JP6502719B2 - Aircraft structure, method of manufacturing aircraft structure, and method of creating design information of aircraft structure - Google Patents

Description

  Embodiments of the present invention relate to an aircraft structure, a method of manufacturing the aircraft structure, and a method of creating design information of the aircraft structure.

  When designing an aircraft structure, it is important to secure the strength of the structure while reducing the weight of the structure. Therefore, in recent years, composite materials such as carbon fiber reinforced plastic (CFRP: Carbon Fiber Reinforced Plastics) and glass fiber reinforced plastics (GFRP: plastic) are used as materials for aircraft structures. There is.

  An aircraft structure composed of a composite material is formed by laminating and heat curing a thin sheet called a prepreg in which reinforcing fibers are impregnated in a resin. For this reason, the strength in the lamination direction of the composite material is smaller than the strength in the direction along the surface of the prepreg. Therefore, in order to secure the strength in the stacking direction of the composite material, a reinforcement material made of titanium, CFRP, or the like may be provided in the aircraft structure formed of the composite material.

  However, the provision of reinforcements leads to an increase in the weight and manufacturing costs of the aircraft structure. Therefore, various attempts have been made to improve the strength of the structure made of the composite material or the composite material itself.

  For example, in order to improve the pull-off strength of a stringer (longitudinal material), which is a typical structure in which securing of strength is important, carbon fibers or interlayer fibers after lamination which are combined to constitute the stringer. A technique has been proposed in which graphite fibers are inserted in a woven or knitted state (see, for example, Patent Document 1).

  According to this technique, since the adhesive applied to the filler formed between the stringer web and the flange is likely to bleed to carbon fibers or graphite fibers, the filler is strongly adhered to the web and the flange. It is described that it can.

  In addition, in order to improve the strength of the composite material itself, a technique has been proposed in which an interlayer reinforcing sheet for improving the toughness of the prepreg before lamination is sandwiched (see, for example, Patent Document 2). As a preferable interlayer reinforcing sheet from the viewpoint of improving toughness, it is described that non-woven fabric, mesh made of thermoplastic resin and particles made of thermoplastic resin are mentioned.

  Similarly, in order to improve the strength of the composite itself, a technique is also proposed in which two types of non-woven fabrics other than fibers and resins are contained in the state of one or more non-woven fabric layers in the prepreg before lamination (for example, Patent Document 3). According to this technology, it is described that the peel strength between the layers of the composite material can be improved.

Japanese Patent Application Publication No. 2011-520690 Japanese Patent Application Laid-Open No. 2003-019763 Japanese Patent Application Publication No. 2014-502569

  An object of the present invention is to further improve the strength of an aircraft structure.

An aircraft structure according to an embodiment of the present invention corresponds to a first plurality of laminated prepregs, and includes at least a first composite portion forming a web of stringers, and a second plurality of laminated prepregs. A second composite part corresponding to the panel to which the stringer is to be attached or a filler for filling the gap formed between the flange of the stringer and the web, the first composite part A second composite part integrated with the part, and a non-woven fabric inserted in at least one place between the filler and the web, between the filler and the panel, and between the filler and the flange ; and it has a.
Further, an aircraft structure according to an embodiment of the present invention includes a first composite portion corresponding to a first plurality of stacked prepregs and a second composite corresponding to a second plurality of stacked prepregs. a wood part, and a second composite section which is integral with the first composite section, and inserted nonwoven between said first composite portion and the second composite section In the portion where the main load applied between the first composite portion and the second composite portion is a shear load, the first nonwoven fabric for increasing the shear strength is inserted. The second non-woven fabric for enhancing the tensile strength is inserted into a portion where the main load applied between the first composite portion and the second composite portion is a tensile load.
In the method of creating design information of an aircraft structure according to an embodiment of the present invention, a step of determining at least a first composite portion constituting at least a web of stringers corresponding to the first plurality of laminated prepregs is determined. And a second composite corresponding to the laminated second plurality of prepregs and constituting a filler to be formed between the panel to which the stringer is to be attached or the flange of the stringer and the web. Determining the second composite part integrated with the first composite part , between the filler and the web, between the filler and the panel, and the filler said first composite section by determining the type of non-woven fabric to be inserted into at least one position, constituting the second composite section and the nonwoven essential between said flange and In which a step of creating a design information of the aircraft structure to.
In the method of creating design information of an aircraft structure according to an embodiment of the present invention, a step of determining a first composite portion corresponding to a first plurality of stacked prepregs ; Determining a second composite part corresponding to a plurality of prepregs, the second composite part being integrated with the first composite part, the first composite part and the first composite part Design information of an aircraft structure having the first composite portion, the second composite portion, and the non-woven fabric as components by determining the type of non-woven fabric inserted between the two composite portions And, if the main load applied between the first composite part and the second composite part is a shear load, the first composite part and the first composite part. The first non-woven fabric can increase shear strength by being inserted between the second composite part and the second composite part. If the main load applied between the first composite portion and the second composite portion is a tensile load while selecting a cloth, the first composite portion and the second composite portion may be selected. The second non-woven fabric can be selected to increase the tensile strength by inserting it between the second and third composite members .
In the method of creating design information of an aircraft structure according to an embodiment of the present invention, a step of determining a first composite portion corresponding to a first plurality of stacked prepregs; Determining a second composite part corresponding to a plurality of prepregs, the second composite part being integrated with the first composite part, the first composite part and the first composite part Design information of an aircraft structure having the first composite portion, the second composite portion, and the non-woven fabric as components by determining the type of non-woven fabric inserted between the two composite portions And forming the first composite material portion and the portion having a shear load as a main load applied between the first composite material portion and the second composite material portion. The shear strength can be increased by inserting it between the second composite part The first composite portion and the second non-woven fabric are selected for the portion where the main load applied between the first composite portion and the second composite portion is a tensile load. A second non-woven fabric whose tensile strength can be enhanced by inserting it between the two composite parts is selected.
The method for manufacturing an aircraft structure according to an embodiment of the present invention further includes the steps of: laminating a first plurality of prepregs to form at least a first laminate forming a web of stringers ; Forming a second laminate that constitutes a filler for filling the gap formed between the panel to which the stringer is to be attached or the flange of the stringer and the web ; Between the first laminate and the second laminate before or before recuring, between the filler and the web, between the filler and the panel, and between the filler and the flange steps and, the first laminate and the second laminate before curing before or re-hardening the nonwoven fabric is inserted to insert a nonwoven fabric in at least one place between the In which a step of curing.
Further, in the method for manufacturing an aircraft structure according to the embodiment of the present invention, a step of laminating a first plurality of prepregs to form a first laminated body, and laminating a second plurality of prepregs to form a second Forming a laminate, inserting a non-woven fabric between the first and second laminates before curing or re-curing, and curing before or re-inserting the non-woven fabric. Curing the first laminate and the second laminate before curing, and a first composite portion obtained as the first laminate after curing and the second after curing In a portion where a main load applied between the second composite part obtained as a laminate of the present invention is a shear load, a first non-woven fabric for increasing the shear strength is inserted, while the first composite part is inserted Main load between the two and the second composite part The portion that the load is a tensile load, is to insert a second nonwoven to improve the tensile strength.

