JP6880706B2 - Aircraft structure - Google Patents

Description

The present invention relates to an aircraft structure composed of a resin, reinforcing fibers and voids.

In recent years, the market demand for aircraft products has been directed toward weight reduction from the viewpoint of improving fuel efficiency, in addition to the conventional improvement of mechanical properties. Therefore, in order to meet the market demand, fiber reinforced plastics, which are lightweight and have excellent mechanical properties, are widely used in aircraft products. On the other hand, products used in the aircraft field are also required to have high combustion characteristics (difficulty in burning) in order to ensure the safety of passengers. That is, in the field of aircraft, materials and structures that are excellent in flammability and mechanical properties while satisfying light weight are indispensable.

Japanese Patent No. 3754313 Japanese Patent Application Laid-Open No. 2014-503964

However, the technology that satisfies all of the lightness, flammability, and mechanical properties is limited, and various efforts are still being made (see Patent Documents 1 and 2). In particular, among the above-mentioned characteristics, light weight is regarded as the most important from the viewpoint of fuel efficiency during flight, and combustibility is regarded as the most important from the viewpoint of passenger safety, and market demand is increasing. Although Patent Document 1 discloses a structure in which a flame-retardant resin is used as the core material, it is presumed that the structure is inferior in mechanical properties because the core material is not a fiber-reinforced resin. Further, Patent Document 2 discloses an aircraft material containing a reinforcing fiber, a polyimide fiber, and a polymer binder fiber. However, since the reinforcing fiber is a glass fiber, the aircraft material is inferior in compressive strength. Inferred. In addition, since several types of fibers are used in combination, product design becomes complicated and the economy is inferior.

The present invention has been made in view of the above problems, and an object of the present invention is to provide an aircraft structure having excellent lightness, flammability, and mechanical properties.

The aircraft structure according to the present invention is an aircraft structure composed of a resin, reinforcing fibers, and voids, and the volume content of the resin is within the range of 2.5% by volume or more and 85% by volume or less. The volume content of the reinforcing fibers is in the range of 0.5% by volume or more and 55% by volume or less, and the voids are in the range of 10% by volume or more and 99% by volume or less in the aircraft structure. When the length of the reinforcing fiber contained is Lf and the orientation angle of the reinforcing fiber in the cross-sectional direction of the aircraft structure is θf, the thickness St of the aircraft structure is a conditional expression: St ≧ Lf ^2. Satisfying (1-cos (θf)) and measuring the combustion time of the aircraft structure after ignition source digestion measured by Advisory Volume (AC) 25.853-1a (1) i is 30 seconds or less. It is a feature.

In the above invention, the aircraft structure according to the present invention has a combustion time of 15 seconds or less after ignition source digestion of the aircraft structure measured by Advisory Circular (AC) 25.853-1a (1) i. It is characterized by that.

The aircraft structure according to the present invention is represented by ^{Ec 1/3} · ρ ^-1 when the flexural modulus of the aircraft structure is Ec and the specific gravity of the aircraft structure is ρ in the above invention. The specific flexural modulus of the aircraft structure is in the range of 3 or more and 20 or less, and the flexural modulus Ec of the aircraft structure is 3 GPa or more.

The aircraft structure according to the present invention is characterized in that, in the above invention, the specific gravity ρ of the aircraft structure is 0.9 g / cm ³ or less.

The aircraft structure according to the present invention is characterized in that, in the above invention, the reinforcing fibers are coated with a resin, and the thickness of the resin is within the range of 1 μm or more and 15 μm or less.

The aircraft structure according to the present invention is characterized in that, in the above invention, the reinforcing fibers are discontinuous, substantially monofilament-like, and randomly dispersed.

The aircraft structure according to the present invention is characterized in that, in the above invention, the orientation angle θf of the reinforcing fibers in the aircraft structure is 3 ° or more.

The aircraft structure according to the present invention is characterized in that, in the above invention, the mass average fiber length of the reinforcing fibers is within the range of 1 mm or more and 15 mm or less.

The aircraft structure according to the present invention is characterized in that, in the above invention, the reinforcing fiber is a carbon fiber.

In the above-mentioned invention, the structure for an aircraft according to the present invention includes at least one kind of resin selected from the group of polyetherimide resin, polyetheretherketone resin, polyimide resin, and phenol resin. It is characterized by.

The aircraft structure according to the present invention is characterized in that, in the above invention, the aircraft structure has a skin layer.

According to the aircraft structure according to the present invention, it is possible to provide an aircraft structure having excellent lightness, flammability, and mechanical properties.

FIG. 1 is a schematic view showing a cross-sectional structure of an aircraft structure according to the present invention. FIG. 2 is a schematic view showing an example of a dispersed state of reinforcing fibers in the reinforcing fiber mat used in the present invention. FIG. 3 is a schematic view showing an example of a cross-sectional structure in the plane direction and the thickness direction of the aircraft structure according to the present invention. FIG. 4 is a diagram showing a portion within 30% from the surface of the structure to the midpoint position in the thickness direction and the remaining portion. FIG. 5 is a diagram showing a modified example of the configuration shown in FIG. FIG. 6 is a schematic view showing an example of a reinforcing fiber mat manufacturing apparatus.

Hereinafter, the aircraft structure according to the present invention will be described.

FIG. 1 is a schematic view showing a cross-sectional structure of an aircraft structure according to the present invention. As shown in FIG. 1, the aircraft structure 1 according to the present invention is composed of a resin 2, reinforcing fibers 3, and voids 4. Examples of aircraft structures include landing gear pods, winglets, spoilers, edges, rudder, elevators, failing, ribs, seats and the like.

Here, examples of the resin 2 include thermoplastic resins and thermosetting resins. Further, in the present invention, the thermosetting resin and the thermoplastic resin may be blended, and in that case, the component occupying an amount exceeding 50% by mass among the components constituting the resin is referred to as the name of the resin. To do.

In one embodiment of the present invention, the resin 2 preferably contains at least one or more thermoplastic resins. Examples of the thermoplastic resin include "polyethylene terephthalate resin (PET), polybutylene terephthalate resin (PBT), polytrimethylene terephthalate resin (PTT), polyethylene naphthalate resin (PEN), polyester resin such as liquid crystal polyester resin, and polyethylene resin ( PE), polyolefin resin such as polypropylene resin (PP), polybutylene resin, polyarylene sulfide resin such as polyoxymethylene resin (POM), polyamide resin (PA), polyphenylene sulfide resin (PPS), polyetherene sulfide resin (PK), poly Etherketone resin (PEK), polyetheretherketone resin (PEEK), polyetherketoneketone resin (PEKK), polyethernitrile resin (PEN), fluororesin such as polytetrafluoroethylene resin, liquid crystal polymer resin (LCP) " Crystalline resins such as "styrene resin, polycarbonate resin (PC), polymethylmethacrylate resin (PMMA), polyvinyl chloride resin (PVC), polyphenylene ether resin (PPE), polyimide resin (PI), polyamideimide resin" (PAI), polyetherimide resin (PEI), polysalphon resin (PSU), polyethersalphon resin, polyarylate resin (PAR) "and other amorphous resins, other phenol resins, phenoxy resins, and polystyrene-based resins. Thermoplastic elastomers such as polyolefin-based, polyurethane-based, polyester-based, polyamide-based, polybutadiene-based, polyisoprene-based, fluorine-based resin, and acrylonitrile-based, and thermoplastic resins selected from these copolymers and modified products. It can be exemplified. Among them, polyolefin is desirable from the viewpoint of lightness of the obtained aircraft structure, polyamide is desirable from the viewpoint of strength, and amorphous resin such as polycarbonate or styrene resin is desirable from the viewpoint of surface appearance, and flammability is desired. From the viewpoint of continuous use, polyetheretherketone resin (PEEK), polyimide resin (PI), and polyetherimide resin (PEI) are desirable, and from the viewpoint of continuous use temperature, polyetheretherketone resin and polyetherimide resin are preferably used. ..

In one embodiment of the present invention, the resin 2 preferably contains at least one type of thermosetting resin. As the thermosetting resin, unsaturated polyester resin, vinyl ester resin, epoxy resin, phenol resin, urea resin, melamine resin, thermosetting polyimide resin, copolymers and modified products thereof, and at least two kinds thereof are used. A blended resin can be exemplified.

