JP4537247B2 - Method for suppressing cracking in composite sandwich panel joints - Google Patents
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- JP4537247B2 JP4537247B2 JP2005106320A JP2005106320A JP4537247B2 JP 4537247 B2 JP4537247 B2 JP 4537247B2 JP 2005106320 A JP2005106320 A JP 2005106320A JP 2005106320 A JP2005106320 A JP 2005106320A JP 4537247 B2 JP4537247 B2 JP 4537247B2
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- 239000002131 composite material Substances 0.000 title claims description 31
- 238000000034 method Methods 0.000 title claims description 12
- 238000005336 cracking Methods 0.000 title description 2
- 239000012763 reinforcing filler Substances 0.000 claims description 43
- 239000010410 layer Substances 0.000 claims description 33
- 239000012792 core layer Substances 0.000 claims description 28
- 239000000463 material Substances 0.000 claims description 20
- 239000003733 fiber-reinforced composite Substances 0.000 claims description 7
- 238000003475 lamination Methods 0.000 claims description 6
- 229920005989 resin Polymers 0.000 claims description 3
- 239000011347 resin Substances 0.000 claims description 3
- 238000010030 laminating Methods 0.000 claims 1
- 239000011162 core material Substances 0.000 description 19
- 229920003002 synthetic resin Polymers 0.000 description 11
- 239000000057 synthetic resin Substances 0.000 description 11
- 238000004519 manufacturing process Methods 0.000 description 7
- 239000000945 filler Substances 0.000 description 6
- 230000008878 coupling Effects 0.000 description 5
- 238000010168 coupling process Methods 0.000 description 5
- 238000005859 coupling reaction Methods 0.000 description 5
- 238000000465 moulding Methods 0.000 description 4
- 230000002787 reinforcement Effects 0.000 description 4
- 238000009864 tensile test Methods 0.000 description 4
- 230000007547 defect Effects 0.000 description 3
- 239000006260 foam Substances 0.000 description 3
- 238000012360 testing method Methods 0.000 description 3
- 239000011800 void material Substances 0.000 description 3
- 229920002430 Fibre-reinforced plastic Polymers 0.000 description 2
- 238000007796 conventional method Methods 0.000 description 2
- 230000032798 delamination Effects 0.000 description 2
- 238000013461 design Methods 0.000 description 2
- 238000010586 diagram Methods 0.000 description 2
- 239000011151 fibre-reinforced plastic Substances 0.000 description 2
- 230000000977 initiatory effect Effects 0.000 description 2
- 238000009413 insulation Methods 0.000 description 2
- 239000011159 matrix material Substances 0.000 description 2
- 229910000838 Al alloy Inorganic materials 0.000 description 1
- 229920000049 Carbon (fiber) Polymers 0.000 description 1
- 239000000853 adhesive Substances 0.000 description 1
- 230000001070 adhesive effect Effects 0.000 description 1
- 239000004917 carbon fiber Substances 0.000 description 1
- 239000004918 carbon fiber reinforced polymer Substances 0.000 description 1
- 230000003111 delayed effect Effects 0.000 description 1
- 238000011161 development Methods 0.000 description 1
- 239000000835 fiber Substances 0.000 description 1
- 238000010438 heat treatment Methods 0.000 description 1
- 238000005304 joining Methods 0.000 description 1
- 238000002844 melting Methods 0.000 description 1
- 230000008018 melting Effects 0.000 description 1
- VNWKTOKETHGBQD-UHFFFAOYSA-N methane Chemical compound C VNWKTOKETHGBQD-UHFFFAOYSA-N 0.000 description 1
- 239000004745 nonwoven fabric Substances 0.000 description 1
- 239000011435 rock Substances 0.000 description 1
- 238000003856 thermoforming Methods 0.000 description 1
- 238000011179 visual inspection Methods 0.000 description 1
- XLYOFNOQVPJJNP-UHFFFAOYSA-N water Substances O XLYOFNOQVPJJNP-UHFFFAOYSA-N 0.000 description 1
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Description
本発明は、航空機の外板などに好適に実施することができる複合材サンドイッチパネル結合部の亀裂発生抑制方法に関する。 The present invention relates to a method for suppressing the occurrence of cracks in a composite material sandwich panel joint, which can be suitably applied to an outer panel of an aircraft.
