CN114524111B - Spacecraft vibration suppression structure and method based on piezoelectric composite material - Google Patents
Spacecraft vibration suppression structure and method based on piezoelectric composite material Download PDFInfo
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- CN114524111B CN114524111B CN202111597006.8A CN202111597006A CN114524111B CN 114524111 B CN114524111 B CN 114524111B CN 202111597006 A CN202111597006 A CN 202111597006A CN 114524111 B CN114524111 B CN 114524111B
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- 239000002131 composite material Substances 0.000 title claims abstract description 28
- 230000001629 suppression Effects 0.000 title claims abstract description 23
- 238000000034 method Methods 0.000 title claims abstract description 9
- 239000000919 ceramic Substances 0.000 claims abstract description 85
- 238000005452 bending Methods 0.000 claims abstract description 44
- 230000010287 polarization Effects 0.000 claims abstract description 30
- 230000000694 effects Effects 0.000 claims abstract description 13
- 239000000463 material Substances 0.000 claims abstract description 10
- 229920000049 Carbon (fiber) Polymers 0.000 claims abstract description 8
- 239000004917 carbon fiber Substances 0.000 claims abstract description 8
- VNWKTOKETHGBQD-UHFFFAOYSA-N methane Chemical compound C VNWKTOKETHGBQD-UHFFFAOYSA-N 0.000 claims abstract description 8
- 230000005684 electric field Effects 0.000 claims description 39
- 230000006698 induction Effects 0.000 claims description 20
- 230000007935 neutral effect Effects 0.000 claims description 6
- 239000000853 adhesive Substances 0.000 claims description 5
- 230000001070 adhesive effect Effects 0.000 claims description 5
- 239000000835 fiber Substances 0.000 claims description 3
- 238000006386 neutralization reaction Methods 0.000 claims description 3
- 230000008901 benefit Effects 0.000 abstract description 10
- 238000010586 diagram Methods 0.000 description 5
- 230000005855 radiation Effects 0.000 description 5
- 230000004323 axial length Effects 0.000 description 4
- 230000008878 coupling Effects 0.000 description 4
- 238000010168 coupling process Methods 0.000 description 4
- 238000005859 coupling reaction Methods 0.000 description 4
- 230000008859 change Effects 0.000 description 2
- 230000006835 compression Effects 0.000 description 2
- 238000007906 compression Methods 0.000 description 2
- 238000010276 construction Methods 0.000 description 2
- 238000005516 engineering process Methods 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 238000012986 modification Methods 0.000 description 2
- 238000005457 optimization Methods 0.000 description 2
- 230000004044 response Effects 0.000 description 2
- 230000007547 defect Effects 0.000 description 1
- 238000006073 displacement reaction Methods 0.000 description 1
- 238000003306 harvesting Methods 0.000 description 1
- 230000000737 periodic effect Effects 0.000 description 1
- 238000000926 separation method Methods 0.000 description 1
- 230000035939 shock Effects 0.000 description 1
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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
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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
- B64G1/228—Damping of high-frequency vibration effects on spacecraft elements, e.g. by using acoustic vibration dampers
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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
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation
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- Aviation & Aerospace Engineering (AREA)
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- Acoustics & Sound (AREA)
- Vibration Prevention Devices (AREA)
Abstract
The invention discloses a spacecraft vibration suppression structure and method based on a piezoelectric composite material, wherein a shell of a spacecraft is circumferentially and uniformly provided with a plurality of even mounting grooves at the position with the maximum strain of a first-order bending vibration mode and the positions with the maximum strain of two second-order bending vibration modes; the mounting grooves are provided with piezoelectric vibration units; the piezoelectric vibration unit comprises a fixed cover and a piezoelectric ceramic piece; the fixed cover is made of carbon fiber composite material, and is covered on the piezoelectric ceramic sheet and is tightly and fixedly connected with the spacecraft shell; two sides of the piezoelectric ceramic sheet form a series closed loop through the fixed cover and the spacecraft shell; the polarization directions of the piezoelectric ceramic plates are all outward or all inward. The invention adopts a passive vibration suppression mode to control the vibration of the spacecraft, utilizes the positive piezoelectric effect of the piezoelectric material to convert the vibration mechanical energy into electric energy, and further converts the electric energy into heat to be dissipated through the self-resistance of the spacecraft shell so as to reduce the vibration, thereby having higher economic benefit and application prospect.
Description
Technical Field
The invention relates to the field of piezoelectric energy harvesting and mechanical vibration control, in particular to a spacecraft vibration suppression structure and method based on a piezoelectric composite material.