Among the above-mentioned thermoplastic resins and thermoplastic resins used for aircraft structures, phenol resins, epoxy resins, unsaturated polyester resins, vinyl ester resins, polyimide resins and the like can be used as particularly desirable resins. Among these, it is preferable to use a phenol resin or a polyimide resin from the viewpoint of having high strength and elastic modulus under a wide range of temperature conditions as the resin.

In the present invention, the phenol resin is a resin obtained by condensing aldehydes with phenols. Examples of the phenol resin include phenol / formalin resin, cresol / formalin resin, resorcinol resin and the like. Examples of the phenols include alkylphenols such as phenol, cresol, xylenol, ethylphenol, butylphenol, nonylphenol or octylphenol, polyhydric phenols such as bisphenol A, bisphenol F, bisphenol S, resorcin or catechol, halogenated phenol and phenylphenol. , Aminophenol, naphthol and the like. Examples of the aldehydes include formaldehyde, acetaldehyde, benzaldehyde, terephthalaldehyde, hydroxybenzaldehyde, paraformaldehyde, hexamethylenetetramine, furfural and trioxane.

Examples of the phenol resin include novolak type phenol resin and resol type phenol resin, which can be used alone or in combination. The novolak-type phenol resin can be obtained by reacting phenols and formaldehydes in the presence of an acid catalyst such as oxalic acid in the same amount or under the condition of excess phenol. When curing a novolak type phenol resin, hexamethylenetetramine is usually used as a curing agent. The blending amount of hexamethylenetetramine is preferably 10 to 20 parts by weight with respect to 100 parts by weight of the novolak type phenol resin. The resole-type phenol resin can be obtained by reacting phenols and formaldehyde in the same amount or under the condition of excess formaldehyde in the presence of a base catalyst such as sodium hydroxide, ammonia, or organic amine. Since the resole-type phenol resin is a self-curing resin, it can be cured without using a catalyst such as hexamethylenetetramine. In the present invention, as the phenol resin, it is preferable to use a resol type phenol resin because it is excellent in coating property on reinforcing fibers and thermosetting properties.

Commercially available phenolic resins include "AV Light (registered trademark)" (manufactured by Asahi Organic Materials Industry Co., Ltd.), "Otalite (registered trademark)" (manufactured by Otarite Co., Ltd.), and "Kouberite (registered trademark). ) ”(Made by Shin-Kobe Electric Co., Ltd.),“ Sumikon PM (registered trademark) ”(Made by Sumitomo Bakelite Co., Ltd.),“ Nikkalite (registered trademark) ”(Made by Nippon Synthetic Chemical Co., Ltd.),“ Stand Light ( "Registered trademark" (manufactured by Hitachi Kasei Kogyo Co., Ltd.), "Fudolite (registered trademark)" (manufactured by Fudo Co., Ltd.), "Macon (registered trademark)" (manufactured by Meiwa Kasei Co., Ltd.), "Phenolite (manufactured by Meiwa Kasei Co., Ltd.)" Examples include "registered trademark" (manufactured by DIC Co., Ltd.) and "Tamanor (registered trademark)" (manufactured by Arakawa Chemical Industry Co., Ltd.). Further, the aircraft structure according to the present invention may contain an impact resistance improver such as an elastomer or a rubber component, and other fillers and additives as long as the object of the present invention is not impaired. Examples of fillers and additives include inorganic fillers, flame retardants, conductivity-imparting agents, crystal nucleating agents, UV absorbers, antioxidants, anti-vibration agents, antibacterial agents, insect repellents, deodorants, and anti-coloring agents. , Heat stabilizers, mold release agents, antistatic agents, plasticizers, lubricants, colorants, pigments, dyes, foaming agents, antifoaming agents, or coupling agents can be exemplified.

The volume content of the resin 2 is in the range of 2.5% by volume or more and 85% by volume or less. When the volume content of the resin 2 is less than 2.5% by volume, the reinforcing fibers 3 in the aircraft structure 1 cannot be bound to each other, and the reinforcing effect of the reinforcing fibers 3 cannot be made sufficient. It is not desirable because the mechanical properties of the aircraft structure, especially the bending properties, cannot be satisfied. On the other hand, when the volume content of the resin 2 is larger than 85% by volume, the amount of the resin is too large and it becomes difficult to form a void structure, which is not desirable.

The reinforcing fibers 3 include metal fibers such as aluminum, brass and stainless steel, insulating fibers such as PAN-based, rayon-based, lignin-based and pitch-based carbon fibers, graphite fibers and glass, aramid, PBO, polyphenylene sulfide and polyester. Examples thereof include organic fibers such as acrylic, nylon and polyethylene, and inorganic fibers such as silicon carbide and silicon nitride. Further, these fibers may be surface-treated. The surface treatment includes a treatment with a coupling agent, a treatment with a sizing agent, a treatment with a binding agent, a treatment with an additive, and the like, in addition to the treatment with a metal as a conductor. Further, one type of these fibers may be used alone, or two or more types may be used in combination. Among them, PAN-based, pitch-based, rayon-based and other carbon fibers having excellent specific strength and specific rigidity are preferably used from the viewpoint of weight reduction effect. Further, from the viewpoint of enhancing the economic efficiency of the obtained aircraft structure, glass fiber is preferably used, and in particular, it is desirable to use carbon fiber and glass fiber in combination from the viewpoint of the balance between mechanical properties and economic efficiency. Furthermore, aramid fibers are preferably used from the viewpoint of enhancing the shock absorption and shapeability of the obtained aircraft structure, and in particular, carbon fibers and aramid fibers should be used in combination from the viewpoint of the balance between mechanical properties and shock absorption. Is desirable. Further, from the viewpoint of increasing the conductivity of the obtained aircraft structure, reinforcing fibers coated with a metal such as nickel, copper or ytterbium can also be used. Among these, PAN-based carbon fibers having excellent mechanical properties such as strength and elastic modulus can be more preferably used.

It is desirable that the reinforcing fibers 3 are discontinuous, substantially monofilament-like, and randomly dispersed. By adopting the reinforcing fiber 3 in such an embodiment, molding into a complicated shape becomes easy when a precursor of a sheet-shaped aircraft structure or an aircraft structure is molded by applying an external force. Further, by adopting the reinforcing fiber 3 in such an embodiment, the voids formed by the reinforcing fiber 3 are densified, and the weak portion at the fiber bundle end of the reinforcing fiber 3 in the aircraft structure 1 can be minimized, which is excellent. In addition to reinforcement efficiency and reliability, isotropic is also imparted. Here, the substantially monofilament means that the reinforcing fiber single yarn is present in the fineness strands of less than 500 threads. More preferably, it is dispersed in a monofilament form.

Here, substantially monofilament-like or monofilament-like dispersion means that the reinforcing fibers 3 arbitrarily selected in the aircraft structure 1 have a two-dimensional contact angle of 1 ° or more. It means that the ratio (hereinafter, also referred to as fiber dispersion rate) is 80% or more, in other words, the number of bundles in which two or more single fibers are in contact with each other and parallel to each other in the aircraft structure 1 is less than 20%. To say. Therefore, here, it is particularly preferable that the mass fraction of the fiber bundle having 100 or less filaments in at least the reinforcing fiber 3 corresponds to 100%.

The two-dimensional contact angle is an angle formed by a single fiber and a single fiber in contact with the single fiber in the case of a discontinuous reinforcing fiber, and is 0 among the angles formed by the contacting single fibers. It is defined as the angle on the acute angle side within the range of ° or more and 90 ° or less. This two-dimensional contact angle will be further described with reference to the drawings. FIG. 2 is a schematic view showing an example of the dispersed state of the reinforcing fibers in the reinforcing fiber mat when observed from the plane direction (FIG. 2 (a)) and the thickness direction (FIG. 2 (b)). With the single fiber 11a as a reference, the single fiber 11a is observed intersecting with the single fibers 11b to 11f in FIG. 2A, but in FIG. 2B, the single fiber 11a is in contact with the single fibers 11e and 11f. Not. In this case, with respect to the reference single fiber 11a, the two-dimensional contact angles are evaluated for the single fibers 11b to 11d, and of the two angles formed by the two contacting single fibers, 0 ° or more and 90 °. The angle A on the acute angle side within the following range.

