US20200010220A1 - Self-Mating Modular Satellite Bus - Google Patents
Self-Mating Modular Satellite Bus Download PDFInfo
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- US20200010220A1 US20200010220A1 US16/028,797 US201816028797A US2020010220A1 US 20200010220 A1 US20200010220 A1 US 20200010220A1 US 201816028797 A US201816028797 A US 201816028797A US 2020010220 A1 US2020010220 A1 US 2020010220A1
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- B64—AIRCRAFT; AVIATION; COSMONAUTICS
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Abstract
Description
- The structural bodies of orbital satellites, space-faring probes, and some spacecraft are typically handmade. These structural bodies, referred to herein as spacecraft buses, are normally machined, cut, bonded, or molded, using very labor-intensive techniques. Each resulting spacecraft bus is semi-custom. Due to the handmade nature of these items, portions of a spacecraft bus tend to be irregular, at least to some extent. Not only are spacecraft buses time-consuming to manufacture, but the cost is considerable as well. Furthermore, the resulting buses are not uniform.
- CubeSats have been recently introduced and comprise satellite devices having a uniform, albeit small, size that allows for a more consistent manufacturing process. Nevertheless, the efficiency of CubeSat production can be improved. Moreover, the small size of CubeSats limits the quantity of internal components as well as the provided functionality.
- A satellite bus, as described herein, includes a plurality of side panels each having a front surface flanked by a first longitudinal edge and a second longitudinal edge, wherein the first longitudinal edge of each side panel is nested with the second longitudinal edge of an adjacent side panel. The individual satellite panels may each include a first flange along the first longitudinal edge and a second flange along the second longitudinal edge, and the first flange superimposed on the second flange forms a joggle.
- A method of producing a modular satellite bus may include forming a plurality of the side panels described above, aligning a first longitudinal edge of each side panel with a second longitudinal edge of an adjacent side panel, such that the first longitudinal edge of each side panel is nested with the second longitudinal edge of the adjacent side panel, and securing the first longitudinal edge of each side panel to the second longitudinal edge of the adjacent side panel.
- Many aspects of the disclosure can be better understood with reference to the following drawings. The components in the drawings are not necessarily to scale, emphasis instead being placed upon clearly illustrating the principles of the present disclosure. Moreover, in the drawings, like reference numerals designate corresponding parts throughout the several views. While several embodiments are described in connection with these drawings, the disclosure is not limited to the embodiments disclosed herein. On the contrary, the intent is to cover all alternatives, modifications, and equivalents.
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FIG. 1 is a perspective view of a self-mating modular satellite bus. -
FIG. 2 is a perspective view of a single side panel of the satellite bus. -
FIG. 3A is a front view of the side panel of the satellite bus. -
FIG. 3B is a top view of the side panel inFIG. 3A . -
FIGS. 3C and 3D are close-up views of the side panel indicated inFIG. 3B . -
FIG. 4A is a front view of the satellite bus inFIG. 1 . -
FIG. 4B is another front view of the satellite bus inFIG. 1 rotated 30° from the view inFIG. 4A . -
FIG. 4C is a top view of the satellite bus inFIG. 4A . -
FIG. 4D is a side view of a top panel of the satellite bus inFIG. 4A . -
FIG. 5 is an exploded view of a self-mating modular satellite bus. - The
satellite bus 20, as shown inFIG. 1 , is the main structural component or framework of a satellite. Thesatellite bus 20 is self-mating in the sense that each of the side panels of the satellite bus fit together without requiring additional support materials. Self-mating can also be referred to herein as self-jigging. Fasteners or adhesives may be used to secure the side panels to one another, but no additional framework materials are required to build thesatellite bus 20. - A
side panel 22 of thesatellite bus 20 is shown inFIG. 2 . Thesatellite bus 20 is modular in the sense that eachside panel 22 is a standardized part that can be formed from a single die or mold, such that eachside panel 22 is virtually identical to the other side panels. - A