BRIEF DESCRIPTION OF THE DRAWINGS The cross-sectional view which shows the structure of the aircraft structure which concerns on embodiment of this invention. Graph showing the Type of Mode I interlaminar fracture toughness G _IC and Mode II interlaminar fracture toughness G _IIC of the nonwoven fabric. The table | surface which shows the property of the nonwoven fabric shown in FIG. The figure which shows the example which changes the kind of nonwoven fabric according to whether the main load concerning a filler is tensile load or shear load. The figure which shows the example which inserts the several non-woven fabric from which a kind differs according to whether the main load concerning a filler is tensile load or shear load in a different position.

  An aircraft structure, a method of manufacturing an aircraft structure, and a method of creating design information of an aircraft structure according to an embodiment of the present invention will be described with reference to the attached drawings.

(Structure of aircraft structure)
FIG. 1 is a cross-sectional view showing the structure of an aircraft structure according to an embodiment of the present invention.

  The aircraft structure 1 is a structure including a spar, a small bone (rib), an outer plate (panel) (also referred to as a skin), and a stringer, which constitute an aircraft, and is a structure formed of a composite material. . FIG. 1 shows, as an example, the case where the aircraft structure 1 is a type I stringer 2 made of a composite material.

  The type I stringer 2 is a stringer having an I-shaped cross section as shown in FIG. That is, the type I stringer 2 has a structure in which two flanges 3 are connected by a web 4. The type I stringer 2 is attached to the panel 5 as a reinforcement. Typically, the flange 3 on the panel 5 side of a type I stringer 2 is secured to the panel 5 by fasteners 6. In addition, as required, a radial filler 7 for reinforcing an R-chamfered portion formed between the flange 3 and the web 4 on the panel 5 side is fixed by the fastener 6 together with the flange 3 on the panel 5 side.

  The type I stringer 2 is manufactured by heat curing after assembling a laminate of a plurality of sheet-like prepregs. Accordingly, the type I stringer 2 has a plurality of composite material layers 2A, 2B, 2C, 2D, 2E, 2F in which the lamination directions of the prepregs, that is, the fiber directions are different from each other. Specifically, as shown in FIG. 1, the type I stringer 2 has six composite material layers 2A, 2B, 2C, 2D, 2E, 2F.

  The first composite material layer 2A is a plate-like composite material layer that forms the upper surface side of the flange 3 on the side away from the panel 5. The second composite material layer 2B is a plate-like composite material layer that forms the lower surface side of the flange 3 on the panel 5 side. The third composite material layer 2C is a composite material layer having an inverted C-shaped cross section forming the lower surface side of the flange 3 on the side away from the panel 5, one side of the web 4 and the upper surface side of the flange 3 on the panel 5 side. It is. The fourth composite layer 2D is a composite layer having a C-shaped cross section forming the lower surface side of the flange 3 on the side away from the panel 5, one side of the web 4 and the upper surface side of the flange 3 on the panel 5 side. is there.

  The third composite layer 2C of inverted C-shaped cross section and the fourth composite layer 2D of C-shaped cross section are combined with each other as the inner left and right sides of the web 4 and the two flanges 3 respectively Ru. The third composite layer 2C and the fourth composite layer 2D are combined on the second composite layer 2B. The first composite layer 2A is combined on the third composite layer 2C and the fourth composite layer 2D.

