CN116573170A - Spacecraft for sensing space debris - Google Patents
Spacecraft for sensing space debris Download PDFInfo
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- CN116573170A CN116573170A CN202310549338.1A CN202310549338A CN116573170A CN 116573170 A CN116573170 A CN 116573170A CN 202310549338 A CN202310549338 A CN 202310549338A CN 116573170 A CN116573170 A CN 116573170A
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- 230000003116 impacting effect Effects 0.000 claims abstract description 12
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Classifications
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- 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
- B64G1/68—Arrangements or adaptations of apparatus or instruments, not otherwise provided for of meteoroid or space debris detectors
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/10—Artificial satellites; Systems of such satellites; Interplanetary vehicles
- B64G1/12—Artificial satellites; Systems of such satellites; Interplanetary vehicles manned
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- 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
- B64G1/222—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles for deploying structures between a stowed and deployed state
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02E—REDUCTION OF GREENHOUSE GAS [GHG] EMISSIONS, RELATED TO ENERGY GENERATION, TRANSMISSION OR DISTRIBUTION
- Y02E10/00—Energy generation through renewable energy sources
- Y02E10/50—Photovoltaic [PV] energy
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- Geophysics And Detection Of Objects (AREA)
Abstract
The embodiment of the application discloses a spacecraft for sensing space debris, which comprises the following components: a bladder; a gas bottle in communication with the bladder; the spacecraft further comprises: an unfolded state and a folded state; when the spacecraft is in a folded state, the spacecraft also comprises a protective shell for wrapping the bag body; when the spacecraft is launched to a preset orbit, the gas bottle is opened to inflate the bag body, so that the spacecraft is converted from a folded state to an unfolded state; the spacecraft further comprises: the sensing structure is attached to the outer side wall of the bag body; a signal modem electrically connected to the sensing structure; when the space debris impacts the spacecraft, the signal modulation and demodulation device obtains the position, the force and the area of the space debris impacting the spacecraft through the sensing structure. The spacecraft provided by the application can acquire the collision frequency, the collision position distribution, the collision area, the damage degree and the like of the manned spacecraft or other spacecraft in different orbits when being impacted by space debris.
Description
Technical Field
The application relates to the technical field of spacecrafts, in particular to a spacecraft for sensing space debris.
Background
Space debris, also known as space debris, refers mainly to waste spacecraft debris left in space during aerospace activities by humans and debris generated by explosions or collisions. With the increasing frequency of human spatial activity, the space debris environment is increasingly worsening. The space monitoring net estimates that there are about 2 to 2.2 tens of thousands of space fragments greater than l0cm in size, about 50 tens of thousands of space fragments greater than 1cm in size, and greater than 1 hundred million space fragments greater than 1mm in size. The space debris weighs about 6300 tons, with about 2700 tons distributed over the low earth orbit. For the next 50 years, the number of fragments will increase at a rate of 10% per year.
For manned spacecraft, the flight time is long, the volume is large, and the risk of encountering space debris impact is highest. If the space debris breaks down the bulkhead of the sealed cabin, the gas in the sealed cabin is leaked, and the life safety of the spacecraft and astronauts is seriously threatened.
At present, the damaged spacecraft due to space debris impact is in an in-orbit state, effective perception and evaluation cannot be performed in a measuring mode and the like, related data can be observed only through limited optical imaging equipment carried by satellites, and both the observation means and the measurement accuracy are limited. In the ultra-high-speed impact test of the ground, the variety and kinetic energy characteristics of a plurality of space fragments cannot be effectively and truly simulated.
Accordingly, to overcome the deficiencies of the prior art, it is desirable to provide a spacecraft for sensing space debris.
Disclosure of Invention
The application aims to provide a spacecraft for sensing space debris, so as to solve the problem that the position, strength, area and damage degree of the spacecraft in an on-orbit state, which are impacted by the space debris, cannot be evaluated.
