Composite material blade

Abstract

A composite material blade (10) comprises: a composite ply (20) that is laminated in the blade thickness direction, that has prescribed orientation angles with respect to the blade length direction, and that constitutes an airfoil part (11) of the composite material blade (10). A surface layer section (32) positioned on the surface layer side of the airfoil part (11) is divided into a first region (32A) and a second region (32B) arranged in the blade length direction. The composite ply (20) includes a first outer ply (23) laminated on the first region (32A), a second outer ply (24) laminated on the second region (32B), and an inner ply (22) laminated closer toward the center of the airfoil part as compared to the surface layer section (32). Orientation angles of the first outer ply (23) include the same value as that of the orientation angles of the inner ply (22), and orientation angles of the second outer ply (24) include a value greater than the orientation angles of the first outer ply (23).

Description

composite wing

[0001] The present disclosure relates to composite wings.

[0002] Reducing fuel consumption in jet engines is a perpetual challenge. To address this challenge, turbofan engines have increased the fan diameter to achieve a higher bypass. However, increasing the bypass size also increases the size of the fan blades, which increases the weight of the engine. For this reason, there is a need for fan blades that are both highly robust and lightweight.

[0003] A composite blade is a blade that has composite plies stacked in the blade thickness direction. Each composite ply is a sheet of carbon fiber reinforced plastic (CFRP) that uses carbon fiber as the reinforcing fiber, which gives the blade strength while reducing its weight. Due to these characteristics, there are increasing opportunities to construct fan blades with composite blades. In this regard, Patent Document 1 discloses a composite blade developed to suppress a decrease in the strength of the blade root.

[0004] Japanese Patent Application Laid-Open No. 2017-194050

[0005] When manufacturing a composite wing, composite plies with a shape that is roughly the same as the overall shape of the wing surface are often stacked. However, to provide the composite wing with resistance to localized impacts such as bird strikes, the thickness of the wing must be increased. However, if composite plies with the above-mentioned shape are stacked uniformly, the thickness of the wing will also increase in areas where impacts are unlikely to occur, resulting in an excessive increase in the overall weight of the wing.

[0006] The present disclosure has been made in consideration of the above circumstances, and aims to provide a composite wing that can ensure impact resistance while suppressing excessive weight increase.

[0007] A composite wing according to one aspect of the present disclosure comprises composite plies stacked in the thickness direction, having a predetermined orientation angle with respect to the length direction, and constituting an airfoil portion of the composite wing, wherein a surface region located on the surface side of the airfoil portion is divided into a first region and a second region arranged in the length direction, and the composite plies include a first outer ply stacked in the first region, a second outer ply stacked in the second region, and an inner ply stacked closer to the center of the airfoil portion than the surface region, wherein the orientation angle of the first outer ply has the same value as the orientation angle of the inner ply, and the orientation angle of the second outer ply has a value larger than the orientation angle of the first outer ply.

[0008] According to the present disclosure, it is possible to provide a composite wing that can ensure impact resistance while suppressing excessive increases in weight.

[0009] Figure 1 is a perspective view of a fan blade, which is an example of a composite airfoil according to an embodiment of the present disclosure. Figure 2 is a diagram illustrating an example of composite ply stacking. Figure 3 is a diagram illustrating an example of composite ply stacking and placement in an airfoil section. Figure 4A is a diagram illustrating an example of composite ply placement at the boundary between a first region and a second region. Figure 4B is a diagram illustrating an example of composite ply placement at the boundary between the first region and a second region.

[0010] A composite blade according to an embodiment of the present disclosure will be described below with reference to the drawings. Common parts in the drawings will be assigned the same reference numerals, and duplicated explanations will be omitted. For ease of explanation, a fan blade 10 will be used as an example of a composite blade according to this embodiment. The fan blade 10 constitutes a fan for an aircraft engine mounted on a turbofan engine (not shown) or the like.

