WO2016088470A1 - 部材 - Google Patents

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Publication number
WO2016088470A1
WO2016088470A1 PCT/JP2015/080002 JP2015080002W WO2016088470A1 WO 2016088470 A1 WO2016088470 A1 WO 2016088470A1 JP 2015080002 W JP2015080002 W JP 2015080002W WO 2016088470 A1 WO2016088470 A1 WO 2016088470A1
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WO
WIPO (PCT)
Prior art keywords
metal
carbon fiber
heat
composite member
fiber
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Ceased
Application number
PCT/JP2015/080002
Other languages
English (en)
French (fr)
Japanese (ja)
Inventor
信幸 神原
阿部　俊夫
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Mitsubishi Heavy Industries Ltd
Original Assignee
Mitsubishi Heavy Industries Ltd
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Mitsubishi Heavy Industries Ltd filed Critical Mitsubishi Heavy Industries Ltd
Priority to US15/531,606 priority Critical patent/US20180281983A1/en
Priority to EP15864708.1A priority patent/EP3214112B1/en
Publication of WO2016088470A1 publication Critical patent/WO2016088470A1/ja
Anticipated expiration legal-status Critical
Ceased legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F28—HEAT EXCHANGE IN GENERAL
    • F28F—DETAILS OF HEAT-EXCHANGE AND HEAT-TRANSFER APPARATUS, OF GENERAL APPLICATION
    • F28F21/00—Constructions of heat-exchange apparatus characterised by the selection of particular materials
    • F28F21/06—Constructions of heat-exchange apparatus characterised by the selection of particular materials of plastics material
    • F28F21/067—Details
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B32—LAYERED PRODUCTS
    • B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
    • B32B5/00—Layered products characterised by the non- homogeneity or physical structure, i.e. comprising a fibrous, filamentary, particulate or foam layer; Layered products characterised by having a layer differing constitutionally or physically in different parts
    • B32B5/02—Layered products characterised by the non- homogeneity or physical structure, i.e. comprising a fibrous, filamentary, particulate or foam layer; Layered products characterised by having a layer differing constitutionally or physically in different parts characterised by structural features of a fibrous or filamentary layer
    • B32B5/12—Layered products characterised by the non- homogeneity or physical structure, i.e. comprising a fibrous, filamentary, particulate or foam layer; Layered products characterised by having a layer differing constitutionally or physically in different parts characterised by structural features of a fibrous or filamentary layer characterised by the relative arrangement of fibres or filaments of different layers, e.g. the fibres or filaments being parallel or perpendicular to each other
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B32—LAYERED PRODUCTS
    • B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
    • B32B5/00—Layered products characterised by the non- homogeneity or physical structure, i.e. comprising a fibrous, filamentary, particulate or foam layer; Layered products characterised by having a layer differing constitutionally or physically in different parts
    • B32B5/22—Layered products characterised by the non- homogeneity or physical structure, i.e. comprising a fibrous, filamentary, particulate or foam layer; Layered products characterised by having a layer differing constitutionally or physically in different parts characterised by the presence of two or more layers which are next to each other and are fibrous, filamentary, formed of particles or foamed
    • B32B5/24—Layered products characterised by the non- homogeneity or physical structure, i.e. comprising a fibrous, filamentary, particulate or foam layer; Layered products characterised by having a layer differing constitutionally or physically in different parts characterised by the presence of two or more layers which are next to each other and are fibrous, filamentary, formed of particles or foamed one layer being a fibrous or filamentary layer
    • B32B5/26—Layered products characterised by the non- homogeneity or physical structure, i.e. comprising a fibrous, filamentary, particulate or foam layer; Layered products characterised by having a layer differing constitutionally or physically in different parts characterised by the presence of two or more layers which are next to each other and are fibrous, filamentary, formed of particles or foamed one layer being a fibrous or filamentary layer another layer next to it also being fibrous or filamentary
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