In order to achieve at least one of the above objects, the present application adopts the following technical scheme:
the present application provides a spacecraft comprising: a bladder;
a gas bottle in communication with the bladder;
the spacecraft further comprises: an unfolded state and a folded state;
before the spacecraft is launched to a preset orbit, the spacecraft is in a folded state;
when the spacecraft is launched to a preset orbit, the gas bottle is opened to inflate the bag body, so that the spacecraft is converted from a folded state to an unfolded state;
the spacecraft further comprises: the sensing structure is attached to the outer side wall of the bag body;
a signal modem electrically connected to the sensing structure;
when the space debris impacts the spacecraft, the signal modulation and demodulation device obtains the position, the force and the area of the space debris impacting the spacecraft through the sensing structure.
Optionally, the capsule comprises: a first airtight film layer;
a self-stiffening layer attached to one side surface of the first airtight layer;
a second airtight film layer attached to a side surface of the self-stiffening layer, which is far from the first airtight film layer;
the sensing structure is attached to the surface of one side, far away from the self-stiffening layer, of the second airtight film layer.
Optionally, the spacecraft further comprises: and the solar cell panel is fixedly connected with one end of the bag body and used for providing electric energy for the spacecraft.
Optionally, the spacecraft further comprises: the power assembly is arranged at one end part of the solar panel, which is far away from the bag body;
the power assembly is used for providing kinetic energy for replacing an operation orbit and an operation posture of the spacecraft.
Optionally, the spacecraft further comprises: a fixed shell which is fixedly connected with one end of the solar panel far away from the capsule body and is provided with an inner cavity;
the power assembly is mounted on the stationary housing.
Optionally, the power assembly includes:
the orbit control engine is arranged at the center of the fixed shell and used for enabling the spacecraft to replace an operation orbit;
and the plurality of attitude control engines are uniformly arranged around the periphery of the rail control engine and are arranged on the fixed shell so as to enable the spacecraft to change the running attitude.
Optionally, the track-controlled engine includes: the first power part is fixed in the inner cavity of the shell; and a first air injection part penetrating to the outer side of the bottom wall of the shell;
the attitude control engine comprises: the second power part is fixed in the inner cavity of the shell; and a second air injection part penetrating to the outer side of the bottom wall of the shell.
Optionally, the bottom wall of the fixing case includes a plurality of assembly holes for cooperating with the gas bottle.
Optionally, the sensing structure is formed by weaving a plurality of fiber lines and fiber fabrics; the optical fiber lines are in a grid structure;
the plurality of fiber optic strands includes: a plurality of equidistant transverse optical fiber lines and a plurality of equidistant longitudinal optical fiber lines which are transversely arranged;
the grid structure is formed by crisscross braiding of a plurality of transverse optical fiber lines and a plurality of longitudinal optical fiber lines;
the signal modulation and demodulation device is respectively and electrically connected with the transverse optical fiber lines and the longitudinal optical fiber lines so as to acquire the position, the force and the area of space debris impacting the spacecraft through optical signals generated by the transverse optical fiber lines and the longitudinal optical fiber lines.
Optionally, the self-hardening layer is made of an aluminum alloy film, and the thickness of the self-hardening layer is 0.01mm to 0.1mm;
the first airtight membrane layer and the second airtight membrane layer are made of polyimide, the thickness of the first airtight membrane layer is 0.1mm-0.2mm, and the thickness of the second airtight membrane layer is 0.1mm-0.2mm
The beneficial effects of the application are as follows:
aiming at the problems existing in the prior art, the application provides the spacecraft for sensing the space debris, the manufacturing cost of the whole spacecraft is greatly reduced through the use of the capsule body, and the whole quality of the spacecraft is also reduced; the folded spacecraft is in a folded state during launching, the volume of the folded spacecraft is less than 10% of that of the folded spacecraft, the launching difficulty is reduced, and the transportation cost of the carrier rocket is also reduced; after reaching the preset orbit, the bag body is inflated through the inflation bottle, so that the bag body is converted into an unfolding state from a folding state, namely, the spacecraft is in the unfolding state, and the manned spacecraft or other spacecraft in the on-orbit state is simulated; when space debris impacts the capsule body, namely, impacts the sensing structure, the signal modulation and demodulation device obtains the position, the force and the area of the space debris impacting the capsule body through the sensing structure, namely, obtains the position, the force and the area of the space debris impacting the spacecraft, the kinetic energy of the space debris can be calculated through signals transmitted through the sensing structure obtained by the signal modulation and demodulation device after impacting, and the calculated effective data are sent to the ground through the telemetry antenna for analysis of scientific researchers. The scientific researchers analyze the damage degree and service life evaluation of the manned spacecraft or other spacecraft through the manned spacecraft or other spacecraft simulated by the capsule body, and guide the optimization of the structural design of the manned spacecraft or other spacecraft. The spacecraft provided by the application can realize the effects that the positioning accuracy is within 5mm and the damage area sensing error is not more than 10% by using the sensing structure; the appearance of various manned spacecrafts or other spacecrafts can be simulated, and the collision frequency, the collision position distribution, the collision area, the damage degree and the like of the manned spacecrafts or other spacecrafts in different orbits can be obtained; moreover, the spacecraft provided by the application has long service life.