図１２は、従来の技術を説明するための航空機の機首１の構造を簡略化して示す斜視図であり、図１３は図１２の切断面線ＸＩＩＩ−ＸＩＩＩから見た一部の拡大断面図である。近年、航空機の機体にフォームコアサンドイッチパネル（Foam Core Sandwich Panel）とも呼ばれる繊維強化複合材サンドイッチパネル２が採用されている。この複合材サンドイッチパネル２は、疲労に強く、断熱性および遮音性にも優れ、一体成形によって重量および部品点数を低減できるという利点があり、ハニカムサンドイッチパネルのように、ハニカムコアへの水の浸入という問題もない優れた構造形態として着目されている。 FIG. 12 is a perspective view showing a simplified structure of the aircraft nose 1 for explaining the prior art, and FIG. 13 is a partially enlarged cross-sectional view taken along section line XIII-XIII of FIG. It is. In recent years, fiber reinforced composite sandwich panels 2, also called foam core sandwich panels, have been employed in aircraft bodies. This composite sandwich panel 2 is resistant to fatigue, has excellent heat insulation and sound insulation properties, and has the advantage of being able to reduce the weight and the number of parts by integral molding. Like the honeycomb sandwich panel, water can enter the honeycomb core. It has attracted attention as an excellent structural form without any problem.
実機への適用にあたっては、たとえば機首１への適用例について述べると、複合材サンドイッチパネル２の構造材３への結合部４が設計上の重要な要素となる。この結合部４の構造は、目視による点検が容易な突合せ結合とし、前記構造材３にハイロックボルトなどによって実現されるファスナ７によって接合されたスプライスプレートとも呼ばれる添え板５を介して、面方向に隣接する２枚の複合材サンドイッチパネル２を、ハイロックボルトなどによって実現される複数のファスナ６によって結合する手法を採用している。このような複合材サンドイッチパネル２をより軽量化するためには、層間剥離の防止を如何にして解決するかが重要な課題とされる。 In application to an actual machine, for example, an application example to the nose 1 will be described. The joint portion 4 of the composite sandwich panel 2 to the structural material 3 is an important design element. The structure of the coupling portion 4 is a butt coupling that can be easily inspected by visual inspection, and a surface direction is provided via a splicing plate 5 called a splice plate joined to the structural material 3 by a fastener 7 realized by a high lock bolt or the like. A method of joining two composite sandwich panels 2 adjacent to each other by a plurality of fasteners 6 realized by a high rock bolt or the like is employed. In order to reduce the weight of such a composite sandwich panel 2, it is an important issue how to prevent the delamination.
他の従来の技術では、加熱成形時にコア材の積層フォーム間に閉じ込められたガスに起因するボイド（泡状の）欠陥の発生を防止し、このボイド欠陥の発生による強度低下を防止して、軽量で高強度の複合繊維強化プラスチック構造体を実現するために、成形型上に、マトリックス樹脂中に低密度付与充填材を内包するコア材の少なくとも２層間に、マトリクス樹脂の成形温度よりも高い融点を有する合成樹脂からなる不織布を挟み、少なくとも片面のほぼ全域に繊維強化プラスチックのプリプレグを配置し、バッグフィルムによって全体を覆って気密化し、バッグフィルム内を真空引きしながら加熱することによって一体成形される複合材サンドイッチパネルおよびその製造方法が提案されている（たとえば、特許文献１参照）。 Other conventional techniques prevent the occurrence of void (foam-like) defects caused by the gas trapped between the laminated foams of the core material during thermoforming, and prevent the strength from being reduced due to the occurrence of these void defects, In order to realize a lightweight and high-strength composite fiber reinforced plastic structure, the molding temperature is higher than the molding temperature of the matrix resin between at least two layers of the core material containing the low density imparting filler in the matrix resin on the mold. Insert a non-woven fabric made of a synthetic resin with a melting point, place a fiber-reinforced plastic prepreg over almost the entire area of at least one side, cover the whole with a bag film, make it airtight, and heat it while evacuating the bag film. A composite sandwich panel and a manufacturing method thereof have been proposed (see, for example, Patent Document 1).
