iridium satellites, resulting in more than 2000 space debris, which would continue to increase substantially if the existing space debris could not be actively and effectively removed, which would seriously affect the space development process of human beings.
At present, more than 20000 space fragments of more than 10cm on orbit are distributed on an earth orbit with different inclination angles of 500 km-1000 km, and the fragments can not enter the earth atmosphere within decades by only attenuating the orbit height by the atmospheric resistance of the fragments, so that compared with a geosynchronous orbit (36000km, the orbit is unique and the resources are scarce), the low orbit space fragments are distributed and dispersed, have the characteristics of high threat and low value, and can be actively removed, but the problems of efficiency and economy for removing the low orbit space fragments are difficult to solve.
The better mode of processing the space debris is to change the height of the orbit which runs by people and reduce the height of the near place of the orbit to be less than 200km, so that the debris is influenced by the atmospheric resistance of the earth to quickly reduce the semimajor axis of the orbit and fall into the atmosphere to be burnt.
Among the various methods proposed at present, the active dragging and derailing method using chemical/electric thrust is the most mature, but the consumption is extremely high; the passive rail descending modes such as installing the air bag resistance sail on the fragments and coating foam to increase the surface-to-mass ratio save the consumption of off-rail propulsion working media, but need to consume installation materials, and the time required by the fragment rail descending is years, so that the probability of secondary collision is undoubtedly and greatly increased.
In addition, working medium consumption-free fragment cleaning methods are actively developed in all countries, and some methods are well assumed but difficult to realize, for example, the fragments are cleaned by using ground/space-based high-energy laser, the basic principle is that a burning product is rapidly expanded to separate from the fragments through high-energy laser burning, and the fragments obtain recoil quantity to reduce orbit.
In the non-working medium consumption type rail transferring and separating method, the electric rope system is a mode which is considered to be the highest in feasibility and realizability at present, the electric rope system is used for adjusting the rail in the earth magnetic field of a low-rail space through the ampere force borne by the electrified rope by collecting charged particles in the space, only power consumption is realized, no working medium is consumed, the reliability of stable operation of the rope system in the space is very outstanding no matter the efficiency of the electric rope system depends on the size of the rope system, and the huge size of thousands of meters to tens of kilometers is large.
Although a plurality of theoretical researches and space tests show that a rope system has a considerable safety factor, the rope system is also experienced as disastrous, and the SED-2 (launched in 1994 and rope system unfolded for 19.7km) in the United states is unfolded to be cut off by space fragments/micrometeors in only 4 days, so that the space safety problem of the rope system is still solved, the problem is solved by adopting a method of separating the space fragments of the geomagnetic energy storage low orbit, but when the space fragments are separated from the orbit by using the geomagnetic energy storage, magnetic moments in other directions are inevitably introduced, and when the additional magnetic moments are excessively accumulated, special unloading is required, so that the energy storage-release action persistence is interrupted.
Disclosure of Invention
Therefore, the embodiment of the invention provides a geomagnetic energy storage low-orbit space debris off-orbit delivery orbit attitude coupling adjustment method, which aims to solve the problem that in the prior art, when the accumulation of additional magnetic moment is too large, special unloading is required, so that the continuity of energy storage-release action is interrupted.
In order to achieve the above object, an embodiment of the present invention provides the following:
a method for adjusting the off-orbit delivery attitude coupling of space debris of a geomagnetic energy storage low orbit comprises the following steps:
step 100, decomposing the whole geomagnetic energy storage low orbit period into an energy accumulation arc section and an energy release arc section, and configuring a momentum wheel energy absorption device on three axes of the spacecraft;
200, respectively determining an energy accumulation arc section and an energy release arc section when the spacecraft accumulates energy around an orbital plane y axis and an orbital plane z axis;
300, after the energy-gathering arc section is determined, for energy-accumulating targets of Y axes or Z axes of different orbit coordinate systems, finishing off-orbit operation of space debris based on a geomagnetic energy-accumulating low-orbit space debris off-orbit control method;
and 400, after the energy release arc section is determined, unloading the momentum wheel energy absorption device on the three shafts of the spacecraft in a mode of generating magnetic moments.
In a preferred embodiment of the present invention, in step 200, the energy-gathering arc segments and the energy-releasing arc segments of the spacecraft around the y-axis of the orbital plane are symmetrically distributed with an angular distance u of 180 ° between the lift points in one orbital period.
As a preferable aspect of the present invention, the energy release arc segments are symmetrically spread around the lift intersection angle distances of 0 °, 90 °, 180 °, and 270 °, and are set to be Δ u0、