Drawings
The following describes the embodiments of the present application in further detail with reference to the drawings.
Fig. 1 shows a schematic overall structure of a spacecraft for sensing space debris in a folded state and with a protective shell according to an embodiment of the application.
Fig. 2 shows a side view of a spacecraft for sensing space debris in a folded state and with a protective shell in an embodiment of the application.
Fig. 3 shows an exploded view of a spacecraft for sensing space debris in a folded state in an embodiment of the application.
Fig. 4 shows a schematic structural view of a spacecraft for sensing space debris in a folded state without a protective shell according to an embodiment of the application.
Fig. 5 shows a schematic structural view of a spacecraft for sensing space debris in an embodiment of the application in an expanded state and with a balloon in a spherical shape.
Fig. 6 shows a schematic structural view of a spacecraft for sensing space debris in an embodiment of the application in an expanded state and with a cylindrical balloon.
Fig. 7 shows a schematic structural view of a spacecraft for sensing space debris in an expanded state and with a capsule in a cube, according to an embodiment of the application.
Fig. 8 shows a schematic structural view of a spacecraft for sensing space debris in an expanded state and with a prismatic table shape of a balloon in an embodiment of the application.
Fig. 9 shows a schematic structural view of a sensing structure located at an outer surface of a spherical capsule in a spacecraft for sensing space debris in an embodiment of the application.
Fig. 10 shows a schematic structural diagram of a sensing structure used in conjunction with a signal modem device in a spacecraft for sensing space debris in an embodiment of the application.
FIG. 11 illustrates a schematic of the positional relationship of a first air-tight membrane layer, a self-stiffening layer, a second air-tight membrane layer, and a sensing structure of a capsule in a spacecraft for sensing space debris in an embodiment of the application.
Detailed Description
In the following description, for purposes of explanation, numerous specific details are set forth in order to provide a thorough understanding of one or more embodiments. It may be evident, however, that such embodiment(s) may be practiced without these specific details.
In the description of the present application, it should be noted that the azimuth or positional relationship indicated by the terms "upper", "lower", etc. are based on the azimuth or positional relationship shown in the drawings, and are merely for convenience of describing the present application and simplifying the description, and are not indicative or implying that the apparatus or element in question must have a specific azimuth, be constructed and operated in a specific azimuth, and thus should not be construed as limiting the present application. Unless specifically stated or limited otherwise, the terms "mounted," "connected," and "coupled" are to be construed broadly, and may be, for example, fixedly connected, detachably connected, or integrally connected; can be mechanically or electrically connected; can be directly connected or indirectly connected through an intermediate medium, and can be communication between two elements. The specific meaning of the above terms in the present application can be understood by those of ordinary skill in the art according to the specific circumstances.
It is further noted that in the description of the present application, relational terms such as first and second, and the like are 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. Moreover, 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. Without further limitation, an element defined by the phrase "comprising one … …" does not exclude the presence of other like elements in a process, method, article, or apparatus that comprises the element.
