JP2015182696A - Method for assembling body structure for spacecraft - Google Patents

Abstract

PROBLEM TO BE SOLVED: To provide a method for assembling a resin body structure for a spacecraft, which reliably protects a spacecraft body against vibrations and impacts at a launch, and which achieves significant weight saving and manufacturing cost reduction.SOLUTION: A body structure comprises shells 14 and 15 made of a PEEK resin, and assumes a truncated octahedron shape. The shells 14 and 15 face each other to make fusion margins 46 face each other. A single fusion-bonding sheet 92 is arranged between the fusion margins 46. The pair of fusion margins 46 are heated and pressed by a pressing machine 80 in a state in which the fusion-bonding sheet 92 is sandwiched between them.

Description

  The present invention relates to a method for assembling a structure of a spacecraft that orbits the earth and other planets represented by artificial satellites or a deep space.

  A space development project has a process in which a spacecraft (typically an artificial satellite) developed for the main purpose of the project is launched into space by a rocket. In order to make the launch of this rocket efficient and meaningful, it is common for an artificial satellite or the like developed for a purpose different from the main purpose to piggyback on the launch of the rocket.

  Such a shared satellite is subject to various restrictions. For example, external size, mass, natural frequency, etc. In particular, the mass of the satellite greatly affects the launch cost, and it is said that a mass increase of only 1 g increases the launch cost by about 100,000 yen. For this reason, artificial satellites are required to be lighter. On the other hand, extremely severe vibrations and impacts are applied to the satellite at launch. In order to prevent damage to the satellite body due to vibration or impact, the satellite body is generally provided with a structure (see, for example, Patent Documents 1 to 3).

JP 2013-184574 A JP 2003-291898 A JP 2003-291899 A

  Conventional satellite structures are generally rectangular parallelepiped, and are configured by fastening a plurality of panels made of thermosetting resin reinforced with aluminum alloy or carbon fiber with bolts. In order to reduce the weight, each panel is subjected to complicated thin-wall processing (for example, cutting with a milling cutter) or a honeycomb sandwich structure is adopted for the panel.

  However, the structure must withstand the vibrations and shocks when the rocket is launched, so there is a limit to its weight reduction. Moreover, although each panel is fastened with a volt | bolt, in such a structure, sufficient rigidity of a structure is hard to be implement | achieved. Furthermore, a large number of bolts are required so that the fastening of each panel is not loosened by vibration at the time of launch. Therefore, the bolt fastening operation becomes enormous and the assembly cost of the artificial satellite increases. In addition, the use of a large number of bolts increases the possibility that one of the bolts will loosen due to vibration or impact at the time of launch, and if one bolt is loosened, there is a risk of resonance with the rocket. Becomes higher.

  The present invention has been made based on such a background, and the purpose of the present invention is to provide a spacecraft that reliably protects the spacecraft body from vibrations and shocks during launch, and that can significantly reduce weight and reduce manufacturing costs. It is to provide an assembly method for realizing the machine structure.

  (1) A method for assembling a spacecraft structure according to the present invention is a spacecraft structure comprising a polyhedron in which a space for housing a spacecraft body is formed by heat-sealing a plurality of resin panels. This is an assembly method. The assembling method includes a first step in which a single thermoplastic resin sheet is disposed between opposing surfaces of a pair of adjacent resin panels among the plurality of resin panels, and the adjacent resin panels. Includes a second step in which the facing surface is pressed at a predetermined pressure along the facing direction and heated to a predetermined temperature by the heating element.

  Since the spacecraft structure is made of resin, a significant weight reduction can be realized as compared with a conventional metal structure. Therefore, when this spacecraft structure is composed of a plurality of resin panels, it is not necessary to design the strength to withstand a large external force when determining the joining means for each resin panel. In this invention, when the resin panels are combined, in a state where a single thermoplastic resin sheet is disposed between adjacent resin panels, the resin panels are pressed while being heated by the heating elements. Is done. At this time, the thermoplastic resin sheet is melted, and the adjacent resin panels are fused.

  (2) The spacecraft structure is a truncated octahedron formed by a pair of resin panels being abutted, and the rectangular fusion allowance formed at the abutting portion of each resin panel is the opposite It is preferable that the resin panel and the thermoplastic resin sheet include a PEEK resin. The heating element preferably includes a rod heater and also serves as a pressing means for pressing the fusion allowance.

