JP2015182694A - Body structure for spacecraft - Google Patents

Abstract

PROBLEM TO BE SOLVED: To provide a 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 12 is made of a resin (CFRP) that is reinforced with carbon fibers. The resin constituting the body structure 12 is a PEEK (polyether ether ketone) resin. The body structure 12 is formed in a truncated octahedron shape. The thickness of the body structure 12 is 2 mm. The body structure 12 comprises shells 14 and 15, and both of them butt against each other.

Description

  The present invention relates to the structure of a spacecraft that orbits the earth and other planets, such as artificial satellites, and a spacecraft that explores 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 its purpose is to reliably protect the spacecraft main body from vibration and impact at the time of launch, realizing a significant weight reduction and a reduction in manufacturing cost. Is to provide a structure.

  (1) A spacecraft structure according to the present invention includes a plurality of thermoplastic resin panels. This spacecraft structure is composed of a polyhedron in which a space for accommodating the spacecraft main body is formed, and each thermoplastic resin panel is continuous or heat-sealed.

  According to this configuration, each thermoplastic resin panel is formed as a molded product. Each thermoplastic resin panel is fixed not by a fastening structure using bolts or the like but by heat fusion. The type of the resin is desirably one that can be used at an environmental temperature not higher than the glass transition temperature of the thermoplastic resin.

  (2) The polyhedron is preferably a truncated octahedron.

  The spacecraft structure has a solar cell panel disposed on the surface thereof. Therefore, when the structure is formed in a polyhedron, each surface needs to have a sufficient area for arranging the solar cell panel. In this configuration, the area of each surface is necessary and sufficient for arranging a general-purpose solar cell panel. Moreover, when the structure is approximated to a sphere and the structure rotates in outer space, it is easy to perform thermal control.

  (3) The polyhedron may be composed of two thermoplastic resin panels having the same shape and arranged to face each other. In this case, each thermoplastic resin panel has four regular hexagonal main plates and three square sub-plates, and the second to fourth main plates are continuous radially around the first main plate, The three sub-plates are heat-processed by a prepreg or semi-preg that is radially continuous around the first main plate and disposed between the second and third main plates and between the third and fourth main plates. It is preferable that

  In this configuration, the thermoplastic resin panel is manufactured by cutting a prepreg or semi-preg into a required shape and subjecting it to heat treatment. Therefore, the thermoplastic resin panel is configured with high dimensional accuracy.

  (4) The thermoplastic resin panel is preferably made of carbon fiber reinforced PEEK resin having a thickness of 1.0 mm to 4.0 mm.

  Among PEEK (polyetheretherketone) resins, one PEEK resin actually used in the industry has a melting point of about 335 ° C and a glass transition temperature of about 250 ° C. It can be used continuously in an environment of 250 ° C. or lower, and can also be used in outer space. Moreover, since the PEEK resin can be injection-molded, the thermoplastic resin panel can be easily manufactured. Further, regarding the mechanical strength of the PEEK resin, the deflection temperature under load is 140 ° C., but by being reinforced with carbon fiber, the deflection temperature under load reaches 300 ° C. or more, and the mechanical strength is sufficient. Therefore, even if the resin panel is thin (1.0 mm to 4.0 mm), the strength and rigidity of the structure can be maintained. In addition, it has excellent chemical resistance and has the advantage that it is not affected by inorganic or organic chemicals other than concentrated sulfuric acid, concentrated nitric acid, and saturated chlorine water.

  (5) It is preferable that a fusion allowance is formed in a portion of the thermoplastic resin panel to be thermally fused.

  In this structure, each thermoplastic resin panel is reliably heat-sealed, and the strength and rigidity of the thermoplastic resin panel are improved.

  (6) It is preferable that a reinforcing member is disposed at a boundary between adjacent surfaces of the thermoplastic resin panel.

  In this configuration, the strength and rigidity of the thermoplastic resin panel are further improved.

  According to the present invention, since the structure is made of carbon fiber reinforced resin (CFRP), it is extremely light and secures good strength and rigidity. Moreover, even if the structure is composed of a plurality of panels, each panel is made of a thermoplastic resin, so that it is not necessary to use bolts for joining them. Therefore, the assembly work of the structure is simple, the assembly cost is reduced, and each panel is not a structure to be fastened, so that it has excellent impact resistance and vibration resistance.

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.

  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 (typically an artificial satellite) 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 spacecraft for communication with deep space, and its outer size is designed to fit within 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.

  A feature of the spacecraft 10 according to the present embodiment is that the structure 12 is provided. This structure 12 is formed in the shape described later from the material described later, and realizes good strength and rigidity as well as vibration resistance and impact resistance as well as a significant weight reduction compared to a conventional satellite structure. Yes.