Figure BDA0002174538280000033
Δuπ，
Figure BDA0002174538280000034
Corresponding to the values of the spread angular distances symmetrically for four angles, respectively, it is clear that Δ u0＝Δuπ，
Figure BDA0002174538280000035
As a preferable scheme of the invention, before the start of the orbit period, the measurement obtains the rotation momentum of the three axes under the corresponding orbit coordinate system which still remains under a certain determined attitude of the spacecraft system
Figure BDA0002174538280000036
The storage capacity of the rotational momentum of the three axes is
Figure BDA0002174538280000032
And distributed.
As a preferable aspect of the present invention, the method of allocating includes the steps of:
step 211, generating a magnetic moment through an international geomagnetic reference field IGRF geomagnetic model and a spacecraft, and performing simulation calculation by using an adjustment strategy of the following formula:
Figure BDA0002174538280000041
wherein, Cmd _ mx、Cmd_my、Cmd_mzRepresenting the desired adjusted magnetic moment magnitude component in each direction in the geomagnetic coordinate system, | m | representing the total magnetic moment magnitude that can be generated,
Figure BDA0002174538280000042
the method comprises the following steps that (1) in the on-orbit flight of the spacecraft, u represents the angular distance of a rising point of the spacecraft in the current flight at the current moment, wherein the geographic latitude corresponds to the subsatellite point of the spacecraft at the current moment;
in any quadrant area of the ascending intersection angle distance of 0-90 degrees, 90-180 degrees, 180-270 degrees and 270-360 degrees, the ascending intersection angle distance is a parabolic curve, so that a value existsIn 0-90 degree quadrant, the moment of Y axis of orbit coordinate system is in uaAnd
Figure BDA0002174538280000043
two points being equal and u within this intervalaAnd ubOne-to-one correspondence is realized;
step 212, in 0-90 degree quadrant, for arbitrary uaAnd u corresponding theretobCalculating the angular distance from u at the intersection point of ascentaTo
Figure BDA0002174538280000044
Cumulative amount on X and Z axes of internal orbital coordinate system
Figure BDA0002174538280000045
Step 213, determine whether or not
Figure BDA0002174538280000046
And at the same time
Figure BDA0002174538280000047
Figure BDA0002174538280000048
k is a predetermined conservative coefficient and is 0.5<k<0.95；
If not, then u is adjustedaAnd ubSuch that the condition is satisfied;
step 214, when the condition is satisfied, outputting uaAnd ubThen Δ u0＝ua，
Figure BDA0002174538280000049
Thereby obtaining the arc segment of energy release in the period of the orbit as
Figure BDA00021745382800000410
The rest are energy storage arc sections.
In a preferred embodiment of the invention, in the charging arc segment, if the stored X-axis rotational momentum at the present moment is present
Figure BDA00021745382800000411
And stored Z-axis rotational momentum
Figure BDA00021745382800000412
Appear
Figure BDA00021745382800000413
Figure BDA00021745382800000414
Or
Figure BDA00021745382800000415
The energy releasing arc segment is immediately switched to, and the current rising point angular distance u is measurednowJudgment unowIn the quadrant and replace ubAnd updating u according to the calculated valueaThereby changing and updating the range of the energy storage arc segment/energy release arc segment in the orbit period.
In a preferred embodiment of the present invention, in step 200, the energy-collecting arc segments and the energy-releasing arc segments of the spacecraft around the orbital plane Z-axis are asymmetrically distributed in one orbital cycle with the angular distance of the lift-intersection as a standard.
As a preferable scheme of the invention, before the start of the orbit period, the three-axis rotation momentum under the corresponding orbit coordinate system is respectively measured and obtained under a certain determined attitude remained in the spacecraft system
Figure BDA0002174538280000051
The storage capacity of the rotational momentum of the three axes is
Figure BDA0002174538280000052
The allocation is made according to the following sub-steps:
step 221, the initial condition is an energy storage arc segment, and an adjustment strategy of the following formula is used:
Figure BDA00021745382800000510
wherein, Cmd _ mx、Cmd_my、Cmd_mzThe method comprises the steps of representing the size component of the magnetic moment which is expected to be adjusted in each direction in a geomagnetic coordinate system, and representing the total producible magnetic moment size in m;
measuring stored X-axis rotational momentum at the present time
Figure BDA0002174538280000053
And stored Y-axis rotational momentum
Figure BDA0002174538280000054
Appear
Figure BDA0002174538280000055
Or
Figure BDA0002174538280000056
Then immediately shifts into an energy release arc segment
Step 222, the energy releasing arc segment continues to
Figure BDA0002174538280000057
And at the same time
Figure BDA0002174538280000058
Figure BDA0002174538280000059
Wherein tau is a preset energy storage coefficient, and is generally 0.05<τ<0.3。
The embodiment of the invention has the following advantages:
according to the invention, the energy storage arc section and the energy release arc section are arranged, so that the additional moment accumulation caused by the complexity of geomagnetic distribution can be released in time, and simultaneously, the geomagnetic distribution characteristic is utilized to complete high-efficiency energy storage to the maximum degree, so that the problems of efficiency and economy of low-orbit space debris are obviously improved.
