CN110562499A - Wall-mounted main frequency adjustable variable-section one-rocket multi-satellite launching moonlet structure - Google Patents
Wall-mounted main frequency adjustable variable-section one-rocket multi-satellite launching moonlet structure Download PDFInfo
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- XAGFODPZIPBFFR-UHFFFAOYSA-N aluminium Chemical compound [Al] XAGFODPZIPBFFR-UHFFFAOYSA-N 0.000 claims description 9
- 229910052782 aluminium Inorganic materials 0.000 claims description 9
- 229910000838 Al alloy Inorganic materials 0.000 claims description 4
- 230000005540 biological transmission Effects 0.000 claims description 4
- 229920000049 Carbon (fiber) Polymers 0.000 claims description 3
- 229910001069 Ti alloy Inorganic materials 0.000 claims description 3
- 239000004917 carbon fiber Substances 0.000 claims description 3
- 239000011229 interlayer Substances 0.000 claims description 3
- 238000003754 machining Methods 0.000 claims description 3
- VNWKTOKETHGBQD-UHFFFAOYSA-N methane Chemical compound C VNWKTOKETHGBQD-UHFFFAOYSA-N 0.000 claims description 3
- 239000011347 resin Substances 0.000 claims description 3
- 229920005989 resin Polymers 0.000 claims description 3
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- 238000009434 installation Methods 0.000 abstract description 14
- 235000015842 Hesperis Nutrition 0.000 abstract description 8
- 235000012633 Iberis amara Nutrition 0.000 abstract description 8
- 108091092919 Minisatellite Proteins 0.000 abstract description 2
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- 229910052741 iridium Inorganic materials 0.000 description 5
- GKOZUEZYRPOHIO-UHFFFAOYSA-N iridium atom Chemical compound [Ir] GKOZUEZYRPOHIO-UHFFFAOYSA-N 0.000 description 5
- 230000008878 coupling Effects 0.000 description 4
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- 238000005859 coupling reaction Methods 0.000 description 4
- 238000012360 testing method Methods 0.000 description 4
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- 239000004411 aluminium Substances 0.000 description 2
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- 241000883964 Ariocarpus retusus Species 0.000 description 1
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- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
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- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
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- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/64—Systems for coupling or separating cosmonautic vehicles or parts thereof, e.g. docking arrangements
- B64G1/641—Interstage or payload connectors
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- 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
- B64G1/64—Systems for coupling or separating cosmonautic vehicles or parts thereof, e.g. docking arrangements
- B64G1/641—Interstage or payload connectors
- B64G1/643—Interstage or payload connectors for arranging multiple satellites in a single launcher
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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
- B64G1/64—Systems for coupling or separating cosmonautic vehicles or parts thereof, e.g. docking arrangements
- B64G1/645—Separators
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Abstract
The invention provides a wall-mounted main frequency adjustable variable-section one-rocket multi-satellite launching minisatellite structure which comprises a satellite, wherein the satellite is connected to a carrier rocket through a wall, and the satellite is of a variable-section structure along the launching direction of the carrier rocket. The invention has the beneficial effects that: the rigidity adaptability to different carrier rockets is improved, the stress of a satellite-rocket connection interface structure is reduced, and the utilization rate of the installation space of the in-satellite equipment is improved.
Description
Technical Field
The invention relates to a small satellite structure, in particular to a wall-mounted main frequency adjustable variable-section one-rocket multi-satellite launching small satellite structure.
