JPS584679B2 - Method for manufacturing composite material wings - Google Patents

Abstract

PURPOSE:To provide a process for producing an aerodynamic wing using a composite material by a metal mold having molding inner surfaces to mold once by eliminating the conventional pressure steps more than twice required to obtian uniform product.

Description

[Detailed description of the invention] The present invention relates to a method of manufacturing a composite material wing.

The method according to the invention is most advantageously used for manufacturing helicopter main rotors, tail steering rotors, aircraft propeller blades, gas turbine engine rotor blades and blower blades, etc., but also general fixed wing aircraft wings. It is also applicable to

Conventionally, these wings were made of composite materials, such as glass fiber, carbon fiber, boron fiber, and organic fiber (Kevlar 49, etc.).
Several methods have been proposed and known in which metal fibers are impregnated with binders such as epoxy resins, polyester resins, and polyimide resins to produce materials with various properties.

For example, U.S. Pat. No. 3,219,123;
No. 7,697 and No. 3,782,856 disclose methods for manufacturing composite wings that include several bonding steps, but these known methods are complicated, labor-intensive, and This is not preferred because it requires repeated heat curing and cooling.

From this point of view, Japanese Patent Publication No. 47-10946 proposes a method of completing the heat curing step in one step.

In this method, the rigid foam material, which has been given a larger predetermined shape, is sandwiched between the half-split shells and bonded while being molded into the predetermined shape using a press mold. However, since the heat-resistant limit temperature of normal thermoplastic resin foam is around 100°C, the upper limit of the molding temperature for composite wings is also around 100°C, and it usually takes about 100°C to complete curing. This resin has disadvantages in that it takes a very long time, has poor productivity, and increases manufacturing costs.

In addition, with this method, only the pressure due to plastic deformation of the thermoplastic resin foam material can be expected as the molding pressure, so the molding pressure is low and bubbles cannot be removed sufficiently, resulting in uneven products and problems with strength. It also has drawbacks.

A thin member such as the outer skin that is in direct contact with the mold is heated early with this method, and the temperature rises sufficiently to cause a curing reaction while the foam core is still at a low temperature and high pressure. Satisfactory results can be obtained with hardening molding, but for parts such as the main spar, which are thick, have a high specific heat, and have a very slow hardening reaction, this method requires that the entire blade be heated to a sufficiently high temperature before applying high pressure. It has the disadvantage of not being able to do anything.

Further, this method of using a foam core has the disadvantage that the main spar geometry is limited to a C-shape as the wing cross-sectional shape, resulting in low torsional rigidity.

Furthermore, the method of compressing by gradually closing the mold while heating the mold has the problem that the glass cloth of the main girder and outer skin protrudes from the mating surface of the mold due to pressure and gets caught, causing dimensional defects. be.

The method disclosed in JP-A-50-16298 solves the above problem by introducing pressure into the internal elastic bag after the mold is closed and pressurizing from inside. Pressure can be applied independently regardless of temperature or size, making it suitable for producing uniform composite materials, but this method is difficult to perform in one step using one mold. be.

In other words, the high pressure supplied to the elastic bag is applied not only to the main girder member but also to the rear core member, and there is a risk of destroying the core member, so the elastic bag is made in a separate mold only for the main girder. There is no choice but to use a method in which the rear assembly, which is processed separately, is then adhesively bonded to the main girder.

Needless to say, such a two-stage process causes problems related to adhesion pretreatment, thermal deformation, and the like.

U.S. Pat. No. 3,713,753 shows a method in which heat curing is completed in one step while applying pressure from inside the product in a mold.

However, while this method applies sufficient pressure to the main spar fibers wrapped directly around the pressurized bag, only partial pressure is applied to the outer skin.

That is, the main spar section pressurized by the pressurizing bag tends to expand into a perfect circular shape, and this main spar section tends to move toward the center of the airfoil section.

This main spar is pressed against the leading edge with a small radius of curvature, and the chordwise position of the main spar is accurately determined.The force that presses the leading edge skin is the reaction force caused by the deformation of the core material inserted at the rear. produced only by.

However, as mentioned earlier, the pressure in the chord direction due to the deformation of the core material is weak at high temperatures, so it is difficult to prevent the main spar from moving toward the center of the airfoil. However, this is disadvantageous because it requires low temperature and long time molding conditions, resulting in an increase in manufacturing costs.