    • C—CHEMISTRY; METALLURGY
    • C08—ORGANIC MACROMOLECULAR COMPOUNDS; THEIR PREPARATION OR CHEMICAL WORKING-UP; COMPOSITIONS BASED THEREON
    • C08J—WORKING-UP; GENERAL PROCESSES OF COMPOUNDING; AFTER-TREATMENT NOT COVERED BY SUBCLASSES C08B, C08C, C08F, C08G or C08H
    • C08J5/00—Manufacture of articles or shaped materials containing macromolecular substances
    • C08J5/04—Reinforcing macromolecular compounds with loose or coherent fibrous material
    • C08J5/0405—Reinforcing macromolecular compounds with loose or coherent fibrous material with inorganic fibres
    • C08J5/042—Reinforcing macromolecular compounds with loose or coherent fibrous material with inorganic fibres with carbon fibres
    • C—CHEMISTRY; METALLURGY
    • C08—ORGANIC MACROMOLECULAR COMPOUNDS; THEIR PREPARATION OR CHEMICAL WORKING-UP; COMPOSITIONS BASED THEREON
    • C08J—WORKING-UP; GENERAL PROCESSES OF COMPOUNDING; AFTER-TREATMENT NOT COVERED BY SUBCLASSES C08B, C08C, C08F, C08G or C08H
    • C08J5/00—Manufacture of articles or shaped materials containing macromolecular substances
    • C08J5/04—Reinforcing macromolecular compounds with loose or coherent fibrous material
    • C08J5/047—Reinforcing macromolecular compounds with loose or coherent fibrous material with mixed fibrous material
    • C—CHEMISTRY; METALLURGY
    • C08—ORGANIC MACROMOLECULAR COMPOUNDS; THEIR PREPARATION OR CHEMICAL WORKING-UP; COMPOSITIONS BASED THEREON
    • C08J—WORKING-UP; GENERAL PROCESSES OF COMPOUNDING; AFTER-TREATMENT NOT COVERED BY SUBCLASSES C08B, C08C, C08F, C08G or C08H
    • C08J5/00—Manufacture of articles or shaped materials containing macromolecular substances
    • C08J5/04—Reinforcing macromolecular compounds with loose or coherent fibrous material
    • C08J5/06—Reinforcing macromolecular compounds with loose or coherent fibrous material using pretreated fibrous materials
    • C—CHEMISTRY; METALLURGY
    • C09—DYES; PAINTS; POLISHES; NATURAL RESINS; ADHESIVES; COMPOSITIONS NOT OTHERWISE PROVIDED FOR; APPLICATIONS OF MATERIALS NOT OTHERWISE PROVIDED FOR
    • C09K—MATERIALS FOR MISCELLANEOUS APPLICATIONS, NOT PROVIDED FOR ELSEWHERE
    • C09K5/00—Heat-transfer, heat-exchange or heat-storage materials, e.g. refrigerants; Materials for the production of heat or cold by chemical reactions other than by combustion
    • C09K5/08—Materials not undergoing a change of physical state when used
    • C09K5/14—Solid materials, e.g. powdery or granular
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B32—LAYERED PRODUCTS
    • B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
    • B32B2262/00—Composition or structural features of fibres which form a fibrous or filamentary layer or are present as additives
    • B32B2262/10—Inorganic fibres
    • B32B2262/106—Carbon fibres, e.g. graphite fibres
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B32—LAYERED PRODUCTS
    • B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
    • B32B2605/00—Vehicles
    • B32B2605/18—Aircraft
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
    • B64C1/40—Sound or heat insulation, e.g. using insulation blankets
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
    • B64C2001/0054—Fuselage structures substantially made from particular materials
    • B64C2001/0072—Fuselage structures substantially made from particular materials from composite materials
    • C—CHEMISTRY; METALLURGY
    • C08—ORGANIC MACROMOLECULAR COMPOUNDS; THEIR PREPARATION OR CHEMICAL WORKING-UP; COMPOSITIONS BASED THEREON
    • C08J—WORKING-UP; GENERAL PROCESSES OF COMPOUNDING; AFTER-TREATMENT NOT COVERED BY SUBCLASSES C08B, C08C, C08F, C08G or C08H
    • C08J2363/00—Characterised by the use of epoxy resins; Derivatives of epoxy resins
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F28—HEAT EXCHANGE IN GENERAL
    • F28F—DETAILS OF HEAT-EXCHANGE AND HEAT-TRANSFER APPARATUS, OF GENERAL APPLICATION
    • F28F13/00—Arrangements for modifying heat-transfer, e.g. increasing, decreasing
    • F28F2013/001—Particular heat conductive materials, e.g. superconductive elements
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F28—HEAT EXCHANGE IN GENERAL
    • F28F—DETAILS OF HEAT-EXCHANGE AND HEAT-TRANSFER APPARATUS, OF GENERAL APPLICATION
    • F28F2255/00—Heat exchanger elements made of materials having special features or resulting from particular manufacturing processes
    • F28F2255/06—Heat exchanger elements made of materials having special features or resulting from particular manufacturing processes composite, e.g. polymers with fillers or fibres
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/40—Weight reduction