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 should be apparent that the drawings in the following description are merely exemplary, and that other embodiments can be derived from the drawings provided by those of ordinary skill in the art without inventive effort.
Fig. 1 is a flowchart of a space debris off-track control method according to embodiment 1 of the present invention;
fig. 2 is a schematic view of the orbital flight inclination angle of embodiment 1 of the present invention.
FIG. 3 is a schematic diagram of magnetic moment and accumulated magnetic torque in an orbital coordinate system under the strategy of equations (1) and (3) in example 2 of the present invention;
fig. 4 is a graph illustrating the cumulative magnetic torque calculated by the strategy of equation (4) in example 2 according to the present invention under different geomagnetic models;
fig. 5 is a flowchart of a delivery track attitude coupling adjustment method according to embodiment 2 of the present invention;
Detailed Description
The present invention is described in terms of particular embodiments, other advantages and features of the invention will become apparent to those skilled in the art from the following disclosure, and it is to be understood that the described embodiments are merely exemplary of the invention and that it is not intended to limit the invention to the particular embodiments disclosed. 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.
Example 1:
as shown in fig. 1 and 2, the invention provides a geomagnetic energy storage low orbit space debris off-orbit control method, which is used for capturing space debris by a spacecraft to perform geomagnetic energy storage, wherein the geomagnetic energy storage comprises the following steps:
step 10, determining an energy storage direction according to the orbit flight inclination angle of the spacecraft;
when the orbit flight inclination angle of the spacecraft is larger than a set angle, accumulating energy around the y axis of the orbit surface;
and when the flight inclination angle of the orbit is smaller than the set angle, accumulating energy around the z axis of the orbit surface.
In this embodiment, the accumulated energy is specifically continuously accelerated in a preset direction, and a magnetic moment needs to be continuously generated in the preset direction, so that the speed in the preset direction is continuously increased as time goes on.
According to the traversing of the height of the 100-km track and the different track inclination angles, the efficiency calculation of energy storage accumulation is carried out in two different modes, the efficiency result is compared to obtain that the boundary point of the inclination angle is at the preferred angle, and the set angle is preferably 49.14 degrees.
Step 20, measuring the current flight attitude of the spacecraft, wherein the parameters comprise the altitude h, the orbit inclination angle, the geographical longitude and latitude of the subsatellite point of the spacecraft, and the component of the geomagnetic field of the current position of the spacecraft in a geomagnetic coordinate system
Figure BDA0002174538280000073
Adjusting the flight attitude of the spacecraft, associating the magnetic moment generating direction of an orthogonal coil fixedly connected to the spacecraft and the accumulated rotating distance direction of the magnetic moment with the flight attitude of the spacecraft, and keeping the associated flight attitude;
the generation of the magnetic moment is realized by using current generated by an orthogonal coil fixedly connected to the spacecraft, is irrelevant to the attitude of the spacecraft, and has the same direction of the generated magnetic moment and the accumulated torque, and is in an orthogonal relation instantaneously.
The process of association can be understood as: the spacecraft can require an attitude angle when executing a specific task, in order to generate a specific accumulation speed on a specific orbital plane, the direction of a delivery mechanism fixedly connected with the spacecraft needs to be aligned, in addition, a strong magnetic moment generating device (orthogonal coil) is fixedly connected with the spacecraft, a resolving strategy for generating the strong magnetic moment is strongly related to the distribution of the geomagnetic field, therefore, preassembly adjustment needs to be carried out according to the attitude of the spacecraft, and the three requirements are determined to be relative relations before delivery.
Step 30, generating strong magnetic moment
Figure BDA0002174538280000071
Is given byAn amount;
when accumulating energy around the Y axis of the track surface
Figure BDA0002174538280000072
The resulting strategy is as follows:
Figure BDA0002174538280000081
when accumulating energy around Z axis of track surface
Figure BDA0002174538280000082
The resulting strategy is as follows:
Figure BDA0002174538280000083
wherein, Cmd _ mx、Cmd_my、Cmd_mzRepresenting the desired adjusted magnetic moment magnitude component in each direction in the geomagnetic coordinate system, | m | representing the total magnetic moment magnitude that can be generated,
Figure BDA0002174538280000084
the method comprises the following steps that (1) in the on-orbit flight of the spacecraft, u represents the angular distance of a rising point of the spacecraft in the current flight at the current moment, wherein the geographic latitude corresponds to the subsatellite point of the spacecraft at the current moment;
step 40, continuously accelerating and rotating the grabbed fragment ends, detecting the accumulated angular momentum value and the relative linear velocity delta V of the grabbed fragment ends of the spacecraft, and calculating whether the difference value V-delta V between the current combination flight velocity V and the relative linear velocity delta V meets the off-orbit requirement;