前述の特許文献１に記載される従来の技術では、コア材の各層間のボイド欠陥の発生という製造上の問題は解消されるが、この複合材サンドイッチパネルに荷重が作用したとき、結合部の耐荷重強度に達する前に、結合部とコア層の傾斜部との間に亀裂が発生し、その亀裂が進展して層間剥離が生じ、終局的には破壊に至ってしまい、十分な耐荷重が得られないという実使用上の問題を有する。 In the conventional technique described in Patent Document 1 described above, the manufacturing problem of void defects occurring between the respective layers of the core material is solved. However, when a load is applied to the composite material sandwich panel, Before reaching the load-bearing strength, a crack occurs between the joint and the inclined portion of the core layer, and the crack progresses to cause delamination, eventually leading to breakage and sufficient load-bearing capacity. There is a problem in practical use that it cannot be obtained.
本発明の目的は、簡単な構成で、亀裂の発生または進展を防止し、耐荷重強度の向上を図ることができる複合材サンドイッチパネル結合部の亀裂発生抑制方法を提供することである。 It is an object of the present invention to provide a method for suppressing the occurrence of cracks in a composite material sandwich panel joint portion that can prevent the occurrence or progress of cracks and improve the load bearing strength with a simple configuration.
本発明は、発泡合成樹脂から成り、厚み方向一方側の第１表面と前記厚み方向他方側の第２表面とが平行な基部に、前記第１表面と第２表面とが相互に近接する方向に傾斜する傾斜部が連なるコア層と、
繊維強化複合材料から成り、コア層の前記第１表面上に一体的に形成される第１スキン層と、
繊維強化複合材料から成り、コア層の前記第２表面上に一体的に形成される第２スキン層とによって、
パネル本体部が形成されるとともに、
前記第１スキン層と第２スキン層とを積層して一体化した結合部が形成される複合材サンドイッチパネル結合部の亀裂発生抑制方法であって、
前記コア層の傾斜部と第１および第２スキン層の積層開始点との間に、硬化後の状態でコア層よりも剛性の高い未硬化のプリプレグから成る補強用充填材を介在させて、前記第１および第２スキン層ならびにコア層と前記補強用充填材とを一体化することを特徴とする複合材サンドイッチパネル結合部の亀裂発生抑制方法である。
The present invention is made of a foamed synthetic resin, and a direction in which the first surface and the second surface are close to each other in a base where the first surface on one side in the thickness direction and the second surface on the other side in the thickness direction are parallel to each other. A core layer with an inclined portion that is inclined to
A first skin layer comprising a fiber reinforced composite material and integrally formed on the first surface of the core layer;
A second skin layer comprising a fiber reinforced composite material and integrally formed on the second surface of the core layer;
A panel body is formed,
A method for suppressing crack occurrence in a composite material sandwich panel joint part in which a joint part in which the first skin layer and the second skin layer are laminated and integrated is formed,
Between the inclined portion of the core layer and the lamination start point of the first and second skin layers, interposing a reinforcing filler made of an uncured prepreg that is stiffer than the core layer in a state after curing , A method for suppressing cracking in a composite material sandwich panel joint, wherein the first and second skin layers and the core layer and the reinforcing filler are integrated.
本発明によれば、コア層の傾斜部と第１および第２スキン層の積層開始点との間に、第１および第２スキン層ならびにコア層に一体化した補強用充填材を介在させるので、複合材サンドイッチパネルに荷重が作用すると、補強用充填材にも荷重が負荷されて分散し、亀裂発生箇所の応力が低減されて、簡単な構成によって、亀裂の発生および進展を抑制することができる。 According to the present invention, the reinforcing filler integrated with the first and second skin layers and the core layer is interposed between the inclined portion of the core layer and the lamination start point of the first and second skin layers. When a load is applied to the composite sandwich panel, the load is also applied to the reinforcing filler to disperse, and the stress at the crack occurrence point is reduced. it can.
また、補強用充填材が未硬化の合成樹脂から成るので、複合材サンドイッチパネルの製造時において、補強用充填材の充填位置、充填範囲、被充填領域の形状などに応じて形態を柔軟に変更することが可能であり、これによって補強用充填材の設計上の自由度を向上することができる。 In addition , since the reinforcing filler is made of uncured synthetic resin, the shape can be flexibly changed according to the filling position of the reinforcing filler, the filling range, the shape of the region to be filled, etc., when manufacturing the composite sandwich panel This can improve the design freedom of the reinforcing filler.