To solve the problems of the prior art, one embodiment of the present application provides a spacecraft for sensing space debris, as shown in fig. 1-11, comprising: a capsule body 1; a gas bottle 2 communicating with the capsule 1; the spacecraft further comprises: an unfolded state and a folded state; before the spacecraft is launched to a preset orbit, the spacecraft is in a folded state; when the spacecraft is in a folded state, the spacecraft further comprises a protective shell 3 for wrapping the bag body 1; when the spacecraft reaches a preset orbit, the gas bottle 2 is opened to inflate the bag body 1, so that the spacecraft is converted from a folded state to an unfolded state; the shape of the bag body 1 in the unfolding state can be adaptively customized according to the test requirement, so that the space debris test requirement of various spacecrafts is met, for example: as shown in figures 5-8, the spherical shape, the cylindrical shape, the square body and the prismatic platform shape can be adopted, the spherical diameter is 1000mm-12000mm, the cylindrical body diameter is 1000mm-2000mm, the length is 3000mm-25000mm, the prismatic platform shape length is 1000mm-25000mm, the side length of the bottom surface is 1000mm-6000mm, and the side length of the square body is 10000mm-16000mm; the spacecraft further comprises: the sensing structure 4 is attached to the outer side wall of the bag body 1, and the sensing structure 4 is a flexible sensing structure; a signal modulation/demodulation device 5 electrically connected to the sensing structure 4; when the space debris impacts the spacecraft, the signal modulation and demodulation device 5 acquires the position, the force and the area of the space debris impacting the spacecraft through the sensing structure 4.
In the embodiment of the application, the use of the capsule body 1 greatly reduces the manufacturing cost of the whole spacecraft and the whole quality of the spacecraft; the folded spacecraft is in a folded state during launching, the volume of the folded spacecraft is less than 10% of that of the folded spacecraft, the launching difficulty is reduced, and the transportation cost of the carrier rocket is also reduced; after reaching the preset orbit, the bag body 1 is inflated through the inflation bottle, so that the bag body 1 is converted into an unfolding state from a folding state, namely, the spacecraft is in the unfolding state, and the manned spacecraft or other spacecraft in the on-orbit state is simulated; when space debris impacts the capsule body 1, namely, impacts the sensing structure 4, the signal modulation and demodulation device 5 obtains the position, the force and the area of the space debris impacting the capsule body 1 through the sensing structure 4, namely, obtains the position, the force and the area of the space debris impacting the spacecraft, the kinetic energy of the space debris can be calculated through the signals transmitted by the sensing structure 4 obtained through the signal modulation and demodulation device 5 after impacting, and the calculated effective data are sent to the ground through the telemetry antenna for analysis by scientific researchers. The scientific research personnel analyze the damage degree and service life evaluation of the manned spacecraft or other spacecraft through the manned spacecraft or other spacecraft simulated by the capsule body 1, and guide the optimization of the structural design of the manned spacecraft or other spacecraft. The spacecraft provided by the application can realize the effects that the positioning accuracy is within 5mm and the damage area sensing error is not more than 10% by using the sensing structure 4; the appearance of various manned spacecrafts or other spacecrafts can be simulated, and the collision frequency, the collision position distribution, the collision area, the damage degree and the like of the manned spacecrafts or other spacecrafts in different orbits can be obtained; moreover, the spacecraft provided by the application has long service life.
In a specific embodiment, the capsule 1 comprises: a first airtight film layer 11; a self-stiffening layer 12 attached to one side surface of the first airtight layer 11; a second airtight layer 13 attached to a side surface of the self-stiffening layer 12 remote from the first airtight layer 11; the sensing structure 4 is attached to a surface of the second airtight layer 13, which is far away from the self-stiffening layer 12. The first airtight membrane layer 11, the self-stiffening layer 12 and the second airtight membrane layer 13 form a membrane layer structure of the capsule body 1, and the first airtight membrane layer 11, the self-stiffening layer 12, the second airtight membrane layer 13 and the sensing structure 4 are sequentially connected from inside to outside through adhesive, and the total thickness of the whole is 3mm-5mm; the adhesive is RTV silicone rubber adhesive solidified at room temperature; the self-stiffening layer is made of an aluminum alloy film, and the thickness of the self-stiffening layer is 0.01mm to 0.1mm; the first airtight membrane layer 11 and the second airtight membrane layer 13 are made of polyimide, the thickness of the first airtight membrane layer 11 is 0.1mm-0.2mm, and the thickness of the second airtight membrane layer 13 is 0.1mm-0.2mm.