step 50, when the difference value V-delta V meets the off-orbit requirement, selecting the direction of the relative linear velocity delta V after the rotation of the grabbed fragment end to be parallel to the direction of the flying velocity V of the assembly, and releasing the grabbed fragment at the moment of opposite direction to finish momentum exchange; if not, step 30 is performed.
The momentum exchange is converted into rotational energy of the system and converted into translational energy, and space debris or a target obtains an off-orbit velocity increment, leaves the existing orbit and enters the atmosphere to be burnt.
In the embodiment 1 of the invention, the moment of rotation of the magnet in the earth space magnetic field is utilized, (for example, a compass can rotate in a continuous and accelerated manner when the direction and the size of the magnet are actively and effectively controlled), when the spacecraft grabs space debris or a target, the grabbed space debris rotates along with the magnet by continuously rotating the grabbed debris in an accelerated manner, and at the moment of loosening, the system rotation energy is converted into translation energy, and the space debris or the target obtain the off-orbit velocity increment, leave the existing orbit and enter the atmosphere.
The embodiment 1 of the invention can be delivered on-orbit without working medium consumption, has low cost and is off-orbit and orbital transfer, and the fact that working medium is not consumed means that long-term on-orbit flight and task development can be realized, and off-orbit operation of a large number of space fragments/targets is implemented, so that the problems of efficiency and economy of low-orbit space fragments are obviously improved.
The embodiment 1 of the invention also has the advantage of flexible in-plane/out-of-plane delivery, and the momentum used by the fragments to be out of orbit can be used for delivering the power-driven orbit change of the main satellite without working medium consumption. The method has high controllability and low dependence of delivery capacity on system scale, and utilizes the energy of the geomagnetic field, but has low requirement on scale, so that the implementation controllability is improved, and the probability of external threats is effectively reduced.
Example 2:
for embodiment 1 of the present invention, further, the basic expression of the rotation moment received by the magnet in the magnetic field is:
L＝m×B (4)
in the formula (4), L is a magnetic moment vector, m is a magnetic moment vector carried by the spacecraft, and B is a magnetic induction vector of the geospatial magnetic field.
For example: in the following fig. 3, the accumulated amounts of magnetic moments and magnetic moments around the X axis, the Y axis, and the Z axis are accumulated while accumulating energy around the Y axis of the orbit coordinate system in one orbit period, where the left, center, and right sides of fig. 3 respectively represent the magnetic moments of the X axis, the Y axis, and the Z axis, and the left, center, and right sides of fig. 3 respectively represent the accumulated amounts of magnetic moments of the X axis, the Y axis, and the Z axis.
Theoretically, only the Y-axis is accumulated, and the others are 0, and the figure corresponds to a dip-dip line, but in the actual situation, the formula (1) is used, the dip-IGRF line is used, the formula (3) is used, and the IGRF-IGRF line corresponds to the IGRF line, and in addition to the expected magnetic moment accumulation around the Y-axis, the magnetic moment accumulation with different degrees of stratification also occurs in the other two axes.
As shown in fig. 4, in one orbit period, while energy accumulation is performed around the Z axis, the accumulated amounts of magnetic moments around the X axis, the Y axis, and the Z axis are theoretically dipole lines, and the accumulated amounts of the X axis and the Y axis are almost 0, while the accumulated amounts of the IGRF lines are all significantly accumulated in the actual case.
In contrast, the problem of accumulation of magnetic moments in other directions is solved by adopting a delivery rail attitude coupling adjustment method.
As shown in fig. 5, a method for adjusting the off-track delivery track attitude coupling of space debris of a geomagnetic energy storage low track is provided, which comprises the following steps:
step 100, decomposing the whole geomagnetic energy storage low orbit period into an energy accumulation arc section and an energy release arc section, and configuring a momentum wheel energy absorption device on three axes of the spacecraft;
200, respectively determining an energy accumulation arc section and an energy release arc section when the spacecraft accumulates energy around an orbital plane y axis and an orbital plane z axis;
in a first aspect of this step 200, the energy concentrating arcs and the energy releasing arcs of the spacecraft around the orbital plane y-axis are symmetrically distributed with an ascending intersection angular distance u of 180 ° in one orbital cycle;
the energy release arc section is symmetrically expanded around the angular distance of the ascending intersection point of 0 degrees, 90 degrees, 180 degrees and 270 degrees and is set as delta u0、
Figure BDA0002174538280000101
Δuπ，
Figure BDA0002174538280000102
Corresponding to the values of the spread angular distances symmetrically for four angles, respectively, it is clear that Δ u0＝Δuπ，
Figure BDA0002174538280000103
Before the start of the orbit period, the three-axis rotation momentum under the corresponding orbit coordinate system is respectively obtained by measurement and still remains under a certain determined attitude of the spacecraft system