In addition, normally for rotary blades, etc., the center of gravity in the chord direction is 1 of the chord length.
Since the characteristics change significantly even if the main spar is moved by a fraction of a percent, the inability to accurately determine the chordwise position of the main spar using this method is a definite drawback.

In view of the above-mentioned prior art, the present invention provides a method for manufacturing a composite material wing that eliminates many of these drawbacks. The belt-shaped member is placed between upper and lower outer skin members in a mold in a state of being wrapped by a member, and both side edges of the band-like member are fixed to the mold, and heat curing is progressed while supplying pressure to the inside of the hollow expansion member. It is in.

That is, in the method of the present invention, the pressurizing hollow expansion member and the main girder member are wrapped in a flexible band-like member, and both side edges of the band-like member are fixed to the mold, so that no leakage from the inside can occur. During pressurization, there is no risk of displacement of the main girder member, and sufficient pressure can be applied to the main girder member.

Furthermore, by arranging the core member at the trailing edge of the blade outside the flexible band member, the force acting on the core member can be kept within a certain limit and its plastic deformation can be prevented from becoming excessive.

Furthermore, the main spar member is formed of resin-impregnated fibers extending in the longitudinal direction of the wing, and the flexible band member is formed at an angle of 90° to the longitudinal direction of the wing.
By forming the resin-impregnated fibers with an angle of 45°, it is possible to maintain sufficient strength in terms of centrifugal force resistance, bending load resistance, and torsional rigidity.

The present invention will be explained in detail below with reference to embodiments shown in the drawings.

FIG. 1 shows a composite material wing to which the present invention is applied.

Although the wing is illustrated as a helicopter main rotor, it is of course equally applicable to other similar applications.

A plurality of rotor blades 1 are mounted to extend radially outward from a rotating rotor hub to generate buoyancy.

The rotor 1 includes a fitting 2 for attachment to the rotor hub, a root 3, a wing tip 4, a longitudinal wing axis 5, a leading edge 6, and a trailing edge 7.

FIG. 2 shows a typical cross section 1--2 of the rotor blade 1 shown in FIG. 1, and illustrates one example manufactured by the method of the present invention.

The rotor blade 1 includes a wear-resistant leading edge cover 8, a blade upper surface skin 9a, a blade lower surface skin 9b, a trailing edge reinforcing member 10, a core 11, a stiffening upper surface material 12a in the flapping direction, a lower surface material 12b, a flexible strip member 13, It includes a main beam member 14, a leading edge counterweight 15, and a foam adhesive 16 disposed as required.

The main girder 14 is made of a synthetic resin reinforced with suitable fibers, such as glass fibers, and for this purpose, as shown in FIG. Bath 1 containing uncured synthetic resin
9 and is impregnated with this resin, and is automatically wound around a mandrel 21 in the form of a flat ring by a group of rollers 20.

Then, the resin-impregnated glass fiber bundle 22 that has been wrapped is
After improving handling properties through appropriate treatment, it is removed from the mandrel 21.

FIG. 4 shows a resin-impregnated glass fiber bundle 22 wound up in this manner, one end loop 23 of which is used as a wing attachment.

Figure 5 shows an example of the configuration of the blade attachment part. Figure a shows the configuration using one loop, and Figures b to d use two loops. In either case, the rotor blade of the rotor hub is used. Install by inserting the mounting bolt into this loop.

In addition, when using one loop shown in FIG. 5a, it is desirable to arrange and connect a metal fitting so as to wrap this loop portion.

In either case, as the resin-impregnated fiber bundle 22 extends from the end loop portion 23 in the longitudinal direction of the blade, it is pressed against the leading edge portion and formed into a lump to constitute the main spar 14 .

A filler material 24 is inserted in the intermediate part from the end loop part to the lump part, and from the loop part to the lump part, that is, the main girder part 1.
4, the structure of the main girder is extremely simple and can be manufactured uniformly and at a very low cost.

FIG. 6 shows an example of a method for manufacturing the flexible strip member 13 used in the present invention. The strip member 13 is made of synthetic resin with high elasticity and high strength fibers such as carbon fibers as reinforcement materials. .

That is, as shown in FIG. 6, the fibers 25 are passed from a reel 26 to a bath 27 containing synthetic resin, impregnated with resin, and then wound around a cylindrical mandrel 28.