[0011] Figure 1 is a perspective view of a fan blade 10. As shown in Figure 1, the fan blade 10 comprises an airfoil portion 11 and a blade root portion 12. The airfoil portion 11 has a suction side 11a, a pressure side 11b, a leading edge 11c, and a trailing edge 11d, and extends from a hub 11e to a tip 11f along the blade length direction SD. The blade root portion 12 is provided on the hub side of the airfoil portion 11 and is formed integrally with the airfoil portion 11. The blade root portion 12 fits into a mounting groove (not shown) of a rotor (not shown) to which the fan blade 10 is attached.

[0012] Various protective members (not shown) are attached to the airfoil portion 11. For example, as shown in FIG. 1, a sheath (cover) 13 is attached to the leading edge 11c with an adhesive. The sheath 13 covers and protects the leading edge 11c. In addition, a tip cap 14 is attached with an adhesive from the tip 11f to the trailing edge 11d. The tip cap 14 protects the tip 11f and the trailing edge 11d.

[0013] The fan blade 10 includes a plurality of composite plies 20. The composite plies 20 are made of fiber-reinforced plastic (FRP) formed in layers. The composite plies 20 are stacked in the blade thickness direction TD as the main structural material of the fan blade 10. The stacked composite plies 20 constitute at least the airfoil portion 11 of the fan blade 10. For example, the composite plies 20 according to this embodiment constitute the airfoil portion 11 and the blade root portion 12.

[0014] The reinforcing fibers of the fiber-reinforced resin according to this embodiment are carbon fibers. The carbon fibers are bundled into threads and uniformly aligned parallel to each other and in a direction of a predetermined azimuth angle. That is, the composite ply 20 according to this embodiment is a unidirectional carbon fiber reinforced plastic (UDCFRP). However, the reinforcing fibers are not limited to carbon fibers as long as they have mechanical strength and flexibility equivalent to those of carbon fibers.

[0015] The resin of the fiber reinforced resin according to this embodiment is a thermosetting resin or a thermoplastic resin. The thermosetting resin is an epoxy resin, a phenolic resin, or a polyimide resin. The thermoplastic resin is a polyether ether ketone or a polyphenylene sulfide. However, the components of the resin are not limited to the above substances.

[0016] In addition to the composite plies 20, fillers (not shown) are provided in the blade root 12. The fillers (not shown) are short composite layers made of the same material as the composite plies 20. The fillers are interposed between the composite plies 20, thereby forming the blade root 12 with a desired cross-sectional shape.

[0017] FIG. 2 is a diagram showing an example of lamination of composite plies 20. As shown in FIG. 2, the composite plies 20 are laminated in the blade thickness direction TD. Each composite ply 20 is defined by the orientation angle of the reinforcing fibers that constitute it. The orientation angle is the angle formed by the extension direction of the reinforcing fibers with respect to the blade length direction SD. In other words, the orientation angle of a composite ply 20 refers to the orientation angle of the reinforcing fibers that form the ply. When comparing orientation angles, their absolute values ​​are used.

[0018] The orientation angles of the laminated composite plies 20 change periodically. For example, as shown in Figure 2, composite ply 20a with an orientation angle of 0°, composite ply 20b with an orientation angle of +45°, and composite ply 20c with an orientation angle of -45° are laminated in this order. Note that the absolute values ​​of the positive and negative orientation angles may be equal.

[0019] Each composite ply 20 is formed, for example, from a prepreg tape 21. In this case, a well-known automated laying device (not shown) employing an AFP (Automated Fiber Placement) method is used to form and lay up each composite ply 20. For example, the automated laying device applies the prepreg tape 21 to the inner surface (not shown) of a mold having a shape complementary to the shape of the outermost surface 16 of the airfoil section 11 while curing it.

[0020] The prepreg tapes 21 are applied in a direction that defines the orientation angle of the composite ply 20, and are arranged in parallel within the range that will form the composite ply 20. The prepreg tapes 21 are an intermediate substrate of a composite material in which reinforcing fibers are impregnated with uncured resin. The reinforcing fibers in the prepreg tape 21 are aligned in the longitudinal direction of the tape.