Figure BDA0002174538280000104
The storage capacity of the rotational momentum of the three axes is
Figure BDA0002174538280000105
And performing distribution, comprising the following sub-steps:
step 211, generating a magnetic moment through an international geomagnetic reference field IGRF geomagnetic model and a spacecraft, and performing simulation calculation by using an adjustment strategy of the following formula:
Figure BDA0002174538280000106
wherein, Cmd _ mx、Cmd_my、Cmd_mzRepresenting the desired adjusted magnetic moment magnitude component in each direction in the geomagnetic coordinate system, | m | representing the total magnetic moment magnitude that can be generated,
Figure BDA0002174538280000107
the method comprises the following steps that (1) in the on-orbit flight of the spacecraft, u represents the angular distance of a rising point of the spacecraft in the current flight at the current moment, wherein the geographic latitude corresponds to the subsatellite point of the spacecraft at the current moment;
in any quadrant area of the ascending intersection angle distance of 0-90 degrees, 90-180 degrees, 180-270 degrees and 270-360 degrees, the ascending intersection angle distance is a parabolic curve, so that a value exists, and in the 0-90 degree quadrant, the moment of the Y axis of the track coordinate system is in uaAnd
Figure BDA0002174538280000111
two points being equal and u within this intervalaAnd ubOne-to-one correspondence is realized;
step 212, in 0-90 degree quadrant, for arbitrary uaAnd u corresponding theretobCalculating the angular distance from u at the intersection point of ascentaTo
Figure BDA0002174538280000112
Cumulative amount on X and Z axes of internal orbital coordinate system
Figure BDA0002174538280000113
Step 213, determine whether or not
Figure BDA0002174538280000114
And at the same time
Figure BDA0002174538280000115
Figure BDA0002174538280000116
k is a predetermined conservative coefficient and is 0.5<k<0.95；
If not, then u is adjustedaAnd ubSo that the condition is satisfied.
Step 214, when the condition is satisfied, outputting uaAnd ubThen Δ u0＝ua，
Figure BDA0002174538280000117
Thereby obtaining the arc segment of energy release in the period of the orbit as
Figure BDA0002174538280000118
The rest are energy storage arc sections;
in the energy accumulating arc segment, if the stored X-axis rotation momentum at the current moment
Figure BDA0002174538280000119
And stored Z-axis rotational momentum
Figure BDA00021745382800001110
Appear
Figure BDA00021745382800001111
Or
Figure BDA00021745382800001112
The energy releasing arc segment is immediately switched to, and the current rising point angular distance u is measurednowJudgment unowIn the quadrant and replace ubAnd updating u according to the calculated valueaThereby changing and updating the range of the energy storage arc section/energy release arc section in the track cycle;
in a second aspect of this step 200, the energy-gathering arcs and the energy-releasing arcs of the spacecraft around the orbital plane Z-axis are asymmetrically distributed within one orbital cycle with the angular distance of the lift-intersection as a standard;
before the start of the orbit period, the three-axis rotational momentum under the corresponding orbit coordinate system is respectively measured and obtained under a certain determined attitude remained in the spacecraft system
Figure BDA00021745382800001113
The storage capacity of the rotational momentum of the three axes is
Figure BDA00021745382800001114
The allocation is made according to the following sub-steps:
step 221, the initial condition is an energy storage arc segment, and an adjustment strategy of the following formula is used:
Figure BDA00021745382800001115
wherein, Cmd _ mx、Cmd_my、Cmd_mzThe method comprises the steps of representing the size component of the magnetic moment which is expected to be adjusted in each direction in a geomagnetic coordinate system, and representing the total producible magnetic moment size in m;
measuring stored X-axis rotational momentum at the present time
Figure BDA0002174538280000121
And stored Y-axis rotational momentum
Figure BDA0002174538280000122
Appear
Figure BDA0002174538280000123
Or
Figure BDA0002174538280000124
Then immediately switching to an energy releasing arc section;
step 222, the energy releasing arc segment continues to
Figure BDA0002174538280000125
And at the same time
Figure BDA0002174538280000126
Figure BDA0002174538280000127
Wherein tau is a preset energy storage coefficient, and is generally 0.05<τ<0.3。
And 300, after the energy-gathering arc section is determined, for energy-accumulating targets of Y axes or Z axes of different orbit coordinate systems, finishing the off-orbit operation of the space debris based on a geomagnetic energy-accumulating low-orbit space debris off-orbit control method.
And 400, after the energy release arc section is determined, unloading the momentum wheel energy absorption device on the three shafts of the spacecraft through the generated magnetic moment.
This embodiment is through setting up energy storage arc section and energy release arc section for the additional moment accumulation that causes because the complexity of earth magnetic distribution can in time release, the utilization of the earth magnetic distribution characteristics of the maximum efficiency simultaneously accomplishes high-efficient energy storage, makes the efficiency nature and the economic nature problem of low track space piece obviously improve.
Although the invention has been described in detail above with reference to a general description and specific examples, it will be apparent to one skilled in the art that modifications or improvements may be made thereto based on the invention. Accordingly, such modifications and improvements are intended to be within the scope of the invention as claimed.