front surface 24 of theside panel 22 is shown inFIG. 3A . Thefront surface 24 is flanked by a firstlongitudinal edge 26 and a secondlongitudinal edge 30. When theside panels 22 are arranged to form thesatellite bus 20, the firstlongitudinal edge 26 of eachside panel 22 is nested with the secondlongitudinal edge 30 of an adjacent side panel. -
FIG. 3B is a top view of theside panel 22 shown inFIG. 3A , illustrating one example of themodular side panel 22 that can be nested with adjacent panels. As shown inFIG. 3B , the firstlongitudinal edge 26 has afirst flange 28 and the secondlongitudinal edge 30 has asecond flange 32. Thefirst flange 28 superimposed on thesecond flange 32 together form a joggle, such that the first andsecond flanges second flanges -
FIGS. 3C and 3D are close-up views of theside panel 22, particularly theflanges FIG. 3B . In this particular example, eachflange longitudinal edge flange front surface 24, but slightly offset with respect to one another. More particularly, theflanges first flange 28 set back from thefront surface 24 and thesecond flange 32 set back from aback surface 34 of theside panel 22. The offset front-and-back spacing of theflanges flanges adjacent side panels 22 to nest together to form a joggle. Theflanges FIGS. 3B, 3C, and 3D , or they may be stepped with complementary stepping or reverse bends that nest to form a joggle, or they may include any other complementary shaping that nests to form a joggle. A joggle, as referred to herein, is a joint between two adjacent pieces/panels. The joggles herein are formed from a projection of a first panel fitting into a notch of a second panel, and vice-versa. - The first
longitudinal edge 26 of eachside panel 22 may be secured to the secondlongitudinal edge 30 of theadjacent side panel 22 using any suitable means. In certain configurations the application of pressure may be sufficient to secure the first and secondlongitudinal edges first flange 28 may include one or more projections or teeth that fit into one or more recesses or notches in thesecond flange 32 to prevent slipping. Alternatively, the first andsecond flanges - A glue or adhesive is another option for securing the first and second
longitudinal edges flanges other flange - Once adhesive has been applied between two adjacent panels via first and second
longitudinal edges - Another technique for securing the first and second
longitudinal edges longitudinal edges - The
satellite bus 20 may include any suitable number ofside panels 22. As shown inFIG. 1 , thesatellite bus 20 may include sixside panels 22 forming a hexagon. When thesatellite bus 20 is in the form of a hexagon, the angle α between thefront surface 24 of eachside panel 22 andflange satellite bus 20 may include fourside panels 22 forming a square with an angle α of 45° between thefront surface 24 of eachside panel 22 andflange satellite bus 20 includes fiveside panels 22 forming a pentagon, the angle α is 36° between thefront surface 24 of eachside panel 22 andflange satellite bus 20 includes sevenside panels 22 forming a heptagon, the angle α is 25.7° between thefront surface 24 of eachside panel 22 andflange satellite bus 20 includes eightside panels 22 forming an octagon, the angle α is 22.5° between thefront surface 24 of eachside panel 22 andflange -
FIGS. 4A-4D show various views of thesatellite bus 20 illustrated inFIG. 1 .FIG. 4A shows a front view of thesatellite bus 20.FIG. 4B shows another front view of thesatellite bus 20 rotated approximately 30° from the view inFIG. 4A . Thesatellite bus 20 may include atop panel 36, best viewed inFIG. 4C , having aflange 38 that attaches thetop panel 36 to each of theside panels 22 along atop edge 40 of eachside panel 22. Specifically, theflange 38 on thetop panel 36 forms a 90° angle with thetop edge 40 of eachside panel 22.FIG. 4D shows a side view of thetop panel 36 separate from thesatellite bus 20, specifically showing theflange 38. Thetop panel 36 can be attached to thetop edges 40 of theside panels 22 using any of the techniques described above for securingadjacent side panels 22 to one another. For example, thetop panel 36 can be secured to thetop edges 40 of theside panels 22 using pressure, adhesive, and/or fasteners.Top panel 36 andside panels 22 might also have corresponding joggle features to create bonding surfaces betweentop panel 36 andside panels 22. Thetop panel 36 may include one ormore apertures 42, which may be used to accommodate an imaging component, such as a camera or for tracking stars, or for any other purpose. - The