The self-stiffening layer 12 can realize the effect of autonomous stiffening of the capsule body 1, and the rigidity and modulus of the capsule body 1 after self stiffening are equivalent to those of the capsule body structure of the manned spacecraft or other spacecraft, so that the capsule body has long-term dimension, can effectively simulate the capsule body structure of the manned spacecraft or other spacecraft to be impacted, and simulate the rigidity and modulus of the manned spacecraft or other spacecraft; because the spacecraft is in a vacuum and weightless state in space, the spacecraft provided by the application can not generate shape change due to air leakage after being impacted by space fragments, and can be used for long-term on-orbit test.
The leak rate of the capsule body 1 is better than 3.0 Pa.L/s, the closed-loop control of the pressure is carried out by a pressure sensor, and the measurement error of the pressure sensor (-35 ℃ to 180 ℃) is +/-6%FS.
In a specific embodiment, the sensing structure 4 is formed by weaving a plurality of fiber optic threads and a fiber fabric 43; the fiber fabric 43 can be aramid fiber, and a plurality of fiber lines and the fiber fabric 43 are uncharged, so that explosion in the in-orbit impact test process can be avoided; the optical fiber lines are in a grid structure; the plurality of fiber optic strands includes: a plurality of equidistant laterally aligned lateral fiber optic strands 41 and a plurality of equidistant longitudinally aligned longitudinal fiber optic strands 42; the grid structure is formed by crisscross braiding a plurality of transverse optical fiber lines 41 and a plurality of longitudinal optical fiber lines 42; the signal modem device 5 is electrically connected to the plurality of transverse optical fiber lines 41 and the plurality of longitudinal optical fiber lines 42, respectively, so as to obtain the position, the force and the area of the space debris striking the spacecraft by the optical signals generated by the plurality of transverse optical fiber lines 41 and the plurality of longitudinal optical fiber lines 42. The distance between two adjacent transverse optical fiber lines 41 is 2cm to 5cm, and the specific size is set according to the actual situation, for example, may be 3cm; the distance between two adjacent longitudinal fiber lines 42 is 2cm to 5cm, and the specific dimension is set according to practical situations, for example, may be 3cm. Because the plurality of transverse optical fiber lines 41 and the plurality of longitudinal optical fiber lines 42 are formed by crisscross braiding, the design of the size greatly improves the perception precision of the optical fiber lines when space debris impacts the spacecraft.
Specifically, each transverse fiber optic line 41 and each longitudinal fiber optic line 42 are numbered therefrom; when the millimeter-sized and above space debris collides with the capsule body 1 in the unfolded state, the space debris can completely penetrate into the inner cavity of the capsule body 1 to form a hole, at least one transverse optical fiber line 41 and one longitudinal optical fiber line 42 can be broken by the hole formed by collision, the position of the collision point is calculated by the modem according to the serial numbers of the broken transverse optical fiber lines 41 and the longitudinal optical fiber lines 42, the area of the hole generated by collision is calculated according to the number of broken optical fibers, and the kinetic energy of the space debris can be calculated through the posture change of the spacecraft recorded by the optical fiber inertial unit after collision.
The sensing structure 4 used by the spacecraft provided by the application is verified in a ground test, so that the positioning accuracy is better than 5mm, the damage area and the sensing error are not more than 10%; the verification can be performed through a secondary light air gun during the verification in the ground test.
In a specific embodiment, the spacecraft further comprises: and the solar panel 6 is fixedly connected with one end of the bag body 1 and is used for providing electric energy for the spacecraft. The solar panel 6 provides electric energy for the components on the whole spacecraft, so that the spacecraft provided by the application can stably run in a preset orbit.