Claims (8)

1. A geomagnetic energy storage low-orbit space debris off-orbit delivery orbit attitude coupling adjustment method is characterized by comprising the following steps of:
step 100, decomposing the whole geomagnetic energy storage low orbit period into an energy accumulation arc section and an energy release arc section, and configuring a momentum wheel energy absorption device on three axes of the spacecraft;
200, respectively determining an energy accumulation arc section and an energy release arc section when the spacecraft accumulates energy around an orbital plane y axis and an orbital plane z axis;
300, after the energy-gathering arc section is determined, for energy-accumulating targets of Y axes or Z axes of different orbit coordinate systems, finishing off-orbit operation of space debris based on a geomagnetic energy-accumulating low-orbit space debris off-orbit control method;
step 400, after the energy release arc section is determined, unloading the momentum wheel energy absorption device of the spacecraft triaxial by generating a magnetic moment;
the geomagnetic energy storage low-orbit space debris off-orbit control method comprises the following steps: snatch space piece through the spacecraft and carry out the earth magnetism energy storage, the earth magnetism energy storage includes following step:
step 10, determining an energy storage direction according to the orbit flight inclination angle of the spacecraft, performing energy storage accumulation around an orbit plane y axis when the orbit flight inclination angle of the spacecraft is larger than a set angle, and performing energy storage accumulation around an orbit plane z axis when the orbit flight inclination angle is smaller than the set angle;
step 20, measuring the current flight attitude of the spacecraft, adjusting the flight attitude of the spacecraft, associating the magnetic moment generating direction of an orthogonal coil fixedly connected to the spacecraft and the accumulated rotating distance direction of the magnetic moment with the flight attitude of the spacecraft, and keeping the associated flight attitude;
step 30, generating strong magnetic moment
Figure FDA0003074121500000011
Is a vector;
step 40, continuously accelerating and rotating the grabbed fragment ends, detecting the accumulated angular momentum value and the relative linear velocity delta V of the grabbed fragment ends of the spacecraft, and calculating whether the difference value V-delta V between the current combination flight velocity V and the relative linear velocity delta V meets the off-orbit requirement;
step 50, when the difference value V-delta V meets the off-orbit requirement, selecting the direction of the relative linear velocity delta V after the rotation of the grabbed fragment end to be parallel to the direction of the flying velocity V of the assembly, and releasing the grabbed fragment at the moment of opposite direction to finish momentum exchange; if not, step 30 is performed.
2. The method according to claim 1, wherein in step 200, the energy-gathering arc segments and the energy-releasing arc segments of the spacecraft around the orbital plane y-axis are symmetrically distributed with a lift-intersection angular distance u of 180 ° in one orbital period.
3. The off-track delivery rail attitude coupling adjustment method for the geomagnetic energy storage low-orbit space debris according to claim 1 or 2, wherein the energy release arc segments are symmetrically expanded around the angular distance of the ascending intersection point of 0 degrees, 90 degrees, 180 degrees and 270 degrees and are set to be delta u0、
Figure FDA0003074121500000021
Δuπ，
Figure FDA0003074121500000022
Corresponding to the values of the spread angular distances symmetrically for four angles, respectively, it is clear that Δ u0＝Δuπ，
Figure FDA0003074121500000023
4. The method for adjusting the coupling of the off-orbit delivery attitude of the fragments in the geomagnetic energy storage low-orbit space according to claim 3, wherein before the start of the orbit period, the rotation momentum of three axes corresponding to the orbit coordinate system is measured and obtained and remained under a certain determined attitude of the spacecraft system
Figure FDA0003074121500000024
And three axesRespectively, the storage capacity of the rotational momentum of
Figure FDA0003074121500000025
And distributed.
5. The off-track delivery rail attitude coupling adjustment method for the space debris of the geomagnetic energy storage low track according to claim 4, wherein the distribution method comprises the following steps:
step 211, generating a magnetic moment through an international geomagnetic reference field IGRF geomagnetic model and a spacecraft, and performing simulation calculation by using an adjustment strategy of the following formula:
Figure FDA0003074121500000026
wherein, Cmd _ mx、Cmd_my、Cmd_mzRepresenting the desired adjusted magnetic moment magnitude component in each direction in the geomagnetic coordinate system, | m | representing the total magnetic moment magnitude that can be generated,
Figure FDA0003074121500000027
the method comprises the following steps that (1) in the on-orbit flight of the spacecraft, u represents the angular distance of a rising point of the spacecraft in the current flight at the current moment, wherein the geographic latitude corresponds to the subsatellite point of the spacecraft at the current moment;
in any quadrant area of the ascending intersection angle distance of 0-90 degrees, 90-180 degrees, 180-270 degrees and 270-360 degrees, the ascending intersection angle distance is a parabolic curve, so that a value exists, and in the 0-90 degree quadrant, the moment of the Y axis of the track coordinate system is in uaAnd 90 ° -ubTwo points being equal and u within this intervalaAnd ubOne-to-one correspondence is realized;
step 212, in 0-90 degree quadrant, for arbitrary uaAnd u corresponding theretobCalculating the angular distance from u at the intersection point of ascentaTo 90-ubCumulative amount on X and Z axes of internal orbital coordinate system
Figure FDA0003074121500000031
Step 213, determine whether or not
Figure FDA0003074121500000032
And at the same time
Figure FDA0003074121500000033
Figure FDA0003074121500000034
k is a predetermined conservative coefficient and is 0.5<k<0.95；
If not, then u is adjustedaAnd ubSuch that the condition is satisfied;
step 214, when the condition is satisfied, outputting uaAnd ubThen Δ u0＝ua，
Figure FDA0003074121500000035
Thereby obtaining the arc segment of energy release in the period of the orbit as
Figure FDA0003074121500000036
The rest are energy storage arc sections.
6. The method of claim 5, wherein in the energy storage arc segment, if the stored X-axis rotational momentum at the current moment is present, the method comprises
Figure FDA0003074121500000037
And stored Z-axis rotational momentum
Figure FDA0003074121500000038
Appear
Figure FDA0003074121500000039
Figure FDA00030741215000000310
Or
Figure FDA00030741215000000311
The energy releasing arc segment is immediately switched to, and the current rising point angular distance u is measurednowJudgment unowIn the quadrant and replace ubAnd updating u according to the calculated valueaThereby changing and updating the range of the energy storage arc segment/energy release arc segment in the orbit period.
7. The method according to claim 1, wherein in step 200, the energy-gathering arc segments and the energy-releasing arc segments of the spacecraft around the orbital plane Z axis are asymmetrically distributed with the angular distance of the lift-intersection point in one orbital period.
8. The method according to claim 7, wherein before the start of the orbit cycle, the rotational momentum of three axes in the corresponding orbital coordinate system is measured to obtain the rotational momentum of the remaining space debris in a certain attitude of the spacecraft system, and the rotational momentum is respectively
Figure FDA0003074121500000041
Figure FDA0003074121500000042
The storage capacity of the rotational momentum of the three axes is
Figure FDA0003074121500000043
The allocation is made according to the following sub-steps:
step 221, the initial condition is an energy storage arc segment, and an adjustment strategy of the following formula is used:
Figure FDA0003074121500000044
wherein, Cmd \umx、Cmd_my、Cmd_mzThe method comprises the steps of representing the size component of the magnetic moment which is expected to be adjusted in each direction in a geomagnetic coordinate system, and representing the total producible magnetic moment size in m;
measuring stored X-axis rotational momentum at the present time
Figure FDA0003074121500000045
And stored Y-axis rotational momentum
Figure FDA0003074121500000046
Appear
Figure FDA0003074121500000047
Or
Figure FDA0003074121500000048
Then immediately switching to an energy releasing arc section;
step 222, the energy releasing arc segment continues to
Figure FDA0003074121500000049
And at the same time
Figure FDA00030741215000000410
Figure FDA00030741215000000411
Wherein tau is a preset energy storage coefficient, and is generally 0.05<τ<0.3。
CN201910774236.3A 2019-08-21 2019-08-21 Off-orbit delivery rail attitude coupling adjustment method for geomagnetic energy storage low-orbit space debris Active CN110510154B (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN201910774236.3A CN110510154B (en) 2019-08-21 2019-08-21 Off-orbit delivery rail attitude coupling adjustment method for geomagnetic energy storage low-orbit space debris