図１は、本発明の実施の一形態の亀裂発生抑制方法が適用された複合材サンドイッチパネル２０の構成を示す一部の断面図である。本実施の形態の亀裂発生抑制方法が適用された複合材サンドイッチパネル２０は、発泡合成樹脂から成り、厚み方向一方側の第１表面２１と前記厚み方向他方側の第２表面２２とが平行な基部２３に、前記第１表面２１と第２表面２２とが相互に近接する方向に傾斜する傾斜部２４が連なるコア層２５と、繊維強化複合材料から成り、コア層２５の前記第１表面２１上に一体的に形成される第１スキン層２６と、繊維強化複合材料から成り、コア層２５の前記第２表面２２上に一体的に形成される第２スキン層２７と、前記コア層２５の傾斜部２４と第１および第２スキン層２６，２７の積層開始点２８との間に介在される補強用充填材２９とを含む。 FIG. 1 is a partial cross-sectional view illustrating a configuration of a composite sandwich panel 20 to which a crack generation suppressing method according to an embodiment of the present invention is applied. The composite material sandwich panel 20 to which the crack generation suppressing method of the present embodiment is applied is made of foamed synthetic resin, and the first surface 21 on one side in the thickness direction and the second surface 22 on the other side in the thickness direction are parallel. The first surface 21 of the core layer 25 is composed of a core layer 25 in which an inclined portion 24 that is inclined in a direction in which the first surface 21 and the second surface 22 are close to each other is connected to the base 23, and a fiber reinforced composite material. A first skin layer 26 integrally formed thereon, a second skin layer 27 made of a fiber reinforced composite material and integrally formed on the second surface 22 of the core layer 25, and the core layer 25 A reinforcing filler 29 interposed between the inclined portion 24 and the lamination starting point 28 of the first and second skin layers 26 and 27.
第１スキン層２６と第２スキン層２７とコア層２５とによって、３層構造のパネル本体部３０が形成されるとともに、第１スキン層２６と第２スキン層２７とが積層して一体化することによって、２層構造の結合部３１が形成される。図１において上方に配置される複合材サンドイッチパネル２０と、下方に配置される複合材サンドイッチパネル２０とは、上方の複合材サンドイッチパネル２０の下部の結合部３１と、下方の複合材サンドイッチパネル２０の上部の結合部３１とが上下に突合わせた状態で、スプライスプレートとも呼ばれる添え板３３にシムあるいは接着材３４を介して複数のファスナ３５によって締結され、相互に結合されている。 The first skin layer 26, the second skin layer 27, and the core layer 25 form a three-layer panel main body 30, and the first skin layer 26 and the second skin layer 27 are laminated and integrated. As a result, the coupling portion 31 having a two-layer structure is formed. In FIG. 1, the composite sandwich panel 20 disposed on the upper side and the composite sandwich panel 20 disposed on the lower side are composed of a lower joint portion 31 of the upper composite sandwich panel 20 and a lower composite sandwich panel 20. In the state where the upper coupling portion 31 is in abutment with the upper and lower portions, they are fastened by a plurality of fasteners 35 via shims or adhesives 34 to a support plate 33 also called a splice plate, and are coupled to each other.
前記添え板３３は、断面が略Ｔ字状の軽量アルミ合金から成り、前述の図１２に示される航空機の機首１の構造材３にファスナ３８によって接合される。前記ファスナ３５，３８は、ハイロックボルト３６およびカラー３７などを含んで構成される。 The accessory plate 33 is made of a lightweight aluminum alloy having a substantially T-shaped cross section, and is joined to the structural member 3 of the aircraft nose 1 shown in FIG. The fasteners 35 and 38 include a high lock bolt 36 and a collar 37.