The array of fibers is arranged to include circumferentially wound layers and ±45° wound layers for torsional stiffness.

The finished resin-impregnated fiber sheet 29 is cut open in the axial direction of the cylindrical mandrel by a cutter 30, and the strip member 13 is cut open in the axial direction of the cylindrical mandrel.
It becomes the material that makes up the.

It is clear that this process is also carried out automatically by simple machinery so that the product is uniform and inexpensive.

Since this band member is expanded under high pressure in the subsequent blade manufacturing process, glass fiber or carbon fiber woven fabric manufactured by ordinary methods has a disadvantage in that its strength is low.

That is, if the woven fibers have a bent part, the bending part is subjected to expansion tension due to high pressure and the load is concentrated, and glass fibers and carbon fibers are extremely susceptible to friction, resulting in a decrease in strength.

Therefore, in the present invention, the sheet 29 is manufactured so that the fibers are not bent by winding the fibers around the cylindrical mandrel 28 as described above, and its strength is improved.

However, in the case of a woven fabric made of a fiber such as Kevlar 49 fiber, which has extremely excellent properties in terms of bending and friction, the woven fabric can be used as a sufficient pressure-resistant member 13 without using the above-mentioned filament winding method. It is possible to satisfy the purpose of

Next, an example of the blade forming method according to the present invention will be explained.

First, as shown in FIG. 7, a resin-impregnated glass cloth 31 is placed on the lower mold 30.

After being heated and hardened, this will constitute the outer skin 9 of the wing, so it must have a part of the torsional rigidity of the wing and the necessary rigidity and strength to maintain the aerodynamic shape of the wing. .

The leading edge of this glass cloth 31 is precisely trimmed on the mold and placed slightly short so that it does not protrude beyond the required airfoil portion.

Further, in order to improve the ease of handling, this glass cloth 31 is subjected to a reaction in advance to be in a semi-hardened state.

Next, on this glass cloth 31, a highly elastic fiber cloth 32 impregnated with resin is placed, the fiber direction of which is aligned parallel to the longitudinal axis direction of the blade.

After being heated and cured, this high elastic fiber cloth 32 constitutes the upper and lower side members 12 for reinforcing stiffness in the flapping direction, thereby reinforcing the stiffness in the flapping direction of the composite wing.

The thickness, width, length, etc. of the upper and lower surface members 12 reinforcing stiffness in the flapping direction are determined by the stiffness requirements of the wing, so they may vary partially and may be unnecessary in some cases. Of course it is possible.

Next, as shown in FIG.
3, the core material 11, and the trailing edge reinforcing material 10 are placed.

A foam adhesive 16 is inserted between the member 13 and the core material 11.

The member 13 is placed precisely with one of its free ends in a defined position within the leading edge cavity 38 of the mold 30 and is first secured to the mold 30 by means of a retaining plate 35 and a mounting bolt 36.

The core material 11 is a hard resin foam material in this embodiment, and is made to be slightly wider than the normal size and to have a wedge-shaped cross section.

The trailing edge reinforcing material 10 is made of resin-impregnated high-strength fibers, and all of the fiber directions are oriented in the longitudinal direction of the blade.

In this embodiment, the resin may be uncured or may have already been molded in another mold.

This trailing edge reinforcing material 10 is arranged to adjust the chordwise bending frequency of the blade and to withstand bending loads in the same direction, and may be arranged as necessary. May be omitted.

Next, after placing the leading edge counterweight 15, the unhardened main girder 14 manufactured by the above method, and the hollow expansion member, that is, the pressurizing bag 33, in order from the leading edge on the belt-shaped member 13, the belt-shaped member 13 is folded over as shown in FIG. Align and fix to the mold using the holding plate 35 and mounting bolts 36.

In this way, the leading edge counterweight 15, the main spar 14, and the pressurizing bag 33 are wrapped in the band member 13.

In this case, the pressurizing bag 33 is in a contracted state and can be easily wrapped with the member 13.

Further, the pressurizing bag 33 has an appropriate wall thickness and also serves to prevent wrinkles from forming on the band member 13.

Thereafter, the upper mold 37 in which the resin-impregnated glass cloth 31 and the high elastic fiber cloth 32 are placed in advance is placed over the lower mold in the same manner as described above.