Background
The spacecraft can experience extremely severe mechanical environment in the launching and power flight stages, and is required to bear loads such as vibration, impact, noise, thermal environment and the like caused by operations such as take-off, interstage (fairing) separation, secondary ignition, shutdown, orbit entering and the like. The vibration of the spacecraft shell seriously affects the working efficiency and safety, and the active control technology has complex structure and complicated control, and the traditional passive control has poor vibration suppression effect and low reliability. According to the invention, the vibration of the spacecraft is controlled by adopting a passive vibration suppression mode through the construction of the composite material, the vibration mechanical energy is converted into electric energy by utilizing the positive piezoelectric effect of the piezoelectric material, and the electric energy is further converted into heat through the self-resistance of the spacecraft shell so as to be conducted outside the system. The invention utilizes the advantages of the carbon fiber composite material such as impact resistance, high temperature resistance, strong wear resistance and the like and the advantages of the piezoelectric material such as quick response, no electromagnetic interference and radiation interference and the like, so that the invention has simple structure and strong reliability, is suitable for the fields of spacecrafts and the like, and has higher economic benefit and application prospect.
Disclosure of Invention
The invention aims to solve the technical problem of providing a spacecraft vibration suppression structure and method based on a piezoelectric composite material aiming at the defects related to the background technology.
The invention adopts the following technical scheme for solving the technical problems:
the spacecraft vibration suppression structure based on the piezoelectric composite material is characterized in that a shell of the spacecraft is cylindrical, the spacecraft vibration suppression structure comprises 2n+4m piezoelectric vibration units, and n and m are natural numbers greater than or equal to 1;
2n mounting grooves are uniformly formed in the circumferential direction of the position with the largest first-order bending vibration modal strain of the shell of the spacecraft, and 2m mounting grooves are uniformly formed in the circumferential direction of the position with the largest second-order bending vibration modal strain of the shell of the spacecraft;
the 2n+4m piezoelectric vibration units are arranged in the 2n+4m mounting grooves of the spacecraft shell in a one-to-one correspondence manner, and each piezoelectric vibration unit comprises a fixed cover and a piezoelectric ceramic plate, wherein each piezoelectric ceramic plate is of a piezoelectric material PZT sheet structure and polarized along the thickness direction of the piezoelectric ceramic plate, and the lower end face of each piezoelectric ceramic plate is electrically connected with the spacecraft shell by being fixed on the bottom wall of the corresponding mounting groove; the fixed cover is made of carbon fiber composite material, covers the piezoelectric ceramic plate and is hermetically and fixedly connected with the spacecraft shell at the position corresponding to the mounting groove; the fixed cover is electrically connected with the piezoelectric ceramic plate and the spacecraft shell, so that two sides of the piezoelectric ceramic plate form a series closed loop through the fixed cover and the spacecraft shell; the polarization directions of piezoelectric ceramic plates in the 2n+4m piezoelectric vibration units are all outward or all inward.
As a further optimization scheme of the spacecraft vibration suppression structure based on the piezoelectric composite material, the piezoelectric ceramic plates in the piezoelectric vibration unit are fixed on the bottom wall of the corresponding mounting groove through conductive adhesive and are electrically connected with the spacecraft shell.
As a further optimization scheme of the spacecraft vibration suppression structure based on the piezoelectric composite material, the fiber direction of the fixed cover in the piezoelectric vibration unit is unidirectional and parallel to the axis of the spacecraft.
The shell of the spacecraft is positioned at the position with the largest first-order bending vibration modal strain at the position with the largest axial length of 1/2, and the positions with the largest second-order bending vibration modal strain are respectively positioned at the positions with the largest axial lengths of 1/4 and 3/4.
The size of the piezoelectric ceramic piece is slightly smaller than that of the corresponding mounting groove, and the piezoelectric ceramic piece is stuck or embedded into the corresponding mounting groove through conductive adhesive; when the piezoelectric ceramic piece deforms in the thickness direction, positive piezoelectric effect occurs to generate an electric field, voltage change occurs on the inner surface and the outer surface of the piezoelectric ceramic piece, so that the electric field is generated, mechanical energy is converted into electric energy, the electric energy is dissipated through self-resistance of the spacecraft shell, and then the electric energy is dissipated in a heat energy mode.