Definitions

  • the present invention relates to a member used for an aircraft or an artificial satellite.
  • Aircraft have heat generating parts such as electronic devices, batteries, and engines.
  • the heat generated in the heat generating part is released (exhaust heat) through a member called a heat channel material.
  • a metal sheet or a jumper wire has been used as a heat flow path material for an aircraft.
  • An example of a heat flow control body having a paper honeycomb core used in an artificial satellite is disclosed in Patent Document 1.
  • An object of an aspect of the present invention is to provide a member that can efficiently release heat generated in a heat generating part such as an aircraft or an artificial satellite.
  • a first composite member having a plastic reinforced with thermally conductive carbon fibers including one or both of metal-coated carbon fibers and pitch-based carbon fibers, and the heat A member is provided in which one end portion of the conductive carbon fiber is disposed in the heat generating portion and the other end portion of the thermally conductive carbon fiber is disposed in the heat radiating portion.
  • the first composite member includes a carbon fiber reinforced plastic reinforced with a thermally conductive carbon fiber including one or both of a metal-coated carbon fiber and a pitch-based carbon fiber.
  • One end portion of the carbon fiber is disposed in the heat generating portion, and the other end portion is disposed in the heat radiating portion. Therefore, the heat generated in the heat generating part is transmitted to the heat conductive carbon fiber and efficiently released to the heat radiating part.
  • the first composite member can be used as a strength member for an aircraft or an artificial satellite. Therefore, the heat of the heat generating portion can be released to the heat radiating portion without separately providing a dedicated heat channel material such as a metal sheet or a jumper wire. Therefore, the heat generated in the heat generating part is efficiently released to the heat radiating part while suppressing an increase in weight.
  • the thermally conductive carbon fiber and the plastic are molded together, the robustness is improved.
  • the method comprises a first composite member having a plastic reinforced with thermally conductive carbon fibers including one or both of metal-coated carbon fibers and pitch-based carbon fibers, wherein the heat There is provided a member in which a central portion of the conductive carbon fiber is disposed in the heat generating portion, and one end portion and the other end portion of the thermally conductive carbon fiber are disposed in the heat radiating portion.
  • the heat generated in the heat generating portion is the heat It is efficiently discharged to the heat dissipation part through the conductive carbon fiber. Therefore, the heat generated in the heat generating part is efficiently released to the heat radiating part while suppressing an increase in weight.
  • the thermally conductive carbon fiber and the plastic are molded together, the robustness is improved.
  • the first composite member includes a prepreg sheet including a plurality of the heat conductive carbon fibers arranged in a parallel direction intersecting the fiber direction.
  • a plurality of stacked bodies may be stacked in a stacking direction that intersects the parallel direction, and the thermal conductivity in the fiber direction may be larger than the thermal conductivity in the parallel direction and the thermal conductivity in the stacking direction.
  • the heat generated in the heat generating portion is suppressed from being transmitted in the parallel direction and the stacking direction, and is efficiently released to the heat radiating portion.
  • the first composite member is a plate-like member, and is disposed on one or both of the front surface and the back surface of the first composite member and reinforced with carbon fiber.
  • the one end part and the other end part of the heat conductive carbon fiber may be exposed.
  • the first composite member is supported by the second composite member, and the strength is maintained. Since each of the one end part and the other end part of the heat conductive carbon fiber is exposed without being covered with the second composite member, the heat generated in the heat generating part is efficiently released from the heat radiating part.
  • the heat generating part may include electronic equipment of the aircraft, and the heat radiating part may include a fuel tank of the aircraft.
  • a member that can efficiently release the heat generated in the heat generating portion is provided.
  • FIG. 1 is a diagram illustrating an example of an aircraft according to the first embodiment.
  • FIG. 2 is a perspective view schematically showing an example of the composite member according to the first embodiment.
  • Drawing 3 is a figure showing typically an example of the manufacturing method of the composite member concerning a 1st embodiment.
  • FIG. 4 is a cross-sectional view illustrating an example of a member according to the first embodiment.
