JP2013107480A - Strength evaluating method of taper part connecting structure of composite material structure, taper part connecting structure of composite material structure, composite material structure, and aircraft fuselage - Google Patents

Abstract

PROBLEM TO BE SOLVED: To appropriately evaluate a strength of a taper part connecting structure applied to a composite material structure.SOLUTION: A method for evaluating the strength of the taper part connecting structure 10, which is applied to the composite material structural 1 configured of a foam core sandwich panel 2 having a taper part 6 and connects a pair of taper parts 6 facing each other, includes a break specifying step for specifying an initial break and its occurrence position CIP in the taper part connecting structure 10, a parameter specifying step for selecting an evaluation parameter for evaluating the strength of the taper part connecting structure 10 in response to the specified initial break, a condition setting step for setting the design condition and load condition of the taper part connecting structure 10, a parameter calculating step for calculating the selected evaluation parameter to the specified generation position CIP of the initial break under the set design condition and load condition, and a comparison step for comparing the calculated evaluation parameter with a predetermined evaluation threshold value.

Description

  The present invention relates to a method for evaluating the strength of a tapered connecting structure applied to a composite structure composed of a foam core sandwich panel. The present invention also relates to a composite structure composed of a foamed core sandwich panel, a tapered connecting structure applied to the composite structure, and an aircraft fuselage to which the composite structure is applied.

  In recent years, the application of composite materials to aircraft fuselage has progressed. In particular, attention has been paid to the characteristics of being capable of integral molding and excellent formability to curved members, and the application of carbon fiber reinforced plastic laminates to aircraft fuselage has become a reality. As a structural mode of a composite material applied to an aircraft fuselage, it is a mainstream to use a conventional structure, that is, a semi-monocoque structure that has been used for many years in an aluminum alloy body. However, since the semi-monocoque structure stretches the frame and the stringer, the weight and the number of parts increase.

  On the other hand, Non-Patent Document 1 proposes to adopt an integrally molded structure using a foamed core sandwich panel as a structural mode of a composite material applied to an aircraft fuselage. The foam core sandwich panel includes a plate-like core made of a foam material, and a pair of face plates that are bonded to both surfaces of the core and sandwich the core. According to Non-Patent Document 1, since the face plate is made of carbon fiber reinforced plastic, it is possible to provide a material that takes advantage of the aforementioned characteristics. In addition, the weight and the number of parts can be reduced as compared with a semi-monocoque structure for sheet metal assembly.

  Since the foam core sandwich panel is provided with a face plate made of carbon fiber reinforced plastic, autoclave molding is required in the manufacturing process. Therefore, when applying foam core sandwich panels to large structures such as aircraft fuselage, considering the dimensional constraints of autoclave equipment, several subdivided panels are manufactured individually and these panels are joined together. It is realistic to obtain a structure using the method of doing.

  Therefore, in Non-Patent Document 1, a so-called tapered end closure type is adopted as a bonding site in the foam core sandwich panel, and a so-called butt splice type is used for bonding between the bonding sites. It has been proposed to adopt. The tapered end closure system is a structural mode in which the core is inclined near the bonding site, and two face plates are stacked to form a laminated plate at the bonding site. The butt splice method is a joint mode in which such binding sites are opposed (or butted together), and the splice plate is bridged over the binding site and bonded to the binding site. In order to bond the splice plate to the binding site, Fasteners such as bolts and nuts are used.

Y. Hirose et al., “Proposal of suppression of delamination for the foam core sandwich panel joint with filler”, Advanced Composite Materials, Vol. 15, No. 3, pp 319-339 (2006).

  In designing a composite structure having a coupling structure employing a tapered end closure system and a butt splice system, delaying the breakage in the coupling structure as much as possible and reducing the overall weight of the composite structure as much as possible. It is important to achieve both. This may be possible if the strength in the bonded structure can be evaluated qualitatively, preferably quantitatively, but at present, methods for evaluating the strength in the bonded structure are well established. Not.

  It goes without saying that the strength of the bond structure depends on the design parameters of the bond structure and the foam core sandwich panel, but there is no definitive knowledge as to which design parameter affects the strength of the bond structure. In other words, the present situation is that the design parameters that are important in appropriately evaluating the bonding strength are not clearly specified. For example, the taper angle of the taper portion is empirically set to the maximum value (for example, about 45 degrees) of the angle range in which pressure can be appropriately applied by autoclave molding, and is set to a half constant, practically. It is not designed in light of the loading conditions in the state in which it is served.

  In addition, each foam core sandwich panel may have a portion where a pair of tapered portions face each other and the tapered portions are continuous with each other. Even in such a case, since the plate thickness changes peculiarly at the portions where the tapered portions are continuous with each other, it is preferable that the strength can be evaluated independently of other portions.

  Therefore, the present invention makes it possible to properly evaluate the strength of the taper portion connection structure applied to the composite material structure when providing a composite material structure composed of a foamed core sandwich panel having a taper portion. It is an object to make it possible to optimally design a taper connection structure, a foam core sandwich panel, a composite structure, and an aircraft fuselage.

  The “taper connection structure” referred to here is a structure that connects a pair of taper portions facing each other. The pair of taper portions may be provided in two separate foam core sandwich panels, or may be provided in a single foam core sandwich panel. In the former case, the tapered portion connecting structure functions as a joint for coupling a pair of tapered portions. In the latter case, the tapered part connecting structure functions as a part included in a single foam core sandwich panel, and a pair of tapered parts are continued in the foam core sandwich panel.