Background
The one-rocket-multi-satellite launching mode that a plurality of small satellites with basically the same size share one carrier rocket provides challenges for the traditional multi-satellite series arrangement mode in the fairing. The multi-satellite series layout has large axial size, the number of launching satellites is limited by the size of a fairing of a carrier rocket, and in addition, the satellites positioned at the high level suffer from a harsher launching mechanical environment due to the high position of the mass center. In order to solve the problem, a mode of launching the small satellites by a multi-satellite parallel layout 'one rocket and multiple satellites' is developed. The connection between the multi-satellite parallel connection satellite structure and the carrier rocket generally comprises two modes of satellite bottom connection and satellite side wall hanging connection. Compared with a side wall hanging type connection mode, the satellite-rocket connection interface in the bottom connection mode generally bears larger load. Therefore, the side wall hanging type connection (hereinafter referred to as wall hanging type) is more suitable for domestic and foreign countries. Wall-mounted satellite structures typically have an entire satellite sidewall connected to a satellite adapter inside a launch vehicle fairing by a pyrotechnic device, such as an explosive bolt. In order to meet the rigidity requirements of different carrier rockets in a one-rocket multi-satellite launching mode of multi-satellite parallel layout and avoid frequency coupling of a satellite structure and the carrier rockets, the design of a wall-mounted type small satellite structure with adjustable dominant frequency is necessary.
1993 to 1998, Iridium system (Iridium) was developed and successfully launched by Motorola in America[1][2]The whole star is designed into a triangular prism uniform section structure, and adopts 'one arrow five stars', 'one arrow seven stars' and 'one arrow two stars'[2][3]Equal emission mode and multi-star parallel layout bottom connection mode. The star configuration allows the launch vehicle fairing internal space utilization (which is the percentage of the satellite volume to the total launch vehicle fairing internal space volume) to reach 75.68% ("one rocket five stars"). Usually the sum of the device itself and the corresponding installation operating space is a cuboid. The cuboid is placed in the internal space of the triangular prism of the satellite, so that the utilization rate of the satellite internal equipment installation space (the percentage of the total volume of the equipment and the volume occupied by the installation operation space in the total space in the satellite) of the satellite is low. In addition to this, the present invention is,The long and thin satellite structure design enables the satellite-rocket connection interface to bear large load under the condition of transverse overload. 1991 to 1999, the limited partnership of globalstar (global-star) initiated by Laura and Gaotong incorporated in America developed globalstar[2]The satellite structure is designed into a trapezoidal prismatic shape with equal sections, the side wall where the short side of the trapezoid is located is hung on a satellite adapter of a carrier rocket, and compared with a bottom connection mode of a triangular prism equal section structure, the structure has the advantage that the satellite-rocket connection interface is reduced by load. However, the wall-mounted connection easily causes frequency coupling between the whole satellite and the carrier rocket, and the rigidity adaptability of the same structure to different carrier rockets is poor. In addition, in the design state of the equal section, the size of the cross section is determined by equipment with large installation space, and equipment with small installation space cannot fully utilize the space in the satellite, so that the utilization rate of the installation space of the equipment in the satellite is low. Literature reference[4]Although a proposal for improving the bottom connection mode of the one-arrow-plus-one-star system into a wall-mounted connection is provided, the domestic application of wall-mounted one-arrow-plus-one-star launching small satellites is not reported in public so far. Earth gravity test satellite Grace jointly developed in 2002 in United states and Germany and successfully launched in 'one-arrow two-star' mode[5][6](ii) a 2011 navigation satellite Galileo IOV successfully launched by ESA with' one arrow two stars[7](ii) a In 2017, the Iridium Next of the Next generation Iridium is successfully emitted in the United states of' one arrow and ten stars[8]. The satellites launched since 2002 have all adopted a wall-mounted structure similar to a global star, with advantages and disadvantages similar to those of a global star.
Based on the analysis of the current situation and existing problems of the structural design in the one-rocket and multi-satellite parallel layout launching mode at home and abroad, how to solve the technical problems that the satellite structure has weak rigidity adaptability to different carrier rockets, the stress of a satellite and rocket connection interface structure is overlarge, and the utilization rate of the in-satellite equipment installation space is low is urgently needed to be solved by technical personnel in the field.