  In this configuration, since the resin panel includes the PEEK resin, the function of the spacecraft structure is not impaired even when exposed to a high temperature environment in outer space. Further, the fusion margins of the resin panels are heated and simultaneously pressurized by a heating element having a rod heater. According to this method, the operation of fusing the resin panel is simple and the operation time is shortened.

  According to this invention, the resin panels are joined by heat fusion, and the spacecraft structure composed of a polyhedron made of resin is assembled. Since the resin panels are bonded to each other by the heat sealing means, the bonded portions are bonded evenly. That is, the spacecraft structure is stable as a structure and exhibits high resistance to external forces (tensile / compressive force, torsion, bending, etc.). In addition, since the spacecraft structure does not have a bolt fastening structure as in the prior art, the assembly work is simple and the assembly cost is greatly reduced.

FIG. 1 is an external perspective view of a spacecraft employing a structure according to an embodiment of the present invention. FIG. 2 is an exploded perspective view of a main part of the spacecraft according to the embodiment of the present invention. FIG. 3 is an exploded perspective view of the structure according to the embodiment of the present invention. FIG. 4 is a developed schematic view of the shell constituting the structure according to one embodiment of the present invention. FIG. 5 is a perspective view of a shell according to an embodiment of the present invention. 6 is a cross-sectional view taken along line VI-VI in FIG. FIG. 7 is a schematic diagram showing a procedure for heat-sealing shells according to an embodiment of the present invention. FIG. 8 is a schematic diagram showing a structure assembling method according to a modification of the present embodiment.

  Hereinafter, preferred embodiments of the present invention will be described with reference to the drawings as appropriate.

  FIG. 1 is an external perspective view of a spacecraft employing a structure according to an embodiment of the present invention. FIG. 2 is an exploded perspective view of the main part of the spacecraft.

  This spacecraft 10 functions as a planet for communication with deep space, and its outer size is designed to fit in a cubic space with a side of 50 cm. The spacecraft 10 has a mass of 50 kg or less and is a so-called micro artificial planet. The spacecraft 10 includes a spacecraft main body 11 and a structure 12 that surrounds the spacecraft main body 11. The spacecraft 10 operates with electric power, and its energy source is sunlight. Therefore, a solar cell (not shown) is mounted on the surface of the structure 12 of the spacecraft 10.

  The spacecraft 10 according to the present embodiment is characterized in that the structure 12 is provided, and that the structure 12 is assembled by a method described later. This structure 12 is formed in the shape described later from the material described later, and realizes a good weight and strength as well as vibration resistance and impact resistance, as well as a significant weight reduction compared to a conventional spacecraft structure. Yes.

  The structure 12 is made of a resin (CFRP) reinforced with carbon fibers. In this embodiment, PEEK (polyether ether ketone) resin is adopted as the resin. However, it is not limited to PEEK resin, and various thermoplastic resins or thermosetting resins can also be employed. In addition, glass fiber reinforced resin (GFRP) containing glass fibers may be employed instead of CFRP. Furthermore, PEEK resins and other resins that are not reinforced with glass fibers and carbon fibers may be employed.

  FIG. 3 is an exploded perspective view of the structure 12.

  The structure 12 defines the outer shape of the spacecraft 10, and in this embodiment, the structure 12 is formed in a truncated octahedron as shown in FIG. The structure 12 includes a pair of shells 14 and 15 (corresponding to “resin panel” recited in the claims), and the truncated octahedron is configured by disposing them in opposition. . As shown in FIG. 3, the shell 14 has a major surface 16 composed of four regular hexagons and a minor surface 17 composed of three squares. That is, the shell 14 has seven faces, and thus the structure 12 is a structure having fourteen faces.

  FIG. 4 is a developed schematic view of the shell 14.