When the upper and lower molds are matched and tightened, there is a very small gap between the upper and lower molds at the engagement portion 34 at the leading edge of the airfoil, and the band member 13 is sandwiched in this gap.

The size of the gap is not so narrow as to bite the member 13, and when the member 13 is securely held and this band-shaped member 13 receives all the pressure applied from the pressure bag in the later process, the member 13 will close the gap. It must be fixed to the extent that it cannot move or fall out.

After the molds 30 and 37 are combined and tightened, a slight pressure of 1 atm or less is applied to the pressurized bag 38, and then the ninth
Turn the front edge downwards as shown and heat gradually.

By heating and pressurizing with the leading edge facing downward, the leading edge counterbalance weight 15, which has a large specific gravity, is fixed at the forefront of the leading edge, and the resin of the main girder 14 flows out, causing the center of gravity to change. This prevents problems such as the shape of the main girder 14 becoming vertically asymmetrical.

A heating means 39 is provided in the mold, and the heat supplied therefrom heats the resin-impregnated glass cloth 31, which is the outer skin, to lower the viscosity of the resin.

Simultaneously with the progress of heating, when the pressure of the pressurizing bag is gradually increased, the pressurizing bag 33 expands and presses the slightly larger core 11 toward the trailing edge via the band member 13.

Due to this pressure, the core 11 moves toward the trailing edge, comes into close contact with the resin-impregnated glass cloth 31, and is compressed.

At this time, since the resin of the glass cloth 31 has fluidity, the pressure applied by the core 11 causes the resin to enter the capillary tubes of the core 11, and the core 11 and the glass cloth 31 are sufficiently bonded.

At the same time, the trailing edge member 10 is also firmly adhered to the glass cloth 31.

As mentioned above, the glass cloth 31 has been subjected to a reaction in advance to make it easier to handle, and is in a semi-hardened state, and the heat from the mold first heats this glass cloth 31, and furthermore, the glass cloth 31 is thin. Therefore, the temperature rises quickly, the curing reaction is completed quickly, and the core 11 and the glass cloth 31 are sufficiently bonded.

If the heating is continued further, the viscosity of the resin decreases to the center of the main girder 14 and becomes a fluid state, but it takes a considerable amount of time for the main girder 14 to complete the curing reaction.

Leading edge counterweight 1 due to pressure transmitted from pressurized bag 33
5 is pressed sufficiently downward, and other gaps are filled with main girder 1.
The glass fiber and resin from step 4 are filled securely.

On the other hand, the pressure of the pressurizing bag 33 further presses the core 11, but as a result of the temperature of the hard foam material also rising, plastic deformation gradually occurs, and a portion of the core 11 is deformed as shown in FIG.

In this state, the strip member 13 is expanded and the pressurized bag 33
Since the pressure due to is no longer transmitted to the core 11,
Further deformation of the core 11 is prevented.

Because the core 11 undergoes plastic deformation at high temperatures, the pressure applied from the core 11 to the glass cloth 31 is lower than in the initial low temperature state, but originally it is not necessary to apply very high pressure to a thin member like the glass cloth 31. Furthermore, as mentioned earlier, the curing reaction of the glass cloth 31 is completed early, so even if the pressure of the core 11 is slightly reduced at high temperatures, sufficiently uniform molding can be achieved. can be done.

On the other hand, the pressure from the pressure bag 38 acts uniformly on the main girder 14, pushes out the air bubbles contained in the previous process, uniformly impregnates the fibers with resin, and causes the leading edge counterweight 15 and the main girder 14 to be uniformly impregnated with resin. Perfect the adhesion.

The pressure that should ultimately be applied to the main girder 14 is a high pressure of 10 atmospheres or more, but because the strip member 13 is fixed at the mold front edge engagement part 34 and the strip member does not expand beyond a certain level. In spite of such high molding pressure, the high pressure does not act on the core 11, and the core 11
will not be destroyed.

Moreover, since the mold is sufficiently fixed in advance, resin and fibers do not fly out from the mold mating surface of the leading edge, thereby preventing dimensional defects from occurring.

Furthermore, the member 13 tries to deform into a perfect circle due to the pressure from the pressurizing bag, but the main girder 14 and the member 13 do not move at all because they are sufficiently fixed by the engagement part 34 at the front edge of the mold. Not only does it not affect the center of gravity position,
The resin-impregnated glass cloth 31 and the high modulus fiber cloth 32 of the wing leading edge are also sufficiently pressurized.