Reference documents:
[1]Maley P D,Pizzicaroli J C.The visual appearance of the Iridium satellites[J].Acta Astronautica 52(2003)629-639
[2] Features of satellite configurations suitable for "one rocket stars" launch are reviewed in international space [ J ]. 2007 (6): 23
[3] Exchange, long conquer No. two change-launch iridium and double star planned rocket [ J ]. space exploration, 2007 (7): 42
[4] Schaviw, old and precious, qianyinging, satellite configuration study of direct injection launched by one starry [ J ] spacecraft engineering, 2012, 21 (1): 43-47
[5]M.Kinnersley，T.Miski，I.Schumacher，P.Freeborn.The rockot launch vehicle-the competitive launch solution for satellite systems[C].21st international communications satellite systems conference and exhibit，AIAA2003-2263
[6]JANES ZB.GRACE(Gravity Recovery and Climate Experiment)[R].2015.5
[7]S.P.Thompson,G.Andersson,W.Davies.The Galleo IOV dispenser system-design,development& verification[C].Proceedings of the 12th European conference on spacecraft structures,materials& environmental testing,20120320-20120323
[8] Spot analysis of flying task of longxuedan, poplars, falcon-9 rocket [ J ]. International
Space, 2017 (2): 76-79
Disclosure of Invention
In order to solve the problems in the prior art, the invention provides a wall-mounted type main frequency adjustable variable cross-section one-rocket multi-satellite launching structure.
The invention provides a wall-mounted main frequency adjustable variable-section one-rocket multi-satellite launching minisatellite structure which comprises a satellite, wherein the satellite is connected to a carrier rocket through a wall, and the satellite is of a variable-section structure along the launching direction of the carrier rocket.
As a further improvement of the invention, the satellite is provided with a satellite joint, and the satellite joint on the satellite is in point-type connection with the carrier rocket through a separation nut.
As a further improvement of the invention, the coordinate system of the satellite is O-XYZ, the satellite comprises a bottom plate, a + X + Y thermal control plate, a-X-Y thermal control plate, a + Y side plate, a-Y side plate, a + Y battery plate, a-Z battery plate, a + X antenna plate and a-X antenna plate, wherein the + X antenna plate and the-X antenna plate are respectively arranged in the middle of the bottom plate, the + X + Y thermal control plate, the-X + Y thermal control plate and the-Y side plate are respectively arranged at one end of the bottom plate, the + X-Y thermal control plate, the-X-Y thermal control plate and the-Y side plate are respectively arranged at the other end of the bottom plate, and the-Z battery plate is erected on the + X antenna plate and the-X antenna plate, the + Y battery plate is erected on the + X + Y thermal control plate, the-X + Y thermal control plate, the + Y side plate, the + X antenna plate and the-X antenna plate, and the-Y battery plate is erected on the + X-Y thermal control plate, the-Y side plate, the + X antenna plate and the-X antenna plate.
As a further improvement of the invention, the base plate, the + Y cell panel, the-Y cell panel and the-Z cell panel are in an isosceles trapezoid frame structure, the-Z cell panel is the top surface of an isosceles trapezoid, the base plate is the bottom surface of the isosceles trapezoid, the + Y cell panel and the-Y cell panel are two inclined planes of the isosceles trapezoid, the length direction of the base plate is parallel to the launching direction of the carrier rocket, and the distance between the-Z cell panel and the base plate is smaller than or far smaller than the length of the base plate.
As a further improvement of the invention, the bottom plate, + X + Y thermal control plate, -X + Y thermal control plate, + X-Y thermal control plate, -X-Y thermal control plate, + Y side plate, -Y side plate, + Y battery board, -Y battery board and-Z battery board are all honeycomb sandwich structure plates.
As a further improvement of the invention, the + Y cell panel, the-Y cell panel and the-Z cell panel are all aluminum honeycomb core sub-sandwich structure plates with carbon fiber reinforced resin panels, and the bottom plate, the + X + Y thermal control plate, the-X + Y thermal control plate, the + X-Y thermal control plate, the-X-Y thermal control plate, the + Y side plate and the-Y side plate are all aluminum panel aluminum honeycomb core sub-sandwich structure plates.