  In order to achieve the above object, the strength evaluation method for a taper portion connection structure of a composite structure according to the present invention is applied to a composite material structure composed of a foamed core sandwich panel having a taper portion, and a pair of facing each other. A method for evaluating the strength of the taper portion connecting structure for connecting the taper portion, according to the initial failure in the taper portion connecting structure and the failure specifying step for specifying the occurrence site, and the specified initial failure, A parameter specifying step for selecting an evaluation parameter for evaluating the strength of the tapered portion connecting structure, a condition setting step for setting design conditions and load conditions of the tapered portion connecting structure, and the set design conditions and load conditions A parameter calculating step for calculating the selected evaluation parameter for the identified initial fracture occurrence site, And having, a comparison step of the evaluation parameter with a predetermined evaluation threshold.

  The generation site of the initial fracture may be a laminated face plate positioned between the pair of taper portions, and the initial fracture may be a crack in the laminated face plate.

  The evaluation parameter may be an energy release rate at the initial fracture occurrence site, and the evaluation threshold may be a fracture toughness value of a material constituting the initial fracture occurrence site.

  The design condition may include a taper angle of the taper portion and a plate thickness of the foam core sandwich panel.

  The pair of taper portions oppose each other in the longitudinal direction of the composite material structure, the taper portion connecting structure connects the pair of taper portions in the longitudinal direction, and the design condition is adjacent to the longitudinal direction. A length between the tapered portion connecting structures or between the tapered portion connecting structure and the end portion in the longitudinal direction of the composite structure may be included.

  The composite material structure is cylindrical, the pair of taper portions are opposed to each other in the circumferential direction of the composite material structure, and the taper portion connection structure connects the pair of taper portions in the circumferential direction. And the design condition may include a radius of the composite structure.

  The composite material structure may be applied to an aircraft fuselage, and the load condition may include a pressurized load due to an external pressure difference between the aircraft fuselage.

  The pair of taper portions are opposed to the longitudinal direction of the aircraft fuselage, the taper portion connecting structure connects the pair of taper portions in the longitudinal direction, and the load condition is the longitudinal direction based on the pressurized load. A tensile load and a pressurized load applied to the inner surface of the aircraft fuselage may be included.

  The pair of taper portions oppose each other in the circumferential direction of the aircraft fuselage, the taper portion connection structure connects the pair of taper portions in the circumferential direction, and the load condition is based on the pressurized load. A circumferential tensile load may be included.

  The method may further include a condition resetting step for changing the design condition based on a comparison result in the comparison step, and the parameter calculation step and the comparison step may be retried after the condition resetting step.

  In the comparison step, when a comparison result that the evaluation parameter calculated immediately before is smaller than the first threshold or smaller than the first ratio or smaller than the first ratio is obtained, in the condition resetting step The design condition may be changed by increasing the taper angle of the tapered portion or decreasing the thickness of the foam core sandwich panel.

  The taper portion connection structure of the composite material structure according to the present invention is applied to a composite material structure formed of a foam core sandwich panel having a taper portion, and is a taper portion connection structure that connects a pair of taper portions facing each other. The pair of taper portions are opposed to each other in the longitudinal direction of the composite material structure, and when the taper angle of the taper portion is θL, 10 degrees ≦ θL ≦ 15 degrees.

  When the length between the taper portion connecting structures adjacent to each other in the longitudinal direction or between the taper portion connecting structure and the end portion in the longitudinal direction of the composite material structure is L, 1000 mm ≦ L ≦ 1300 mm It may be.

  The taper part connection structure of the composite material structure according to the present invention is applied to a cylindrical composite material structure formed of a foamed core sandwich panel having a taper part, and connects a pair of taper parts facing each other. In the connecting structure, the pair of tapered portions are opposed to each other in the circumferential direction of the composite structure, and when the taper angle of the tapered portion is θC, 10 degrees ≦ θC <30 degrees.

  A composite material structure according to the present invention is a cylindrical composite material structure composed of a foam core sandwich panel having a taper portion, and a longitudinal connection structure that connects a pair of taper portions opposed to each other in the longitudinal direction. A circumferential connection structure that connects a pair of taper portions facing each other in the circumferential direction, and a taper angle of the taper portion in the longitudinal connection structure is a taper of the taper portion in the circumferential connection structure Near the angle.

  At least one of the longitudinal connection structure and the circumferential connection structure forms a joint of two separate foam core sandwich panels, and the joint is an outer edge of the two foam core sandwich panels. You may couple | bond together the taper part formed in each.

  The joint may include a splice plate that spans the pair of tapered portions, and a fastener that fixes the splice plate to each of the pair of tapered portions.

  The aircraft fuselage according to the present invention is constituted by the composite material structure.

  According to the present invention, when providing a composite material structure including a foam core sandwich panel having a taper portion, it is possible to appropriately evaluate the strength of the taper portion connection structure applied to the composite material structure. it can. Thereby, the design of the taper connection structure, the foam core sandwich panel, the composite structure and the aircraft fuselage can be optimized.

It is the perspective view which showed the example of arrangement | positioning of the connection structure which concerns on the case where the cylindrical composite material structure comprised by a foam core sandwich panel is applied to an aircraft fuselage as an example. It is sectional drawing of the connection structure cut | disconnected and shown along the II-II line | wire of FIG. It is a flowchart which shows the procedure of the strength evaluation method of the taper part connection structure of the composite material structure which concerns on this invention. It is a figure which shows the FEM model of the longitudinal direction connection structure shown in FIG. It is a conceptual diagram which shows the calculation method of an energy release rate. It is a conceptual diagram which shows the load conditions applied in the FEM model shown in FIG. It is a graph which shows the correlation of the taper angle of a longitudinal direction connection structure, and an energy release rate. It is a graph which shows the correlation with the taper angle and energy release rate of a longitudinal direction connection structure, the correlation with a panel length and an energy release rate, and the correlation with the taper angle of a circumferential direction connection structure and an energy release rate. It is a conceptual diagram which shows an example of contrast with an evaluation parameter and an evaluation threshold value.