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN201910774236.3A CN110510154B (en) 2019-08-21 2019-08-21 Off-orbit delivery rail attitude coupling adjustment method for geomagnetic energy storage low-orbit space debris

Publications (2)

Publication Number Publication Date
CN110510154A CN110510154A (en) 2019-11-29
CN110510154B true CN110510154B (en) 2021-07-02

Family

ID=68626948

Family Applications (1)

Application Number Title Priority Date Filing Date
CN201910774236.3A Active CN110510154B (en) 2019-08-21 2019-08-21 Off-orbit delivery rail attitude coupling adjustment method for geomagnetic energy storage low-orbit space debris

Country Status (1)

Country Link
CN (1) CN110510154B (en)

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN111114852B (en) * 2019-12-09 2021-04-23 上海航天控制技术研究所 Space rolling target four-dimensional perception capturing device and method
CN113075592B (en) * 2021-03-12 2022-05-24 中国科学院力学研究所 Coaxial contra-rotating geomagnetic energy storage and release delivery ground experiment system and method

Family Cites Families (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS491439B1 (en) * 1968-09-27 1974-01-14
US4114841A (en) * 1977-02-22 1978-09-19 Rca Corporation Magnetic torquing system for changing the spin rate of an orbiting satellite
US4489383A (en) * 1981-10-16 1984-12-18 Rca Corporation Closed-loop magnetic roll/yaw control system for high inclination orbit satellites
US6942186B1 (en) * 2001-03-07 2005-09-13 Star Technology And Research, Inc. Method and apparatus for propulsion and power generation using spinning electrodynamic tethers
JP5755836B2 (en) * 2009-12-04 2015-07-29 株式会社Ihi Space debris removal method
CN101767657A (en) * 2009-12-30 2010-07-07 哈尔滨工业大学 Electro-dynamic tether based satellite deorbit device and method thereof
GB2482481B (en) * 2010-08-02 2012-06-13 Lockheed Martin Uk Insys Ltd Apparatus, method and system
JP6473960B2 (en) * 2014-06-13 2019-02-27 国立研究開発法人宇宙航空研究開発機構 Space debris orbital descent method, orbital descent system, and satellite orbit conversion method, orbit conversion system
CN105966643B (en) * 2016-05-20 2018-04-17 中国西安卫星测控中心 A kind of space junk low cost sweep-out method using ground electromagnetic launcher
CN107719708B (en) * 2017-09-27 2020-09-08 西北工业大学深圳研究院 Space debris capturing and removing device and method
CN109250158B (en) * 2018-08-02 2020-04-14 西北工业大学 Multi-space robot-based space failure target despinning and capturing integrated device and method
CN109799835B (en) * 2019-01-17 2020-10-09 北京理工大学 Optimal derailing method for rope system towing of space debris
CN109969433B (en) * 2019-04-01 2020-06-19 中国人民解放军国防科技大学 Space debris batch removing system based on low-earth orbit satellite

Also Published As

Publication number Publication date
CN110510154A (en) 2019-11-29

Similar Documents

Publication Publication Date Title
CN110481815B (en) Transmission type contra-rotating geomagnetic energy storage-release delivery system and method
CN110160407A (en) A kind of carrier rocket grade is settled in an area scope control system
CN112512924B (en) Removal of orbital space debris from near-earth orbit
Hoyt et al. Cislunar tether transport system
CN110510152B (en) Low-orbit geomagnetic energy storage-release delivery system
CN110510154B (en) Off-orbit delivery rail attitude coupling adjustment method for geomagnetic energy storage low-orbit space debris
CN103778823A (en) Suspension device and microgravity experimental method applied to interior of space capsule
US8336826B2 (en) Propulsion systems and methods utilizing smart propellant
CN109969433A (en) Space debris batch removing system based on low-earth orbit satellite
CN105511493B (en) A kind of low rail constellation dispositions method based on martian atmosphere auxiliary
CN110510153B (en) Off-orbit control method for geomagnetic energy storage low-orbit space debris
CN110510157B (en) Low-orbit geomagnetic energy storage ground experiment system and method
CN109583041A (en) A kind of Craft Orbit design method
Zagórski Modeling disturbances influencing an Earth-orbiting satellite
Hou et al. Electromagnetic-launch-based method for cost-efficient space debris removal
Smitherman Jr et al. Space elevators
Williams et al. Magnetospheric multiscale mission attitude dynamics: Observations from flight data
Nan et al. Global 4-D trajectory optimization for spacecraft
Ward et al. Orbital docking dynamics
Williams et al. Magnetospheric Multiscale mission micrometeoroid/orbital debris impacts
Bolonkin Innovations and New Technologies (v. 2)
CN113608539B (en) Space rolling target non-contact racemization method based on double-star electromagnetic formation satellite
Hewen et al. Space Environment Modeling and Deorbiting Efficiency Analysis for Electro-dynamic Tether System
CN108959827A (en) The design method of polar region suspension railway based on electronic sail
CN111776249A (en) Lunar surface low-altitude aircraft taking lunar surface soil as projection working medium

Legal Events

Date Code Title Description
PB01 Publication
PB01 Publication
SE01 Entry into force of request for substantive examination
SE01 Entry into force of request for substantive examination
GR01 Patent grant
GR01 Patent grant