satellite bus 20 may also include abottom panel 44, shown inFIGS. 4A and 4B . Like thetop panel 36, thebottom panel 44 may also include a flange (not shown) that attaches thebottom panel 44 to each of theside panels 22 along abottom edge 46 of eachside panel 22. Much like the configuration of theflange 38 on thetop panel 36, the flange on the base forms a 90° angle with thebottom edge 46 of eachside panel 22. Thebottom panel 44 can be attached to thebottom edges 46 of theside panels 22 using any of the techniques described above for securing thetop panel 36 to thetop edges 40 of theside panels 22, such as using pressure, adhesive, and/or fasteners.Bottom panel 44 andside panels 22 might also have corresponding joggle features to create bonding surfaces betweenbottom panel 44 andside panels 22. Thebottom panel 44 may include an adapter, such as another flange, for mounting the satellite to a rocket stage or deployer. - The
bottom panel 44 may include one or more apertures to support propulsion components. These propulsion components can include external nozzles, engines, grids, electrodes, or other suitable propulsion extensions fromsatellite bus 20. A lower chamber might be formed using a correspondinginternal deck 28 to separate propulsion components from avionics, power, and sensing components. - Overall, the
satellite bus 20 may be comparable in size to a CubeSat form factor, or thesatellite bus 20 may be larger or smaller than a CubeSat. Consequently, thesatellite bus 20 may be launched in similar types of rockets used to launch CubeSats. In particular, thesatellite bus 20 may have an overall height (y, z), 40 cm, depicted inFIG. 4B , between about 35 cm and about 45 cm, or between about 20 cm and about 35 cm, or between about 45 cm and about 75 cm. Theindividual side panels 22 may have a height (y, z), 40 cm, depicted inFIG. 4A , between about 30 cm and about 45 cm, or between about 20 cm and about 35 cm, or between about 45 cm and about 75 cm. Thesatellite bus 20 may have a diameter (w), 45.7 cm, depicted inFIG. 4C , between about 350 cm and about 45.7 cm, or between about 45.7 cm and about 75 cm, or between about 75 cm and about 100 cm. Thesatellite bus 20, on its own, may weigh between about 500 grams and about 1000 grams, or between about 750 grams and about 1500 grams, or between about 1500 grams and about 6000 grams. It should be understood that satellite bus 60 can have other dimensions and weights. - The
satellite bus 20 may either be hollow to house a single chamber of internal components, or the satellite bus may include one ormore decks 48 or other externals, as shown inFIG. 5 to house more than one chamber of internal components. Each of theside panels 22 may have one or more molded features formed into theback surface 34 facing the inside of thesatellite bus 20. The molded features may be designed to hold one ormore decks 48 or other externals in place. The molded features may include one or more tabs, fasteners, bonded areas, embossments, or shelves, either for securing adeck 48 or for any other purpose. - The
deck 48 may be a propulsion deck or an avionics deck, for example. The avionics deck may include a sun sensor, a star tracker, radio frequency (RF) transceivers, optical transceivers, reaction wheels, wire harnesses, power bus, internal heat radiator connections, and/or any other avionics features. Thedeck 48 can be attached to theback surface 34 of each of theside panels 22 using any of the aforementioned features on the back surfaces 34 of theside panels 22. Thesatellite bus 20 may also include internal heat radiator connections, such as one or more thermal straps, secured to the inside of thesatellite bus 20 to conduct heat from inside the satellite to the exterior of the satellite through one or more apertures inside panels 22,top panel 36, orbottom panel 44. - According to certain examples, each of the
side panels 22 may have a mounting flange on thefront surface 24. The mounting flange can be used to attach a solar panel or a window to thefront surface 24 of therespective side panel 22. Using a flange to mount the solar panels or windows results in fewer layers in thesatellite bus 20, thereby reducing the overall weight of thesatellite bus 20. - One of the advantages of the
satellite bus 20 is that the design is modular, so the structure is formed by panels, namelyside panels 22, a top panel, and abottom panel 44. Theside panels 22 can all be formed from the same manufacturing die. Theside panels 22 can have predetermined features to hold decks and other equipment, in accordance with a user's specifications. The predetermined features may be obtained either by using a single die that includes the desired features for each of theside panels 22, or by modifying a single die to include, for example, a molded feature, tabs, fasteners, bonded areas, embossments, or shelves. Consequently, thesatellite bus 20 can be made available as a kit and assembled by a user. Alternatively, thesatellite bus 20 can be manufactured according to user specifications, in modular form, and subsequently transferred to the user for final assembly. The internal features of the satellite may be pre-assembled and mounted inside thesatellite bus 20 before either the user or the manufacturer seals thesatellite bus 20 using associated joggles and flanges, thereby forming a satellite. - A method of producing the