  Then, a void having a substantially inverted triangle cross section is generated between the first composite material layer 2A, the third composite material layer 2C, and the fourth composite material layer 2D. Therefore, a bar-like fifth composite material layer 2E having a substantially inverted triangle cross section is provided in a space generated between the first composite material layer 2A, the third composite material layer 2C, and the fourth composite material layer 2D. Be

  Similarly, a void having a substantially triangular cross section is generated between the second composite layer 2B, the third composite layer 2C, and the fourth composite layer 2D. Therefore, a rod-like sixth composite material layer 2F having a substantially triangular cross section is provided in the space generated between the second composite material layer 2B, the third composite material layer 2C, and the fourth composite material layer 2D. .

  Thus, the type I stringer 2 has a plurality of composite material layers 2A, 2B, 2C, 2D, 2E, 2F corresponding to a plurality of laminated prepregs. The same is true for other non-flat plate structures such as stringer and panel integral structures, spars or ribs.

  The filler for filling the space generated between the web 4 and the flange 3 as in the fifth composite material layer 2E and the sixth composite material layer 2F is also called noodle. When heat curing is performed simultaneously with the type I stringer 2 combined with the panel 5, the second composite material layer 2B may not be provided. In that case, noodles are provided to fill the air gaps that occur between the web 4, the flanges 3 and the panels 5.

  The type I stringer 2 needs to be designed so as not to break when a load in the direction of pulling away from the panel 5 is applied. When a load in the direction of pulling away from the panel 5 is applied to the type I stringer 2, the portion to which the load is locally applied is the periphery of the noodle on the panel 5 side. When a load in the direction of pulling away from the panel 5 is applied to the type I stringer 2, the load at which breakage occurs around the noodle is called a pull-off load. The type I stringer 2 is designed so that the pull-off load is a load that is sufficiently large relative to the load that is actually assumed. The same applies to other stringers such as T-type stringers and Z-type stringers used under similar conditions. In other words, it is important to design the stringer so that the pull-off load is as large as possible while suppressing the increase in weight.

  Therefore, a non-woven fabric 8 for improving strength is inserted around the sixth composite material layer 2F where stress is concentrated locally when a load in a direction to be separated from the panel 5 is applied to the I-type stringer 2. Be done. The non-woven fabric 8 is a sheet-like material in which fibers are entangled without weaving. In particular, it was confirmed by the test that inserting the non-woven fabric 8 having thermoplasticity is preferable from the viewpoint of improving the strength.

  The non-woven fabric 8 is provided between the sixth composite material layer 2F and the second composite material layer 2B, between the sixth composite material layer 2F and the third composite material layer 2C, and between the sixth composite material layer 2F and It can be inserted in at least one place between the fourth composite layer 2D. Insertion of the appropriate type of non-woven fabric 8 can improve the strength against local tensile stress and shear stress generated in the type I stringer 2.

  The insertion position of the non-woven fabric 8 into the type I stringer 2 is not limited to the periphery of the sixth composite material layer 2F, and may be between other adjacent composite material layers 2A, 2B, 2C, 2D, and 2E. In addition, the non-woven fabric 8 may be locally inserted into the respective composite material layers 2A, 2B, 2C, 2D and 2E. Furthermore, the strength of the aircraft structure is improved by inserting the non-woven fabric at the required position between at least two adjacent composite layers not only for the type I stringer 2 but also for other aircraft structures composed of composites. be able to. It is also possible to improve the strength of the desired aircraft structure by inserting the non-woven fabric locally at the required position inside the composite layer in the same stacking direction.

  That is, an arbitrary aircraft structure is integrated with a first composite portion corresponding to the first plurality of laminated prepregs and a second composite portion corresponding to the second plurality of laminated prepregs. Can be organized and organized. The mechanical properties of the aircraft structure can then be improved by inserting the non-woven fabric between the first composite part and the second composite part. That is, the non-woven fabric can be locally inserted at a position where the load of the composite material layer constituting the aircraft structure is concentrated.

  Generally, after combining a laminate of a plurality of prepregs, a composite structure is manufactured by heat curing, but a laminate of a prepreg after curing and a laminate of a prepreg before curing are produced. Composite structures may also be produced by combining and re-curing. For example, when manufacturing a structure having a structure in which the stringer is attached to the panel, after curing the laminate of the stringer, the adhesive surface of the stringer after curing is adhered onto the panel before curing with a film adhesive. In some cases, heat curing may occur. In this case, the stringer will be recured. Therefore, the non-woven fabric is inserted between the laminates of adjacent prepregs before curing or re-curing at the time of manufacture.

  Further, in the case where a non-woven fabric is locally inserted into the composite material layer in the same lamination direction, a non-woven fabric is inserted after laminating a predetermined number of prepregs, and then a predetermined number of prepregs are laminated again and then heat cured. Become. Therefore, after curing, the first composite portion corresponding to the first plurality of prepregs, and the second composite corresponding to the second plurality of prepregs laminated in the same stacking direction as the first plurality of prepregs. The non-woven fabric is inserted between the material part.

  On the other hand, as described above, in the case of combining a plurality of laminates having different prepreg lamination directions and inserting a non-woven fabric between the laminates, after curing, the first composite corresponding to the first plurality of prepregs The non-woven fabric is inserted between the material portion and the second composite portion corresponding to the second plurality of prepregs stacked in the stacking direction different from the first plurality of prepregs.

  When the aircraft structure is a stringer, such as a Type I stringer 2, the aircraft structure comprises a first composite section constituting at least a web of the stringer and a flange or web of the panel or stringer to which the stringer is to be attached And a second composite part that constitutes a filler that fills the gap formed between the two. Thus, the non-woven fabric can be inserted into at least one place between the filler and the web, between the filler and the panel, and between the filler and the flange. Of course, it is also possible to insert the non-woven fabric locally into the inside of the web or the inside of the flange as required.