  Hereinafter, embodiments of the present invention will be described with reference to the drawings. Throughout the drawings, the same or corresponding elements are denoted by the same reference numerals, and detailed description thereof is omitted.

[Aircraft fuselage, composite structure, foam core sandwich panel, taper connection structure]
FIG. 1 is a perspective view of an aircraft fuselage shown as an example of a composite structure according to an embodiment of the present invention. An aircraft fuselage 1 shown in FIG. 1 is a composite material structure formed by sequentially joining a plurality of segments 2 made of a foam core sandwich panel, and is formed in a substantially cylindrical shape as a whole. The number of segments 2 in the circumferential direction and the number of segments 2 in the longitudinal direction are not particularly limited. In the present embodiment, as an example, the four segments 2 are sequentially coupled in the circumferential direction, and the two segments 2 are sequentially coupled in the longitudinal direction. The structure of the aircraft fuselage 1 shown in FIG. 1 is a monolithic structure of sandwich panels and does not include a frame and stringers. Since the segment 2 is composed of a single foam core sandwich panel, in the following description, the foam core sandwich panel is described with reference numeral 2 as in the case of the segment.

2 is a cross-sectional view taken along the line II-II in FIG. As shown in FIG. 2, the foam core sandwich panel 2 has a plate-shaped core 3 and a pair of face plates 5 and 4 provided on the front surface 3b and the back surface 3a of the core 3, respectively. Sandwiched between a pair of face plates 4 and 5. The core 3 is manufactured from a foam material such as polyetherimide. The pair of face plates 4 and 5 are made of carbon fiber reinforced plastic (hereinafter referred to as CFRP) and are configured by laminating prepregs using carbon fiber twill fabric as a reinforcing material. In this case, the face plates 4 and 5 can be CFRP laminated plates formed by laminating a total of 16 layers of prepregs with [{(+45, −45) / (0.90)} ₄ ] _sym . In addition, among the total of 16 prepregs, the resin volume content of the (0, 90) layer and (+45, -45) layer adjacent to the core is larger than the content of the other 14 layers. Alternatively, the layer thickness may be thicker than the other 14 layers. A resin film (not shown) is interposed at the interface between the core 3 and the face plate 4, thereby improving the bonding reliability of the pair of face plates 4 and 5 to the core 3.

  A so-called tapered end closure system is applied to the foam core sandwich panel 2 according to the present embodiment. That is, the tapered portion 6 is provided at the edge portion of the core 3. In the tapered portion 6, the surface 3 b is smoothly continuous with a portion near the center of the core 3. Therefore, the tapered portion 6 has a tapered surface 7 that is inclined from the back surface 3a toward the front surface 3b and connects the back surface 3a to the front surface 3b. Hereinafter, an angle formed by the surface 3b and the tapered surface 7 is defined as a “taper angle θ”. Depending on the portion of the taper connecting structure 10, the suffix C or L may be added to θ.

  The face plate 4 is continuously provided from the back surface 3 a to the tapered surface 7. The pair of face plates 4 and 5 are overlapped at a pointed end portion 8 where the surface 3b and the tapered surface 7 intersect. On the inner side of the pointed end portion 8, a laminated face plate portion 9 configured by laminating a pair of face plates 4 and 5 is provided. In this way, the tapered portion 6 is surrounded by the laminated face plate portion 9 and is closed by the laminated face plate portion 9.

  In the manufacture of the foam core sandwich panel 2 having such a structure, a prepreg is laminated after a resin film is attached to the core 3. The assembly produced in this way is autoclaved. Thereby, a structure in which the core 3 is sandwiched between the pair of face plates 4 and 5 made of carbon fiber reinforced plastic can be obtained. When the surface 3b is set on a jig, the taper angle θ is preferably set to 45 degrees or less. If it sets in this way, the part provided on the taper surface 7 among the prepregs which comprise the faceplate 4 can be pressurized favorably, and intensity | strength can be ensured.

  The aircraft fuselage 1 is provided with a tapered portion connection structure 10 for connecting the segments 2 adjacent to each other in the circumferential direction or the longitudinal direction. The taper part connection structure 10 is a structure for connecting a pair of taper parts 6 facing each other, and a so-called butt splice system is applied to the taper part connection structure 10 according to the present embodiment.

  Two adjacent segments 2 are arranged in such a manner that the taper portion 6 and the laminated face plate portion 9 surrounding the taper portion 6 face each other. The mutual surface 3b is directed to the same side, and the laminated face plate portions 9 are close to each other. Then, the inverted trapezoidal concave portion 11 is formed by the face plate 4 and the laminated face plate portion 9 of one segment 2 and the face plate 4 and the laminated face plate portion 9 of the other segment 2. A flat splice plate 12 is accommodated in the recess 11. The splice plate 12 is bridged between one laminated face plate portion 9 and the other laminated face plate portion 9 and is coupled to each laminated face plate portion 9 using fasteners 13 such as bolts and nuts. The splice plate 12 can be made of an aluminum alloy, and the fastener 13 can be made of a titanium alloy.