  • FIG. 5 is a cross-sectional view illustrating an example of a member according to the second embodiment.
  • FIG. 6 is a plan view showing an example of a main wing of an aircraft according to the third embodiment.
  • FIG. 7 is a diagram schematically showing an example of the shear tie according to the third embodiment.
  • FIG. 8 is a cross-sectional view showing an example of a shear tie according to the third embodiment.
  • FIG. 1 is a diagram illustrating an example of an aircraft 1 according to the present embodiment.
  • the aircraft 1 includes a fuselage 2, a main wing 3, a horizontal tail 4, a vertical tail 5, an engine 6, a fuel tank 7, a cockpit 8, an electronic device 9, and a battery 10. It is equipped with.
  • At least a part of the fuselage 2, the main wing 3, the horizontal tail 4, and the vertical tail 5 is formed of a composite material.
  • the composite material includes carbon fiber reinforced plastic (CFRP), which is a plastic reinforced with carbon fiber.
  • CFRP carbon fiber reinforced plastic
  • the composite material may include glass fiber reinforced plastic (GFRP), which is a plastic reinforced with glass fiber.
  • the fuselage 2, the main wing 3, the horizontal tail 4, and the vertical tail 5 may be formed of a metal such as an aluminum alloy (duralumin).
  • At least a part of the members of the aircraft 1 has a metal-coated carbon fiber reinforced plastic that is a plastic reinforced with metal-coated carbon fiber (MC).
  • Plastics reinforced with metal-coated carbon fibers are also called MC-CFRP.
  • the metal-coated carbon fiber is a thermally conductive carbon fiber having thermal conductivity.
  • FIG. 2 is a perspective view schematically showing an example of the composite member 11 including the metal-coated carbon fiber reinforced plastic according to the present embodiment.
  • the composite member 11 has a metal-coated carbon fiber reinforced plastic 14 which is a plastic 13 reinforced with metal-coated carbon fibers 12.
  • the metal coated carbon fiber 12 includes a carbon fiber 15 and a metal 16 covered with the carbon fiber 15.
  • the composite member 11 has a plurality of metal-coated carbon fibers 12.
  • the metal coated carbon fiber 12 is long in the first direction.
  • a plurality of metal-coated carbon fibers 12 are arranged in parallel in a second direction orthogonal to the first direction.
  • the plurality of metal-coated carbon fibers 12 are arranged in a third direction orthogonal to the first direction and the second direction.
  • the longitudinal direction (first direction) of the metal-coated carbon fiber 12 is appropriately referred to as a fiber direction.
  • the direction in which the plurality of metal-coated carbon fibers 12 are arranged is appropriately referred to as a parallel direction.
  • a direction (third direction) in which the plurality of metal-coated carbon fibers 12 are arranged is appropriately referred to as a lamination direction.
  • the plurality of metal-coated carbon fibers 12 are arranged with an interval in each of the parallel direction and the stacking direction.
  • a plastic 13 is disposed between the plurality of metal-coated carbon fibers 12.
  • the plastic 13 includes an epoxy resin.
  • the diameter of the carbon fiber 15 of the metal-coated carbon fiber 12 is, for example, 5 ⁇ m or more and 10 ⁇ m or less.
  • the surface of the carbon fiber 15 is covered with a metal 16.
  • the heat conductivity of the metal 16 is higher than the heat conductivity of the carbon fiber 15.
  • the thermal conductivity of the plastic 13 is lower than the thermal conductivity of the metal 16 and the thermal conductivity of the carbon fiber 15. That is, the thermal conductivity of the plastic 13 is lower than the thermal conductivity of the metal-coated carbon fiber 12.
  • the metal 16 is nickel.
  • the metal coated carbon fiber 12 is a nickel coated carbon fiber.
  • the metal 16 may be at least one of gold, silver, and copper.
  • FIG. 3 is a diagram schematically illustrating an example of a method for manufacturing the composite member 11 according to the present embodiment.
  • the metal-coated carbon fiber 12 is produced.
  • the metal-coated carbon fiber 12 is manufactured by coating a metal 16 on a carbon fiber 15 having a diameter of about 5 ⁇ m to 10 ⁇ m.
  • nickel is coated as the metal 16.
  • the metal 16 may be coated with at least one of gold, silver, and copper.
  • a plurality of metal-coated carbon fibers 12 are arranged in a parallel direction and are hardened with a plastic 13 such as an epoxy resin.
  • the plurality of metal-coated carbon fibers 12 are hardened with a plastic 13 in an aligned state without being twisted.
  • the parallel direction is a direction in which a plurality of metal-coated carbon fibers 12 arranged in the fiber direction are arranged.
  • the fiber direction and the parallel direction are orthogonal to each other.
  • the sheet-like member including the plastic 13 and the plurality of metal-coated carbon fibers 12 arranged in the parallel direction and fixed by the plastic 13 is referred to as a prepreg sheet 17.