  In addition, since tensile stress and shear stress are locally applied to the filler of the stringer, inserting the non-woven fabric also into the inside of the filler leads to an increase in the pull-off load of the stringer. The filler can be formed by piling up the prepreg in a spiral shape. Therefore, it is possible to insert the non-woven fabric spirally between the spiral prepregs. Such a non-woven fabric inserted filler can be produced, for example, by overlapping and rolling a prepreg and a non-woven fabric.

(Design method of aircraft structure)
Next, a method of creating design information of an aircraft structure including determination of the type of non-woven fabric will be described.

  When creating design information of an aircraft structure as exemplified in FIG. 1, the first corresponding to the first plurality of prepregs stacked based on the size such as the board thickness and cross-sectional area of the aircraft structure A first composite part constituting one composite part, for example, a web, a flange or the like is determined. Also, a second composite part, such as a filler, integrated with the first composite part is determined. The second composite portion is a composite portion corresponding to the stacked second plurality of prepregs. The stacking direction of the second plurality of prepregs may be the same as or different from the stacking direction of the first plurality of prepregs.

  Next, the type of non-woven fabric inserted between the first composite part and the second composite part is determined. This makes it possible to create design information of an aircraft structure having improved mechanical strength, which includes the first composite portion, the second composite portion, and the non-woven fabric as components. Then, design information of the created aircraft structure can be provided to a design client, a designer, or the like.

  The type of non-woven fabric can be determined depending on whether the main load applied between the first composite portion and the second composite portion is a tensile load or a shear load. Whether the main load applied between the first composite part and the second composite part is a tensile load or a shear load can be determined by conducting stress analysis of the aircraft structure .

  For example, if the aircraft structure is an I-type stringer 2 as shown in FIG. 1, the flange 3 on the side away from the panel 5 may be formed on the I-type stringer 2 by stress analysis using FEM (Finite Element Method) or the like. The stress distribution can be determined when a tensile load is applied in a direction away from the panel 5. Thereby, between the sixth composite material layer 2F and the second composite material layer 2B which are fillers, between the sixth composite material layer 2F and the third composite material layer 2C, and the sixth composite material layer Each tensile stress and each shear stress between 2F and the fourth composite material layer 2D can be determined. And, between the sixth composite material layer 2F and the second composite material layer 2B, between the sixth composite material layer 2F and the third composite material layer 2C, and between the sixth composite material layer 2F and the fourth It can be determined whether the main load between the composite material layer 2D and the composite material layer 2D is a tensile load or a shear load. In addition, in particular, locations where tensile stress or shear stress is concentrated can also be identified.

  If the aircraft structure is a structure that receives the pressure of air such as a main wing, the first composite portion and the first composite part that configure the aircraft structure by executing stress analysis under the condition that loads the pressure distribution of air. The tensile stress and the shear stress between the two composite parts may be determined.

  Moreover, if the destructive test of the aircraft structure is performed, it is possible to confirm the location where stress concentration occurs, without necessarily performing the numerical analysis for obtaining the stress distribution. In fact, a fracture test is performed by applying a tensile load to a T-type stringer in which a filler is filled with voids by combining two symmetrical composite layers with an inverted L-shaped cross section and a flat composite layer. Did.

  As a result, in the case of a T-type stringer having a thickness of 20 ply laminated prepregs, it was confirmed that the origin of breakage is inside the filler. Thereafter, when the tensile load was increased, it was confirmed that cracks were generated in the inside of the web and the surface on the flange side of the filler, and the cracks gradually developed.

  On the other hand, in the case of a T-type stringer having a thickness of 56 ply laminated prepregs, peeling of the composite material occurred at the R-chamfered portion of the inverted L-shaped composite material layer. Thereafter, when the tensile load was increased, new peeling occurred between the filler and the flat composite layer separately from the peeling at the R-chamfered portion.

  Thus, the failure mode of the aircraft structure changes depending on the size, cross-sectional area, etc. of each part of the aircraft structure. In addition to the distribution and orientation of the applied external load, it is believed that the failure mode of the aircraft structure will also change due to factors that affect the strength, such as the location and type of fasteners. That is, the position where the load is locally applied and whether the main load is a tensile load or a shear load depend on the size and structure of the aircraft structure and the distribution and direction of the external load applied to the aircraft structure. Change.

  Even when stress analysis of an aircraft structure is performed, it is considered that if the analysis model is appropriately prepared, it is possible to obtain the same result as the destructive test. That is, it is considered that the analysis model changes according to the distribution and direction of the external load applied to the aircraft structure and the size and structure of the aircraft structure, and the result of the stress analysis also changes.

  When stress analysis or failure test of such an aircraft structure is performed, it is possible to determine an appropriate non-woven fabric type and an appropriate insertion position of the non-woven fabric based on the results of the stress analysis or failure test. That is, it is sufficient to insert the non-woven fabric at a position where the load is concentrated. Specifically, the plurality of composite parts to be integrated are determined, and the insertion position of the non-woven fabric into the plurality of composite parts based on the result of the destructive test or stress analysis of the aircraft structure, ie, between the composite parts And by determining the insertion position of the non-woven fabric into at least one of the inside of the composite part, it is possible to create design information of the aircraft structure having the plurality of composite parts and the non-woven fabric as components.