The method for measuring the two-dimensional contact angle is not particularly limited, and for example, a method of observing the orientation of the reinforcing fibers 3 from the surface of the aircraft structure 1 can be exemplified. In this case, by polishing the surface of the aircraft structure 1 to expose the reinforcing fibers 3, the reinforcing fibers 3 can be more easily observed. Further, a method of taking an orientation image of the reinforcing fiber 3 by performing X-ray CT transmission observation can also be exemplified. In the case of the reinforcing fiber 3 having high X-ray transparency, if the reinforcing fiber 3 is mixed with the tracer fiber or the reinforcing fiber 3 is coated with the tracer chemical, the reinforcing fiber 3 is observed. It is desirable because it makes it easier to do. If measurement is difficult by the above method, the resin component of the aircraft structure 1 is burned down at a high temperature by a heating furnace or the like, and then the reinforcing fibers 3 taken out using an optical microscope or an electron microscope are used as reinforcing fibers. A method of observing the orientation of 3 can be exemplified.

The fiber dispersion rate is measured by the following procedure based on the above-mentioned observation method. That is, the two-dimensional contact angles with all the single fibers (single fibers 11b to 11d in FIG. 2) in contact with the randomly selected single fibers (single fibers 11a in FIG. 2) are measured. This is performed for 100 single fibers, and the ratio is calculated from the ratio of the total number of all single fibers whose two-dimensional contact angles have been measured to the number of single fibers having a two-dimensional contact angle of 1 ° or more.

Further, it is particularly desirable that the reinforcing fibers 3 are randomly dispersed. Here, the fact that the reinforcing fibers 3 are randomly dispersed means that the arithmetic mean value of the two-dimensional orientation angles of the arbitrarily selected reinforcing fibers 3 in the aircraft structure 1 is within the range of 30 ° or more and 60 ° or less. Say something. The two-dimensional orientation angle is an angle formed by the single fiber of the reinforcing fiber 3 and the single fiber intersecting the single fiber, and is 0 ° or more among the angles formed by the intersecting single fibers. It is defined as the angle on the acute angle side within the range of 90 ° or less.

This two-dimensional orientation angle will be further described with reference to the drawings. In FIGS. 2A and 2B, with the single fiber 11a as a reference, the single fiber 11a intersects with other single fibers 11b to 11f. Here, the crossing means a state in which the reference single fiber is observed intersecting with other single fibers in the two-dimensional plane to be observed, and the single fibers 11a and the single fibers 11b to 11f are not necessarily in contact with each other. It does not have to be, and the state of being observed at the same time when projected and viewed is no exception. That is, when looking at the reference single fiber 11a, all of the single fibers 11b to 11f are the evaluation targets of the two-dimensional orientation angles, and in FIG. 2A, the two single fibers whose two-dimensional orientation angles intersect are Of the two angles to be formed, the angle A on the acute angle side is within the range of 0 degrees or more and 90 degrees or less.

The method for measuring the two-dimensional orientation angle is not particularly limited, but for example, a method for observing the orientation of the reinforcing fibers 3 from the surface of the component can be exemplified, and the same means as the above-mentioned method for measuring the two-dimensional contact angle. Can be taken. The average value of the two-dimensional orientation angles is measured by the following procedure. That is, the average value of the two-dimensional orientation angles with all the single fibers (single fibers 11b to 11f in FIG. 2) intersecting with the randomly selected single fibers (single fibers 11a in FIG. 2) is measured. .. For example, when there are a large number of other single fibers intersecting one single fiber, the arithmetic mean value measured by randomly selecting 20 other intersecting single fibers may be substituted. This measurement is repeated a total of 5 times with respect to another single fiber, and the arithmetic mean value is calculated as the arithmetic mean value of the two-dimensional orientation angle.

Since the reinforcing fibers 3 are substantially monofilament-like and randomly dispersed, the performance provided by the reinforcing fibers 3 dispersed in the substantially monofilament shape described above can be maximized. Further, in the aircraft structure 1, isotropic properties can be imparted to the mechanical properties. From this point of view, it is desirable that the fiber dispersion ratio of the reinforcing fiber 3 is 90% or more, and the closer it is to 100%, the more desirable. Further, it is desirable that the arithmetic mean value of the two-dimensional orientation angle of the reinforcing fiber 3 is within the range of 40 ° or more and 50 ° or less, and it is desirable that the angle approaches the ideal angle of 45 °.

On the other hand, as an example in which the reinforcing fibers 3 do not take the form of a non-woven fabric, there are a sheet base material, a woven base material, a non-crimp base material, etc. in which the reinforcing fibers 3 are arranged in one direction. In these forms, since the reinforcing fibers 3 are regularly and densely arranged, the voids 4 in the aircraft structure 1 are reduced, the impregnation of the resin 2 becomes extremely difficult, and an unimpregnated portion is formed. , The choice of impregnation means and resin type may be greatly restricted.

The form of the reinforcing fiber 3 may be either a continuous reinforcing fiber having the same length as that of the aircraft structure 1 or a discontinuous reinforcing fiber having a finite length cut to a predetermined length. From the viewpoint that the resin 2 can be easily impregnated and the amount thereof can be easily adjusted, the discontinuous reinforcing fiber is desirable.

The volume content of the reinforcing fiber 3 is in the range of 0.5% by volume or more and 55% by volume or less. When the volume content of the reinforcing fiber 3 is less than 0.5% by volume, the reinforcing effect derived from the reinforcing fiber 3 cannot be made sufficient, which is not desirable. On the other hand, when the volume content of the reinforcing fibers 3 is larger than 2.5% by volume, the volume content of the resin 2 with respect to the reinforcing fibers 3 is relatively small, so that the reinforcing fibers 3 in the aircraft structure 1 are used with each other. It is not desirable because the reinforcing effect of the reinforcing fiber 3 cannot be made sufficient and the mechanical properties of the aircraft structure 1, particularly the bending property, cannot be satisfied.

The reinforcing fiber 3 is coated with the resin 2, and it is desirable that the thickness of the resin 2 is within the range of 1 μm or more and 15 μm or less. The state of coating of the reinforcing fibers 3 coated with the resin 2 is such that the shape stability of the aircraft structure 1 is as long as the intersections of the single fibers of the reinforcing fibers 3 constituting the aircraft structure 1 are covered. It is sufficient from the viewpoint of ease of thickness control and degree of freedom, but in a more desirable embodiment, it is desirable that the resin 2 is coated with the above-mentioned thickness around the reinforcing fibers 3. This state means that the surface of the reinforcing fiber 3 is not exposed by the resin 2, in other words, the reinforcing fiber 3 forms an electric wire-like film by the resin 2. As a result, the aircraft structure 1 further has shape stability and sufficient expression of mechanical properties. Further, the coating state of the reinforcing fibers 3 coated with the resin 2 does not have to be coated with all of the reinforcing fibers 3, and the shape stability of the aircraft structure 1 according to the present invention, the flexural modulus, and the like. It may be within the range that does not impair the bending strength.

It is desirable that the mass average fiber length of the reinforcing fiber 3 is within the range of 1 mm or more and 15 mm or less. As a result, the reinforcing efficiency of the reinforcing fibers 3 can be increased, and excellent mechanical properties can be given to the aircraft structure 1. If the mass average fiber length of the reinforcing fibers 3 is less than 1 mm, the voids 4 in the aircraft structure 1 cannot be efficiently formed, so that the specific gravity may increase. In other words, the desired thickness may be obtained even though the mass is the same. It is not desirable because it becomes difficult to obtain the aircraft structure 1. On the other hand, when the mass average fiber length of the reinforcing fiber 3 is longer than 15 mm, the reinforcing fiber 3 is likely to be bent by its own weight in the aircraft structure 1, which is not desirable because it becomes a factor that hinders the development of mechanical properties. .. For the mass average fiber length, the resin component of the aircraft structure 1 was removed by a method such as burning or elution, 400 fibers were randomly selected from the remaining reinforcing fibers 3, and the length was measured up to 10 μm units, and these were measured. Can be calculated as the average length of.