  • Step C a plurality of manufactured prepreg sheets 17 are stacked in the stacking direction.
  • the stacking direction is a direction in which a plurality of prepreg sheets 17 are stacked.
  • the stacking direction is orthogonal to the fiber direction and the parallel direction.
  • the laminate of the prepreg sheet 17 is heat-treated at a high temperature and a high pressure in a heating and pressurizing device called an autoclave. Thereby, the composite member 11 which is a laminated body of the several prepreg sheet 17 is manufactured.
  • the metal-coated carbon fibers 12 of the plurality of prepreg sheets 17 are all arranged in the same direction, so-called unidirectional lamination.
  • the metal-coated fibers 12 of the first prepreg sheet 17 are arranged in the first direction, and the metal-coated fibers 12 of the second prepreg sheet 17 that overlap the first prepreg sheet 17 are the first of the first prepreg sheet 17.
  • a so-called cross-ply lamination may be employed that is arranged in a second direction that intersects the direction.
  • FIG. 4 is a cross-sectional view showing an example of a member (heat channel material) 20 for the aircraft 1 according to the present embodiment.
  • the member 20 includes a composite member 11 and a composite member 21 connected to the composite member 11.
  • the member 20 is used for at least a part of the fuselage 2, the main wing 3, the horizontal tail 4, and the vertical tail 5.
  • a member 20 is manufactured by laminating a prepreg sheet containing metal-coated carbon fibers 12 and a prepreg sheet containing carbon fibers not coated with metal, and subjecting the laminate to heat and pressure treatment in an autoclave. May be.
  • one end portion 12 ⁇ / b> A of the metal-coated carbon fiber 12 is arranged in the heat generating portion of the aircraft 1 with respect to the fiber direction.
  • the other end 12 ⁇ / b> B of the metal-coated carbon fiber 12 is arranged in the heat radiating portion of the aircraft 1 in the fiber direction.
  • the heating unit of the aircraft 1 includes, for example, at least one of the electronic device 9, the battery 10, and the engine 6 of the aircraft 1. Further, the heat generating part includes a housing of the electronic device 9.
  • the heat dissipation unit of the aircraft 1 includes, for example, the fuel tank 7 of the aircraft 1.
  • the heat radiating portion of the aircraft 1 may be an external space of the aircraft 1 (a space facing the outer surface of the fuselage 2).
  • the composite member 11 is a plate-like member.
  • the composite member 21 is disposed on each of the front surface and the back surface of the composite member 11.
  • the composite member 21 has a carbon fiber reinforced plastic including a plastic reinforced with carbon fiber.
  • the carbon fiber of the carbon fiber reinforced plastic of the composite member 21 is not coated with metal.
  • the thermal conductivity of the composite member 21 in the fiber direction is lower than the thermal conductivity of the composite member 11 in the fiber direction.
  • the thermal conductivity of the composite member 21 in the parallel direction is lower than the thermal conductivity of the composite member 11 in the fiber direction.
  • the thermal conductivity of the composite member 21 in the stacking direction is lower than the thermal conductivity of the composite member 11 in the fiber direction.
  • thermal conductivity of the composite member 21 in the parallel direction may be equal to the thermal conductivity of the composite member 11 in the parallel direction, or may be lower than the composite member 11 in the parallel direction.
  • the thermal conductivity of the composite member 21 in the stacking direction may be equal to the thermal conductivity of the composite member 11 in the stacking direction, or may be lower than the composite member 11 in the stacking direction.
  • the composite member 21 is not disposed on each of the one end portion 12A and the other end portion 12B of the metal coated carbon fiber 12. Each of the one end portion 12A and the other end portion 12B of the metal coated carbon fiber 12 is exposed. One end portion 12A of the metal-coated carbon fiber 12 is in contact with the heat generating portion. Note that the one end portion 12A of the metal-coated carbon fiber 12 may be opposed to the heat generating portion via a gap. The other end portion 12B of the metal coated carbon fiber 12 is in contact with the heat radiating portion. The other end 12B of the metal-coated carbon fiber 12 may be opposed to the heat radiating part via a gap.
  • the composite member 21 is disposed on the front surface of the composite member 11 and may not be disposed on the back surface of the composite member 11.
  • the composite member 21 is disposed on the back surface of the composite member 11 and may not be disposed on the surface of the composite member 11.
  • the composite member 21 may not be arranged on both the front surface and the back surface of the composite member 11.
  • the metal coated carbon fiber 12 is arranged in the fiber direction of the member 20.