In the working state of the spacecraft, the frequency of the first-order bending vibration mode, the second-order bending vibration mode and the third-order bending vibration mode is relatively low, so that the normal working and the safety of the spacecraft are seriously affected by the vibration of the spacecraft. The composite material structure is suitable for the working environment of the spacecraft, and the energy is converted, dissipated and conducted outside the system by utilizing the electromechanical coupling characteristic of the piezoelectric material when the system is deformed due to vibration, and the vibration amplitude is reduced by adding a channel mode of conducting the energy outside the system because the piezoelectric ceramic sheet has the advantages of strong electromechanical coupling characteristic, corresponding rapidness, electromagnetic interference resistance, radiation resistance and the like and the carbon fiber composite material has the advantages of strong tensile property, high temperature resistance, strong wear resistance, impact resistance, strong fixing property and the like.
The invention also discloses a vibration suppression method of the spacecraft vibration suppression structure based on the piezoelectric composite material, which comprises the following steps:
when the spacecraft shell is in a stable state, the upper surface and the lower surface of the piezoelectric ceramic plate in each piezoelectric vibration unit are electrically neutral;
when the spacecraft shell is excited to generate first-order bending vibration, second-order bending vibration and third-order bending vibration, the spacecraft shell deforms, the piezoelectric ceramic piece of the piezoelectric vibration unit subjected to compressive stress stretches along the polarization direction, a positive piezoelectric effect is generated to generate an induction electric field, and the direction of the induction electric field is opposite to the polarization direction along the thickness direction; the piezoelectric ceramic plate of the piezoelectric vibration unit subjected to tensile stress is shortened along the polarization direction, and a positive piezoelectric effect is generated to generate an induced electric field, wherein the direction of the induced electric field is the same as the polarization direction along the thickness direction; the piezoelectric ceramic plates generating the induction electric field are all subjected to neutralization and dissipation through the self resistance of the spacecraft shell, and then are dispersed in the form of heat energy, so that the vibration amplitude of the spacecraft shell is reduced.
Compared with the prior art, the technical scheme provided by the invention has the following technical effects:
1. the invention adopts a composite structure, and the outer wall is made of carbon fiber composite material, so that the invention has the advantages of impact resistance, high temperature resistance, wear resistance, high strength, light weight and the like;
2. the piezoelectric material has good vibration suppression effect on first-order bending vibration, second-order bending vibration and third-order bending vibration, and adopts a passive vibration suppression method, so that the piezoelectric material has the advantages of simple structure, strong reliability, quick response, no electromagnetic and radiation interference and the like;
3. according to the invention, through the construction of the composite material structure, the self-resistivity of the spacecraft is utilized, and the vibration is passively controlled, so that the structure is simple, and the lightweight, shock-resistant, radiation-resistant and high-temperature-resistant requirements of the spacecraft are met.
Drawings
FIG. 1 is a schematic diagram of the structure of the present invention;
FIG. 2 is a schematic cross-sectional view of the present invention;
FIG. 3 is a schematic diagram of an equivalent circuit formed by matching a spacecraft shell, a piezoelectric ceramic piece and a fixed cover;
FIG. 4 is a schematic diagram of the stress state of the piezoelectric ceramic plate in the first-order bending vibration mode according to the present invention;
FIG. 5 is a schematic diagram of the stress state of the piezoelectric ceramic plate in the second-order bending vibration mode;
fig. 6 is a schematic diagram of a stress state of the piezoelectric ceramic sheet in a third-order bending vibration mode.
In the figure, a shell of a 1-spacecraft, a 2-piezoelectric ceramic piece and a 3-fixed cover.
Detailed Description
The technical scheme of the invention is further described in detail below with reference to the accompanying drawings:
this invention may be embodied in many different forms and should not be construed as limited to the embodiments set forth herein. Rather, these embodiments are provided so that this disclosure will be thorough and complete, and will fully convey the scope of the invention to those skilled in the art. In the drawings, the components are exaggerated for clarity.