The void 4 in the present invention refers to a space formed by the reinforcing fibers 3 coated with the resin 2 forming a columnar support, which are overlapped or intersected with each other. For example, when an aircraft structure precursor in which the reinforcing fibers 3 are pre-impregnated with the resin 2 is heated to obtain an aircraft structure, the reinforcing fibers 3 are raised due to the melting or softening of the resin 2 due to the heating, resulting in voids. 4 is formed. This is based on the property that the internal reinforcing fibers 3 that have been compressed by pressurization in the aircraft structure precursor are raised by the raising force derived from the elastic modulus. Further, the content of the void 4 in the aircraft structure 1 is within the range of 10% by volume or more and 99% by volume or less. If the content of the voids 4 is less than 10% by volume, the specific gravity of the aircraft structure 1 becomes high and the lightness cannot be satisfied, which is not desirable. On the other hand, when the content of the voids 4 is larger than 99% by volume, in other words, the thickness of the resin 2 coated around the reinforcing fibers 3 becomes thin, so that the reinforcing fibers 3 in the aircraft structure 1 become thin. It is not desirable because the mechanical properties are low due to insufficient reinforcement. The upper limit of the content of the void 4 is preferably 97% by volume. In the present invention, the total volume content of the resin 2, the reinforcing fibers 3, and the voids 4 constituting the aircraft structure 1 is 100% by volume.

When the length of the reinforcing fiber 3 is Lf and the orientation angle of the reinforcing fiber 3 in the cross-sectional direction of the aircraft structure 1 is θf, the thickness St of the aircraft structure 1 is the conditional expression: St ≧ Lf ² · (1-). cos (θf)) is satisfied. When the thickness St of the aircraft structure 1 does not satisfy the above conditional expression, the reinforcing fibers 3 in the aircraft structure 1 are bent, or the balance between the aircraft structure 1 having the desired thickness and the fiber length is obtained. Indicates that is inferior. As a result, the aircraft structure 1 is inferior in the degree of freedom in thickness design because the characteristics of the introduced reinforcing fibers 3 cannot be fully exhibited, and further, among the mechanical properties of the aircraft structure 1, it is reinforced. The property of utilizing the tensile strength and tensile elastic modulus of the fiber 3 is not desirable because the straightness of the reinforcing fiber 3 is lost and an efficient reinforcing effect cannot be obtained. In the above conditional expression, the balance between the flexural modulus and the specific flexural modulus, which are the characteristics of the aircraft structure 1 formed by the length Lf of the reinforcing fiber 3 and the orientation angle θf thereof, can be obtained, and for aircraft. Depending on the fiber length in the structure 1 and its orientation angle, it is easy to deform in the state before solidification or hardening during the molding process, and it is easy to mold the desired aircraft structure 1. Therefore, the aircraft structure 1 The thickness St is preferably in the range of 2% or more and 20% or less, and particularly preferably in the range of 5% or more and 18% or less. The units used in the conditional expression are St [mm], Lf [mm], and θf [°].

Here, for the length Lf of the reinforcing fibers 3, 400 fibers were randomly selected from the remaining reinforcing fibers 3 after removing the resin component of the aircraft structure 1 by a method such as burning or elution, and the length was 10 μm. It can be measured up to the unit and calculated as the mass average fiber length calculated from those lengths. Further, the orientation angle θf of the reinforcing fiber 3 in the cross-sectional direction of the aircraft structure 1 is the degree of inclination of the aircraft structure 1 with respect to the cross-sectional direction, in other words, the degree of inclination of the reinforcing fiber 3 with respect to the thickness direction. is there. The larger the value, the more it stands and tilts in the thickness direction, and it is given in the range of 0 ° or more and 90 ° or less. That is, by setting the orientation angle θf of the reinforcing fibers 3 within such a range, the reinforcing function in the aircraft structure 1 can be more effectively exhibited. The upper limit of the orientation angle θf of the reinforcing fiber 3 is not particularly limited, but it is preferably 60 ° or less, and further 45 ° or less in view of the development of the flexural modulus when the structure for aircraft 1 is used. Is more desirable. Further, when the orientation angle θf of the reinforcing fibers 3 is less than 3 °, the reinforcing fibers 3 in the aircraft structure 1 are in a planar, in other words, two-dimensionally oriented state, so that the thickness of the aircraft structure 1 is increased. It is not desirable because the degree of freedom of the aircraft is reduced and the lightness cannot be satisfied. Therefore, the orientation angle θf of the reinforcing fibers 3 is preferably 3 ° or more.

The orientation angle θf of the reinforcing fibers 3 can be measured based on the observation of the vertical cross section of the aircraft structure 1 with respect to the plane direction. FIG. 3 is a schematic view showing an example of a cross-sectional structure of the aircraft structure according to the present invention in the plane direction (FIG. 3 (a)) and the thickness direction (FIG. 3 (b)). In FIG. 3A, the cross sections of the reinforcing fibers 3a and 3b are approximated to an elliptical shape for simplification of measurement. Here, the cross section of the reinforcing fiber 3a has a small elliptical aspect ratio (= elliptical major axis / elliptical minor axis), whereas the cross section of the reinforcing fiber 3b has a large elliptical aspect ratio. On the other hand, according to FIG. 3B, the reinforcing fibers 3a have an inclination substantially parallel to the thickness direction Y, and the reinforcing fibers 3b have an inclination of a certain amount with respect to the thickness direction Y. In this case, for the reinforcing fibers 3b, the angle θx formed by the plane direction X of the aircraft structure 1 and the fiber main axis (long axis direction in the ellipse) α is substantially equal to the out-of-plane angle θf of the reinforcing fibers 3b. On the other hand, with respect to the reinforcing fiber 3a, there is a large discrepancy between the angle θx and the angle indicated by the orientation angle θf, and it cannot be said that the angle θx reflects the orientation angle θf. Therefore, when the orientation angle θf is read from the cross section perpendicular to the plane direction of the aircraft structure 1, the detection accuracy of the orientation angle θf can be improved by extracting those having an elliptical aspect ratio of the fiber cross section of a certain value or more.

As an index of the elliptical aspect ratio to be extracted, when the cross-sectional shape of the single fiber is close to a perfect circle, that is, the fiber aspect ratio in the cross section perpendicular to the long direction of the reinforcing fiber is 1.1 or less, the elliptical aspect ratio A method of measuring the angle formed by the X direction and the fiber principal axis α for the reinforcing fibers 3 having an aspect ratio of 20 or more and adopting this as the orientation angle θf can be used. On the other hand, when the cross-sectional shape of the single fiber is elliptical or cocoon-shaped and the fiber aspect ratio is larger than 1.1, the reinforcing fiber 3 having a larger elliptical aspect ratio is focused on and the orientation angle θf is measured. It is better to have an elliptical aspect ratio of 30 or more when the fiber aspect ratio is 1.1 or more and less than 1.8, and an elliptical aspect ratio of 40 when the fiber aspect ratio is 1.8 or more and less than 2.5. As described above, when the fiber aspect ratio is 2.5 or more, the reinforcing fiber 3 having an elliptical aspect ratio of 50 or more may be selected and the orientation angle θf may be measured.

Regarding the flammability of the aircraft structure according to the present invention, the burning time can be measured according to FAR25.853 Appendix F. In addition, the superiority or inferiority of the burning time is based on the Advisory Circular (AC) 25.853-1a (1) i "Flamability Requirements for Aircraft After Ignition Second Aircraft" published by Federal Aviation Adminition. The burning time of the structure 1 is 30 seconds or less, preferably 15 seconds or less. If the burning time of the aircraft structure is longer than 30 seconds, it means that the combustibility is poor, that is, it burns, which is not desirable from the viewpoint of safety. There is no limitation on the short burning time of the aircraft structure because the shorter the burning time, the better the combustibility. The above-mentioned FAR25.853 Appendix F of the US Federal Aviation Regulations (FAR) is to fix the test piece with the long side vertical, and apply a flame to the center of the bottom of the test piece from directly below using a burner to burn it. Measured and evaluated. In such an evaluation method, the distance from the burner to the test piece is 19 mm, and the time during which the test piece is roasted for 60 seconds and then the flame is emitted from the test piece is measured as the burning time.