第１および第２スキン層２６，２７を構成する材料としては、炭素繊維強化合成樹脂（略称ＣＦＲＰ）が用いられ、具体的には、異なる繊維方向を有する複数のプリプレグが用いられる。またコア層２５を構成する材料としては、独立気泡発泡の硬質合成樹脂が用いられる。さらに補強用充填材を構成する材料としては、硬化後の状態でコア層２５の剛性よりも高い剛性を有する材料、たとえば前記第１および第２スキン層２６，２７の材料と同一材料、すなわち未硬化のプリプレグが用いられる。これらの第１および第２スキン層２６，２７ならびにコア層２５は、前記各材料に補強用充填材２９の材料を充填した状態で、成形型上に乗載して、たとえばオートクレーブ内で２〜３時間、加熱されて一体化される。 As a material constituting the first and second skin layers 26 and 27, a carbon fiber reinforced synthetic resin (abbreviation CFRP) is used, and specifically, a plurality of prepregs having different fiber directions are used. Further, as a material constituting the core layer 25, a hard synthetic resin with closed cell foam is used. Further, the material constituting the reinforcing filler is a material having rigidity higher than that of the core layer 25 in the cured state, for example, the same material as the material of the first and second skin layers 26 and 27, that is, not yet formed. A curable prepreg is used. The first and second skin layers 26 and 27 and the core layer 25 are mounted on a mold in a state in which the respective materials are filled with the material of the reinforcing filler 29, and are, for example, 2 to 2 in an autoclave. It is heated and integrated for 3 hours.
図２は、本件発明者による引張り試験で用いた供試体４１および歪ゲージＡ〜Ｄの取り付け位置を示す断面図であり、図３は引張り試験によって得られた荷重と歪との関係を示すグラフである。本件発明者は、従来の複合材サンドイッチパネル２０を模擬した供試体４１を作成し、この供試体４１に引張り試験機によって引張り力Ｆを与えて引張り試験を行い、破壊モードおよび破壊荷重を確認した。その結果、試験荷重Ｆ＝３９ｋＮでコア先端部４６に図４に示されるように長さＬ＝７ｍｍの亀裂が生じ、試験荷重Ｆ＝１１２ｋＮで、結合部３１で第１および第２スキン層２６，２７が破断した。 FIG. 2 is a cross-sectional view showing the mounting position of the specimen 41 and strain gauges A to D used in the tensile test by the present inventors, and FIG. 3 is a graph showing the relationship between the load and strain obtained by the tensile test. It is. The present inventor created a specimen 41 simulating the conventional composite sandwich panel 20, and applied the tensile force F to the specimen 41 with a tensile tester to perform a tensile test to confirm the fracture mode and the fracture load. . As a result, a crack having a length L = 7 mm as shown in FIG. 4 is generated at the core tip 46 at a test load F = 39 kN, and the first and second skin layers 26 at the joint 31 at a test load F = 112 kN. , 27 broke.
次に、図４に示される複合材サンドイッチパネル２０の傾斜部２４を含む供試体モデル４５を作成し、この供試体モデル４５を利用して、図５に示されるコア先端部４６に補強用充填材２９が挿入されないコア先端部４６付近の供試体モデル４７と、図６に示される補強用充填材２９を挿入したコア先端部４６付近の供試体モデル４８とを作成し、単位荷重を負荷した状態を模擬したＦＥＭ(Finite-Element Method)解析を実施した。 Next, a specimen model 45 including the inclined portion 24 of the composite sandwich panel 20 shown in FIG. 4 is created, and the core tip 46 shown in FIG. A specimen model 47 near the core tip 46 where the material 29 is not inserted and a specimen model 48 near the core tip 46 inserted with the reinforcing filler 29 shown in FIG. An FEM (Finite-Element Method) analysis simulating the state was performed.
図７は、補強用充填材２９を挿入しない場合の供試体モデル４７のせん断応力の発生状態を示す図であり、図８は補強用充填材２９を挿入した場合の供試体モデル４８のせん断応力の発生状態を示す図である。前記ＦＥＭ解析の結果、補強用充填材２９を挿入しない場合は、図７に示されるように、亀裂発生想定位置Ｐでのせん断応力の最大値τｍａｘ１は、１．７×１０−２ＭＰａであった。これに対して、補強用充填材２９を挿入した場合は、図８に示されるように、亀裂発生想定位置Ｐでのせん断応力の最大値τｍａｘ２は、１．１×１０−２ＭＰａであり、３５％低減することを確認した。 FIG. 7 is a diagram showing the state of occurrence of shear stress in the specimen model 47 when the reinforcing filler 29 is not inserted, and FIG. 8 is the shear stress of the specimen model 48 when the reinforcing filler 29 is inserted. It is a figure which shows the generation | occurrence | production state of. As a result of the FEM analysis, when the reinforcing filler 29 is not inserted, as shown in FIG. 7, the maximum value τmax1 of the shear stress at the crack initiation position P is 1.7 × 10 −2 MPa. It was. On the other hand, when the reinforcing filler 29 is inserted, as shown in FIG. 8, the maximum value τmax2 of the shear stress at the assumed crack generation position P is 1.1 × 10 −2 MPa, A reduction of 35% was confirmed.