  Also, select a non-woven fabric with an effect of improving tensile strength as the composite material layer in which the tensile load is dominant, and select a non-woven fabric with an improvement effect in the shear strength as the composite material layer in which the shear load is dominant. It means that it is good.

FIG. 2 is a graph showing Mode I interlaminar fracture toughness value G _IC and Mode II interlaminar fracture toughness value G _{IIC according} to the type of nonwoven fabric, and FIG. 3 is a table showing the properties of the nonwoven fabric shown in FIG.

The vertical axis of the bar graph in FIG. 2A indicates the mode I interlaminar fracture toughness value G _IC, and the vertical axis of the bar graph in FIG. 2B indicates the mode II interlaminar fracture toughness value G _IIC . In addition, the numbers described in each bar graph indicate the average value of a plurality of data measured under the condition corresponding to the bar graph, and the width of the line segment described in each bar graph corresponds to the condition corresponding to the bar graph The width between the maximum value and the minimum value of a plurality of measured data is shown.

As shown in FIG. 3, a non-woven fabric selected from non-woven fabrics specified by product numbers LNS0030, LNS1225, G5025, G4020 and RD13-0003 manufactured by Toha Tech Co., Ltd. is inserted between the composite layers and the Mode I interlaminar fracture toughness value is shown. G _IC and Mode II interlaminar fracture toughness G _{II C} were measured.

The non-woven fabric specified by the product No. LNS 0030 is a non-woven fabric mainly composed of polyamide and having a basis weight of 30 g / m ² , a melting point of 115 ° C., and a melt viscosity at 160 ° C. of 650 Pa · s. The nonwoven fabric specified by product No. LNS 1225 is a nonwoven fabric mainly composed of polyamide and having a basis weight of 25 g / m ² , a melting point of 115 ° C., and a melt viscosity at 160 ° C. of 1100 Pa · s. The nonwoven fabric specified by product No. G5025 is a nonwoven fabric mainly composed of polyester having a basis weight of 25 g / m ² , a melting point of 150 ° C., and a melt viscosity at 190 ° C. of 300 Pa · s. The nonwoven fabric specified by product No. G4020 is a nonwoven fabric mainly composed of polyester having a basis weight of 20 g / m ² , a melting point of 143 ° C., and a melt viscosity at 190 ° C. of 590 Pa · s. The nonwoven fabric specified by product number RD 13-0003 is a nonwoven fabric mainly composed of copolyester having a basis weight of 15 g / m ² , a melting point of 95 ° C. to 103 ° C., and a melt viscosity at 160 ° C. of 230 Pa · s. Incidentally, the basis weight is the weight per unit area, usually the weight per 1 m ² is displayed as basis weight.

The leftmost data in the bar graphs of FIGS. 2A and 2B are data measured without inserting a non-woven fabric between the composite layers. Other data are data measured by inserting a non-woven fabric between composite layers. As for mode II interlaminar fracture toughness value G _IIC , data was acquired by changing the number of non-woven fabrics to be inserted.

According to FIG. 2 (A), the nonwoven fabric mainly composed of polyamide specified by product number LNS0030, the nonwoven fabric mainly composed of polyamide specified by product number LNS1225, the nonwoven fabric mainly composed of polyester specified by product number G5025 Even if any of the non-woven fabric composed mainly of copolyester specified by the part number RD13-0003 is inserted between the composite layers, the Mode I interlaminar fracture toughness value G _IC compared to the case where the non-woven fabric is not inserted between the composite layers. Can be confirmed. In particular, when a non-woven fabric composed mainly of a polyamide specified by part number LNS 0030 is inserted between the composite layers, it can be confirmed that the increase in Mode I interlaminar fracture toughness value G _IC is large as compared with the case where other non-woven fabric is inserted. .

On the other hand, according to FIG. 2 (B), when non-woven fabric is not inserted between composite layers if three non-woven fabrics made mainly of copolyester specified by product number RD13-0003 are laminated and inserted between composite layers. It can be confirmed that the mode II interlaminar fracture toughness value G _IIC is increased as compared with the above.

As a result of the above, when a non-woven fabric composed mainly of a copolyester specified by part number RD13-0003 is inserted between the composite layers, both Mode I interlaminar fracture toughness value G _IC and Mode II interlaminar fracture toughness value G _IIC can be obtained. It turns out that it can improve. However, in order to increase the mode I interlaminar fracture toughness value G _IC , it is understood that it is more effective to insert a non-woven fabric mainly composed of the polyamide specified by the product number LNS 0030 into the composite material layer.

A large Mode I interlaminar fracture toughness value G _IC means that the strength to a mode I tensile load is large. On the other hand, the fact that the mode II interlaminar fracture toughness value G _IIC is large means that the strength against the mode II shear load is large.

  For this reason, it is possible to use a first nonwoven fabric mainly composed of copolyester as the first raw material, regardless of whether the main load applied between the first composite part and the second composite part is a tensile load or a shear load. The strength of the aircraft structure can be improved by inserting it between the composite part and the second composite part. Therefore, the part where the main load applied between the first composite part and the second composite part is a tensile load, and the load between the first composite part and the second composite part Even in the case where a portion where the main load to be performed is a shear load is mixed, the same kind of non-woven fabric can be used if non-woven fabric containing co-polyester as a main raw material is used. Therefore, for example, a non-woven fabric composed mainly of copolyester can be inserted into and around the noodle filler.