When the flexural modulus of the aircraft structure 1 is Ec and the specific gravity of the aircraft structure 1 is ρ, the specific flexural modulus of the aircraft structure 1 represented as ^{Ec 1/3} · ρ ^{-1 is 3 or more.} , 20 or less. When the specific flexural modulus of the aircraft structure 1 is less than 3, the flexural modulus is high and the specific gravity is also high, which is not desirable because the desired weight reduction effect cannot be obtained. On the other hand, when the specific flexural modulus of the aircraft structure 1 is larger than 20, it indicates that the flexural modulus is low although the weight reduction effect is sufficient, and the shape desired as the aircraft structure 1 It is not desirable because it is difficult to hold the structure and the flexural modulus of the aircraft structure 1 itself is inferior. Generally, the specific flexural modulus of steel or aluminum is 1.5 or less, which is a region of the specific flexural modulus extremely superior to those of these metal materials. Furthermore, it is 3 or more, more preferably 5 or more, which exceeds 2.3, which is the general specific bending elastic modulus of the carbon fiber reinforced resin composite material, which is focused on the weight reduction effect.

The flexural modulus Ec of the aircraft structure 1 is preferably 3 GPa or more, preferably 6 GPa or more. When the flexural modulus Ec of the aircraft structure 1 is less than 3 GPa, the range of use as the aircraft structure 1 is limited, which is not desirable. Further, in order to facilitate the design of the aircraft structure 1, it is desirable that the flexural modulus has isotropic property. The upper limit of the flexural modulus is not limited, but in general, in a structure composed of reinforcing fibers and resin, the upper limit can be a value calculated from the elastic moduli of the reinforcing fibers and the resin, which are the constituents thereof. In the aircraft structure according to the present invention, whether the aircraft structure is used alone or in combination with other members, the bending elastic modulus of the aircraft structure itself is used to form the member. In order to design and put it into practical use, 50 GPa is desirable.

The specific gravity ρ of the aircraft structure 1 ^{is preferably 0.9 g / cm 3} or less. When the specific gravity ρ of ^{the aircraft structure 1 is larger than 0.9 g / cm 3} , it means that the mass of the aircraft structure 1 increases, and as a result, the mass of the product increases. It is not desirable because it becomes. There is no limit to the lower limit of the specific gravity, but in general, in a structure composed of reinforcing fibers and resin, the lower limit can be a value calculated from the volume ratios of the reinforcing fibers, the resin, and the voids, which are the constituents thereof. .. In the aircraft structure according to the present invention, whether the aircraft structure is used alone or in combination with other members, the specific gravity of the aircraft structure itself is determined by the reinforcing fibers used or the like. ^{Although it depends on the resin, 0.03 g / cm 3} or more is a desirable range from the viewpoint of maintaining the mechanical properties of the aircraft structure.

In the aircraft structure according to the present invention, an aircraft member can be manufactured with an arbitrary thickness, but the following structure can be adopted from the viewpoint of achieving both light weight and flexural rigidity and improving flammability. Specifically, the porosity in the portion within 30% from the surface of the aircraft structure 1 to the midpoint position in the thickness direction is within the range of 0% by volume or more and less than 10% by volume, and the voids in the remaining portion. It is desirable that the ratio is within the range of 10% by volume or more and 99% by volume or less. The smaller the porosity, the better the mechanical properties, and the larger the porosity, the better the lightness. In other words, when the aircraft structure 1 is made of the same material, the porosity in the portion within 30% from the surface of the aircraft structure 1 to the midpoint position in the thickness direction is 0 volume. % Or more and less than 10% by volume ensures the mechanical properties of the aircraft structure 1, and the porosity of the remaining portion is within the range of 10% by volume or more and 99% by volume or less to ensure lightweight characteristics. It is desirable because it can be satisfied. The upper limit of the porosity of the remaining portion is preferably 97% by volume.

In the present invention, the thickness of the aircraft structure 1 can be obtained from the shortest distance connecting one point on the surface for which the thickness is desired and the surface on the back side thereof. The midpoint in the thickness direction means the midpoint of the thickness of the aircraft structure 1. The portion within 30% from the surface of the aircraft structure to the midpoint position in the thickness direction is for aircraft when the distance between the surface of the aircraft structure 1 and the midpoint in the thickness direction is 100%. It means a portion including a distance of 30% from one surface of the structure 1. The remaining portion here is a portion within 30% from one surface of the aircraft structure 1 to the midpoint position in the thickness direction and from the other surface of the aircraft structure 1 to the midpoint position in the thickness direction. It means the remaining part excluding the part within 30% of. As shown in FIG. 4, the portion R1 within 30% from the surface of the aircraft structure 1 to the midpoint position in the thickness direction and the remaining portion R2 exist at different positions in the thickness direction of the aircraft structure 1. It may be present at different positions in the plane direction as shown in FIG.

It is desirable that the reinforcing fiber 3 in the present invention takes the form of a non-woven fabric from the viewpoint of easiness of impregnating the reinforcing fiber 3 with the resin 2. Further, since the reinforcing fiber 3 has a non-woven fabric-like form, in addition to the ease of handling of the non-woven fabric itself, impregnation is easy even in the case of a thermoplastic resin generally having a high viscosity. It is desirable because it can be done. Here, the non-woven fabric-like form refers to a form in which the strands and / or monofilaments of the reinforcing fibers 3 are dispersed in a plane without regularity, and is a chopped strand mat, a continuance strand mat, a papermaking mat, a carding mat, or an airlaid mat. Etc. (hereinafter, these are collectively referred to as a reinforcing fiber mat).

As a method for manufacturing a reinforcing fiber mat constituting an aircraft structure 1, for example, there is a method in which reinforcing fibers 3 are dispersed in advance in a strand and / or substantially monofilament form to manufacture a reinforcing fiber mat. As a method for manufacturing the reinforcing fiber mat, a dry process such as an air-laid method in which the reinforcing fibers 3 are dispersed and sheeted by an air flow, a carding method in which the reinforcing fibers 3 are mechanically combed to form a sheet, and the like are used for reinforcement. As a known technique, a wet process by the Radrite method in which fibers 3 are stirred in water to make paper can be mentioned. As a means for bringing the reinforcing fiber 3 closer to a monofilament shape, in the dry process, a method of providing a fiber-spreading bar, a method of further vibrating the fiber-spreading bar, a method of finening the eyes of the card, and a method of rotating the card are used. An example of the method of adjustment can be given. In the wet process, a method of adjusting the stirring conditions of the reinforcing fibers 3, a method of diluting the reinforcing fiber concentration of the dispersion liquid, a method of adjusting the viscosity of the dispersion liquid, a method of suppressing vortex flow when transferring the dispersion liquid, etc. Can be exemplified. In particular, it is desirable that the reinforcing fiber mat is manufactured by a wet method, and the reinforcing fiber 3 of the reinforcing fiber mat can be manufactured by increasing the concentration of the input fiber or adjusting the flow velocity (flow rate) of the dispersion liquid and the speed of the mesh conveyor. The ratio can be easily adjusted. For example, by slowing the speed of the mesh conveyor with respect to the flow velocity of the dispersion liquid, the orientation of the fibers in the obtained reinforcing fiber mat becomes difficult to be oriented in the picking direction, and a bulky reinforcing fiber mat can be manufactured. The reinforcing fiber mat may be composed of a single reinforcing fiber 3, and the reinforcing fiber 3 may be mixed with a powder-shaped or fiber-shaped matrix resin component, or the reinforcing fiber 3 may be mixed with an organic compound or an inorganic compound. , Reinforcing fibers 3 may be sealed with a resin component.

Further, the reinforcing fiber mat may be impregnated with the resin 2 in advance to serve as a structural precursor for an aircraft. As a method for producing the structure precursor for an aircraft according to the present invention, a method is used in which a reinforcing fiber mat is impregnated with a reinforcing fiber mat by applying pressure to the reinforcing fiber mat in a state of being heated to a temperature higher than the temperature at which the resin 2 is melted or softened. It is desirable from the viewpoint of ease of manufacture. Specifically, a method of melt-impregnating the laminate in which the resin 2 is arranged from both sides in the thickness direction of the reinforcing fiber mat can be preferably exemplified. On the other hand, in the method of melting or softening the resin, in addition to heating, the resin component is diluted with a soluble solvent to reduce the apparent viscosity, and then impregnated into the reinforcing fiber mat, and then only the solvent is used. Can be exemplified by a method of removing the solvent by volatilization or evaporation.