  As shown in the figure, the shell 14 includes main plates 18, 19, 20, and 21 (forming the main surface 16) and sub plates 22, 23, and 24 (forming the sub surface 17). The main plates 18 to 21 have a regular hexagonal shape. Main plates 19, 20, and 21 are connected to outer sides 25, 26, and 27 of the main plate 18, respectively. Further, sub plates 22, 23, and 24 are continuous with the outer sides 28, 29, and 30 of the main plate 18, respectively. Then, the main plates 19 to 21 and the sub plates 22 to 24 are bent along the outer sides 25 to 30 of the main plate 18 to form the vessel-shaped shell 14 shown in FIG.

  In this embodiment, the shell 14 is formed using a prepreg or a semi-preg. That is, a vessel-like shell 14 is formed by heating and pressing a sheet (prepreg or semi-preg) in which carbon fiber is impregnated with PEEK resin, for example, using a required mold. In this case, since the single prepreg or semi-preg is a thin sheet, a plurality of prepregs or semi-pregs are laminated to form a shell 14 having a thickness of 2 mm. Note that the thickness of the shell 14 can be set to 1.0 mm to 4.0 mm.

  A hole 31 is provided in the main plate 19 of the shell 14. The shape of the hole 31 is a regular hexagon, whereby a band portion 32 protruding inward is formed on each side of the main plate 19. This belt part 32 is a part to which the spacecraft main body 11 is attached. In the present embodiment, the flange 47 (see FIGS. 1 and 2) of the spacecraft main body 11 is fastened to the band portion 32 by bolts or the like. A window 33 is provided on the main plate 20. The window 33 is formed by making a hole in the main plate 20. The window 33 has the same shape as the hole 31 provided in the main plate 19 and is used as a work window when the spacecraft 10 is assembled. A lid 13 (see FIG. 1) is attached to the window 33. The lid 13 is fastened with a bolt or the like, or bonded by thermal fusion. Carbon fiber reinforced resin can be adopted as the material of the lid 13. However, when radio wave transmission is required, glass fiber reinforced resin is used as the material of the lid 13.

  FIG. 5 is a perspective view of the shell 14 and shows the inner wall surface of the shell 14. 6 is a cross-sectional view taken along the line VI-VI in FIG.

  The shell 14 has the four main plates 18 to 21 and the three sub plates 22 to 24 as described above. Other main plates 19 to 21 and sub plates 22 to 24 are arranged around the main plate 18 (see FIG. 4), but the surfaces adjacent to each other are bent at the boundary portion. In the present embodiment, a reinforcing sheet 45 is provided along this boundary portion. The reinforcing sheet 45 is formed in an elongated strip shape and covers the boundary portion. The reinforcing sheet 45 may be made of PEEK resin reinforced with carbon fibers. The reinforcing sheet 45 may be omitted.

  Since the shell 15 has the same configuration as the shell 14, the description thereof is omitted. As shown in FIG. 3, the shell 12 is abutted against the shell 14 to form the structure 12. Sites where adjacent shells 14 and 15 are abutted (corresponding to “butting portions” described in the claims), that is, outer sides 34, 35, 36 of the main plate 19, outer sides 37 of the main plate 20, 38, 39, the outer sides 40, 41, 42 of the main plate 21, the outer side 43 of the sub-plate 22, the outer side 44 of the sub-plate 23 and the outer side 45 of the sub-plate 24 are heat-sealed to the opposing shell 15, It is fixed. In the present embodiment, a fusion allowance 46 is formed on the outer sides 34 to 45. By joining the shell 14 and the shell 15, the fusion allowances 46 come into contact with each other. By fusing the fusion allowance 43, the operation of joining the shell 14 and the shell 15 becomes easy, and the shells 14 and 15 are firmly fused. However, the fusion allowance 43 may be omitted.

  FIG. 7 is a schematic diagram showing a procedure in which the shell 14 and the shell 15 are heat-sealed.

  When the shell 14 and the shell 15 are abutted as described above (see FIG. 3), the fusion allowances 46 facing each other face each other. FIG. 7 shows this state. One fusion sheet 92 is disposed between the opposing surfaces 91 of the pair of fusion allowances 46. The fusion sheet 92 is made of PEEK resin (prepreg) reinforced with carbon fibers. However, when radio wave permeability is supplied to the spacecraft 10, glass fibers that do not exhibit electrical conductivity may be employed instead of carbon fibers. Further, the fusion sheet 92 can be made of PEEK resin or other resin that is not reinforced with fibers. The thickness of the fusion sheet 92 is 0.2 mm in this embodiment. However, the thickness of the fusion sheet 92 is not limited to this.