As a further improvement of the invention, the satellite and rocket joint is machined and formed by adopting aluminum alloy, and the satellite and rocket joint is solidified in the bottom plate 3 in a pre-embedded mode.
As a further improvement of the invention, 3 rigidity adjusting screws are arranged near each satellite-rocket joint along the Y direction, the rigidity adjusting screws have different Y-direction supporting spans corresponding to the bottom plate in a fully or partially disassembled state, the rigidity adjusting screws adopt M5 titanium alloy screws, the rigidity adjusting screws are not glued and loose-proof during rigidity adjustment, and the glue is glued and loose-proof after the final state is determined.
As a further development of the invention, the satellite is symmetrical about the ZOX reference plane, and the floor has equipment mounted thereon.
As a further improvement of the invention, the satellites are six and evenly distributed around the circumference of the carrier rocket at intervals.
The invention has the beneficial effects that: by the scheme, the rigidity adaptability to different carrier rockets is improved, the stress of a satellite-rocket connection interface structure is reduced, and the utilization rate of an in-satellite equipment installation space is improved.
Drawings
Fig. 1 is a view of a rocket six-star layout rocket coordinate system Z1 of a wall-mounted main frequency adjustable variable cross-section one-rocket multi-star launching moonlet structure.
FIG. 2 is a six-arrow-six-star layout rocket coordinate system X1 orientation view.
Fig. 3 is a view in direction a of an arrow six-star layout hiding the remaining 5 stars.
Fig. 4 is a B-direction view of an arrow six-star layout hiding the remaining 5 stars.
fig. 5 is a composition diagram of a satellite.
Fig. 6 is a view of the satellite in the direction + X.
Fig. 7 is a view of the satellite in the + Y direction.
Figure 8 is a C-C cross-sectional view of the satellite.
Figure 9 is a D-D cross-sectional view of the satellite.
Figure 10 is an E-E cut-away view of a satellite and different size equipment placement diagram.
Fig. 11 is a dominant frequency tunable design of a satellite.
FIG. 12 is a graph of a fit of dominant frequencies to span of a satellite.
Detailed Description
The invention is further described with reference to the following description and embodiments in conjunction with the accompanying drawings.
As shown in fig. 1 and fig. 2, a wall-mounted main-frequency adjustable cross-section one-rocket multi-satellite launching structure comprises a satellite 1, wherein the satellite 1 is launched by adopting a one-rocket six-satellite layout, and a launch coordinate system O1-X1Y1Z1 of a carrier rocket 2. Fig. 3 and 4 hide the other five satellites 1 and reserve one satellite 1, and the satellite and rocket joint 3 on the satellite 1 is in point-type connection with the carrier rocket 2 through the connecting and separating nut 4, so that the satellite 1 is connected to the carrier rocket 2 in a wall-mounted manner.
The composition of the satellite 1 is shown in fig. 5, and each component is named as follows under the body coordinate system O-XYZ of the satellite 1: a bottom plate 5, + X + Y thermal control plate 6, -X + Y thermal control plate 7, + X-Y thermal control plate 8, -X-Y thermal control plate 9, + Y lateral plate 10, -Y lateral plate 11, + Y panel 12, -Y panel 13, -Z panel 14, + X antenna plate 15, -X antenna plate 16, and a structural connecting piece 17.