As equipment for realizing each of the above methods, a compression molding machine or a double belt press can be preferably used. In the case of the batch type, it is the former, and productivity can be improved by using an intermittent press system in which two or more machines for heating and cooling are arranged in parallel. In the case of the continuous type, it is the latter, and continuous processing can be easily performed, so that continuous productivity is excellent.

When manufacturing the aircraft structure 1 according to the present invention, it is preferable to adopt the method manufactured by at least the following steps [1] and [2] from the viewpoint of ease of manufacture.

Step [1]: A step of applying pressure in a state of being heated to a temperature higher than the temperature at which the resin 2 melts or softens, and impregnating the reinforcing fiber mat with the resin 2 to prepare a structural precursor for an aircraft.
Step [2]: A step of expanding the aircraft structure precursor by adjusting the thickness in a heated state.

The step [2] is a step of expanding the aircraft structure precursor obtained in the step [1] by adjusting the thickness in a heated state. At this time, when the resin 2 constituting the aircraft structure 1 is a thermoplastic resin, the heating temperature can be determined by giving a sufficient amount of heat to melt or soften the thickness of the manufactured aircraft structure 1 and control the thickness of the manufactured aircraft structure 1. It is preferable from the viewpoint of the production speed, and specifically, it is preferable that the temperature is 10 ° C. or higher with respect to the melting temperature and the temperature is not higher than the thermal decomposition temperature of the thermoplastic resin. Further, when a thermosetting resin is used as the resin 2, it is possible to provide a sufficient amount of heat to melt or soften the thermosetting resin raw material before forming a crosslinked structure and curing the structure 1 for an aircraft. It is preferable from the viewpoint of thickness control and manufacturing speed.

The thickness control method does not depend on the method as long as the heated aircraft structure precursor can be controlled to the desired thickness, but a method of constraining the thickness using a metal plate or the like, or applying the thickness to the aircraft structure precursor. A method of controlling the thickness according to the pressure applied is exemplified as a preferable method from the viewpoint of ease of production. As the equipment for realizing the above method, a compression molding machine or a double belt press can be preferably used. In the case of the batch type, it is the former, and productivity can be improved by using an intermittent press system in which two or more machines for heating and cooling are arranged in parallel. In the case of the continuous type, it is the latter, and continuous processing can be easily performed, so that continuous productivity is excellent.

Examples of the reinforcing fiber mat that does not take the form of a non-woven fabric include a sheet base material, a woven base material, and a non-crimp base material in which the reinforcing fibers 3 are arranged in one direction. In these forms, since the reinforcing fibers 3 are regularly and densely arranged, there are few voids in the reinforcing fiber mat, and the thermoplastic resin does not form a sufficient anchoring structure. Joining ability is reduced. Further, when the resin 2 is a thermoplastic resin, impregnation becomes extremely difficult, an unimpregnated portion is formed, and the choice of impregnation means and resin type is greatly restricted.

In the present invention, a sheet-like intermediate group in which the aircraft structure 1 or the aircraft structure precursor is used for the core layer and the continuous reinforcing fibers 3 are impregnated with the resin as long as the characteristics of the present invention are not impaired. It is also possible to make a sandwich structure in which the material is used for the skin layer. Here, the continuous reinforcing fibers 3 are continuous in at least one direction with a length of 100 mm or more, and an aggregate in which a large number of the continuous reinforcing fibers 3 are arranged in one direction, a so-called reinforcing fiber bundle, is the total length of the sandwich structure. It is continuous over. As a form of the sheet-like intermediate base material composed of continuous reinforcing fibers 3, a cloth composed of a reinforcing fiber bundle composed of a large number of continuous reinforcing fibers 3 and a large number of continuous reinforcing fibers 3 are arranged in one direction. Reinforcing fiber bundles (unidirectional fiber bundles), unidirectional cloths composed of these unidirectional fiber bundles, and the like. The reinforcing fiber 3 may be composed of a plurality of fiber bundles having the same form, or may be composed of a plurality of fiber bundles having different forms. The number of reinforcing fibers constituting one reinforcing fiber bundle is usually 300 to 48,000, but it is preferably 300 to 24,000, more preferably, considering the production of prepreg and the production of cloth. The number is 1,000 to 12,000.

In order to control the flexural modulus, a form in which the reinforcing fibers 3 are laminated by changing the direction is preferably used. In particular, in order to efficiently increase the elastic modulus and strength of the sandwich structure, it is desirable to use continuous reinforcing fibers (referred to as UD) in which the fiber bundles are aligned in one direction.

Hereinafter, the present invention will be described in more detail with reference to Examples.

(1) Volume content of reinforcing fibers in aircraft structures Vf
After measuring the mass Ws of the aircraft structure, the aircraft structure was crushed, the volume Vs was measured according to the JIS K0061 method (2001), and the volume Vs was calculated from the following equation.

Vf (volume%) = 100 / (Vs / ρf + (100-Vs) / ρf)
ρf: Density of reinforcing fibers [g / cm ³ ]
ρr: Resin density [g / cm ³ ]

(2) Bending test of aircraft structure A test piece was cut out from the aircraft structure, and the flexural modulus was measured according to the ISO178 method (1993). As the test piece, when the arbitrary direction is the 0 ° direction, a test piece cut out in four directions of + 45 °, -45 °, and 90 ° is prepared, and the number of measurements n = 5 in each direction, and the arithmetic mean. The value was defined as the flexural modulus Ec. As a measuring device, an "Instron (registered trademark)" 5565 type universal material testing machine (manufactured by Instron Japan Co., Ltd.) was used. From the obtained results, the specific flexural modulus of the aircraft structure was calculated by the following formula.

Specific flexural modulus = Ec ^1/3 / ρ

(3) Orientation angle θf of reinforcing fibers in aircraft structures
A small piece having a width of 25 mm was cut out from the aircraft structure, embedded in epoxy resin, and polished so that the vertical cross section in the sheet thickness direction became the observation surface to prepare a sample. The sample was magnified 400 times with a laser microscope (VK-9510 manufactured by KEYENCE CORPORATION), and the cross-sectional shape of the fiber was observed. The observation image is developed on general-purpose image analysis software, the individual fiber cross sections visible in the observation image are extracted using the program built into the software, an ellipse inscribed in the fiber cross section is provided, and the shape of the fiber cross section is approximated. (Hereafter referred to as fiber ellipse). Further, the angle formed by the X-axis direction and the major axis direction of the fiber ellipse was determined for the fiber ellipse having an aspect ratio of 20 or more represented by the major axis length α / minor axis length β of the fiber ellipse. By repeating the above operation for the observation samples extracted from different parts of the structure, the orientation angles of a total of 600 reinforcing fibers were measured, and the arithmetic mean value was obtained as the orientation angle θf of the reinforcing fibers.

(4) Specific gravity of aircraft structure ρ
A test piece was cut out from the aircraft structure, and the apparent density of the aircraft structure was measured with reference to JIS K7222 (2005). The dimensions of the test piece were 100 mm in length and 100 mm in width. The length, width, and thickness of the test piece were measured with a micrometer, and the volume V of the test piece was calculated from the obtained values. Moreover, the mass M of the cut-out test piece was measured with an electronic balance. By substituting the obtained mass M and volume V into the following equation, the specific gravity ρ of the aircraft structure was calculated.

ρ [g / cm ³ ] = 10 ³ x M [g] / V [mm ³ ]

(5) Volume content of voids in the aircraft structure A test piece is cut out from the aircraft structure to a length of 10 mm and a width of 10 mm, and the cross section is scanned electron microscope (SEM) (S-4800 type manufactured by Hitachi High-Technologies Corporation). ), And 10 places were photographed at equal intervals from the surface of the aircraft structure at a magnification of 1000 times. For each image to determine the area A _a void in the image. Moreover, to calculate the porosity by dividing the area A _a void in the area of the entire image. The volume content of the voids in the aircraft structure was determined by arithmetic mean from the porosities of a total of 50 locations taken at 10 locations each with 5 test pieces.