modular satellite bus 20 includes forming theside panels 22, which may be achieved using a die or a set of male and female dies, or any other suitable casting or molding process. More particularly, the die can be filled with a composite material, such as laser-cut carbon, pre-impregnated fiberglass, carbon, resin, and combinations thereof. The die can then be compressed, using a jig, belts, hydraulic press, or any other suitable pressure-inducing device, to form at least one compressed composite structure in the shape of aside panel 22. After the material has been compressed in the die, any excess material extending from the die can be trimmed. Depending on the material used, it may be necessary or at least beneficial to heat or otherwise cure the resultingside panel 22 either before or after removing theside panel 22 from the die. Vacuum chambers might be employed during formation of side panels to ensure desired curing of the associated composite material. Other manufacturing techniques may be used to form theside panels 22 as well, such as additive manufacturing or 3D printing. - As described above, the
satellite bus 20 may have any reasonable number ofside panels 22, such as four, five, six, seven, or eight. Eachside panel 22 can be formed in the same manner using the same die, such that all of the resultingside panels 22 in anysatellite bus 20 are identical. - Another advantage of the
satellite bus 20 design is that only a simple modification of the process is needed to vary the number ofside panels 22 in a resultingsatellite bus 20. As described above, theside panels 22 joined together have an angle between them, which is two times the angle α between thefront surface 24 of eachside panel 22 andflange side panels 22 in thesatellite bus 20. By simply changing the angle α in the die or during additive manufacturing or other process used to form theside panels 22, essentially any polygonal shape ofsatellite bus 20 can be formed. Thus, the same die, albeit with modified flange or joggle angles α, can be used to formsatellite buses 20 of virtually any polygonal shape. - In order to form a
square satellite bus 20 having fourside panels 22, the flange or joggle angle α can be set at 45° between thefront surface 24 of eachside panel 22 andflange pentagonal satellite bus 20 having fiveside panels 22, the flange or joggle angle α can be set at 36° between thefront surface 24 of eachside panel 22 andflange hexagonal satellite bus 20 having sixside panels 22, the flange or joggle angle α can be set at 30° between thefront surface 24 of eachside panel 22 andflange heptagonal satellite bus 20 having sevenside panels 22, the flange or joggle angle α can be set at 25.7° between thefront surface 24 of eachside panel 22 andflange octagonal satellite bus 20 having eightside panels 22, the flange or joggle angle α can be set at 22.5° between thefront surface 24 of eachside panel 22 andflange - The
top panel 36 and thebottom panel 44 can each be formed according to the same processes as theside panels 22, such as die-forming or additive manufacturing. One or more holes or apertures may be stamped into thetop panel 36, thebottom panel 44, and/or theside panels 22 for various reasons, such as to accommodate propulsion components, sensors, cameras, solar arrays, antennas, star trackers, and the like. - Additionally, one or more molded features may be formed into the
back surface 34 of each of theside panels 22. These features may be designed to hold one ormore decks 48 or other externals in place inside thesatellite bus 20. For example, one or more tabs, fasteners, bonded areas, embossments, or shelves may be molded into or attached to theback surface 34 of eachside panel 22. - Furthermore, a mounting flange may be molded into or otherwise attached to the
front surface 24 of each of theside panels 22. The mounting flange can be used to attach a solar panel or a window to thefront surface 24 of therespective side panel 22, which may be done either before or after theside panels 22 are secured to one another. - Once each of the