On the other hand, a plurality of non-woven fabrics of different types may be inserted between the first composite part and the second composite part, or an optimal non-woven fabric selected from a plurality of non-woven fabrics of different types as the first composite It can also be inserted between the material part and the second composite part. Therefore, the number of candidates can be increased by inserting various non-woven fabrics between composite layers and measuring the mode I interlaminar fracture toughness value G _IC and the mode II interlaminar fracture toughness value G _IIC . That is, not only the non-woven fabric listed in FIG. 3 but also other non-woven fabric can be used.

  FIG. 4 is a view showing an example in which the type of non-woven fabric is changed depending on whether the main load applied to the filler is a tensile load or a shear load.

  In general, when the thickness of the composite layer constituting the flange or the web is thin, the failure mode of the composite layer is mode I. That is, when the thickness of the composite layer is thin, tensile stress is dominant. Conversely, when the thickness of the composite layer is large, the failure mode of the composite layer becomes mode II. That is, when the thickness of the composite layer is large, shear stress is dominant.

Therefore, as shown in FIG. 4A, when the thickness of the composite material layer 11 to be combined with the filler 10 is thin, the nonwoven fabric 12 _I having high tensile strength is placed between the filler 10 and the composite material layer 11. It is appropriate to insert. For this reason, for example, a non-woven fabric having high tensile strength and mainly containing polyamide specified by product number LNS 0030 shown in FIG. 3 can be inserted between the filler 10 and the composite material layer 11.

Conversely, as shown in FIG. 4B, when the thickness of the composite material layer 11 to be combined with the filler 10 is large, the nonwoven fabric 12 _II having high shear strength is placed between the filler 10 and the composite material layer 11. It is appropriate to insert in For this reason, for example, it is possible to insert three layers of non-woven fabric composed mainly of a copolyester specified by part number RD13-0003 shown in FIG. 3 and insert it between the filler 10 and the composite material layer 11.

  Thus, if the main load applied between the first composite portion and the second composite portion is a shear load, the first composite portion and the second composite portion The first non-woven fabric can be selected such that the shear strength can be increased by inserting it in between. On the other hand, when the main load to be applied between the first composite portion and the second composite portion is a tensile load, between the first composite portion and the second composite portion The second non-woven fabric can be selected such that the tensile strength can be increased by the insertion.

  FIG. 5 is a view showing an example in which a plurality of non-woven fabrics of different types are inserted at different positions depending on whether the main load applied to the filler 10 is a tensile load or a shear load.

  As shown in FIG. 5, a tensile load and a shear load may be applied to different surfaces of the filler 10 as main loads, respectively. In addition, the main load applied to the inside of the filler 10 is often a tensile load.

Therefore, non-woven fabric 12 _II having high shear strength can be inserted between the filler 10 and the composite material layer 11 A in which shear load is dominant. For example, non-woven fabrics composed mainly of a copolyester specified by part number RD13-0003 shown in FIG. 3 can be laminated in three plies and inserted between the filler 10 and the composite material layer 11A.

On the other hand, non-woven fabric 12 _I having high tensile strength can be inserted between the inside of the filler 10 where the tensile load is dominant and between the filler 10 and the composite material layer 11B. For example, a non-woven fabric whose main raw material is a polyamide specified by product number LNS 0030 shown in FIG. 3 can be inserted between the inside of filler 10 and between filler 10 and composite material layer 11B.

  Thus, for the portion where the main load applied between the first composite portion and the second composite portion is a shear load, the first composite portion and the second composite portion The first non-woven fabric can be selected such that the shear strength can be increased by inserting it in between. On the other hand, for a portion where the main load applied between the first composite portion and the second composite portion is a tensile load, the portion between the first composite portion and the second composite portion is used. The second non-woven fabric can be selected such that the tensile strength can be increased by the insertion.

  That is, different types of non-woven fabrics can be inserted at different positions depending on the direction of the main load. In this case, the strength of the aircraft structure can be further improved as compared with the case of inserting the same kind of non-woven fabric. Also within the filler 10, according to the direction of the main load, it is possible to insert a non-woven fabric of the same or a different type as the non-woven fabric inserted around the filler 10.

  As described above, the direction of the main load, ie, whether the main load is a tensile load or a shear load, changes depending on the dimensions such as the thickness of each member of the aircraft structure. For example, if the aircraft structure is a stringer, the direction of the main load is different for each portion such as the web, flange, peripheral portion of the R-chamfer and filler. In addition, the position at which concentration of tensile load or shear load occurs also changes depending on dimensions such as the thickness of each member constituting the aircraft structure.

  Therefore, the relationship between the dimensions such as the plate thickness of each member constituting the aircraft structure and the direction of the main load applied to each member can be checked and recorded as reference information. Alternatively, the dimensions such as the thickness of each member constituting the aircraft structure and the type of non-woven fabric to be inserted may be directly associated with each other and stored as reference information. In this case, the type of nonwoven fabric to be inserted and the position to be inserted can be associated and stored as reference information.

  That is, it is possible to store, as reference information, the relationship between dimensions such as the thickness of each member constituting the aircraft structure, the type of non-woven fabric to be inserted and the position at which the non-woven fabric is to be inserted. Then, only by referring to the reference information, it is possible to determine the type and insertion position of the non-woven fabric without performing stress analysis or a destructive test based on dimensions such as plate thickness of each member constituting the aircraft structure. Become.