(6) Thickness of resin coated with reinforcing fibers in aircraft structure A test piece is cut out from the aircraft structure to a length of 10 mm and a width of 10 mm, and the cross section is scanned by a scanning electron microscope (SEM) (manufactured by Hitachi High-Technologies Corporation). It was observed with an S-4800 type), and arbitrary 10 points were photographed at a magnification of 3000 times. The coating thickness of the resin coated on the reinforcing fibers was measured from any 50 places where the cross section of the reinforcing fibers in the obtained image was cut. As the thickness of the resin coated with the reinforcing fibers, the arithmetic mean value of the measurement results at these 50 locations was used.

(7) Evaluation of Combustibility of Aircraft Structure A test piece was cut out from the aircraft structure, and the combustion time was measured according to FAR25.853 Appendix F. The test piece was fixed with its long side vertical, and the center of the bottom of the test piece was lit from directly below using a burner. The distance from the burner to the test piece was 19 mm, and the time during which the test piece was roasted for 60 seconds and then the flame was emitted from the test piece was measured as the burning time to determine the flammability. The superiority or inferiority of the burning time was based on Advisory Circular (AC) 25.853-1a (1) i. The number of measurements was n = 3, and the arithmetic mean value of the burning time was used.

[Carbon fiber 1]
A copolymer containing polyacrylonitrile as a main component was spun, fired, and surface-oxidized to obtain continuous carbon fibers having a total number of single yarns of 12,000. The characteristics of this continuous carbon fiber were as shown below.

Single fiber diameter: 7 μm
Relative density: 1.8
Tensile strength: 4600 MPa
Tensile modulus: 220 GPa

[PEI resin]
A resin film having a ^{grain size of 63 g / m 2} made of an amorphous thermoplastic polyetherimide resin (“ULTEM” (registered trademark) 1010R manufactured by SABIC Innovative Plastics Co., Ltd.) was used. The characteristics of the obtained resin sheet are shown in Table 1.

[Phenol resin]
A liquid novolak type phenol resin (“PHENOLITE” (registered trademark) 5592 manufactured by DIC Corporation) was used.

[PP resin]
100g of unmodified polypropylene resin ("Prime Polypro" (registered trademark) J105G manufactured by Prime Polymer Co., Ltd.) and 20% by mass of acid-modified polypropylene resin ("Admer" QB510 manufactured by Mitsui Chemicals, Inc.) A sheet of / m ^{2 was prepared.} The characteristics of the obtained resin sheet are shown in Table 1.

[PA resin]
A resin film having a ^{basis weight of 124 g / m 2} made of nylon 6 resin (“Amilan” (registered trademark) CM1021T manufactured by Toray Industries, Inc.) was produced. The characteristics of the obtained resin sheet are shown in Table 1.

[PPS resin]
A resin non-woven fabric having a ^{basis weight of 147 g / m 2} made of polyphenylene sulfide resin (“Trelina” (registered trademark) M2888 manufactured by Toray Industries, Inc.) was produced. The characteristics of the obtained resin sheet are shown in Table 1.

[Reinforced fiber mat 1]
Carbon fiber 1 was cut to 6 mm with a cartridge cutter to obtain chopped carbon fiber. A dispersion liquid having a concentration of 0.1% by mass consisting of water and a surfactant (polyoxyethylene lauryl ether (trade name) manufactured by Nakaraitex Co., Ltd.) was prepared, and this dispersion liquid and chopped carbon fiber were used. A reinforcing fiber mat was manufactured using the reinforcing fiber mat manufacturing apparatus shown in FIG. The manufacturing apparatus shown in FIG. 6 is provided with a cylindrical container having a diameter of 1000 mm having an opening cock at the bottom of the container as a dispersion tank, and a linear transport portion (inclination angle of 30 °) connecting the dispersion tank and the paper making tank. There is. A stirrer is attached to the opening on the upper surface of the dispersion tank, and chopped carbon fibers and a dispersion liquid (dispersion medium) can be charged through the opening. The papermaking tank is a tank provided with a mesh conveyor having a papermaking surface having a width of 500 mm at the bottom, and a conveyor capable of transporting a carbon fiber base material (papermaking base material) is connected to the mesh conveyor. Papermaking was carried out with the carbon fiber concentration in the dispersion liquid being 0.05% by mass. The paper-made carbon fiber base material was dried in a drying oven at 200 ° C. for 30 minutes to obtain a reinforcing fiber mat 1. The basis weight obtained was 50 g / m ² . The characteristics of the obtained reinforcing fiber mat 1 are shown in Table 2.

[Reinforced fiber mat 2]
A reinforcing fiber mat 2 was obtained in the same manner as the reinforcing fiber mat 1 except that the carbon fiber 1 was cut to 0.5 mm with a cartridge cutter to obtain chopped carbon fibers. The characteristics of the obtained reinforcing fiber mat 2 are shown in Table 2.

(Example 1)
Reinforced fiber mat 1 as a reinforcing fiber mat, PEI resin as a resin sheet, [resin sheet / resin sheet / reinforcing fiber mat / resin sheet / resin sheet / reinforcing fiber mat / resin sheet / resin sheet / reinforcing fiber mat / resin sheet / A laminate in which the resin sheet / reinforcing fiber mat / resin sheet / resin sheet] was arranged in this order was produced. Next, an aircraft structure was obtained through the following steps (I) to (V). In the obtained aircraft structure, gaps with reinforcing fibers as columnar supports were confirmed by cross-sectional observation. The characteristics of the obtained aircraft structure are shown in Table 3.

(I) The laminate is placed in a press molding die cavity preheated to 280 ° C. and the die is closed.
(II) Next, after holding for 120 seconds, a pressure of 5 MPa is applied to hold for another 60 seconds.
(III) After the step (II), the mold cavity is opened, a metal spacer is inserted at the end thereof, and the thickness at the time of obtaining the aircraft structure is adjusted to 1.2 mm.
(IV) After that, the mold cavity is fastened again, and the cavity temperature is cooled to 50 ° C. while maintaining the pressure.
(V) Open the mold and take out the aircraft structure.

(Example 2)
Reinforced fiber mat 1 as a reinforcing fiber mat, PEI resin as a resin sheet, [Resin sheet / Reinforced fiber mat / Resin sheet / Resin sheet / Reinforced fiber mat / Resin sheet / Reinforced fiber mat / Resin sheet / Resin sheet / Reinforced fiber mat A structure for an aircraft was obtained in the same manner as in Example 1 except that the laminates arranged in the order of [/ resin sheet] were produced. The characteristics of the obtained aircraft structure are shown in Table 3.

(Example 3)
The phenol resin was added to acetone until the mass of the phenol resin became 20% to prepare a resin mixture. The resin mixture was applied to the reinforcing fiber mat 1 until the resin mixture reached 95% of the total weight, and the mixture was heated at 50 ° C. for 8 hours to evaporate acetone to obtain an aircraft structure. The characteristics of the obtained aircraft structure are shown in Table 3.

(Example 4)
A laminate was prepared in which the reinforcing fiber mat 1 was arranged as the reinforcing fiber mat and the PEI resin was arranged as the resin sheet in the order of [resin sheet / resin sheet / reinforcing fiber mat / resin sheet / reinforcing fiber mat / resin sheet / resin sheet]. A structure for an aircraft was obtained in the same manner as in Example 1 except for the above. The characteristics of the obtained aircraft structure are shown in Table 3.

(Example 5)
Reinforced fiber mat 1 as a reinforcing fiber mat, PEI resin as a resin sheet, [resin sheet / resin sheet / reinforcing fiber mat / resin sheet / resin sheet / reinforcing fiber mat / resin sheet / resin sheet / reinforcing fiber mat / resin sheet / Reinforcing fiber mat / resin sheet / resin sheet / reinforcing fiber mat / resin sheet / resin sheet / reinforcing fiber mat / resin sheet / resin sheet] Obtained an aircraft structure. The characteristics of the obtained aircraft structure are shown in Table 3.