As shown in fig. 1, the invention discloses a spacecraft vibration suppression structure based on a piezoelectric composite material, wherein a shell of a spacecraft is cylindrical, the spacecraft vibration suppression structure comprises 2n+4m piezoelectric vibration units, and n and m are natural numbers greater than or equal to 1;
2n mounting grooves are uniformly formed in the circumferential direction of the position with the largest first-order bending vibration modal strain of the shell of the spacecraft, and 2m mounting grooves are uniformly formed in the circumferential direction of the position with the largest second-order bending vibration modal strain of the shell of the spacecraft;
as shown in fig. 2, the 2n+4m piezoelectric vibration units are arranged in the 2n+4m mounting grooves of the spacecraft shell in a one-to-one correspondence manner, and each piezoelectric vibration unit comprises a fixed cover and a piezoelectric ceramic plate, wherein the piezoelectric ceramic plates are of a sheet structure of piezoelectric material PZT and polarized along the thickness direction of the piezoelectric ceramic plates, and the lower end surfaces of the piezoelectric ceramic plates are electrically connected with the spacecraft shell by being fixed on the bottom walls of the corresponding mounting grooves; the fixed cover is made of carbon fiber composite material, covers the piezoelectric ceramic plate and is hermetically and fixedly connected with the spacecraft shell at the position corresponding to the mounting groove; the fixed cover is electrically connected with the piezoelectric ceramic plate and the spacecraft shell, so that two sides of the piezoelectric ceramic plate form a series closed loop through the fixed cover and the spacecraft shell, as shown in fig. 3; the polarization directions of piezoelectric ceramic plates in the 2n+4m piezoelectric vibration units are all outward or all inward.
The piezoelectric ceramic plates in the piezoelectric vibration unit are fixed on the bottom wall of the corresponding mounting groove through conductive adhesive and are electrically connected with the spacecraft shell.
The fiber direction of the fixed cover in the piezoelectric vibration unit is unidirectional and parallel to the axis of the spacecraft.
The shell of the spacecraft is positioned at the position with the largest first-order bending vibration modal strain at the position with the largest axial length of 1/2, and the positions with the largest second-order bending vibration modal strain are respectively positioned at the positions with the largest axial lengths of 1/4 and 3/4.
The size of the piezoelectric ceramic piece is slightly smaller than that of the corresponding mounting groove, and the piezoelectric ceramic piece is stuck or embedded into the corresponding mounting groove through conductive adhesive; when the piezoelectric ceramic piece deforms in the thickness direction, positive piezoelectric effect occurs to generate an electric field, voltage change occurs on the inner surface and the outer surface of the piezoelectric ceramic piece, so that the electric field is generated, mechanical energy is converted into electric energy, the electric energy is dissipated through self-resistance of the spacecraft shell, and then the electric energy is dissipated in a heat energy mode.
In the working state of the spacecraft, the frequency of the first-order bending vibration mode, the second-order bending vibration mode and the third-order bending vibration mode is relatively low, so that the normal working and the safety of the spacecraft are seriously affected by the vibration of the spacecraft. The composite material structure is suitable for the working environment of the spacecraft, and the energy is converted, dissipated and conducted outside the system by utilizing the electromechanical coupling characteristic of the piezoelectric material when the system is deformed due to vibration, and the vibration amplitude is reduced by adding a channel mode of conducting the energy outside the system because the piezoelectric ceramic sheet has the advantages of strong electromechanical coupling characteristic, corresponding rapidness, electromagnetic interference resistance, radiation resistance and the like and the carbon fiber composite material has the advantages of strong tensile property, high temperature resistance, strong wear resistance, impact resistance, strong fixing property and the like.
The invention also discloses a vibration suppression method of the spacecraft vibration suppression structure based on the piezoelectric composite material, which comprises the following steps:
when the spacecraft shell is in a stable state, the upper surface and the lower surface of the piezoelectric ceramic plate in each piezoelectric vibration unit are electrically neutral;
when the spacecraft shell is excited to generate first-order bending vibration, second-order bending vibration and third-order bending vibration, the spacecraft shell deforms, the piezoelectric ceramic piece of the piezoelectric vibration unit subjected to compressive stress stretches along the polarization direction, a positive piezoelectric effect is generated to generate an induction electric field, and the direction of the induction electric field is opposite to the polarization direction along the thickness direction; the piezoelectric ceramic plate of the piezoelectric vibration unit subjected to tensile stress is shortened along the polarization direction, and a positive piezoelectric effect is generated to generate an induced electric field, wherein the direction of the induced electric field is the same as the polarization direction along the thickness direction; the piezoelectric ceramic plates generating the induction electric field are all subjected to neutralization and dissipation through the self resistance of the spacecraft shell, and then are dispersed in the form of heat energy, so that the vibration amplitude of the spacecraft shell is reduced.