[0021] When the automatic laying device lays the composite ply 20, the automatic laying device applies prepreg tape 21 onto the already formed composite ply 20 to form a new composite ply 20. The longitudinal direction (stretching direction) of the applied prepreg tape 21 defines the orientation angle of the composite ply 20.

[0022] Figure 3 is a diagram showing an example of the stacking and arrangement of the composite ply 20 in the airfoil section 11. Figures 4A and 4B are diagrams showing an example of the arrangement of the composite ply 20 at the boundary 34 (35) between the first region 32A and the second region 32B.

[0023] As shown in Figure 3, the interior of the airfoil section 11 is divided into a core layer section 31 and a surface layer section 32, which are arranged in the thickness direction TD as laminated sections of the composite ply 20. The surface layer sections 32 are set on both sides of the core layer section 31 in the thickness direction TD (i.e., on the pressure side and suction side of the airfoil section 11). The surface layer section 32 is further divided into a first region (first portion) 32A and a second region (second portion) 32B, which are arranged in the spanwise direction SD. In other words, the surface layer section 32 is divided into three regions along the spanwise direction SD.

[0024] The composite ply 20 described above is laid over each of the core region 31, the first region 32A of the surface region 32, and the second region 32B of the surface region 32. Specifically, the composite ply 20 includes an inner ply 22, a first outer ply 23, and a second outer ply 24. The inner ply 22 is laid over the core region 31, the first outer ply 23 is laid over the first region 32A of the surface region 32, and the second outer ply 24 is laid over the second region 32B of the surface region 32. Although laid over at different positions, these plies are made of the same material, and their orientation angles have values ​​that correspond to the respective regions.

[0025] The core layer region 31 is located toward the center of the airfoil portion 11 in the thickness direction TD. The core layer region 31 is distributed from the hub 11e to the tip 11f. The thickness of the core layer region 31 along the thickness direction TD accounts for 80% to 90% of the thickness of the airfoil portion 11 from the center surface 15 to the outermost surface 16 in the thickness direction TD.

[0026] As described above, the inner ply 22 is laminated on the core layer region 31. The inner ply 22 extends from the hub 11e to the tip 11f in the airfoil portion 11. The inner ply 22 accounts for the majority of the members constituting the airfoil portion 11, and forms the framework of the airfoil portion 11. The orientation angle of the inner ply 22 is set to 0°, −45°, or +45° for each ply (see FIG. 2), and plies with these different orientation angles are laminated. Plies with orientation angles of ±45° impart mechanical strength in the chord direction CD that cannot be obtained with plies with an orientation angle of 0° alone.

[0027] The surface layer region 32 is located closer to the surface of the airfoil portion 11 than the core layer region 31 in the blade thickness direction TD. In other words, the surface layer region 32 is set from the boundary 33 with the core layer region 31 to the outermost surface 16 of the airfoil portion 11. Like the core layer region 31, the surface layer region 32 is also distributed from the hub 11e to the tip 11f. The thickness of the surface layer region 32 along the blade thickness direction TD is 10% to 20% of the thickness from the center plane 15 to the outermost surface 16 of the airfoil portion 11 (i.e., half the length of the thickness of the airfoil portion 11).

[0028] The first region 32A of the surface layer section 32 is set on both sides (hub side and tip side) of the second region 32B in the blade span direction SD. That is, the first region 32A is set in the range from the hub 11e to a boundary 34 with the second region 32B, and in the range from a boundary 35 with the second region 32B to the tip 11f.

[0029] The first outer ply 23 is laminated in the first region 32A. The orientation angle of the first outer ply 23 includes the same value as the orientation angle of the inner ply 22. The order in which the first outer plies 23 with different orientation angles are laminated is the same as that of the inner ply 22. That is, the orientation angle of each first outer ply 23 is also set to 0°, −45°, or +45° (see FIG. 2 ), and plies with these different orientation angles are laminated.