(Example 6)
Reinforcing fiber mat 1 was arranged as a reinforcing fiber mat, and PA resin was arranged as a resin sheet in the order of [resin sheet / reinforcing fiber mat / resin sheet / reinforcing fiber mat / resin sheet / reinforcing fiber mat / resin sheet / reinforcing fiber mat]. An aircraft structure was obtained in the same manner as in Example 1 except that a laminate was produced. The characteristics of the obtained aircraft structure are shown in Table 3.

(Example 7)
Except for producing a laminate in which reinforcing fiber mat 1 is arranged as a reinforcing fiber mat and PPS resin is arranged as a resin sheet in the order of [resin sheet / reinforcing fiber mat / resin sheet / reinforcing fiber mat / resin sheet / reinforcing fiber mat]. , An aircraft structure was obtained in the same manner as in Example 1. The characteristics of the obtained aircraft structure are shown in Table 3.

(Comparative Example 1)
Reinforced fiber mat 1 as a reinforcing fiber mat, PEI resin as a resin sheet, [resin sheet / resin sheet / reinforcing fiber mat / resin sheet / resin sheet / reinforcing fiber mat / resin sheet / resin sheet / reinforcing fiber mat / resin sheet / Resin sheet / Reinforced fiber mat / Resin sheet / Resin sheet / Resin sheet / Reinforced fiber mat / Resin sheet / Resin sheet / Reinforced fiber mat / Resin sheet / Resin sheet / Reinforced fiber mat / Resin sheet / Resin sheet / Reinforced fiber mat / A structure for an aircraft was obtained in the same manner as in Example 1 except that a laminate in which [resin sheet / resin sheet] was arranged was produced. The characteristics of the obtained aircraft structure are shown in Table 4.

(Comparative Example 2)
Reinforced fiber mat 1 as a reinforcing fiber mat, PEI resin as a resin sheet, [Reinforced fiber mat / Reinforced fiber mat / Reinforced fiber mat / Reinforced fiber mat / Reinforced fiber mat / Reinforced fiber mat / Reinforced fiber mat / Reinforced fiber mat / Reinforced Fiber mat / Reinforced fiber mat / Reinforced fiber mat / Reinforced fiber mat / Resin sheet / Reinforced fiber mat / Resin sheet / Reinforced fiber mat / Reinforced fiber mat / Reinforced fiber mat / Reinforced fiber mat / Reinforced fiber mat / Reinforced fiber mat / Reinforced An aircraft structure was obtained in the same manner as in Example 1 except that a laminate in which fiber mat / reinforcing fiber mat / reinforcing fiber mat / reinforcing fiber mat / reinforcing fiber mat / reinforcing fiber mat] was arranged in this order was produced. .. The characteristics of the obtained aircraft structure are shown in Table 4.

(Comparative Example 3)
Example 1 and Example 1 except that a laminate was prepared in which the reinforcing fiber mat 1 was arranged as the reinforcing fiber mat and the PEI resin was arranged as the resin sheet in the order of [resin sheet / resin sheet / reinforcing fiber mat / resin sheet / resin sheet]. In the same way, an aircraft structure was obtained. The characteristics of the obtained aircraft structure are shown in Table 4.

(Comparative Example 4)
Reinforcing fiber mat 2 as a reinforcing fiber mat, PEI resin as a resin sheet, [resin sheet / resin sheet / reinforcing fiber mat / resin sheet / reinforcing fiber mat / resin sheet / reinforcing fiber mat / resin sheet / resin sheet] in this order. The arranged laminate was prepared. Next, an aircraft structure was obtained in the same manner as in Example 1 except that the structure was obtained through the steps (I) to (V) in Example 1. The characteristics of the obtained aircraft structure are shown in Table 4.

(Comparative Example 5)
A laminate was prepared in which the reinforcing fiber mat 1 was arranged as the reinforcing fiber mat and the PP resin was arranged as the resin sheet in the order of [resin sheet / resin sheet / reinforcing fiber mat / resin sheet / reinforcing fiber mat / resin sheet / resin sheet]. A structure for an aircraft was obtained in the same manner as in Example 1 except for the above. The characteristics of the obtained aircraft structure are shown in Table 4.

〔Consideration〕
In this embodiment, the thickness St of the aircraft structure satisfies the conditional expression: St ≧ Lf ² · (1-cos (θf)), thereby achieving a balance between the specific flexural modulus and the absolute value of the flexural modulus. It is clear that it is excellent. It is also clear that it exhibits extremely excellent characteristics in the burning time showing flammability. It was predicted that this is not only due to the flammability of the resin, but also that the performance is particularly improved by coating the reinforcing fibers with the resin. Further, the same can be said for Example 3 in which the resin type is changed. On the other hand, in Comparative Example 1, the reinforcing fiber mat and the resin were the same as in Example 1, but the specific flexural modulus could not be satisfied due to the absence of voids. In Comparative Example 2, the volume ratio of the resin type and the voids was adjusted, but the balance between the volume ratio of the reinforcing fiber mat and the resin was poor, and not only the flexural modulus was low, but also the flammability was satisfied. could not. It is presumed that this is because the resin coating around the reinforcing fibers was not formed. In Comparative Example 3, it was difficult to obtain a desired shape. It was presumed that this was because the volume content of the reinforcing fiber mat was low. In Comparative Example 4, the reinforcing fibers used for the reinforcing fiber mat had a short length. As a result, the specific flexural modulus could not be satisfied. It is presumed that this is because the conditional expression: St ≧ Lf ² · (1-cos (θf)) could not be satisfied, so that the reinforcing fibers in the aircraft structure were unevenly biased. To. In Comparative Example 5, the resin type was changed to a resin having a lower specific density, but the combustion characteristics could not be satisfied due to the low flammability derived from the PP resin.

1 Aircraft structure 2 Resin 3 Reinforcing fiber 4 Void

Claims (9)

An aircraft structure consisting of resin, reinforcing fibers, and voids.
The volume content of the resin is in the range of 2.5% by volume or more and 85% by volume or less.
The volume content of the reinforcing fiber is in the range of 0.5% by volume or more and 55% by volume or less.
The gap is 10% by volume or more, is contained in said aircraft structure at a ratio in the range of 9 7 vol% or less,
When the length of the reinforcing fiber is Lf and the orientation angle of the reinforcing fiber in the cross-sectional direction of the aircraft structure is θf, the thickness St of the aircraft structure is the conditional expression: St ≧ Lf ² · (1-). Satisfying cos (θf))
Advisory Circular (AC) 25.853-1a (1 ) der burning time less than 30 seconds after the ignition source digestion of aircraft structure to be measured by the i is,
The reinforcing fiber is coated with a resin, and the thickness of the resin is within the range of 1 μm or more and 15 μm or less.
An aircraft structure, wherein the resin contains at least one resin selected from the group of polyetherimide resin, polyetheretherketone resin, polyimide resin, and phenol resin. The aircraft structure according to claim 1, wherein the combustion time after the ignition source digestion of the aircraft structure measured by Advisory Circular (AC) 25.853-1a (1) i is 15 seconds or less. body. When the flexural modulus of the aircraft structure is Ec and the specific gravity of the aircraft structure is ρ, the specific flexural modulus of the aircraft structure represented by ^{Ec 1/3} · ρ ^{-1 is 3 or more.} , 20 or less, and the aircraft structure according to claim 1 or 2, wherein the aircraft structure has a flexural modulus Ec of 3 GPa or more. The aircraft structure according to any one of claims 1 to 3, wherein the specific gravity ρ of the aircraft structure is 0.9 g / cm ^{3 or less.} The aircraft structure according to any one of claims 1 to 4 , wherein the reinforcing fibers are discontinuous, substantially monofilament-like, and randomly dispersed. The aircraft structure according to any one of claims 1 to 5 , wherein the orientation angle θf of the reinforcing fibers in the aircraft structure is 3 ° or more. The aircraft structure according to any one of claims 1 to 6 , wherein the mass average fiber length of the reinforcing fibers is within the range of 1 mm or more and 15 mm or less. The aircraft structure according to any one of claims 1 to 7 , wherein the reinforcing fiber is a carbon fiber. The aircraft structure according to any one of claims 1 to 8 , wherein the aircraft structure has a skin layer.

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