  The structure 12 is made of a thermoplastic 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 can be adopted. In addition, glass fiber reinforced resin (GFRP) containing glass fibers may be employed instead of CFRP. Furthermore, PEEK resin and other thermoplastic 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 is formed in a truncated octahedron as shown in FIGS. 1 to 3 in this embodiment. The structure 12 includes a pair of shells 14 and 15 (corresponding to the “thermoplastic resin panel” recited in the claims), and the truncated octahedron is configured by opposingly arranging them. ing. As shown in the figure, the shell 14 has four main surfaces 16 made of regular hexagons and three sub-surfaces 17 made of 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, the shell 14 is formed by heating and pressurizing a sheet (prepreg or semi-preg) in which carbon fiber is impregnated with PEEK resin using, for example, a required mold. In this case, since the single prepreg or semi-preg is a thin sheet (thickness 0.05 mm to 0.5 mm), 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.

  As shown in FIG. 4, 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. 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 (corresponding to a “reinforcing member” described in the claims) is provided along the 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. The parts where the shell 14 is abutted with the shell 15, that is, the outer sides 34, 35, 36 of the main plate 19, the outer sides 37, 38, 39 of the main plate 20, the outer sides 40, 41, 42 of the main plate 21, The outer side 43, the outer side 44 of the sub-plate 23, and the outer side 45 of the sub-plate 24 are thermally fused and fixed to the opposing shell 15.

  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 46, 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 46 may be omitted.

  The structure 12 according to the present embodiment includes shells 14 and 15 configured as molded articles made of a thermoplastic resin, and exhibits a truncated octahedron. Therefore, the structure 12 is a light-weight structure with excellent strength and rigidity, and excellent vibration resistance and impact resistance. As shown in FIG. 2, the spacecraft main body 11 is inserted from the hole 31 of the shell 14 and exposed from the hole 31 of the shell 15. The flange 47 is fixed to the hole 31 of the shell 15 and fixed to the spacecraft main body 11 (see FIG. 1). As this fixing means, fastening by bolts is typically adopted.

  Further, since the reinforcing sheet 45 is provided at the boundary portion of each surface constituting the structure 12, the rigidity of the structure 12 is significantly increased. 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 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.

  The shells 14 and 15 constituting the structure 12 are formed from a prepreg or semi-preg 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 this embodiment, the structure 12 is formed in a 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.

  Although the present invention is applied to a structure for artificial satellites as a typical spacecraft, a space shuttle, a return vehicle, a planetary spacecraft, a deep space spacecraft, or the like is assumed as a spacecraft.

DESCRIPTION OF SYMBOLS 10 ... Spacecraft 11 ... Spacecraft main body 12 ... Structure 14 ... Shell 15 ... Shell 16 ... Main surface 17 ... Subsurface 18 ... Main plate 19 ... Main plate 20 ... Main plate 21 ... Main plate 22 ... Sub plate 23 ... Sub plate 24 ... Sub plate 25 ... Outer side 26 ... Outer side 27 ... Outer side 28 ... Outer side 29 ... Outer side 30 ... Outer side 31 ... Hole 32 ... Belt part 33 ... Window 34 ... Outer side 35 ... Outer side
36 ... Outside 37 ... Outside 38 ... Outside 39 ... Outside 40 ... Outside 41 ... Outside 42 ... Outside 43 ... Fusion allowance

Claims (6)

It consists of a plurality of thermoplastic resin panels, and is a sphere or polyhedron in which a space for housing the spacecraft body is formed,
Each thermoplastic resin panel is a spacecraft structure that is continuously or thermally fused.   The structure for a spacecraft according to claim 1, wherein the polyhedron is a truncated octahedron. The polyhedron consists of two thermoplastic resin panels having the same shape and arranged opposite to each other,
Each thermoplastic panel is
It has four regular hexagonal main plates and three square sub-plates, and the second to fourth main plates are continuously arranged radially around the first main plate. The prepreg or semi-preg that is radially continuous around the main plate and disposed between the second and third main plates and between the third and fourth main plates is heat-processed. The structure for spacecraft described.   The spacecraft structure according to any one of claims 1 to 3, wherein the thermoplastic resin panel is made of a carbon fiber reinforced PEEK resin having a thickness of 1.0 mm to 4.0 mm.   The spacecraft structure according to any one of claims 1 to 4, wherein a fusion allowance is formed in a portion of the thermoplastic resin panel to be thermally fused. The spacecraft structure according to any one of claims 1 to 5, wherein a reinforcing member is disposed at a boundary between adjacent surfaces of the thermoplastic resin panel.

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