  As shown in FIG. 1 using parenthesis reference numerals, the structure 10C for connecting the segments 2 adjacent to each other in the circumferential direction and the structure 10L for connecting the segments 2 adjacent to each other in the longitudinal direction are also used. The tapered portion connecting structure 10 having the structure shown in FIG. 2 is applied. In the following description, the structure 10C is referred to as a “circumferential connection structure 10C”, and the structure 10L is referred to as a “longitudinal connection structure 10L”.

  Although the circumferential connection structure 10C and the longitudinal connection structure 10L have the same structural style, when the aircraft fuselage 11 is put into practical use, the mode of loads acting on the connection structures 10C and 10L is different. Since the load conditions are different, if the design conditions are not individually optimized, inconveniences such as the progress of one of the fatigues becoming extremely fast or slow will occur.

  As will be described in detail later, when a cylindrical composite material structure having the tapered portion connecting structure 10 is applied to the aircraft fuselage 1, the design conditions of the segment 2 and the tapered portion connecting structure 10 include the fuselage radius, the face plates 4, 5 and The thickness of the core 3, the longitudinal dimension of the segment 2 (hereinafter referred to as “panel length L”), and the taper angle θ are included. In the past, at least the taper angle θ is only set to be a constant value in light of only the restrictions at the time of autoclave molding, and is not set in light of the load conditions in a state of practical use. .

  In the taper part connection structure 10 according to the present embodiment, the panel length L and the taper angle θ are set to values obtained by performing the strength evaluation method described later, and thus are necessary for the taper part connection structure 10. It is possible to ensure the strength.

  As a specific example, the aircraft fuselage 1 is for a small passenger aircraft, the radius is about 2.5 to 3.0 m (for example, 2.87 m), and the thickness of the core 3 is about 30 to 35 mm (for example, 34 mm). When the thickness of the face plates 4 and 5 is about 6 to 7 mm (for example, 6.24 mm), the panel length L is set to 1000 to 1300 mm in consideration of the strength required for the longitudinal connection structure 10L. The taper angle θL of the longitudinal connection structure 10L is set to 10 to 30 degrees, preferably around 15 degrees, in consideration of the strength required for the longitudinal connection structure 10L. The taper angle θC of the circumferential connection structure 10C is set to 10 to 30 degrees, preferably around 15 degrees, in consideration of the strength required for the circumferential connection structure 10C. If the taper angle θL of the longitudinal connection structure 10L is set to an angle in the vicinity of the taper angle θC of the circumferential connection structure 10C, it is beneficial to simplify the manufacture of the segment 2. If both are set to around 15 degrees, the necessary strength can be obtained, which is beneficial.

  Hereinafter, after describing the strength evaluation method of the taper portion connecting structure 10, the validity of setting the design conditions as described above will be described.

[Strength evaluation method]
FIG. 3 is a flowchart showing the procedure of the strength evaluation method according to the embodiment of the present invention. As shown in FIG. 3, first, the initial breakage and the site where the taper connection structure 10 is generated are specified (step S <b> 1). Next, an evaluation parameter for evaluating the strength of the tapered portion connecting structure 10 is selected according to the identified initial fracture (step S2). Next, the design condition and load condition of the taper part connection structure 10 are set (step S3). Next, under the set design conditions and load conditions, the selected evaluation parameters are calculated for the identified initial fracture occurrence site (step S4). Next, the calculated evaluation parameter is compared with a predetermined evaluation threshold value (step S5), thereby evaluating the strength of the tapered portion connecting structure 10.

  It is also possible to optimize the design using the strength evaluation method. That is, as a result of comparing the evaluation parameter with the evaluation threshold value, when the evaluation that the strength is insufficient or the evaluation that the strength is excessive is made (step S6: NO), the design condition of the tapered portion connecting structure 10 is Changed (step S3 retry). An example of the case where the evaluation that the strength is insufficient is made when the evaluation parameter calculated immediately before is equal to the evaluation threshold value. As an example of the case where the evaluation that the intensity is excessive is performed, when the evaluation parameter calculated immediately before is smaller than the evaluation threshold and exceeds the first deviation, the evaluation parameter calculated immediately before is compared with the evaluation threshold. And the case where the ratio is smaller than the first ratio. As described above, the evaluation parameter may be compared with the evaluation threshold based on addition / subtraction, or may be compared with the evaluation threshold based on multiplication / division. The first deviation is a concept including zero, and the first ratio is a concept including 1. In other words, when the evaluation parameter is simply smaller than the evaluation threshold, the design condition may be changed because the strength is insufficient.

  After the retry of step S3, the evaluation parameter is calculated in the same manner as the previous time under the design condition changed this time and the previously set load condition (retry of step S4). Next, the evaluation parameter calculated this time is re-compared with the evaluation threshold value, thereby re-evaluating the strength of the tapered portion connecting structure 10 (step S5 retry). If it is evaluated that the strength is sufficient (step S6: YES), the design condition is determined.

  As described above, according to the strength evaluation method according to the present embodiment, the initial fracture and the generation site thereof are specified in the tapered portion connecting structure 10, the evaluation parameter is selected according to the initial fracture, and the set design condition and Evaluation parameters under load conditions are compared with an evaluation threshold. By following such a procedure, the strength can be properly evaluated at a site that is identified or assumed to have an initial failure, and the design with the strength required for the entire structure can be appropriately performed. it can.

  Next, under the assumption that a cylindrical composite structure having the taper portion connecting structure 10 is applied to the aircraft fuselage 1, the above method will be specifically described along the procedure.