  However, when all dimensions of each member constituting the aircraft structure are used as parameters, the number of parameters is extremely large. For this reason, at least a part of the size of the aircraft structure and the main load applied between the first composite portion and the second composite portion constituting the aircraft structure is a tensile load or a shear load It is possible to create reference information representing the relationship with the subject. Alternatively, reference information may be created that represents the relationship between the size of at least a portion of the aircraft structure and the appropriate type of nonwoven. This makes it possible to easily determine the type of non-woven fabric with reference to the reference information. That is, it becomes possible to insert a thermoplastic non-woven fabric having high strength in the direction of the main load between the appropriate composite layers such as the web, the flange, the peripheral portion of the R-chamfer and the filler.

(Method of manufacturing aircraft structure)
Next, a method of manufacturing an aircraft structure designed by the above-described method will be described.

  When manufacturing an aircraft structure, a laminate of at least two prepregs to be integrated is created. That is, the first plurality of prepregs are stacked to form a first stacked body. On the other hand, the second plurality of prepregs are stacked to form a second stacked body. The stacking direction of the second plurality of prepregs may be the same as or different from the stacking direction of the first plurality of prepregs.

  As a specific example, if the aircraft structure is an I-type stringer 2 shown in FIG. 1, 6 layers corresponding to 6 composite material layers 2A, 2B, 2C, 2D, 2E, 2F can be obtained by laminating a plurality of prepregs. A laminate of two prepregs is made. A non-woven fabric may be inserted inside the laminate of the prepreg corresponding to the sixth composite material layer 2F which is a noodle filler in which the load is concentrated.

  Next, assembly of the first laminate and the second laminate is performed with insertion of the non-woven fabric between the first laminate and the second laminate. That is, when assembling the laminate of the prepreg before curing, a thermoplastic nonwoven fabric whose type is previously determined according to the direction of the load is inserted at an appropriate position between the laminates of the prepreg.

  As a specific example, if the aircraft structure is type I stringer 2 shown in FIG. 1, six prepreg laminates respectively corresponding to six composite material layers 2A, 2B, 2C, 2D, 2E, 2F can be combined. . At this time, the non-woven fabric 8 is inserted around the laminate of the prepreg corresponding to the sixth composite material layer 2F which is a noodle filler in which the load is concentrated. The nonwoven fabric 8 may be covered on the sixth composite material layer 2F after the adhesive is applied.

  Next, the first laminate and the second laminate of the prepreg before curing assembled with the non-woven fabric inserted are cured. Specifically, the first laminate and the second laminate of the prepreg are carried into the autoclave device in a bagged state. Then, the first laminate and the second laminate are heat-cured under pressure by autoclave treatment.

  This makes it possible to obtain an aircraft structure after curing. For example, as shown in FIG. 1, the six composite material layers 2A, 2B, 2C, 2D, 2E, 2F are integrated in a state where the non-woven fabric 8 is inserted around the sixth composite material layer 2F which is a noodle filler. Type I stringer 2 can be produced.

  As described above, some parts such as stringers may be cured and then combined with the panel and recured. Therefore, the nonwoven fabric can be inserted between the first laminate and the second laminate before curing or before recuring. Then, the first laminate and the second laminate can be cured before curing or recuring in which the non-woven fabric is inserted.

  Further, in the case where the non-woven fabric is locally inserted between the laminates having the same prepreg lamination direction, the non-woven fabric may be inserted during the lamination of the prepreg. That is, the nonwoven fabric may not necessarily be inserted after the lamination of the prepreg. For example, after a first plurality of prepregs are laminated to produce a first laminate, a nonwoven fabric is overlaid on the first laminate, and a second plurality of prepregs are laminated on the nonwoven fabric to form a second laminate. A laminate may be produced.

(effect)
According to the above method of manufacturing an aircraft structure and a method of creating design information of an aircraft structure, a panel reinforced with stringers by locally inserting a non-woven fabric at the position of the composite layer where the load is concentrated The strength of the aircraft structure composed of the composite material can be improved as compared with the prior art. For example, if the aircraft structure is a stringer, the stringer pull-off strength can be improved. For this reason, in some cases, the radial filler 7 can be omitted in the type I stringer 2 illustrated in FIG.

  In particular, the number of prepregs laminated to construct an aircraft structure may reach up to 80 ply. For this reason, as conventionally proposed, when the non-woven fabric is inserted between the prepreg itself and all the prepregs, the plate thickness of the structure is increased.

  On the contrary, in the above-mentioned method, the non-woven fabric is selectively inserted between specific composite layers in which the load is concentrated locally. Therefore, it is possible to prevent an increase in the thickness of the aircraft structure. In addition, the required non-woven fabric can ensure the required strength of the aircraft structure. For this reason, compared with the case where a nonwoven fabric is inserted between the prepreg itself and all the prepregs, the usage-amount of a nonwoven fabric can also be suppressed. In other words, it is possible to easily provide the required strength to the aircraft structure while configuring the aircraft structure using the conventional prepreg in which the non-woven fabric is not inserted.

  Although the specific embodiments have been described above, the described embodiments are merely examples and do not limit the scope of the invention. The novel methods and apparatus described herein may be embodied in various other ways. Also, various omissions, substitutions and changes in the form of the methods and apparatus described herein may be made without departing from the spirit of the invention. The appended claims and their equivalents are intended to cover such different forms and modifications as fall within the scope and spirit of the invention.