In a specific embodiment, the spacecraft further comprises: the power assembly is arranged at one end part of the solar panel 6 far away from the bag body 1; the power assembly is used for providing kinetic energy for replacing an operation orbit and an operation posture of the spacecraft. The solar panel 6 provides electric energy for the power assembly, and when the spacecraft needs to change the running orbit or the running gesture, the power assembly provides kinetic energy for the spacecraft, so that the orbit or the gesture of the spacecraft is changed.
Specifically, the spacecraft further comprises: a fixed shell 9 with an inner cavity, which is fixedly connected with one end of the solar panel 6 far away from the capsule body 1; the power assembly is mounted on the stationary housing 9. The power assembly includes: a rail engine installed at the center of the fixed housing 9 for replacing the spacecraft with a running orbit; the plurality of attitude control engines, which are arranged on the fixed shell 9 and uniformly surround the periphery of the orbit control engine, are used for enabling the spacecraft to change the running attitude, and are preferably four. The spacecraft further comprises a controller for controlling the on or off of the track control engine and the plurality of gesture control engines; when the spacecraft needs to change the running orbit, ground scientific researchers transmit an orbit transfer signal to a controller, and the controller controls an orbit control engine to start so as to enable the spacecraft to be changed into a preset orbit; when the running posture of the spacecraft needs to be replaced, the controller controls the posture control engine to start so as to enable the spacecraft to change the running posture, wherein the running posture of the spacecraft changes according to the direction of the sun, so that the solar cell panel 6 can absorb sunlight at any time, and sufficient electric energy is ensured.
Further, the rail-controlled engine includes: a first power section (not shown) secured within the housing interior; and a first air injection portion 72 penetrating to the outside of the bottom wall of the housing; when the controller receives the orbit transfer signal, the first power part is controlled to be opened so as to enable the first air injection part 72 to inject air, provide kinetic energy for space air and enable the spacecraft to realize orbit transfer; the attitude control engine comprises: a second power section 72 secured within the housing interior; and a second air injection part 82 penetrating to the outside of the bottom wall of the housing; when the operation posture of the spacecraft needs to be changed, the controller controls the second power part 72 to be opened so that the second air injection part 82 injects air to provide kinetic energy for space air, and the spacecraft realizes the change of the operation posture.
In one embodiment, the bottom wall of the fixing case 9 includes a plurality of assembly holes 10 for cooperating with the gas bottle 2. In this way, the mass of the whole spacecraft can be reduced, and the gas bottle 2 can be further fixed; as shown in fig. 1, the number of the gas bottles 2 is preferably four, and the gas bottles uniformly encircle the periphery of the track-controlled engine; in practical application, the gas bottle 2 can also provide gas for the rail control engine and the attitude control engine, so that the rail control engine and the attitude control engine can realize normal gas injection in practical use.
The service life of the spacecraft provided by the application is 5-6 years, the total weight is not more than 200kg, the weight of the manned spacecraft or other spacecraft is more than 1000kg, the weight of the manned spacecraft or other spacecraft is far less than that of the manned spacecraft or other spacecraft, the transportation quality of the carrier rocket is reduced, and the launching cost is saved; .
It should be understood that the foregoing examples of the present application are provided merely for clearly illustrating the present application and are not intended to limit the embodiments of the present application, and that various other changes and modifications may be made therein by one skilled in the art without departing from the spirit and scope of the present application as defined by the appended claims.
Claims (10)
1. A spacecraft for sensing space debris, comprising: a bladder;
a gas bottle in communication with the bladder;
the spacecraft further comprises: an unfolded state and a folded state;
before the spacecraft is launched to a preset orbit, the spacecraft is in a folded state;
when the spacecraft is launched to a preset orbit, the gas bottle is opened to inflate the bag body, so that the spacecraft is converted from a folded state to an unfolded state;
the spacecraft further comprises: the sensing structure is attached to the outer side wall of the bag body;
a signal modem electrically connected to the sensing structure;
when the space debris impacts the spacecraft, the signal modulation and demodulation device obtains the position, the force and the area of the space debris impacting the spacecraft through the sensing structure.