  • a plastic 13 is disposed between the plurality of metal-coated carbon fibers 12 in each of the parallel direction and the stacking direction of the members 20.
  • the thermal conductivity of the member 20 in the fiber direction is larger than the thermal conductivity of the member 20 in the parallel direction and the thermal conductivity of the member 20 in the stacking direction.
  • the heat of the heat generating part is absorbed by the metal coated carbon fiber 12 from the one end part 12A.
  • the heat absorbed by the metal-coated carbon fiber 12 is transmitted through the metal-coated carbon fiber 12 and released (exhaust heat) from the other end 12B.
  • the thermal conductivity of the member 20 in the parallel direction and the thermal conductivity of the member 20 in the stacking direction are smaller than the thermal conductivity of the member 20 in the fiber direction. Therefore, the heat of the metal coated carbon fiber 12 moves exclusively in the fiber direction. It is suppressed that the heat
  • the composite member 21 is arrange
  • the heat transfer coefficient of each composite member 21 in the fiber direction, the parallel direction, and the stacking direction is smaller than the heat transfer coefficient of the composite member 11 in the fiber direction. Therefore, the heat of the metal coated carbon fiber 12 is suppressed from being released from the surface of the composite member 21.
  • the composite member 11 includes the metal-coated carbon fiber reinforced plastic 14 reinforced with the metal-coated carbon fiber 12, and one end portion 12 ⁇ / b> A of the metal-coated carbon fiber 12 is the aircraft 1.
  • the other end portion 12 ⁇ / b> B of the metal-coated carbon fiber 12 is disposed in the heat radiating portion of the aircraft 1.
  • the metal 16 of the metal coated carbon fiber 12 has a high thermal conductivity. Therefore, the heat generated in the heat generating part is transmitted to the metal coated carbon fiber 12 and is efficiently released to the heat radiating part.
  • the composite member 11 can be used as a strength member of the aircraft 1. Therefore, the heat of the heat generating part can be released to the heat radiating part without providing a dedicated heat channel material such as a metal sheet or a jumper wire as in the prior art. Therefore, the heat generated in the heat generating part of the aircraft 1 is efficiently released to the heat radiating part while suppressing an increase in the weight of the aircraft 1.
  • the composite member 11 is formed by laminating a plurality of prepreg sheets 17 including a plurality of metal-coated carbon fibers 12 arranged in a parallel direction intersecting the fiber direction in a lamination direction intersecting the fiber direction and the parallel direction. Includes laminates.
  • the thermal conductivity of the member 20 in the fiber direction is larger than the thermal conductivity of the member 20 in the parallel direction and the thermal conductivity of the member in the stacking direction.
  • anisotropy is imparted to the thermal conductivity, and heat generated in the heat generating portion is suppressed from being transmitted in the parallel direction and the stacking direction, and is efficiently released to the heat radiating portion.
  • heat is transferred to the member or device by the member 20 having anisotropy in heat transfer coefficient. It is suppressed.
  • the composite member 11 is a plate-like member
  • the member 20 includes a composite member 21 including a carbon fiber reinforced plastic disposed on one or both of the front surface and the back surface of the composite member 11.
  • the composite member 11 is supported by the composite member 21, and the strength is maintained.
  • the thermal conductivity of the composite member 21 is smaller than the thermal conductivity of the composite member 11 in the fiber direction. Therefore, when there is a member or device that is not desired to be heated in at least one of the parallel direction and the stacking direction of the member 20, the composite member 21 suppresses heat from being transmitted to the member or device.
  • each of the one end portion 12A and the other end portion 12B of the metal-coated carbon fiber 12 is not covered with the composite member 21 or the like and exposed. Since the one end portion 12A is exposed, the heat generated in the heat generating portion is efficiently absorbed by the metal 16 of the metal-coated carbon fiber 12 through the one end portion 12A. Since the other end portion 12B is exposed, the heat generated in the heat generating portion and moved through the metal 16 of the metal-coated carbon fiber 12 is efficiently released to the heat radiating portion via the other end portion 12B. Thus, in this embodiment, since each of the one end portion 12A and the other end portion 12B of the metal coated carbon fiber 12 is exposed, the heat generated in the heat generating portion is efficiently released from the heat radiating portion.
  • the heat generating portion of the aircraft 1 includes the electronic device 9 of the aircraft 1.
  • the heat radiating unit of the aircraft 1 includes the fuel tank 7 of the aircraft 1.