The + X antenna plate 15 and the-X antenna plate 16 are respectively arranged in the middle of the bottom plate 5, the + X + Y thermal control plate 6, the-X + Y thermal control plate 7 and the + Y side plate 10 are respectively arranged at one end of the bottom plate 5, the + X-Y thermal control plate 8, the-X-Y thermal control plate 9 and the-Y side plate 11 are respectively arranged at the other end of the bottom plate 5, the-Z battery plate 14 is erected on the + X antenna plate 15 and the-X antenna plate 16, the + Y battery plate 12 is erected on the + X + Y thermal control plate 6, the-X + Y thermal control plate 7, the + Y side plate 10, the + X antenna plate 15 and the-X antenna plate 16, and the-Y battery plate 13 is erected on the + X-Y thermal control plate 8, the-X-Y thermal control plate 9, the-Y side plate 11, the + X antenna plate 15, On the X antenna board 16, the base plate 5, the + Y battery board 12, the-Y battery board 13 and the-Z battery board 14 are in an isosceles trapezoid frame structure, the-Z battery board 14 is the top surface of an isosceles trapezoid, the base plate 5 is the bottom surface of the isosceles trapezoid, the + Y battery board 12 and the-Y battery board 13 are two inclined surfaces of the isosceles trapezoid, the length direction of the base plate 5 is parallel to the launching direction of the launch vehicle 2, and the distance between the-Z battery board 14 and the base plate 5 is smaller than or far smaller than the length of the base plate 5.
In order to improve the structural rigidity and reduce the weight, the rest structural plates except the + X antenna plate 15 and the-X antenna plate 16 are all designed into honeycomb sandwich structural plates; the + Y cell panel 12, -Y cell panel 13 and-Z cell panel 14 need to meet the insulating installation requirement of the solar cell piece, and a carbon fiber reinforced resin panel aluminum honeycomb core sub-interlayer structural plate is adopted; the bottom plate 5, + X + Y heat control board 6, + X-Y heat control board 7, -X + Y heat control board 8, -X-Y heat control board 9, + Y curb plate 10, -Y curb plate 11 have heat conduction and deformation matching requirements, adopt the aluminium panel aluminium honeycomb core sub-sandwich structure board; the + X antenna plate 15 and the-X antenna plate 16 adopt a structure and load function integrated design scheme according to the antenna load design requirement; the structural connecting piece 17 is formed by machining aluminum alloy, and the structural connecting piece 17 assembles all parts into an integral star structure; the satellite-rocket joint 3 is also formed by machining aluminum alloy and then solidified in the bottom plate 3 in a pre-embedded mode.
During satellite launching, the force transmission path is as follows: the equipment load arranged on the bottom plate 5 is directly transmitted to the satellite-rocket joint 3; the rest equipment loads are transmitted to the bottom plate 5 through the + Y side plate 10, the-Y side plate 11, the + X + Y thermal control plate 6, the-X + Y thermal control plate 7, the + X-Y thermal control plate 8, the-X-Y thermal control plate 9, the + X antenna plate 15 and the-X antenna plate 16, and finally transmitted to the satellite-rocket joint 3.
As the size of the one-rocket multi-satellite parallel launching satellite along the axial direction of the carrier rocket is far larger than the size of the one-rocket multi-satellite parallel launching satellite along the transverse direction of the carrier rocket, the distance from the mass center of the wall-mounted satellite to the satellite-rocket connecting interface is far smaller than that of the bottom connecting satellite by combining the force transmission path analysis of the satellite 1 in the satellite launching process. Further analysis shows that the satellite 1 with the same gravity and the same size can reduce the load of the satellite rocket joint 3 by 67.7 percent compared with the connection of the wall-mounted type and the carrier rocket 2 with the bottom.
As shown in fig. 6, the satellite is symmetric about the ZOX reference plane. The cross section in the Y-direction width range of the-Z panel 14 from the reference plane ZOX to the + Y and-Y directions is a trapezoid as shown in fig. 8, and this part of the space is used for installation and operation of devices such as a storage tank and an antenna, and the space utilization rate of this part of the devices is 100% (the space utilization rate of the devices refers to the percentage of the total space in the satellite space that is the sum of the volume of the devices themselves and the volume occupied by the installation and operation space). As shown in fig. 7, 8 and 9, the cross section of the region outside the Y-direction width of the-Z panel 14 from the ZOX reference plane to the + Y and-Y directions is a rectangle of varying height, the height of the rectangle decreasing with increasing distance from the ZOX reference plane.