During this time, the foaming adhesive 16 fills the gap between the core 11, the strip member 13, and the resin-impregnated glass cloth 31, and the foaming pressure applies the resin-impregnated glass cloth 31 to the parts of the resin-impregnated glass cloth 31 that are not pressurized by the core 11 or the member 13. At the same time as applying pressure, core 1
1 and member 13 are bonded together, and the external air force transmitted from the trailing edge of the wing is transmitted to the main spar 14.

After curing is completed, the pressure bag 33 can be easily removed by releasing the pressure from the pressure bag 33.

Next, the wing 1 is taken out from the mold, and the leading and trailing edges are cut to size using a diamond cutter or the like.

Conventionally, resin-impregnated fibers that make up the outer skin and main girder were often trimmed with scissors or a cutter before being heated and cured. It was impossible to cut to exact dimensions, which sometimes resulted in large variations in wing weight and properties.

In the method of the present invention, as described above, after the molding process, the product is trimmed by an automated machine, so the dimensions, weight, various properties, etc. are accurate, and costs can be reduced due to improved working conditions.

Next, another embodiment of the present invention will be described with reference to FIG.

In this example, a honeycomb core 40 is used instead of the hard foaming agent core 11 in the previous example.

One feature of the composite material blade is that, except for the wear-resistant leading edge cover 8 and the leading edge counterweight 15, all of the blades are made of non-metallic materials and are resistant to corrosion. Therefore, the honeycomb core 40 is also made of non-metallic material, such as Nomex. It is preferable that it be made of FRP or other materials.

In the present invention, exactly the same steps as in the previous example can be used.

That is, since the honeycomb core 40 is made somewhat large in size, it bites into the uncured resin-impregnated glass cloth 31 when the molds 30 and 37 are closed, and applies sufficient pressure to the glass cloth.

However, since the honeycomb core has a nearly wedge-shaped triangular shape, it produces an effect of moving toward the leading edge of the wing, which may reduce the pressure on the glass cloth 31. 16 and the belt-like member 13, pressure from the pressurizing bag 33 acts to counteract the above-mentioned tendency.

However, as long as the honeycomb core 40 remains in its proper position, the member 13 will not expand and move excessively and crush the honeycomb core 40.

This is because the band member 13 is fixed at the mold front edge engaging portion 34 and will not expand beyond a certain level.

Therefore, according to the method of the present invention, using a honeycomb core,
The resin-impregnated glass cloth 31 for the outer skin and the honeycomb core 40 are sufficiently bonded, and sufficient molding pressure is applied to the main girder 14, and moreover, no more than necessary pressure is applied to the honeycomb core 40, so that it is extremely uniform. Therefore, it is possible to provide a rotor blade that is sufficiently adhesively bonded.

The most important advantage is that the above-mentioned molding process can be completed in one step, which reduces manufacturing costs and prevents dimensional errors and poor adhesion.

Further, unlike the previous example, since a honeycomb core with good heat resistance is used, the curing temperature can be raised to complete heat curing in a short time, which has the advantage of reducing manufacturing costs.

Next, a third embodiment of the present invention is shown in FIGS. 12 and 13.

This embodiment is suitable for cases where a weight for balance adjustment is installed near the blade tip 4 of the blade 1, or when an adjustment weight is installed near the center of the blade to adjust the natural frequency of the blade 1. This is a structural example, and there is essentially no difference in the mounting method of the two adjustment weights, and the illustrated embodiment shows a structural example in which the adjustment weight is provided near the wing tip.

In the figure, a balance adjustment weight 43 is disposed within a blade inertia moment adding weight 42 at the leading edge of the blade.

The rotor blades require a large moment of inertia to spin when the engine is stopped and to safely descend and land.
For this reason, an additional weight 42 is attached to the wing tip to increase the moment of inertia.

Therefore, for this additional weight 42, in addition to a corrosion-resistant alloy such as titanium alloy, stainless steel, or tungsten, a material in which tungsten powder or lead bullets are bonded with epoxy resin, rubber, or the like may be used.

Even if the position of the main spar 14 changes drastically as in these embodiments, there is almost no change in the blade manufacturing method described so far.

Attaching the additional weight 42 as described above does not require any special equipment or metal fittings, and since sufficient pressure is applied, the adhesive connection between the main girder 14 and the additional weight 42 is perfect.