  The fusion allowance 46 and the fusion sheet 92 are sandwiched by a press device 95. The press device 95 includes a lower die 96 and an upper die 97, and the upper die 97 and the lower die 96 come in contact with each other. The upper die 97 and the lower die 96 incorporate a rod heater 89 (corresponding to “heating element” described in claims). When the rod heater 89 generates heat, the upper die 97 and the lower die 96 are heated. When the upper die 97 and the lower die 96 are clamped while the rod heater 89 generates heat, the fusion allowance 46 and the fusion sheet 92 are pressed in the direction of the arrow 98. Since the fusion sheet 92 is a thermoplastic resin sheet, it is melted by being heated, and as a result, the adjacent fusion allowances 46 face each other in the direction of the arrow 98 and are fused.

  The structure 12 according to this embodiment includes shells 14 and 15 configured as resin molded products. Therefore, the structure 12 is light and secures good strength and rigidity, and also has excellent vibration resistance and vibration resistance. Moreover, since the said reinforcement sheet 45 is provided in the boundary part of each surface which comprises the structure 12, the intensity | strength and rigidity of the structure 12 increase significantly. Furthermore, since the shells 14 and 15 are fixed not by fastening with bolts or the like but by heat fusion, the assembly work of the structure 12 is simple and the assembly cost is greatly reduced.

  In addition, since the shells 14 and 15 are resin molded products, the surface of the structure 12 becomes extremely smooth. Combined with the fact that the thermal conductivity of PEEK resin is much lower than that of metal, the structure 12 coated with an appropriate paint suppresses the absorption of solar heat (infrared radiation) in outer space, and the spacecraft body 11 Temperature change is suppressed. When the structure is made of metal as in the prior art, the surface of the structure has to be made smooth to suppress the absorption of solar heat, but this processing cost is enormous. Compared to this, in this embodiment, the problem of the processing cost does not occur.

  By the way, the power source of the spacecraft 10 is electric power. For this purpose, a solar cell panel (not shown) is arranged on the structure 12. In this embodiment, the structure 12 is a truncated octahedron. Therefore, if the structure 12 is a polyhedron, the main plates 18 to 21 and the like can secure a sufficient area for arranging the solar cell panel. Moreover, since the truncated octahedron approximates a sphere as a structure, thermal control is facilitated when the structure 12 rotates in outer space. Further, the shell 14 and the shell 15 are bonded via the outer sides 34 to 42. That is, since the corrugated joint surface goes around the structure 12, there is an advantage that the torsional rigidity of the structure 12 becomes extremely high.

In particular, in this embodiment, the structure 12 is made of PEEK resin having a thickness of 1.0 mm to 4.0 mm. In this embodiment, the carbon fiber contained in the PEEK resin is a spread yarn fabric of 100 g / m ² · 20 ply, the fiber volume ratio is 41%, the fiber elastic modulus is 40 tf / mm ² , and the PEEK resin is 390 °. C, pressed for a predetermined time under the condition of 30 kgf / cm ² . As a result, the structure 12 is extremely light and has good strength and rigidity. Although the thickness of the structure 12 is set to 2 mm in the present embodiment as described above, the thickness is appropriately changed depending on the mechanical strength and rigidity required for the structure 12.

  Further, the shells 14 and 15 constituting the structure 12 are formed from a prepreg cut into a shape as shown in FIG. Therefore, the shells 14 and 15 are formed with high dimensional accuracy, and as a result, the mechanical strength of the structure 12 is finished as designed.

  In the present embodiment, the shell 14 and the shell 15 are heat-sealed through the single fusion sheet 92 as described above, and the joint portions are bonded evenly. Therefore, the structure 12 is stable as a structure and exhibits high resistance to external forces (tensile / compressive force, torsion, bending, etc.). Moreover, since the structure 12 does not have a bolt fastening structure as in the prior art, the assembling work is simplified and the assembling cost is greatly reduced.

  FIG. 8 is a schematic view showing a method for assembling the structure 12 according to a modification of the present embodiment. As shown in the figure, in this modification, a simple handy type press 80 is used.