  • FIG. 5 is a cross-sectional view showing an example of how to use the member 20.
  • the heat generating portion of the aircraft 1 may be disposed in the center portion of the member 20 between the one end portion 12A and the other end portion 12B. That is, the center part of the metal-coated carbon fiber 12 may be arranged in the heat generating part with respect to the fiber direction, and the one end part 12A and the other end part 12B of the metal-coated carbon fiber 12 may be arranged in the heat radiating part.
  • each of the one end portion 12 ⁇ / b> A and the other end portion 12 ⁇ / b> B is disposed in the heat radiating portion of the aircraft 1.
  • the heat generated in the heat generating part is released from each of the one end part 12A and the other end part 12B. Moreover, when one end part 12A and the other end part 12B of the metal coated carbon fiber 12 are exposed, heat generated in the heat generating part is efficiently released from the heat radiating part.
  • FIG. 6 shows a plan view of the main wing 3 of the aircraft 1.
  • a shear tie (structural material) 31 to be described later is disposed on at least a part of the libline 29.
  • FIG. 7 is a diagram schematically showing the positional relationship between the outer plate 30 and the shear tie 31.
  • the sheer tie 31 is a member that couples the stringer, the rib, and the like with the outer plate 30.
  • the shear tie 31 is formed by the composite member 21 including the carbon fiber reinforced plastic and the composite member 11 including the metal-coated carbon fiber reinforced plastic 14.
  • the outer plate 30 and the sheer tie 31 are fixed by a fastener 51.
  • the outer plate 30 and the shear tie 31 are fixed by connecting a collar (nut) 37 to the tip of the fastener 51.
  • a washer 41 and a spacer 42 are disposed between the collar 37 and the shear tie 31.
  • the collar 37, the washer 41, and the spacer 42 are covered with a cap 44.
  • the cap 44 is disposed so as to be in close contact with the shear tie 31.
  • the outer plate 30 includes a carbon fiber reinforced plastic layer 32, a glass fiber reinforced plastic layer 34, and a copper paint layer 39.
  • FIG. 8 is a cross-sectional view showing an example of the shear tie 31 according to the present embodiment.
  • the sheer tie 31 is formed by the composite member 21 including the carbon fiber reinforced plastic and the composite member 11 including the metal-coated carbon fiber reinforced plastic 14.
  • the composite member 11 is sandwiched between the composite members 21.
  • the composite member 11 is provided in a part of the shear tie 31.
  • one end portion 12 ⁇ / b> A of the metal-coated carbon fiber 12 of the composite member 11 is disposed at the lower end portion of the shear tie 31.
  • the other end portion 12 ⁇ / b> B of the metal coated carbon fiber 12 of the composite member 11 is disposed at the upper left end portion of the shear tie 31.
  • the upper right end portion of the shear tie 31 is formed by the composite member 21.
  • the heat generating part of the aircraft 1 is arranged at the lower end of the sheer tie 31.
  • the heat radiating portion of the aircraft 1 is disposed at the upper left end portion of the shear tie 31.
  • the other end 12B of the metal-coated carbon fiber 12 of the composite member 11 may be disposed at the upper right end of the shear tie 31, and the heat dissipation portion of the aircraft 1 may be disposed.
  • the other end portion 12B of the metal-coated carbon fiber 12 of the composite member 11 and the heat dissipating portion of the aircraft 1 may be disposed on both the upper right end portion and the left upper end portion of the shear tie 31.
  • One end portion 12A of the metal-coated carbon fiber 12 of the composite member 11 and the heat generating portion of the aircraft 1 may be disposed on one or both of the upper right end portion and the left upper end portion of the shear tie 31.
  • the other end portion 12 ⁇ / b> B of the metal-coated carbon fiber 12 of the composite member 11 may be disposed at the lower end portion of the shear tie 31, and the heat radiating portion of the aircraft 1 may be disposed.
  • the composite member 11 and the composite member 21 may be bent or processed into an arbitrary shape (three-dimensional shape).
  • pitch-based carbon fibers may be disposed in place of the metal-coated carbon fibers 12 or together with the metal-coated carbon fibers 12.
  • the pitch-based carbon fiber is a thermally conductive carbon fiber having a higher thermal conductivity than at least the PAN carbon fiber.
  • the member 20 may be used as a structural member of an artificial satellite.
  • the fiber direction, the parallel direction, and the stacking direction are orthogonal to each other.
  • the fiber direction and the parallel direction may intersect at an angle of, for example, 80 degrees or more and 100 degrees or less.
  • the fiber direction and the lamination direction may intersect at an angle of, for example, 80 degrees or more and 100 degrees or less.
  • the parallel direction and the stacking direction may intersect at an angle of not less than 80 degrees and not more than 100 degrees.