[Initial destruction and identification of the site of its occurrence]
In order to specify the failure mode of the initial fracture of the taper portion connecting structure 10 and the generation site thereof, it is preferable to perform a strength test using a model. It should be noted that the model may of course be the actual size, may be a size obtained by reducing the actual size three-dimensionally, or may be cut in the direction perpendicular to the plane of FIG. As indicated by reference numeral “CIP” (that is, crack generation position) in FIG. 2, as a result of the inventor conducting a strength test using a model, the initial fracture occurrence site of the taper connection structure 10 is a laminated face plate. The inside of the part 9 is a part about 5 mm away from the pointed part 8 of the core 3, and the failure mode was found to be a crack (delamination) of the part. In this strength test, a tensile load is applied to the tapered portion connecting structure 10 in the direction in which the pair of tapered portions 6 face each other until breakage occurs, and the taper angle θ is set to 30 degrees.

  In order to obtain confirmation of this knowledge, it is preferable to perform numerical analysis incidentally using an FEM model or the like. FIG. 4 illustrates an FEM model for the longitudinal connection structure 10L. The model illustrated in FIG. 4 is a two-dimensional model in a cross section cut along the line II-II illustrated in FIG. 1, and targets a portion having no shape change in the circumferential direction. A four-node plane strain element is used for the face plate, core, and resin film.

  Prior to performing the FEM analysis, the specifications (that is, design conditions) of the taper portion connection structure 10 and the load conditions applied to the taper portion connection structure 10 such as the dimensions and materials of the face plate, the core, the resin film, etc. Must be set. As will be described later, the FEM model can be used not only in the step of specifying initial fracture, but also in the step of setting design conditions and load conditions and the step of calculating evaluation parameters.

In the step of identifying the initial fracture and the site where the fracture occurs, the specifications and load conditions of the face plate, core, and resin film in the FEM model are set to be the same as the specifications of the model used in the strength test. For example, the thickness of the core is 34 mm, the thickness of each face plate is 6.24 mm, the thickness of the laminated face plate portion is twice that, the thickness of the resin film is 0.254 mm, and the taper angle θ is 30 degrees. The face plate is made of a total of 16 layers of prepreg UT500 / # 135 (reinforcement volume content V _f = 56%) with [{(+45, -45) / (0. 90)} ₄ ] _sym . A CFRP laminate is used, and the thickness of one prepreg layer is 0.38 mm. However, in each of the pair of face plates, the reinforcing material volume content is 46% and the layer thickness is 0.46 mm for the (0, 90) layer and (+45, -45) layer adjacent to the core. . Table 1 shows the material constants of the core 3, the resin film, and the four patterns of prepreg. In Table 1, E is the longitudinal elastic modulus, G is the transverse elastic modulus, and ν is the Poisson's ratio.

  In this step, in light of the fact that the failure mode of the initial failure is a crack (delamination) of the CFRP laminate, the site where the initial failure occurred by the strength test using the FEM model (that is, the laminated face plate part) The energy release rate is calculated at a part 5 mm away from the apex. The energy release rate is a change in the total energy that occurs as the crack progresses. When the energy release rate becomes equal to the fracture toughness value, cracks (delamination) develop. Therefore, if the calculated energy release rate is the same value as the fracture toughness value of the material of the laminated faceplate part, the strength that the fracture mode of the initial fracture is a crack and the occurrence site is inside the laminated faceplate part Confirmation will be obtained in the test results.

As a result of analysis by the present inventors, the energy release rate at the site where the initial fracture is assumed to occur by the strength test may be substantially the same value (1.53 kJ / m ² ) as the fracture toughness value of the material of the laminated faceplate part. understood. Therefore, the initial fracture in the taper part connection structure 10 and the generation site thereof were identified with certainty.

FIG. 5 is a conceptual diagram showing a method for calculating the energy release rate. Although the calculation method of an energy release rate is not specifically limited, The crack closure method (Crack closure method) can be employ | adopted suitably. The crack closure method is a calculation method based on the idea that the energy release rate can be obtained by tracing the crack growth process in reverse since the energy release rate is a change in the total energy that accompanies the crack growth. is there. Energy release rate G is the energy release rate G _I with respect to the opening displacement (mode I), optionally in defined as the sum of the energy release rate G _II to shear displacement (mode II). That is, the x-direction component and the y-direction component of the nodal force at the node j near the crack tip are F _xj and F _yj, and the x-direction component of the relative displacement between the node (j-1) and the node (j-1) ′ and When the y-direction component is U _{x (j-1)} -U _{x (j-1) '} , U _{y (j-1)} -U _{y (j-1)'} , the energy release rate G _I for the aperture displacement, The energy release rate G _{II with} respect to the shear displacement and the energy release rate G defined as the sum of these are calculated from the following equations (1) to (3).

[Selection of evaluation parameters and evaluation thresholds]
In consideration of the process of specifying the initial fracture, in this embodiment, the energy release rate at the laminated face plate portion where the crack (delamination) that is the initial fracture occurs is selected as the evaluation parameter, and the strength is evaluated. If the fracture toughness value of the material of the laminated face plate portion is adopted as the evaluation threshold value compared with the evaluation parameter, it can be said that it is appropriate for evaluating the strength of the tapered portion connecting structure 10. When the failure mode of the initial failure is not a crack (delamination) of the CFRP laminate, the evaluation parameter and the evaluation threshold may be selected differently according to this.

[Load conditions and design conditions]
FIG. 6 is a schematic diagram showing load conditions. During the flight of the aircraft, the cabin is pressurized. Therefore, in the present embodiment, a pressurized load (a load based on the pressure difference outside the body) is considered as a main load added to the tapered portion connecting structure 10. It can be considered that a circumferential tensile load (Hoop load) is applied to the circumferential connecting structure 10C based on the pressurized load. It can be considered that a longitudinal tensile load and a pressurized load from the inside of the body are added to the longitudinal connecting structure 10L based on the pressurized load. The longitudinal tensile load can be obtained by multiplying the circumferential tensile load by 1/2.