As shown in fig. 4, when the spacecraft is excited to generate first-order bending vibration, the cycle steps are as follows:
step A.1), the spacecraft shell is in a stable state, and the upper surface and the lower surface of the piezoelectric ceramic plate are electrically neutral; when the spacecraft is subjected to load changes such as impact and the like to generate first-order bending vibration, the two sides of the spacecraft shell start to generate relative displacement;
step A.2), the spacecraft shell deforms, one side (upper side in the figure) is stressed by pressure, so that the piezoelectric ceramic plate stretches along the polarization direction, a positive piezoelectric effect occurs to generate an induced electric field, and the direction of the induced electric field is opposite to the polarization direction along the thickness direction; the other side (lower side in the figure) receives tensile stress, so that the piezoelectric ceramic plate is shortened along the polarization direction, a positive piezoelectric effect is generated to generate an induced electric field, and the direction of the induced electric field is the same as the polarization direction along the thickness direction; because the upper surface and the lower surface of the piezoelectric ceramic plate are connected in series through the spacecraft shell and the fixed cover, the generated electric energy is consumed through the self resistance of the spacecraft shell and is dissipated in the form of heat energy;
step A.3), the spacecraft shell is restored to a stable state, and no induction electric field is generated at the moment;
step A.4), the deformation direction of the spacecraft shell is opposite to that of the step A.2), and the piezoelectric ceramic plate deforms in the thickness direction, so that the piezoelectric ceramic plate on one side subjected to tensile stress generates an induction electric field with the same polarization direction; the piezoelectric ceramic piece at one side which receives the compressive stress generates an induced electric field opposite to the polarization direction; the electric energy generated by the piezoelectric ceramic plate is consumed through the self resistance of the spacecraft shell and is emitted in the form of heat energy;
and (A.5) recovering the spacecraft shell to a stable state, wherein no induction electric field exists.
As shown in fig. 5, when the spacecraft is excited to generate second-order bending vibration, the cycle steps are as follows:
step B.1), the spacecraft shell is in a stable state, and the upper surface and the lower surface of the piezoelectric ceramic plate are electrically neutral; when the spacecraft is subjected to load changes such as impact and the like to generate second-order bending vibration, the spacecraft shell starts to deform;
step B.2), the spacecraft shell deforms, the deformation is maximum at the maximum strain position of the second-order bending vibration, namely the full length position of 0.25 and the full length position of 0.75, the piezoelectric ceramic plate is respectively stressed by compression stress and tensile stress, the deformation is generated in the thickness direction, the piezoelectric ceramic plate stressed by compression stress generates an induction electric field opposite to the polarization direction, and the piezoelectric ceramic plate stressed by tensile stress generates an induction electric field identical to the polarization direction; the piezoelectric ceramic plate at the position with the largest first-order bending vibration strain (namely the position with the full length of 0.5) is not deformed or is less deformed in the thickness direction; the electric energy generated by the piezoelectric ceramic plate is consumed through the self resistance of the spacecraft shell and is emitted in the form of heat energy;
step B.3), the spacecraft shell is restored to a stable state, and no induction electric field is generated at the moment;
step B.4), the deformation direction of the spacecraft shell is opposite to that of the step B.2), the piezoelectric ceramic plate at the position of the maximum strain of the second-order flexural vibration generates deformation in the thickness direction, the piezoelectric ceramic plate at the side subjected to tensile stress generates an induced electric field in the same direction as the polarization direction, and the piezoelectric ceramic plate at the side subjected to compressive stress generates an induced electric field in the opposite direction to the polarization direction; the electric energy generated by the piezoelectric ceramic plate is consumed through the self resistance of the spacecraft shell and is emitted in the form of heat energy;
and B.5), recovering the spacecraft shell to a stable state, and at the moment, no induction electric field exists.
As shown in fig. 6, when the spacecraft is excited to generate third-order bending vibration, the periodic steps are as follows:
step C.1), the spacecraft shell is in a stable state, and the upper surface and the lower surface of the piezoelectric ceramic plate are electrically neutral; when the spacecraft is subjected to load changes such as impact and the like to generate third-order bending vibration, the spacecraft shell starts to deform;
step C.2), the spacecraft shell deforms, the deformation amount is maximum at the trough and the crest of the third-order bending vibration at the maximum strain position of the first-order bending vibration and the second-order bending vibration, and the piezoelectric ceramic plates deform in the thickness direction; the piezoelectric ceramic piece subjected to compressive stress generates an induced electric field opposite to the polarization direction, and the piezoelectric ceramic piece subjected to tensile stress generates an induced electric field identical to the polarization direction; the electric energy generated by the piezoelectric ceramic plate is consumed through the self resistance of the spacecraft shell and is emitted in the form of heat energy;
step C.3), the spacecraft shell is restored to a stable state, and no induction electric field is generated at the moment;
step C.4), the deformation direction of the spacecraft shell is opposite to that of the step C.2), the piezoelectric ceramic plate deforms in the thickness direction, the piezoelectric ceramic plate on the side subjected to tensile stress generates an induction electric field with the same polarization direction, and the piezoelectric ceramic plate on the side subjected to compressive stress generates an induction electric field with the opposite polarization direction; the electric energy generated by the piezoelectric ceramic plate is consumed through the self resistance of the spacecraft shell and is emitted in the form of heat energy;
and C.5), recovering the spacecraft shell to a stable state, and at the moment, no induction electric field exists.