[0030] The second region 32B of the surface layer region 32 is set between the first region 32A on the tip side and the first region 32A on the hub side. That is, the second region 32B is set in the range from the boundary 34 to the boundary 35. In this way, the surface layer region 32 is divided into three regions along the spanwise direction SD.

[0031] The second outer ply 24 is laminated on the second region 32B. The orientation angle of the second outer ply 24 includes a value larger than the orientation angles of the first outer ply 23 and the inner ply 22. That is, all of the second outer plies 24 may have an orientation angle larger than the orientation angles of the first outer ply 23 and the inner ply 22, or some of the second outer plies 24 may have an orientation angle larger than the orientation angles of the first outer ply 23 and the inner ply 22.

[0032] 4A, the boundary 34 (35) between the first region 32A and the second region 32B extends substantially in the chord direction CD. Meanwhile, the prepreg tapes 21 constituting the first outer ply 23 and the second outer ply 24 intersect with the chord direction CD at angles corresponding to the orientation angles of the respective plies.

[0033] When an automatic laminating device is used, the end of the prepreg tape 21 is cut in a direction perpendicular to the stretching direction (longitudinal direction) of the prepreg tape 21. Therefore, the edge 23a of the first outer ply 23 stretches in a zigzag pattern along the boundary 34 (35). The same applies to the second outer ply 24.

[0034] When the first outer ply 23 and the second outer ply 24 are arranged in the same layer, they may overlap at a boundary 34 (35) (see FIG. 4A ) or not (see FIG. 4B ). In the latter case, a gap 27 is formed between the edge 23 a of the first outer ply 23 and the edge 24 a of the second outer ply 24.

[0035] In either of the two states described above, the boundary 34 (35) between the first region 32A and the second region 32B is offset in the spanwise direction SD for each layer. That is, the overlapping region of the first outer ply 23 and the second outer ply 24 is offset in the spanwise direction SD for each layer, or the gap 27 between the first outer ply 23 and the second outer ply 24 is offset in the spanwise direction SD for each layer. For example, the above-mentioned region or gap 27 is alternately located at one of two locations in the spanwise direction SD. This prevents a local increase or decrease in the blade thickness at the boundary 34 (35). Furthermore, because the seams between the first outer ply 23 and the second outer ply 24 are offset, a decrease in strength at the boundary 34 (35) against bending can be suppressed.

[0036] Alternatively, the first outer ply 23 and the second outer ply 24 may be layered so that they alternate between overlapping and non-overlapping at the boundary 34 (35). In this case, the position of the boundary 34 (35) between each layer in the spanwise direction may be the same, or may be shifted in the spanwise direction SD for each layer. In these cases, a local increase in blade thickness at the boundary 34 (35) can also be avoided. Furthermore, because the region where the first outer ply 23, the second outer ply 24, or the first outer ply 23 and the second outer ply 24 overlap is located above the gap 27, a decrease in strength at the boundary 34 (35) against bending can be suppressed.

[0037] As shown in Figure 3, the second outer ply 24 may include a plurality of pairs of plies 26. Each pair of plies 26 is laminated in the thickness direction TD and has positive and negative orientation angles with equal absolute values. In this case, the cross elasticity effect occurring in each pair of plies 26 can be canceled. For example, uniform bending stiffness can be obtained along the chord direction CD.

[0038] As shown in FIG. 3 , the composite ply 20 may include at least one third outer ply 25. The third outer ply 25 has a constant orientation angle and extends from the hub 11e to the tip 11f of the airfoil portion 11 via the first region 32A and the second region 32B. For example, the third outer ply 25 has an orientation angle of 45° and extends from the hub 11e to the tip 11f. The third outer ply 25 is interposed, for example, between two adjacent first outer plies 23 and between two adjacent second outer plies 24. This increases the bending rigidity of the surface layer region 32 along the spanwise direction SD.