The regions of the Y-direction width of the-Z panel 14 from the ZOX datum in the + Y and-Y directions shown in figure 10 have large devices 18 and small devices 19 which occupy the volume within the satellite (including the corresponding installation operating space) as shown in the shaded area of figure 10. Considering the equipment space utilization alone from the ZOX datum in the + Y and-Y directions outside the Y-width of the-Z panel 14, the variable cross-section design shown in figure 10 with the cross-section varying along the full star Y direction corresponds to an equipment space utilization of 74.2%. If the heights of the cross sections of the areas outside the Y-direction width range of the Z-Z battery plate 14 from the ZOX reference plane to the + Y direction and the-Y direction in the figure 10 are all the heights within the Y-direction width range of the Z-Z battery plate 14, namely the cross sections of the areas outside the Y-direction width range of the Z-Z battery plate 14 from the ZOX reference plane to the + Y direction and the-Y direction are designed to be the uniform sections which are not changed along the whole star Y direction, the utilization rate of the equipment space corresponding to the scheme is 49.1%. Therefore, the space utilization rate is improved by 51.1% by adopting the variable cross-section design relative to the equal cross-section design, and the fundamental reason is that the variable cross-section design can be combined with the sizes of different equipment on the satellite, so that the equipment with large size is placed at the part with large cross section, and the equipment with small size is placed at the part with small cross section.
In order to improve the adaptability of the small satellite to different launch environments of the carrier rocket, the rigidity adaptability of the satellite structure to different carrier rockets needs to be improved, the frequency coupling between the satellite structure and the carrier rocket is avoided, and the adoption of the satellite structure dominant frequency adjustable design is the most effective measure for solving the problems. The specific implementation mode is as follows: as shown in fig. 11, 3 stiffness adjusting screws 20 are arranged near each satellite-rocket joint 3 along the Y direction, and the stiffness adjusting screws 20 have different Y-direction supporting spans L corresponding to the bottom plate in a fully or partially removed state (the stiffness adjusting screws 20 adopt a satellite structure to connect a commonly used M5 titanium alloy screw, the screw is not glued and loose-proof during stiffness adjustment, and is glued and loose-proof after a final state is set). Analogy the transverse vibration frequency f of the simply supported beam at two ends is in negative correlation with the span l, and the first-order frequency f of the satellite in the X directionxZ-direction first order frequency fzinversely related to the support span L. In addition, the support span L also influences the first-order frequency f of the satellite Y direction by influencing the support boundary rigidityyThe smaller L, the better the support rigidity, fyThe larger, therefore fyAnd is also inversely related to L. By passingObtaining the whole satellite modal analysis results at different spans L, finding out the first order frequency f of the satellite as shown in FIG. 12x、fy、fzThe relationship with L. The corresponding rigidity adjusting screw 20 is disassembled to obtain the range of L between 0.471 and 0.960m and the corresponding fx、fy、fzThe adjustable ranges are respectively 24.9-32.9 Hz, 26.3-40.1 Hz and 43.4-55.9 Hz, and the whole satellite main frequency is adjustable on the premise of not increasing the weight.
The satellite 1 passes finite element simulation analysis verification, ground vibration mechanics test verification and actual flight verification.
According to the wall-mounted type main frequency adjustable variable-section one-rocket multi-satellite launching structure, the height dimension of the whole satellite in the direction vertical to the bottom plate 5 is designed to be far smaller than the length dimension of the bottom plate 5, the length direction of the bottom plate 5 is parallel to the axis direction of the carrier rocket 2, the whole satellite is connected with the carrier rocket 2 through the bottom plate 5, and then the whole satellite is connected to the cylindrical surface of the satellite adapter of the carrier rocket 2 in a wall-mounted mode. Meanwhile, a plurality of satellites can be connected in a wall-mounted manner according to the available space in the fairing of the carrier rocket, and the one-rocket multi-satellite launching with the satellites uniformly distributed along the circumferential direction of the cylindrical surface of the satellite adapter is realized.