It will be understood by those skilled in the art that, unless otherwise defined, all terms (including technical and scientific terms) used herein have the same meaning as commonly understood by one of ordinary skill in the art to which this invention belongs. It will be further understood that terms, such as those defined in commonly used dictionaries, should be interpreted as having a meaning that is consistent with their meaning in the context of the prior art and will not be interpreted in an idealized or overly formal sense unless expressly so defined herein.
While the foregoing is directed to embodiments of the present invention, other and further details of the invention may be had by the present invention, it should be understood that the foregoing description is merely illustrative of the present invention and that no limitations are intended to the scope of the invention, except insofar as modifications, equivalents, improvements or modifications are within the spirit and principles of the invention.
Claims (4)
1. The spacecraft vibration suppression structure based on the piezoelectric composite material is characterized in that a spacecraft shell is cylindrical and comprises 2n+4m piezoelectric vibration units, and n and m are natural numbers greater than or equal to 1;
2n mounting grooves are uniformly formed in the circumferential direction of the spacecraft shell at the position with the maximum first-order bending vibration modal strain, and 2m mounting grooves are uniformly formed in the circumferential direction of the position with the maximum second-order bending vibration modal strain;
the 2n+4m piezoelectric vibration units are arranged in the 2n+4m mounting grooves of the spacecraft shell in a one-to-one correspondence manner, and each piezoelectric vibration unit comprises a fixed cover and a piezoelectric ceramic plate, wherein each piezoelectric ceramic plate is of a piezoelectric material PZT sheet structure and polarized along the thickness direction of the piezoelectric ceramic plate, and the lower end face of each piezoelectric ceramic plate is electrically connected with the spacecraft shell by being fixed on the bottom wall of the corresponding mounting groove; the fixed cover is made of carbon fiber composite material, covers the piezoelectric ceramic plate and is hermetically and fixedly connected with the spacecraft shell at the position corresponding to the mounting groove; the fixed cover is electrically connected with the piezoelectric ceramic plate and the spacecraft shell, so that two sides of the piezoelectric ceramic plate form a series closed loop through the fixed cover and the spacecraft shell; the polarization directions of piezoelectric ceramic plates in the 2n+4m piezoelectric vibration units are all outward or all inward.
2. The spacecraft vibration suppression structure based on the piezoelectric composite material according to claim 1, wherein the piezoelectric ceramic plates in the piezoelectric vibration unit are fixed on the bottom wall of the corresponding mounting groove through conductive adhesive and are electrically connected with the spacecraft shell.
3. The structure for suppressing vibration of a spacecraft based on a piezoelectric composite material according to claim 1, wherein the fiber direction of the stationary cover in the piezoelectric vibration unit is unidirectional and parallel to the axis of the spacecraft.
4. The method for suppressing the vibration of the spacecraft vibration suppression structure based on the piezoelectric composite material according to claim 1, which is characterized by comprising the following steps:
when the spacecraft shell is in a stable state, the upper surface and the lower surface of the piezoelectric ceramic plate in each piezoelectric vibration unit are electrically neutral;
when the spacecraft shell is excited to generate first-order bending vibration, second-order bending vibration and third-order bending vibration, the spacecraft shell deforms, the piezoelectric ceramic piece of the piezoelectric vibration unit subjected to compressive stress stretches along the polarization direction, a positive piezoelectric effect is generated to generate an induction electric field, and the direction of the induction electric field is opposite to the polarization direction along the thickness direction; the piezoelectric ceramic plate of the piezoelectric vibration unit subjected to tensile stress is shortened along the polarization direction, and a positive piezoelectric effect is generated to generate an induced electric field, wherein the direction of the induced electric field is the same as the polarization direction along the thickness direction; the piezoelectric ceramic plates generating the induction electric field are all subjected to neutralization and dissipation through the self resistance of the spacecraft shell, and then are dispersed in the form of heat energy, so that the vibration amplitude of the spacecraft shell is reduced.
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