また、補強用充填材２９を挿入しない場合は、図７に示されるように、コア先端部４６付近のせん断応力τ１は、１．５×１０−６ＭＰａであった。これに対して、補強用充填材２９を挿入した場合は、図８に示されるように、コア先端部４６付近のせん断応力τ２は、６．９×１０−４ＭＰａに増加した。この現象は、発泡コアより剛性が高い補強用充填材２９がコア先端部４６と結合部３１との間に存在することによって、補強用充填材２９が荷重を負担して、亀裂先端部の応力が減少したことを示している。この亀裂先端部の応力低下により、亀裂の発生および進展を遅延させることが確認された。このように補強用充填材２９の剛性については、コア層２５以上のせん断剛性があれば、十分なせん断応力の緩和が期待できることが判明した。 Further, when the reinforcing filler 29 was not inserted, as shown in FIG. 7, the shear stress τ1 near the core tip 46 was 1.5 × 10 −6 MPa. On the other hand, when the reinforcing filler 29 was inserted, as shown in FIG. 8, the shear stress τ2 near the core tip 46 increased to 6.9 × 10 −4 MPa. This phenomenon is caused by the fact that the reinforcing filler 29 having higher rigidity than the foamed core exists between the core tip 46 and the coupling portion 31, so that the reinforcing filler 29 bears a load and stress at the crack tip. Indicates a decrease. It was confirmed that the stress reduction at the crack tip delayed the generation and propagation of the crack. As described above, it has been found that the rigidity of the reinforcing filler 29 can be expected to sufficiently relax the shear stress if it has a shear rigidity higher than that of the core layer 25.
図９は、挿入される補強用充填材の材料剛性を説明するためのグラフであり、図１０は補強用充填材２９の範囲を説明するための供試体モデルを示す図であり、図１１は補強用充填材２９のせん断応力への影響の範囲を示すグラフである。補強用充填材２９の範囲について評価したところ、コア先端部４６から１０ｍｍ程度の範囲内に充填すれば、十分なせん断応力の緩和が期待できることを確認した。補強用充填材２９の材料選定や充填範囲Ｘの決定は、安価な合成樹脂を、限定した範囲で充填すれば十分な効果が得られることがわかる。 FIG. 9 is a graph for explaining the material rigidity of the reinforcing filler to be inserted, FIG. 10 is a diagram showing a specimen model for explaining the range of the reinforcing filler 29, and FIG. It is a graph which shows the range of the influence with respect to the shear stress of the filler 29 for reinforcement. When the range of the reinforcing filler 29 was evaluated, it was confirmed that sufficient shear stress relaxation could be expected if it was filled within a range of about 10 mm from the core tip 46. It can be seen that the material selection of the reinforcing filler 29 and the determination of the filling range X are sufficiently effective if an inexpensive synthetic resin is filled in a limited range.
本実施の形態によれば、コア層２５の傾斜部２４と第１および第２スキン層２６，２７の積層開始点２８との間に、第１および第２スキン層２６，２７ならびにコア層２５に一体化した補強用充填材２９を介在させるので、複合材サンドイッチパネル２０に荷重が作用すると、補強用充填材２９にも荷重が負荷されて分散し、亀裂発生箇所の応力が低減されて、簡単な構成によって、亀裂の発生および進展を抑制することができる。 According to the present embodiment, the first and second skin layers 26 and 27 and the core layer 25 are provided between the inclined portion 24 of the core layer 25 and the lamination start point 28 of the first and second skin layers 26 and 27. Since the reinforcing filler 29 integrated in the middle is interposed, when a load acts on the composite sandwich panel 20, the load is also applied to the reinforcing filler 29 and dispersed, and the stress at the crack occurrence point is reduced. With a simple configuration, the occurrence and development of cracks can be suppressed.