  In this embodiment and the modification, the thickness of the fusion allowance 46 is about 2 mm. That is, the thickness of the fusion allowance 46 is thin, and sufficient heat fusion is possible even with the gripping force of a human hand. In this modification, the press machine 80 includes a pair of operation arms 81 and 82. Each operation arm 81 and 82 is operated by an operator by hand. The operation arms 81 and 82 are crossed and connected via a support shaft 85. The operation arms 81 and 82 are rotatable about the support shaft 85, and when the operator operates the base end portions 86 of the operation arms 81 and 82, the distal end portions 87 are brought into contact with and separated from each other. Press stands 83 and 84 are provided at the distal end portion 87. The press stands 83 and 84 are made of a material having a high thermal conductivity such as a metal, and are fastened to the tip portion 87 by a bolt 79. Rod heaters 89 are built in the press stands 83 and 84, respectively. A power supply device (not shown) is connected to the rod heater 89, and the amount of heat generated by the rod heater 89 is controlled by this power supply device. Since the press tables 83 and 84 have high thermal conductivity as described above, the temperature of the press tables 83 and 84 rises evenly when the rod heater 89 generates heat.

As shown in the figure, when the shell 14 and the shell 15 are abutted, the fusion allowance 46 is opposed. The fusion sheet 92 is sandwiched between the fusion allowances 46, and the fusion allowances 46 are pressed. Specifically, the operator holds the operation arms 81 and 82 and presses the press tables 83 and 84 against the fusion allowance 46. Pressing machine 80 by the pressing force at this time ^{is 10kgf / cm 2 ~20kgf / cm 2} . With such a pressing force, the handy-type press 80 according to the present modification is sufficiently exerted. The heating temperature of the fusion sheet 92 is 385 ° C. to 405 ° C., and the heating and pressing time is 5 minutes to 15 minutes. In any condition, the fusion allowance 46 is uniformly fused over the entire area of the facing surface 91, and the shell 14 and the shell 15 are easily and reliably joined.

  By the way, in this embodiment, the structure 12 is formed in the truncated octahedron. However, the outer shape of the structure 12 is not limited to the truncated octahedron, and various polyhedrons can be adopted. However, as described above, when a general-purpose solar cell panel is employed, a plane portion on which this can be arranged is necessary. Further, if a solar cell panel that can be installed on a curved surface can be adopted, the outer shape of the structure 12 may be a sphere.

  Further, in the present embodiment, the rod heater 89 is adopted as the heating element that generates heat when the fusion allowance 46 is heat-sealed, but a nichrome wire or an enamel wire may be adopted instead.

DESCRIPTION OF SYMBOLS 10 ... Spacecraft 11 ... Spacecraft body 12 ... Structure 14 ... Shell 15 ... Shell 18 ... Main plate 19 ... Main plate 20 ... Main plate 21 ... Main plate 22 ..Sub-plate 23 ... Sub-plate 24 ... Sub-plate 31 ... Hole 32 ... Strip 33 ... Window 46 ... Fusion allowance 80 ... Press machine 81 ... Operation Arm 82 ... Operating arm 83 ... Press stand 84 ... Press stand 85 ... Support shaft 86 ... Base end 87 ... Front end 89 ... Rod heater 91 ... Opposite surface 92 ... Fusing sheet 95 ... Press machine 96 ... Upper die 97 ... Lower die

Claims (2)

A method of assembling a spacecraft structure comprising a polyhedron in which a space for housing a spacecraft body is formed by heat-sealing a plurality of resin panels,
A first step in which a single thermoplastic resin sheet is disposed between the opposing surfaces of a pair of adjacent resin panels among the plurality of resin panels;
A method of assembling a spacecraft structure, comprising: a second step in which the adjacent resin panels are pressed at a predetermined pressure along a direction in which the facing surfaces face each other and heated to a predetermined temperature by a heating element. The spacecraft structure is a truncated octahedron formed by a pair of resin panels being abutted,
The rectangular fusion margin formed at the abutting portion of each resin panel forms the facing surface,
The resin panel and the thermoplastic resin sheet contain PEEK resin,
2. The method of assembling a spacecraft structure according to claim 1, wherein the heating element includes a rod heater and also serves as a pressing unit that presses the fusion allowance.

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