In order to improve the utilization rate of the equipment space, the satellite adopts a variable cross-section structure design. The large-size equipment is placed at the part with the large section size, and the small-size equipment is placed at the part with the small section size, so that the area of the structural plate can be reduced, the structural weight can be reduced, and the rigidity of the whole star can be improved.
The local rigidity of the structure on the main force transmission path has great influence on the main frequency of the satellite, so that the adjustable design of the main frequency of the satellite can be realized by adjusting the rigidity of the structure. The specific adjustment measures are that a plurality of rigidity adjusting connecting screws are arranged near each satellite-rocket joint, and the support span and the local rigidity of the bottom plate are changed by adopting different connecting states, so that the main frequency of the whole satellite can be adjusted.
The invention provides a wall-mounted main frequency adjustable variable cross-section one-rocket multi-satellite launching structure, which has the following advantages:
(1) By adopting a wall-mounted structure configuration and a point type satellite-rocket connection mode, the stress of a satellite-rocket connection interface structure can be reduced.
(2) For the same structure, the connection state of the rigidity adjusting screw 20 near the satellite-rocket joint 3 can be adjusted to realize simple, convenient and quick adjustment of the main frequency of the whole satellite, so that the main frequency requirements of different carrier rockets on the satellite can be met, the risk of frequency coupling between a newly-researched satellite structure and the carrier rocket is reduced, and great convenience is brought to design, production and test of the whole satellite.
(3) The structural design that the cross section is changed along the launching direction of the carrier rocket is adopted, the utilization rate of the equipment space is improved, and then the volume of the whole satellite is reduced, and the structural weight is lightened.
(4) Through test combination analysis, the satellite structure can achieve that the first-order frequency adjustable ranges of the whole satellite X, Y, Z direction are 24.9-32.9 Hz, 26.3-40.1 Hz and 43.4-55.9 Hz respectively, the stress of the satellite-rocket connection interface structure in a bottom connection mode is reduced by about 67.7%, and the utilization rate of the in-satellite equipment installation space is improved to 74.2% from 49.1% of the equal-section satellite structure.
(5) The structure of the wall-mounted satellite, the point-type satellite-rocket connection mode, the main frequency adjustable structure design and the variable cross-section structure design are combined, and the obtained satellite structure can be popularized to a small satellite constellation launched by one rocket and multiple satellites.
The invention provides a wall-mounted main frequency adjustable variable-section one-rocket multi-satellite launching moonlet structure, belongs to the field of spacecraft structures, and is particularly suitable for one-rocket multi-satellite launching below 500kg and moonlet structures with basically consistent sizes.
The foregoing is a more detailed description of the invention in connection with specific preferred embodiments and it is not intended that the invention be limited to these specific details. For those skilled in the art to which the invention pertains, several simple deductions or substitutions can be made without departing from the spirit of the invention, and all shall be considered as belonging to the protection scope of the invention.
Claims (10)
1. A wall-hanging adjustable variable cross-section one-arrow-more-star transmission small satellite structure of dominant frequency, its characterized in that: the satellite is connected to a carrier rocket through a wall, and is of a variable cross-section structure along the launching direction of the carrier rocket.
2. The wall mountable dominant frequency tunable cross-section one-arrow-plus-one-star launching moonlet structure of claim 1, wherein: the satellite is provided with a satellite joint, and the satellite joint on the satellite is in point-type connection with the carrier rocket through a separation nut.