[0039] The third outer ply 25 may be provided as a ply that forms the outermost surface 16 of the airfoil section 11. The third outer ply 25 that forms the outermost surface 16 covers the boundaries 34, 35 between the first outer ply 23 and the second outer ply 24, and prevents peeling (turning up) of the edge 23 a of the first outer ply 23 (see FIG. 4B ) and the edge 24 a of the second outer ply 24 (see FIG. 4B ).

[0040] The second region 32B is located at a position relatively susceptible to localized impacts from foreign objects. Therefore, the second region 32B is located, for example, near the center (midspan) of the airfoil section 11. As shown in the lower part of Figure 3, the second outer ply 24 is provided in the second region 32B, and the orientation angles of most of the second outer ply 24 are set to values ​​larger than the orientation angles (absolute values) of the first outer plies 23 provided on both sides of the second outer ply 24 in the spanwise direction SD.

[0041] When a foreign object strikes the fan blade 10 rotating at high speed, the impact often occurs on the airfoil portion 11. In this case, a torsional moment due to the impact of the foreign object is generated in the airfoil portion 11. This torsional moment deforms the airfoil portion 11, causing delamination or cracks within the airfoil portion 11, or peeling or falling off of protective members such as the sheath 13.

[0042] Furthermore, the rigidity of the airfoil portion against a foreign object collision (i.e., a localized impact) is more likely to be affected by the fiber structure on the surface layer than by the fiber structure on the center layer. Therefore, in this embodiment, the second outer ply 24 is provided in the second region 32B. The orientation angle of most of the second outer ply 24 is larger than the orientation angles of the inner ply 22 and the first outer ply 23. In other words, the reinforcing fibers in the second region 32B are more inclined (oriented) toward a cross section perpendicular to the spanwise direction SD than the reinforcing fibers in the first region 32A. Furthermore, this cross section is approximately parallel to the main direction of the torsional moment. Therefore, deformation due to the torsional moment is suppressed in the second region 32B. This prevents delamination and peeling of the protective member, improving the impact resistance of the airfoil portion 11.

[0043] In addition, the first region 32A is laminated with a first outer ply 23 having the same orientation angle as the inner ply 22. The portion of the airfoil 11 where the first region 32A is set is a portion that is relatively less likely to be struck by foreign objects or has relatively high rigidity due to positional factors. Therefore, instead of providing a ply with an orientation angle that takes the above-mentioned torsional moment into consideration, a composite ply 20 (i.e., the first outer ply 23) with the same orientation angle as the inner ply 22 and high durability against tensile stress generated by centrifugal force is provided.

[0044] In this embodiment, the orientation angle of the composite ply 20 in the surface layer region 32 is set according to each position of the airfoil section 11 along the blade span direction SD, and the composite ply is laminated with that orientation angle. Therefore, impact resistance can be ensured without excessively increasing the blade thickness or weight.

[0045] The orientation angle of the second outer ply 24 may increase stepwise from the center side of the airfoil 11 toward the surface side. In other words, the orientation angle of the second outer ply 24 may decrease stepwise from the surface side of the airfoil 11 toward the center. For example, as shown in the lower part of FIG. 3 , the absolute value of the orientation angle is set stepwise to 40°, 50°, and 60° from the center side of the airfoil 11 toward the surface side. By gradually increasing the orientation angle, the occurrence of an extreme difference in stiffness between the surface region 32 and the core region 31 and the resulting occurrence of damage such as delamination can be suppressed. Furthermore, stress in a direction perpendicular to the fiber direction (hereinafter referred to as fiber-direction perpendicular stress) is dispersed, reducing in-plane shear stress. This suppresses the occurrence of resin cracking when a foreign object strikes the airfoil.

[0046] The minimum value of the orientation angle (absolute value) of the second outer ply 24 may be smaller than the maximum value of the orientation angle (absolute value) of the inner ply 22. In other words, the minimum value of the angle formed by the reinforcing fibers of the second outer ply 24 with the spanwise direction SD may be smaller than the maximum value of the angle formed by the reinforcing fibers of the inner ply 22 with the spanwise direction SD. In this case, for example, as shown in the lower part of Figure 3, the absolute values ​​of the orientation angles of the two plies of the second outer ply 24 that are closest to the inner ply 22 are set to, for example, 40°.