  In the present embodiment, the aircraft fuselage 1 is for a small passenger aircraft and has a radius of 2.87 m. The pressure inside the fuselage is 0.78 atm (atmospheric altitude 2000 m), and the pressure outside the fuselage is 0.26 atm (atmospheric altitude 10000 m). Based on the fuselage radius and the fuselage external pressure set in this way, the circumferential tensile load acting on the circumferential connecting structure 10C and the longitudinal tensile load acting on the longitudinal connecting structure 10L are calculated. The

  The design conditions include the fuselage radius, panel length, core plate thickness, face plate thickness, and taper angle. The fuselage radius is almost constant if the general specifications of the aircraft to be applied are determined. Will be handled. The panel length varies greatly depending on the number of segments arranged in the longitudinal direction even if the general specifications of the aircraft are determined. The plate thickness of the core, the plate thickness of the face plate, and the taper angle are design conditions that can be set and changed relatively freely within the design category of the aircraft fuselage 1 regardless of the general specifications of the aircraft.

  Needless to refer to the drawings, the plate thickness of the core and the plate thickness of the face plate are factors that affect the strength of the tapered portion connecting structure (that is, the energy release rate inside the laminated face plate portion). If the plate thickness is increased, the energy release rate of the laminated face plate portion is reduced, and the strength of the tapered portion connecting structure 10 is improved. Therefore, the characteristics of the panel length and the taper angle, that is, the influence of the panel length and the taper angle on the strength of the tapered portion connecting structure will be described with reference to the drawings.

  FIG. 7 is a graph showing the correlation between the taper angle and the energy release rate inside the laminated face plate portion of the longitudinal connection structure. The horizontal axis in FIG. 7 represents the taper angle, and the vertical axis represents the energy release rate. FIG. 7 shows a line representing the case where only the longitudinal tensile load is considered as the load condition and a line representing the case where both the longitudinal tensile load and the pressurized load are considered as the load condition. . Comparing the two lines, it can be seen that the energy release rate considering the pressurized load is larger than the energy release rate not considering the pressurized load. This is evidence that the pressurized load has an influence on the strength of the longitudinal connection structure, and it is highly appropriate to consider the pressurized load as a load condition.

  Referring to the line representing the case where both the tensile load and the pressurized load in the longitudinal direction are considered, the energy release rate changes as the taper angle changes. This is evidence that the taper angle is a factor affecting the strength of the longitudinal connection structure. Specifically, the energy release rate decreases as the taper angle decreases from 45 degrees. The energy release rate takes the minimum value when the taper angle takes a value near 15 degrees. As the taper angle decreases from 15 degrees, the energy release rate increases.

  FIG. 8 is a graph showing the influence of the panel length and taper angle on the energy release rate. As shown in FIG. 8, as the panel length increases, the energy release rate inside the laminated face plate portion of the longitudinal connection structure increases. This is evidence that the panel length is a factor affecting the strength of the longitudinal connection structure.

  FIG. 8 also shows the correlation between the taper angle of the circumferential connection structure and the energy release rate inside the laminated face plate portion of the circumferential connection structure. As shown in FIG. 8, also in the circumferential connection structure, the energy release rate changes with the change of the taper angle. This is evidence that the taper angle is a factor affecting the strength of the circumferential connecting structure. In the circumferential connection structure, unlike the longitudinal connection structure, the energy release rate monotonously decreases as the taper angle decreases. Specifically, while the taper angle decreases from 45 degrees to 30 degrees, the energy release rate gradually decreases. However, as the taper angle θ decreases from 30 degrees, the energy release rate decreases until then. Compared to, it becomes smaller rapidly.

  As described above, the taper angle is a factor that affects the strength of the tapered portion coupling structure in both the longitudinal direction coupling structure and the circumferential direction coupling structure. It can be seen that in the longitudinal connection structure, the energy release rate takes the minimum value when the taper angle is around 15 degrees. On the other hand, in the circumferential connection structure, the energy release rate monotonously decreases as the taper angle decreases, but the tendency to decrease increases when the taper angle is less than 30 degrees. In the longitudinal connection structure, the panel length is a factor affecting the strength of the connection structure, and the energy release rate decreases as the panel length decreases.

  In the past, since such knowledge did not exist, the taper angle and the panel length were treated like constants. On the other hand, the present inventors based on this knowledge, the taper angle and the panel length are the design conditions that should be handled like variables when trying to optimize the design of the taper connection structure from the viewpoint of strength. The idea is that the design can be optimized by changing the design conditions. Based on the knowledge and idea, design conditions are set together with load conditions.

[Calculation of evaluation parameters]
When the design condition and the load condition are set, the energy release rate inside the laminated face plate portion is calculated using the FEM model. As described above, the crack closure method can be suitably used as the energy release rate calculation method.

[Comparison between evaluation parameter and evaluation threshold]
The calculated energy release rate is compared with the fracture toughness value of the material of the laminated faceplate portion. FIG. 9 is a graph for explaining the procedure of strength evaluation and design condition change. The horizontal axis represents design conditions such as the taper angle, and the vertical axis represents the energy release rate. Since the fracture toughness value is constant regardless of the design conditions of the tapered portion connecting structure, the line representing the fracture toughness value extends in parallel with the horizontal axis in the coordinate system of FIG. In this coordinate system, the calculated energy release rate is plotted in correspondence with the set design conditions. The degree of deviation in the vertical axis direction from the line representing the fracture toughness value to the plot reflects the strength evaluation.