3. The wall mountable dominant frequency tunable cross-section one-arrow-plus-one-star launching moonlet structure of claim 2, wherein: the coordinate system of the satellite is O-XYZ, the satellite comprises a bottom plate, + X + Y thermal control plate, -X + Y thermal control plate, + X-Y thermal control plate, -X-Y thermal control plate, + Y side plate, -Y battery plate, + Y battery plate, -Z battery plate, + X antenna plate, -X antenna plate, wherein the + X antenna plate and the X antenna plate are respectively arranged in the middle of the bottom plate, the + X + Y thermal control plate, -X + Y thermal control plate and the + Y side plate are respectively arranged at one end of the bottom plate, the + X-Y thermal control plate, -X-Y thermal control plate and the-Y side plate are respectively arranged at the other end of the bottom plate, the-Z battery plate is erected on the + X antenna plate and the-X antenna plate, and the + Y battery plate is erected on the + X + Y thermal control plate, The X-Y solar panel is erected on the + X-Y thermal control board, the-Y side board, the + X antenna board and the-X antenna board.
4. the wall mountable dominant frequency tunable cross-section one-arrow-plus-one-star launching moonlet structure of claim 3, wherein: the base plate, the + Y battery plate, the-Y battery plate and the-Z battery plate are of an isosceles trapezoid frame structure, the-Z battery plate is the top surface of the isosceles trapezoid, the base plate is the bottom surface of the isosceles trapezoid, the + Y battery plate and the-Y battery plate are two inclined planes of the isosceles trapezoid, the length direction of the base plate is parallel to the launching direction of the carrier rocket, and the distance between the-Z battery plate and the base plate is smaller than or far smaller than the length of the base plate.
5. The wall mountable dominant frequency tunable cross-section one-arrow-plus-one-star launching moonlet structure of claim 3, wherein: the bottom plate, + X + Y heat control board, -X + Y heat control board, + X-Y heat control board, -X-Y heat control board, + Y curb plate, -Y curb plate, + Y panel, -Z panel are honeycomb sandwich structure boards.
6. The wall mountable dominant frequency tunable cross-section one-arrow-plus-one-star launching moonlet structure of claim 3, wherein: the positive Y battery board, the negative Y battery board and the positive Z battery board are all carbon fiber reinforced resin panel aluminum honeycomb core sub-interlayer structural boards, and the bottom board, the positive X + Y thermal control board, the negative X + Y thermal control board, the positive X-Y thermal control board, the negative X-Y thermal control board, the positive Y side board and the negative Y side board are all aluminum panel aluminum honeycomb core sub-interlayer structural boards.
7. The wall mountable dominant frequency tunable cross-section one-arrow-plus-one-star launching moonlet structure of claim 3, wherein: the satellite-rocket joint is formed by machining aluminum alloy, and is solidified in the bottom plate 3 in a pre-embedded mode.
8. The wall mountable dominant frequency tunable cross-section one-arrow-plus-one-star launching moonlet structure of claim 7, wherein: and 3 rigidity adjusting screws are arranged near each satellite-rocket joint along the Y direction, the rigidity adjusting screws have different Y-direction supporting spans corresponding to the bottom plate in a fully or partially disassembled state, the rigidity adjusting screws adopt M5 titanium alloy screws, the rigidity adjusting screws are not glued and loose-proof during rigidity adjustment, and the screws are glued and loose-proof after the final state is determined.
9. The wall mountable dominant frequency tunable cross-section one-arrow-plus-one-star launching moonlet structure of claim 3, wherein: the satellite is symmetrical about the ZOX reference plane, and the devices are mounted on the bottom plate.
10. The wall mountable dominant frequency tunable cross-section one-arrow-plus-one-star launching moonlet structure of claim 1, wherein: six satellites are uniformly distributed around the carrier rocket at intervals in the circumferential direction.
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| CN201910800680.8A CN110562499B (en) | 2019-08-28 | 2019-08-28 | Wall-mounted main frequency adjustable variable-section one-rocket multi-satellite launching moonlet structure |
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| CN201910800680.8A CN110562499B (en) | 2019-08-28 | 2019-08-28 | Wall-mounted main frequency adjustable variable-section one-rocket multi-satellite launching moonlet structure |
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| CN110562499B (en) | 2022-09-09 |
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