また、補強用充填材２９がコア層２５よりも剛性の高い合成樹脂から成るので、補強材充填材２９による分担される荷重を大きくして、亀裂の発生および進展の抑制範囲を大きくして、耐荷重特性を向上することができる。 In addition, since the reinforcing filler 29 is made of a synthetic resin having rigidity higher than that of the core layer 25, the load shared by the reinforcing filler 29 is increased, and the range in which cracks are generated and progressed is increased, The load bearing characteristics can be improved.
さらに、補強用充填材２９が未硬化の合成樹脂から成るので、複合材サンドイッチパネル２０の製造時において、補強用充填材２９の充填位置、充填範囲、被充填領域の形状などに応じて形態を柔軟に変更することが可能であり、これによって補強用充填材２９の設計上の自由度を向上することができる。 Further, since the reinforcing filler 29 is made of an uncured synthetic resin, the shape of the reinforcing filler 29 according to the filling position, the filling range, the shape of the region to be filled, etc., when the composite sandwich panel 20 is manufactured. It is possible to change flexibly, and thereby the degree of freedom in designing the reinforcing filler 29 can be improved.
本発明の実施の他の形態では、前記補強用充填材２９は、硬化した合成樹脂から成ってもよい。このような構成によれば、補強用充填材２９が硬化した合成樹脂から成るので、複合材サンドイッチパネル２０の製造時において、補強用充填材２９の形態が一定であり、これによって補強用充填材２９の充填作業上の取り扱いが容易であり、製造作業の作業性が向上され、複合材サンドイッチパネル２０の生産性を向上することができる。 In another embodiment of the present invention, the reinforcing filler 29 may be made of a cured synthetic resin. According to such a configuration, since the reinforcing filler 29 is made of a cured synthetic resin, the form of the reinforcing filler 29 is constant when the composite sandwich panel 20 is manufactured, thereby the reinforcing filler. 29 is easy to handle in the filling operation, the workability of the manufacturing operation is improved, and the productivity of the composite sandwich panel 20 can be improved.
本発明の実施のさらに他の形態では、図１の仮想線５１，５２で示されるように、前述と同様なプレプリグから成る幅ｂ＝約６ｃｍの補助充填材をコア先端部の両側に成形前に貼着し、前述と同様にオートクレーブによって加熱成形して硬化させて一体形成してもよい。この場合には、亀裂発生想定位置のせん断応力τは１．７６×１０−３ｋｇ／ｍｍ２から１．７２×１０−３ｋｇ／ｍｍ２に変化し、約２％低減することができる。このような低減量はわずかであるが、前述の補強用充填材と併用することによって、せん断応力を微調整するために利用することができる。 In still another embodiment of the present invention, as indicated by phantom lines 51 and 52 in FIG. 1, an auxiliary filler made of a prepreg similar to the above and having a width b = about 6 cm is formed on both sides of the core tip before molding. In the same manner as described above, the film may be formed by heating and curing with an autoclave and then integrally formed. In this case, the shear stress τ at the assumed crack generation position changes from 1.76 × 10 −3 kg / mm 2 to 1.72 × 10 −3 kg / mm 2, and can be reduced by about 2%. Although such a reduction amount is slight, it can be used to finely adjust the shear stress by using it together with the reinforcing filler described above.