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  • Chemical & Material Sciences (AREA)
  • Engineering & Computer Science (AREA)
  • Materials Engineering (AREA)
  • Organic Chemistry (AREA)
  • Chemical Kinetics & Catalysis (AREA)
  • Polymers & Plastics (AREA)
  • Medicinal Chemistry (AREA)
  • Health & Medical Sciences (AREA)
  • Manufacturing & Machinery (AREA)
  • Thermal Sciences (AREA)
  • Mechanical Engineering (AREA)
  • Physics & Mathematics (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Inorganic Chemistry (AREA)
  • Combustion & Propulsion (AREA)
  • General Engineering & Computer Science (AREA)
  • Moulding By Coating Moulds (AREA)
  • Laminated Bodies (AREA)
  • Reinforced Plastic Materials (AREA)
PCT/JP2015/080002 2014-12-03 2015-10-23 部材 Ceased WO2016088470A1 (ja)

Priority Applications (2)

Application Number Priority Date Filing Date Title
US15/531,606 US20180281983A1 (en) 2014-12-03 2015-10-23 Member
EP15864708.1A EP3214112B1 (en) 2014-12-03 2015-10-23 Heat passage member

Applications Claiming Priority (2)

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JP2014-245346 2014-12-03
JP2014245346A JP6550230B2 (ja) 2014-12-03 2014-12-03 部材

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EP (1) EP3214112B1 (enExample)
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WO (1) WO2016088470A1 (enExample)

Cited By (1)

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EP3409578A1 (en) * 2017-05-30 2018-12-05 The Boeing Company A thermal management system, a composite wing, and a composite wing spar

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JP7149577B2 (ja) * 2018-10-15 2022-10-07 有限会社ヒロセ金型 炭素繊維強化樹脂成形品の製造方法、及び炭素繊維強化樹脂成形品
EA202191198A1 (ru) 2018-11-01 2021-09-14 Этомос Ньюклиар Энд Спейс Корпорейшн Система ребер радиатора, выполненных из углеродного волокна
US11117346B2 (en) * 2019-07-18 2021-09-14 Hamilton Sundstrand Corporation Thermally-conductive polymer and components
JP7430248B2 (ja) * 2020-04-03 2024-02-09 日本製鉄株式会社 蓄電デバイス構造体及び蓄電デバイス構造体の放熱方法
CN114474557B (zh) * 2021-12-31 2023-11-10 富联裕展科技(深圳)有限公司 金属塑胶结合件及其形成方法、电子产品壳体

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EP3409578A1 (en) * 2017-05-30 2018-12-05 The Boeing Company A thermal management system, a composite wing, and a composite wing spar
US10723437B2 (en) 2017-05-30 2020-07-28 The Boeing Company System for structurally integrated thermal management for thin wing aircraft control surface actuators
EP3792175A1 (en) * 2017-05-30 2021-03-17 The Boeing Company A thermal management system, a composite wing, and a composite wing spar
US11273900B2 (en) 2017-05-30 2022-03-15 The Boeing Company System for structurally integrated thermal management for thin wing aircraft control surface actuators

Also Published As

Publication number Publication date
US20180281983A1 (en) 2018-10-04
JP6550230B2 (ja) 2019-07-24
EP3214112B1 (en) 2021-04-14
EP3214112A1 (en) 2017-09-06
EP3214112A4 (en) 2017-11-22
JP2016108398A (ja) 2016-06-20

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