  In other words, if the plot is on the line representing the fracture toughness value, it will cause the crack to develop under the set design conditions and load conditions. When the plot is below the line representing the fracture toughness value in the vertical axis direction, crack growth can be suppressed under the set design conditions and load conditions, so it should be evaluated that there is a margin in strength. Is the situation. However, when the plot is significantly below the fracture toughness value (for example, when the energy release rate indicated by the plot is smaller than the fracture toughness value by a certain deviation ΔG), for the set load condition The strength may be excessive. In response to this, it is possible to easily recognize that there is still room for weight reduction by, for example, simplifying the structure even if the strength is slightly reduced.

  According to the strength evaluation method according to the present embodiment, as described above, the strength can be appropriately evaluated in a site that is specified or assumed to cause initial fracture, and the strength required for the entire structure is provided. Design can be done properly.

  Then, the occurrence site of the initial failure in the taper portion connecting structure 10 is inside the laminated face plate portion, and the initial failure is specified as a crack (delamination) of the laminated face plate portion made of the CFRP laminated plate. The evaluation parameter and the evaluation threshold can be appropriately selected based on the specific result. For this reason, strength can be evaluated appropriately. In particular, when initial fracture is specified as described above, the evaluation parameter is the energy release rate at the initial fracture occurrence site, and the evaluation threshold is the fracture toughness value of the material constituting the initial fracture occurrence site. The strength can be properly evaluated.

[Change design conditions]
As described above, when the energy release rate is significantly lower than the fracture toughness value (for example, when the energy release rate is smaller than the fracture toughness value by a certain deviation ΔG), the strength is slightly reduced. However, since the progress of cracks can be suppressed, the design conditions can be changed within such a range. In changing the design conditions in the direction of decreasing the strength, the plate thickness can be reduced, the panel length can be increased, and the taper angle can be increased. If the plate thickness is reduced, the weight can be reduced. As the panel length increases, the number of segments in the longitudinal direction can be reduced. When the taper angle is increased, the size of the taper portion in the longitudinal direction is reduced, and the taper portion can be manufactured relatively easily. Of course, it is possible to reduce the panel length or taper angle while reducing the plate thickness.

  Conversely, when the energy release rate is equal to the fracture toughness value, cracks may develop unless the strength is improved. Therefore, the panel length can be shortened or the taper angle can be reduced in order to improve the strength. In order to improve strength, there is no need to immediately depend on the option of increasing the plate thickness or adding a strength member like the semi-monocoque structure of sheet metal assembly, and improving the strength while avoiding the increase in weight as much as possible be able to.

  As a result of following the above procedure, as described above, the panel length L is set to 1000 to 1300 mm, and the taper angle θL of the longitudinal connecting structure 10L is set to 10 to 30 degrees, preferably around 15 degrees. When the taper angle θC of the circumferential connection structure 10C is set to 10 to 30 degrees, preferably around 15 degrees that is near the taper angle θL of the longitudinal connection structure, the necessary strength is ensured, and the taper portion It is possible to achieve both delaying the breakage in the connecting structure as much as possible and reducing the weight of the entire aircraft fuselage 1 as much as possible.

[Modification]
From the foregoing description, many modifications and other embodiments of the present invention are obvious to one skilled in the art. Accordingly, the foregoing description should be construed as illustrative only and is provided for the purpose of teaching those skilled in the art the best mode of carrying out the invention. The details of the structure and / or function may be substantially changed without departing from the spirit of the invention.

  For example, the pressurized load is considered as the load condition, but other aerodynamic loads may be considered. In addition, after setting one design condition, set two or more load conditions assumed for practical use, and calculate the evaluation parameters under each load condition under the design condition of the one pattern. All calculated evaluation parameters may be compared with an evaluation threshold value.

  Moreover, in the said embodiment, although the composite structure comprised by the foam core sandwich panel which has a taper part was cylindrical shape, other shapes, such as flat form and a rectangular parallelepiped shape, may be sufficient. In addition, the composite structure is applied to an aircraft fuselage for a small passenger aircraft. However, the composite structure may be applied to an aircraft fuselage for other purposes, and may be suitably applied to the top structure of a high-speed railway vehicle. it can.

  Moreover, in the said embodiment, a taper part is provided only in the edge part of the foam core sandwich panel, and the taper part connection structure functions as a joint which couple | bonds two foam core sandwich panels. However, a pair of taper portions facing each other may be provided in a single foam core sandwich panel. In this case, a taper part connection structure functions as a part contained in a single foam core sandwich panel, and makes a pair of taper parts continue in the foam core sandwich panel. With regard to such a tapered portion connection structure, the design can be optimized by performing the strength evaluation method according to the embodiment.

  When the present invention provides a composite structure comprising a foam core sandwich panel having a taper portion, the strength of the taper portion connection structure applied to the composite structure can be appropriately evaluated. The composite material composed of the foam core sandwich panel having the taper portion, which has a remarkable effect of being able to optimize the design of the taper portion connection structure, the foam core sandwich panel, the composite material structure, and the aircraft fuselage. It is beneficial to apply it to structures.