1 Aircraft Structure 2 Type I Stringer 2A, 2B, 2C, 2D, 2E, 2F Composite Layer 3 Flange 4 Web 5 Panel 6 Fastener 7 Radial Filler 8 Non-woven 10 Filler 11, 11A, 11B Composite Layer 12 _I , 12 _II Non-woven

Claims (11)

A first composite portion corresponding to the first plurality of laminated prepregs and constituting at least a web of stringers ;
A second composite part corresponding to a second plurality of laminated prepregs and constituting a filler for filling a gap formed between a panel to which the stringer is to be attached or a flange of the stringer and the web A second composite part integrated with the first composite part;
And the nonwoven fabric to be inserted in at least one place between the filler between the web, and between the filler and the filler and the panel and the flange,
An aircraft structure having a . A first composite portion corresponding to the first plurality of laminated prepregs;
A second composite part corresponding to the stacked second plurality of prepregs, wherein the second composite part is integrated with the first composite part;
Non-woven fabric inserted between the first composite part and the second composite part
I have a,
For the portion where the main load applied between the first composite portion and the second composite portion is a shear load, a first non-woven fabric for increasing shear strength is inserted, while the first non-woven fabric is inserted. An aircraft structure in which a second non-woven fabric for enhancing tensile strength is inserted in a portion where a main load applied between the composite portion and the second composite portion is a tensile load . The aircraft structure according to claim 1 or 2 , wherein a non-woven fabric of the same type or a different type as the non-woven fabric is inserted inside the filler. Determining a first composite section corresponding to the first plurality of laminated prepregs and constituting at least a web of stringers ;
A second composite part corresponding to a second plurality of laminated prepregs and constituting a filler for filling a gap formed between a panel to which the stringer is to be attached or a flange of the stringer and the web Determining a second composite part to be integrated with the first composite part;
The first composite part by determining the type of non-woven fabric inserted in at least one place between the filler and the web, between the filler and the panel, and between the filler and the flange ; Creating design information of an aircraft structure having the second composite part and the non-woven fabric as components;
A method of creating design information for an aircraft structure having: Determining a first composite section corresponding to the first plurality of laminated prepregs;
Determining a second composite part corresponding to the second plurality of laminated prepregs, the second composite part being integrated with the first composite part;
The first composite portion, the second composite portion, and the non-woven fabric are configured by determining the type of non-woven fabric inserted between the first composite portion and the second composite portion. Creating design information of an aircraft structure to be an element;
Have
When the main load applied between the first composite portion and the second composite portion is a shear load, the first composite portion and the second composite portion When the first non-woven fabric whose shear strength can be enhanced by inserting between the two is selected while the main load applied between the first composite material part and the second composite material part is a tensile load A method of creating design information of an aircraft structure for selecting a second non-woven fabric whose tensile strength can be enhanced by inserting it between the first composite portion and the second composite portion. Determining a first composite section corresponding to the first plurality of laminated prepregs;
Determining a second composite part corresponding to the second plurality of laminated prepregs, the second composite part being integrated with the first composite part;
The first composite portion, the second composite portion, and the non-woven fabric are configured by determining the type of non-woven fabric inserted between the first composite portion and the second composite portion. Creating design information of an aircraft structure to be an element;
Have
With respect to a portion where the main load applied between the first composite portion and the second composite portion is a shear load, the portion between the first composite portion and the second composite portion is used. The first non-woven fabric whose shear strength can be enhanced by inserting between the two, while the main load applied between the first composite part and the second composite part is the tensile load And a method of creating design information of an aircraft structure in which a second non-woven fabric whose tensile strength can be enhanced by inserting it between the first composite portion and the second composite portion is selected. The method for creating design information of an aircraft structure according to claim 5 or 6 , wherein stress analysis of the aircraft structure is performed to determine whether the main load is a tensile load or a shear load. The method for creating design information of an aircraft structure according to any one of claims 4 to 7, wherein the insertion position of the non-woven fabric is determined based on the result of a destructive test or a stress analysis of the aircraft structure. Relationship between the size of at least one part of the aircraft structure and whether the main load applied between the first composite part and the second composite part is a tensile load or a shear load Create reference information that represents the relationship between the size of at least one part of the aircraft structure and the appropriate type of the non-woven fabric, and determine the type of the non-woven fabric with reference to the reference information Item 4. A method of creating design information of an aircraft structure according to Item 4 . Laminating a first plurality of prepregs to form a first laminate comprising at least a web of stringers;
The second plurality of prepregs are laminated to form a second laminate that constitutes a filler for filling the gap formed between the panel to which the stringer is to be attached or the flange of the stringer and the web. Step and
Between the first laminate and the second laminate before curing or before recuring, between the filler and the web, between the filler and the panel, and between the filler and the flange Inserting the non-woven fabric in at least one place between
Curing the first laminate and the second laminate before curing or recuring in which the non-woven fabric is inserted;
A method of manufacturing an aircraft structure comprising: Laminating a first plurality of prepregs to form a first laminate;
Laminating a second plurality of prepregs to form a second laminate;
Inserting a non-woven fabric between the first laminate and the second laminate before curing or before recuring;
Curing the first laminate and the second laminate before curing or recuring in which the non-woven fabric is inserted;
I have a,
The main load applied between the first composite part obtained as the first laminate after curing and the second composite part obtained as the second laminate after curing is shear load The first non-woven fabric to increase the shear strength is inserted in the part where the shear strength is increased, and the part in which the main load applied between the first composite part and the second composite part is the tensile load is inserted. , A method of manufacturing an aircraft structure in which a second non-woven fabric for increasing tensile strength is inserted .

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