2. The spacecraft for sensing space debris of claim 1, wherein the bladder comprises: a first airtight film layer;
a self-stiffening layer attached to one side surface of the first airtight layer;
a second airtight film layer attached to a side surface of the self-stiffening layer, which is far from the first airtight film layer;
the sensing structure is attached to the surface of one side, far away from the self-stiffening layer, of the second airtight film layer.
3. The spacecraft for sensing space debris of claim 1, further comprising: and the solar cell panel is fixedly connected with one end of the bag body and used for providing electric energy for the spacecraft.
4. A spacecraft for sensing space debris according to claim 3, further comprising: the power assembly is arranged at one end part of the solar panel, which is far away from the bag body;
the power assembly is used for providing kinetic energy for replacing an operation orbit and an operation posture of the spacecraft.
5. The spacecraft for sensing space debris of claim 4, further comprising: a fixed shell which is fixedly connected with one end of the solar panel far away from the capsule body and is provided with an inner cavity;
the power assembly is mounted on the stationary housing.
6. The spacecraft for sensing space debris of claim 5, wherein said power assembly comprises:
the orbit control engine is arranged at the center of the fixed shell and used for enabling the spacecraft to replace an operation orbit;
and the plurality of attitude control engines are uniformly arranged around the periphery of the rail control engine and are arranged on the fixed shell so as to enable the spacecraft to change the running attitude.
7. The spacecraft for sensing space debris of claim 6, wherein said orbit control engine comprises: the first power part is fixed in the inner cavity of the shell; and a first air injection part penetrating to the outer side of the bottom wall of the shell;
the attitude control engine comprises: the second power part is fixed in the inner cavity of the shell; and a second air injection part penetrating to the outer side of the bottom wall of the shell.
8. The spacecraft for sensing space debris of claim 5, wherein the bottom wall of the containment vessel includes a plurality of mounting holes for use with the gas bottle.
9. The spacecraft for sensing space debris of claim 1, wherein said sensing structure is formed of a plurality of fiber optic strands and a fabric weave; the optical fiber lines are in a grid structure;
the plurality of fiber optic strands includes: a plurality of equidistant transverse optical fiber lines and a plurality of equidistant longitudinal optical fiber lines which are transversely arranged;
the grid structure is formed by crisscross braiding of a plurality of transverse optical fiber lines and a plurality of longitudinal optical fiber lines;
the signal modulation and demodulation device is respectively and electrically connected with the transverse optical fiber lines and the longitudinal optical fiber lines so as to acquire the position, the force and the area of space debris impacting the spacecraft through optical signals generated by the transverse optical fiber lines and the longitudinal optical fiber lines.
10. The spacecraft for sensing space debris according to claim 2, wherein the self-stiffening layer is made of an aluminum alloy film with a thickness of 0.01mm to 0.1mm;
the first airtight membrane layer and the second airtight membrane layer are made of polyimide, the thickness of the first airtight membrane layer is 0.1mm-0.2mm, and the thickness of the second airtight membrane layer is 0.1mm-0.2mm.
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| Publication number | Priority date | Publication date | Assignee | Title |
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| CN117302552A (en) * | 2023-10-16 | 2023-12-29 | 中国科学院力学研究所 | Modularized extraterrestrial space capsule structure with high folding and unfolding ratio |
| CN117302552B (en) * | 2023-10-16 | 2024-06-07 | 中国科学院力学研究所 | Modularized extraterrestrial space capsule structure with high folding and unfolding ratio |
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Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
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| CN117302552A (en) * | 2023-10-16 | 2023-12-29 | 中国科学院力学研究所 | Modularized extraterrestrial space capsule structure with high folding and unfolding ratio |
| CN117302552B (en) * | 2023-10-16 | 2024-06-07 | 中国科学院力学研究所 | Modularized extraterrestrial space capsule structure with high folding and unfolding ratio |
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