1 Aircraft fuselage (composite structure)
2 segments (foamed core sandwich panel)
3 Core 4, 5 Face plate 6 Tapered portion 7 Tapered surface 8 Pointed end portion 9 Laminated face plate portion 10 Tapered portion connecting structure 10C Circumferential direction connecting structure 10L Longitudinal direction connecting structure 11 Recess 12 Splice plate 13 Fastener

Claims (18)

A method for evaluating the strength of a taper connection structure applied to a composite structure composed of a foam core sandwich panel having a taper, and connecting a pair of tapers facing each other,
Destructive identification step for identifying the initial fracture in the tapered portion connecting structure and its occurrence site;
A parameter specifying step for selecting an evaluation parameter for evaluating the strength of the tapered portion connecting structure according to the specified initial fracture;
A condition setting step for setting a design condition and a load condition of the tapered portion connecting structure;
A parameter calculation step for calculating a selected evaluation parameter for the identified initial fracture occurrence site under the set design conditions and load conditions;
A comparison step of comparing the calculated evaluation parameter with a predetermined evaluation threshold value, and a strength evaluation method of the tapered portion connection structure of the composite structure.   2. The tapered portion connection of the composite structure according to claim 1, wherein the occurrence site of the initial failure is a laminated face plate positioned between the pair of tapered portions, and the initial failure is a crack in the laminated face plate. Structural strength evaluation method.   3. The composite structure according to claim 1, wherein the evaluation parameter is an energy release rate at a site where the initial fracture occurs, and the evaluation threshold is a fracture toughness value of a material forming the site where the initial fracture occurs. Strength evaluation method of body taper part connection structure.   The strength evaluation method for a tapered portion connecting structure of a composite material structure according to any one of claims 1 to 3, wherein the design condition includes a taper angle of the tapered portion and a plate thickness of the foamed core sandwich panel. The pair of taper portions are opposed to the longitudinal direction of the composite structure, and the taper portion connection structure connects the pair of taper portions in the longitudinal direction.
The design condition includes a length between the taper portion connecting structures adjacent to each other in the longitudinal direction, or a length between the taper portion connecting structure and an end portion in the longitudinal direction of the composite structure. The strength evaluation method of the taper part connection structure of the composite material structure according to any one of claims 1 to 4. The composite material structure is cylindrical, the pair of taper portions are opposed to each other in the circumferential direction of the composite material structure, and the taper portion connection structure connects the pair of taper portions in the circumferential direction. And
The strength evaluation method for a tapered portion connection structure of a composite material structure according to any one of claims 1 to 5, wherein the design condition includes a radius of the composite material structure.   The composite structure according to any one of claims 1 to 6, wherein the composite structure is applied to an aircraft fuselage, and the load condition includes a pressurized load due to an external pressure difference between the aircraft fuselage. Strength evaluation method of body taper part connection structure. The pair of taper portions are opposed to the longitudinal direction of the aircraft fuselage, the taper portion connection structure connects the pair of taper portions in the longitudinal direction,
The taper part connection structure of the composite material structure according to claim 7, wherein the load condition includes the longitudinal tensile load based on the pressurized load and a pressurized load applied to an inner surface of the aircraft fuselage. Strength evaluation method. The pair of taper portions oppose each other in a circumferential direction of the aircraft fuselage, and the taper portion connection structure connects the pair of taper portions in the circumferential direction;
The strength evaluation method for a taper portion connection structure of a composite material structure according to claim 7, wherein the load condition includes a tensile load in the circumferential direction based on the pressurized load. Further comprising a condition resetting step for changing the design condition based on the comparison result in the comparison step;
10. The strength evaluation method for a tapered portion connection structure of a composite structure according to claim 1, wherein the parameter calculation step and the comparison step are retried after the condition resetting step. 11.   In the comparison step, when a comparison result that the evaluation parameter calculated immediately before is smaller than the first threshold or smaller than the first ratio or smaller than the first ratio is obtained, in the condition resetting step The taper portion connection of the composite structure according to claim 10, wherein the design condition is changed by increasing a taper angle of the taper portion or decreasing a plate thickness of the foam core sandwich panel. Structural strength evaluation method. Applied to a composite material structure composed of a foam core sandwich panel having a taper portion, a taper portion connection structure for connecting a pair of taper portions facing each other,
The pair of taper portions are opposed to the longitudinal direction of the composite structure,
A taper portion connecting structure of a composite structure, wherein 10 ° ≦ θL ≦ 15 °, where the taper angle of the taper portion is θL.   When the length between the taper portion connecting structures adjacent to each other in the longitudinal direction or between the taper portion connecting structure and the end portion in the longitudinal direction of the composite material structure is L, 1000 mm ≦ L ≦ 1300 mm The taper part connection structure of the composite material structure according to claim 12. It is applied to a cylindrical composite structure composed of a foam core sandwich panel having a taper part, and is a taper part connection structure for connecting a pair of taper parts facing each other,
The pair of taper portions are opposed to the circumferential direction of the composite structure,
A taper portion connection structure of a composite structure, wherein 10 ° ≦ θC <30 °, where the taper angle of the taper portion is θC. A cylindrical composite structure composed of a foam core sandwich panel having a tapered portion,
A longitudinal connection structure that connects a pair of tapered portions opposed to each other in the longitudinal direction;
A circumferential connection structure that connects a pair of tapered portions facing in the circumferential direction, and
The composite material structure, wherein a taper angle of the taper portion in the longitudinal connection structure is in the vicinity of a taper angle of the taper portion in the circumferential connection structure. At least one of the longitudinal connection structure and the circumferential connection structure forms a joint of two separate foam core sandwich panels,
The composite structure according to claim 15, wherein the joint joins the tapered portions formed on the respective outer edges of the two foam core sandwich panels.   The composite material structure according to claim 16, wherein the joint includes a splice plate spanned between the pair of taper portions, and a fastener for fixing the splice plate to each of the pair of taper portions.   An aircraft fuselage comprising the composite structure according to any one of claims 15 to 17.

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