side panels 22 for aparticular satellite bus 20 has been formed, theside panels 22 can be aligned with the firstlongitudinal edge 26 of eachside panel 22 aligned with the secondlongitudinal edge 30 of an adjacent side panel, such that the firstlongitudinal edge 26 of eachside panel 22 is nested with the secondlongitudinal edge 30 of the adjacent side panel. In certain examples, the first and secondlongitudinal edges side panels 22 in place while securing theside panels 22 to one another. However,side panels 22 are self-jigging in that each side panel nests into one or more adjacent side panels without additional tooling or jigs. As described above, the first and secondlongitudinal edges flanges longitudinal edges - Any internal features of the satellite may be manufactured separately, pre-assembled, and mounted inside the
satellite bus 20, particularly using the features that are molded or attached to theback surface 34 of each of theside panels 22 to secure the internal features in place, during assembly of thesatellite bus 20. Such internal features may include, for example, one ormore decks 48, such as a propulsion deck or an avionics deck, and corresponding features of the decks, such as a sun sensor, a star tracker, a radio, reaction wheels, and/or a wire harness. One or more thermal straps may also be secured to the inside of thesatellite bus 20 during assembly. - After the
side panels 22 are secured to one another, thetop panel 36 can be attached to each of theside panels 22 along thetop edge 40 of eachside panel 22. Also, thebottom panel 44 can be attached to each of theside panels 22 along thebottom edge 46 of eachside panel 22. - The included descriptions and figures depict specific embodiments to teach those skilled in the art how to make and use the best mode. For the purpose of teaching inventive principles, some conventional aspects have been simplified or omitted. Those skilled in the art will appreciate variations from these embodiments that fall within the scope of the disclosure. Those skilled in the art will also appreciate that the features described above can be combined in various ways to form multiple embodiments. As a result, the invention is not limited to the specific embodiments described above, but only by the claims and their equivalents.
Claims (20)
Priority Applications (2)
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PCT/US2018/046584 WO2020009711A1 (en) | 2018-07-06 | 2018-08-14 | Self-mating modular satellite bus |
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US16/028,797 US10538341B1 (en) | 2018-07-06 | 2018-07-06 | Self-mating modular satellite bus |
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US16/028,797 Active US10538341B1 (en) | 2018-07-06 | 2018-07-06 | Self-mating modular satellite bus |
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Cited By (5)
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US20210354859A1 (en) * | 2020-05-18 | 2021-11-18 | The Boeing Company | Additively manufactured satellite |
CN114435627A (en) * | 2022-02-23 | 2022-05-06 | 航天科工空间工程发展有限公司 | Satellite structure and satellite assembly method |
US11794927B2 (en) | 2019-08-28 | 2023-10-24 | The Boeing Company | Additively manufactured spacecraft panel |
US11802606B2 (en) | 2020-05-18 | 2023-10-31 | The Boeing Company | Planate dynamic isolator |
US11878819B2 (en) | 2020-12-17 | 2024-01-23 | The Boeing Company | Satellite thermal enclosure |
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EP3424823B1 (en) * | 2016-02-29 | 2021-08-18 | Canon Denshi Kabushiki Kaisha | Artificial satellite |
US11242161B1 (en) * | 2018-05-24 | 2022-02-08 | David Michael White | Cube-shaped primary structure module |
FR3104546A1 (en) * | 2019-12-17 | 2021-06-18 | Airbus Defence And Space Sas | ASSEMBLY PROCESS OF A PLURALITY OF EQUIPMENT ON A SATELLITE STRUCTURE AND A SATELLITE STRUCTURE CARRYING A PLURALITY OF EQUIPMENT |
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US11794927B2 (en) | 2019-08-28 | 2023-10-24 | The Boeing Company | Additively manufactured spacecraft panel |
US20210354859A1 (en) * | 2020-05-18 | 2021-11-18 | The Boeing Company | Additively manufactured satellite |
US11802606B2 (en) | 2020-05-18 | 2023-10-31 | The Boeing Company | Planate dynamic isolator |
US11827389B2 (en) * | 2020-05-18 | 2023-11-28 | The Boeing Company | Additively manufactured satellite |
US11878819B2 (en) | 2020-12-17 | 2024-01-23 | The Boeing Company | Satellite thermal enclosure |
CN114435627A (en) * | 2022-02-23 | 2022-05-06 | 航天科工空间工程发展有限公司 | Satellite structure and satellite assembly method |
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US10538341B1 (en) | 2020-01-21 |
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