For example, if the mounting portion of the additional weight 42 is molded into a cavity and the additional weight 42 is coated with adhesive and inserted into the cavity after the wing molding is completed, sufficient pressure will be applied to the adhesive. Not only does it not work, but the thickness of the adhesive is unstable due to the dimensional error between the cavity and the additional weight 42, and the reliability of the adhesive strength is significantly reduced, making it useless.

At the wing tip, the wear-resistant metal cover 8 is required to be disposed thickly and deeply, but in the present invention, instead of forming the wear-resistant metal cover 8 deeply, the wear-resistant metal plate 41 is placed as shown in FIG. can be placed in upper and lower parts and glued together at the same time when forming the wing.

By arranging it in this way, after the forming process shown in FIG. 12 is completed, both wear metal plates 41 as shown in FIG. 13,
The additional weight 42 can be mechanically coupled with rivets 44, bolts, etc., and arranged so as to ensure safety in the event that the adhesive is destroyed.

A wear-resistant metal cover 8 is bonded to cover the heads of the rivets and the like to prevent corrosion and wear.

In this way, only the leading edge, which is severely worn, can be replaced, and since a thick plate can be processed, the replacement time can be extended, which is economical.

As described above, it is now possible to complete the adhesive bonding of various rotor blade components from the blade root to the blade tip except for the wear-resistant leading edge cover 8 in one heating process, and the pressure is evenly applied. It works well enough to ensure a reliable bond.

As described above in detail with respect to the embodiments, the present invention has the following effects.

In other words, as a result of wrapping the main girder and pressurized bag in a band-shaped member and fixing the band-shaped member at the front edge, the cross-sectional area is large and thick like the main girder, which is important for strength and requires high uniformity. A high, uniform pressure is applied from the pressurized bag to the parts to be treated,
On the other hand, for parts that are thin and do not require high pressure, such as the outer skin, or parts that cannot withstand high pressure, such as hard foam cores and honeycomb cores, the molding effect is that no more pressure than necessary is applied. It is now possible to complete molding in one step, which previously required two or more heating steps.
In addition, productivity is increased by molding in one step, which allows uniform products to be obtained.As a result, one set of molds is required for mass production on the scale of conventional small aircraft, which reduces mold manufacturing costs. Because all products are manufactured using a set of molds,
The variation in quality has become extremely small.

In addition, when there were two or more heating and cooling steps in the past, there was a risk of mismatch with the mold due to curing reactions and thermal shrinkage,
Adhesion failure occurred due to variations in surface treatment, but
According to the present invention, this problem has been completely solved, and a homogeneous and safe composite blade can now be provided.

In addition, as a result of fixing the strip member at the front edge, problems such as oversize due to fibers getting caught in the mold mating surface are eliminated, the dimensions and baking soda are uniform, and the position of the main girder and the strip member is accurate. This has the advantage that the center of gravity in the chord direction does not vary, and that pressure also acts on the leading edge skin, making it possible to manufacture a uniform composite blade.

In addition, the pressure-resistant members are covered with a wear-resistant leading edge metal cover and outer skin, and are supported by the main girder, making it possible to use highly elastic and high-strength fibers such as carbon fibers, which are vulnerable to lightning strikes and local loads. The fiber direction is ±45 with respect to the longitudinal direction of the blade.
By arranging the blades at 10°, the torsional rigidity was increased and the stability of the wing was improved.

Furthermore, since the strip member covers the main girder and is D-shaped, torsional rigidity is further increased.

Because the band-shaped member is subjected to high pressure, it is preferable to manufacture it by a filament winding molding method in which the fibers do not bend, rather than using a conventional woven fabric, which reduces costs and provides uniformity through mechanization.

The pressurizing method using a pressurizing bag does not require expensive initial investment like hot press, and is suitable for prototype development.It also allows temperature and pressure conditions to be applied independently, making it possible to produce uniform and stable products. Since only the female mold needs to be manufactured, the initial investment is small and repairs, modifications, etc. are easy.

In addition, the main girders can be grouped together at the front edge and arranged in a block, so
The center of gravity moves forward, the leading edge counterweight is reduced, and the area of the main spar increases by that weight, increasing strength margin and safety.In addition, the filament winding type main spar forming method is used. It was possible to reduce manufacturing costs through mechanization and achieve uniformity.