２０ 複合材サンドイッチパネル
２１ 第１表面
２２ 第２表面
２３ 基部
２４ 傾斜部
２５ コア層
２６ 第１スキン層
２７ 第２スキン層
２８ 積層開始点
２９ 補強用充填材
３０ パネル本体部
３１ 結合部
３３ 添え板
３５ ファスナ
４１ 供試体
４５，４７，４８ 供試体モデル
４６ コア先端部
Ｐ 亀裂発生想定位置
Ａ〜Ｄ 歪ゲージ
DESCRIPTION OF SYMBOLS 20 Composite material sandwich panel 21 1st surface 22 2nd surface 23 Base 24 Inclined part 25 Core layer 26 1st skin layer 27 2nd skin layer 28 Lamination start point 29 Reinforcing filler 30 Panel body part 31 Joint part 33 Attachment plate 35 Fastener 41 Specimen 45, 47, 48 Specimen model 46 Core tip P Crack initiation position A to D Strain gauge
Claims (1)
繊維強化複合材料から成り、コア層の前記第１表面上に一体的に形成される第１スキン層と、
繊維強化複合材料から成り、コア層の前記第２表面上に一体的に形成される第２スキン層とによって、
パネル本体部が形成されるとともに、
前記第１スキン層と第２スキン層とを積層して一体化した結合部が形成される複合材サンドイッチパネル結合部の亀裂発生抑制方法であって、
前記コア層の傾斜部と第１および第２スキン層の積層開始点との間に、硬化後の状態でコア層よりも剛性の高い未硬化のプリプレグから成る補強用充填材を介在させて、前記第１および第２スキン層ならびにコア層と前記補強用充填材とを一体化することを特徴とする複合材サンドイッチパネル結合部の亀裂発生抑制方法。 A slope made of a synthetic foamed resin, in which a first surface on one side in the thickness direction and a second surface on the other side in the thickness direction are parallel to each other, and the first surface and the second surface are inclined in a direction close to each other. A core layer of continuous parts,
A first skin layer comprising a fiber reinforced composite material and integrally formed on the first surface of the core layer;
A second skin layer comprising a fiber reinforced composite material and integrally formed on the second surface of the core layer;
A panel body is formed,
A method for suppressing crack occurrence in a composite material sandwich panel joint part in which a joint part is formed by laminating and integrating the first skin layer and the second skin layer,
Between the inclined portion of the core layer and the lamination start point of the first and second skin layers, interposing a reinforcing filler made of an uncured prepreg that is stiffer than the core layer in a state after curing , A method for suppressing the occurrence of cracks in a composite material sandwich panel joint, wherein the first and second skin layers, the core layer, and the reinforcing filler are integrated.
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| FR2905739B1 (en) * | 2006-09-08 | 2008-11-07 | Airbus France Sas | PANEL ASSEMBLY AND METHOD OF MOUNTING PANEL ASSEMBLY |
| DE102007003275B4 (en) * | 2007-01-23 | 2013-11-28 | Airbus Operations Gmbh | Shell element as part of an aircraft fuselage |
| JP4928403B2 (en) * | 2007-09-27 | 2012-05-09 | 社団法人日本航空宇宙工業会 | Structure for preventing delamination of sandwich panels |
| DE102007055233A1 (en) * | 2007-11-20 | 2009-05-28 | Airbus Deutschland Gmbh | Coupling device for joining fuselage sections, combination of a coupling device and at least one fuselage section and method for producing the coupling device |
| JP5357586B2 (en) * | 2009-03-19 | 2013-12-04 | ジャパンマリンユナイテッド株式会社 | FRP molded product |
| GB201012552D0 (en) | 2010-07-23 | 2010-11-17 | Bae Systems Plc | Aircraft thermal insulation |
| JP5875147B2 (en) * | 2011-11-21 | 2016-03-02 | 学校法人金沢工業大学 | Composite structure and aircraft fuselage |
| DE102012111022A1 (en) | 2012-11-15 | 2014-06-26 | Airbus Operations Gmbh | Reinforced vehicle structure part, vehicle and process |
| WO2015008536A1 (en) * | 2013-07-18 | 2015-01-22 | 日産自動車株式会社 | Molded fiber-reinforced composite material, method for producing same, and panel material |
| EP2993124B1 (en) * | 2014-09-08 | 2019-04-03 | Airbus Operations GmbH | Preventing cracks at bolted or riveted joints of aircraft structural parts |
| KR102213059B1 (en) * | 2015-01-06 | 2021-02-05 | 에스케이케미칼 주식회사 | Composite material panel for ship and menufacturing method thereof |
| JP2018168893A (en) * | 2017-03-29 | 2018-11-01 | 積水化成品工業株式会社 | Joint structure |
| FR3101388B1 (en) * | 2019-10-01 | 2022-02-25 | Airbus Operations Sas | ASSEMBLY COMPRISING A STRUCTURE, TWO PANELS AND A WATERPROOFING SYSTEM |
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