Conventionally, it was difficult to accurately cut the outer skin and band-shaped members in an unhardened state, but with the method of the present invention, the cutting can be performed automatically with a diamond cutter after hardening, resulting in accurate dimensions and improved workmanship. This reduced costs and made it possible to obtain an accurate chordwise center of gravity position.

Furthermore, the method of the present invention has sufficient adaptability to any changes in the cross-sectional shape of the blade or to changes in the component parts, etc., and fully takes advantage of the molding freedom of composite materials, resulting in an aerodynamically sophisticated and efficient method. It has the advantage of providing wings.

Note that the method of the present invention can be modified in various ways without departing from the spirit thereof.

[Brief explanation of the drawing]

Fig. 1 is a perspective view showing the entire composite material wing manufactured by the method of the present invention, Fig. 2 is a sectional view taken along the line ■-■ of Fig. 1, and Fig. 3 shows the main spar manufactured by the filament winding method. FIG. 4 is a perspective view showing an uncured main spar material manufactured by the filament winding method; FIG. 5 is a perspective view showing an example of the root structure of a composite material wing according to the present invention; FIG. 6 is a perspective view showing an example of manufacturing the material of the band-shaped member used in the method of the present invention by the filament winding method, and FIGS. 7, 8, 9, and 10 show blade manufacturing according to the present invention. Cross-sectional diagram showing the steps of the method, No. 11
The figure is a sectional view showing an example using a honeycomb core, and FIGS. 12 and 13 are sectional views showing other examples. 11...Core, 13...Band-shaped member, 1
4...Main girder, 33...Pressure bag, 3
0...lower mold, 31...outer skin, 37...
...Upper mold.

Claims (1)

[Scope of Claims] 1. In a method for manufacturing an aerodynamic wing using an outer skin of a composite material made of resin-impregnated fibers using upper and lower molds having molding inner surfaces corresponding to almost the entire outer surface thereof, the main spar member and the hollow inflatable member are placed in a mold between the upper and lower outer skin members while being wrapped by a flexible band-like member, so that the main spar member, the hollow inflatable member, and the core member are lined up in this order from the leading edge side of the wing. A core member is placed, both side edges of the band-like member are sandwiched between the upper and lower molds, and the core member is fixed, and the resin of the outer skin is heated and hardened while supplying pressure to the inside of the hollow expansion member. A method for manufacturing a composite material wing, characterized by: 2. In the above item 1, the flexible band member is initially placed in a relaxed state in a mold, the core member is placed outside the band member, and when pressure is applied to the hollow expandable member, the flexible band member is placed in a mold. A composite material characterized in that the forerunner core member is pressed in the chord direction until the flexible band member is in tension to cause plastic deformation thereof, and then a pressurizing force is applied to the main spar. Wing manufacturing method. 3. In the above item 2, the core member has a wedge-shaped cross-sectional shape, is arranged at the trailing edge of the airfoil, and is pushed toward the trailing edge by the inflatable member via the flexible band-like member, during which the core member A method for manufacturing a composite material wing, characterized in that the wing is tightly bonded to the other material. 4. The method for manufacturing a composite material wing according to item 2, wherein the core member is formed of a foamed resin material. 5. The method for manufacturing a composite material wing according to item 1, wherein both side edges of the flexible band member are fixed to a mold in an overlapping manner at a leading edge of the wing. 6. The method for manufacturing a composite material wing according to item 1 above, characterized in that the main spar member including the resin-impregnated fibers wound into a flat ring shape is arranged so that all the fibers extend in the longitudinal direction of the wing. 7. In the above item 1, the flexible band-like member is manufactured by cutting resin-impregnated fibers wound into a cylindrical shape along a longitudinal line, and the direction of the fibers forming the flexible band-like member is: A method for manufacturing a composite material wing, comprising: a part wound at an angle of 90° to the longitudinal axis of the wing; and a part wrapped at an angle of 45° to the longitudinal axis of the wing. 8. The method for manufacturing a composite material wing according to item 1, wherein the main spar member is arranged in a block at the leading edge of the wing. 9. The method for manufacturing a composite material wing according to item 1 above, characterized in that heating and curing is carried out with the leading edge facing downward. 10. The method for manufacturing a composite material wing according to item 2 above, characterized in that the core member is manufactured to be slightly larger than the final molded size.