[0047] When the orientation angle of the second outer ply 24 is gradually decreased from the surface layer side toward the center of the airfoil section 11, the orientation angle of the second outer ply 24 approaches ±45°, the orientation angle of the inner ply 22, as the second outer ply 24 approaches the core layer region 31. On the other hand, when composite plies having the same orientation angle are stacked on top of each other, stress perpendicular to the fiber direction and in-plane shear stress tend to increase. Therefore, in this embodiment, in order to improve the strength in the spanwise direction SD in the surface layer region 32 while avoiding a value close to that of the inner ply 22 having an orientation angle of ±45°, the absolute value of the orientation angle of the second outer ply 24 closer to the center is set to a value smaller than 45°. Setting the orientation angle in this manner can suppress fiber breakage and resin cracking in the surface layer region 32 near the boundary with the core layer region 31.

[0048] The second outer ply 24 may be arranged so that it gradually approaches the tip 11f of the airfoil portion 11 and gradually moves away from the hub 11e of the airfoil portion 11 as it approaches the surface of the airfoil portion 11. For example, as shown in the lower part of Figure 3, the orientation angle of the second outer ply 24 gradually increases from 40° to 60°, with the larger the orientation angle, the closer the second outer ply 24 is to the tip 11f. This forms a tip-side transition portion 28. Also, the orientation angle of the second outer ply 24 gradually increases from 40° to 60°, with the larger the orientation angle, the farther the second outer ply 24 is from the hub 11e. This forms a hub-side transition portion 29.

[0049] By providing a transition portion 28 on the tip side, the first outer ply 23 including the composite ply 20 with an orientation angle of 0° and the second outer ply 24 with an orientation angle larger than that of the first outer ply 23 coexist within the same range in the blade span direction SD.

[0050] The composite ply 20 with an orientation angle of 0°, which is the first outer ply 23, suppresses deformation of the airfoil 11 at the tip side due to bending moments caused by loads in the spanwise direction SD when a foreign object strikes, thereby suppressing damage such as delamination at the tip side due to the bending moments. Meanwhile, in the transition section 28, a high proportion of the second outer plies 24 have relatively large orientation angles. The second outer plies 24 with relatively large orientation angles suppress deformation of the airfoil 11 due to the torsional moment and bending of the airfoil 11 caused by loads in the chordwise direction CD, thereby suppressing damage such as delamination originating from the impact point of a foreign object. In other words, the tip side of the airfoil 11 is prone to twisting in the chordwise direction CD due to foreign object strikes and bending of the airfoil 11 caused by loads in the chordwise direction CD, and damage due to the latter bending is particularly a concern. The second outer ply 24 with a relatively large orientation angle suppresses twisting and bending of the airfoil 11.

[0051] Similar to the tip-side transition region 28, the first outer ply 23 and the second outer ply 24 also coexist within the same range in the spanwise direction SD in the hub-side transition region 29. Therefore, also in the transition region 29, the composite ply 20 with an orientation angle of 0°, which is the first outer ply 23, suppresses deformation of the airfoil portion 11 due to the bending moment generated when a foreign object strikes, and suppresses damage such as delamination on the hub side due to the bending moment.

[0052] In the transition region 29, a high proportion of the second outer plies 24 have relatively small orientation angles among the orientation angles set for the second outer plies 24. When a bending moment due to a load in the blade length direction SD occurs on the airfoil 11, the internal stress (bending stress) caused by the moment is greater in the hub-side transition region 29 than in the tip-side transition region 28. Therefore, in the transition region 29, the proportion of the second outer plies 24 with relatively small orientation angles (i.e., those with a strong tendency to be oriented in the blade length direction SD) is increased. As a result, the second outer plies 24 in the transition region 29, together with the hub-side first outer ply 23, suppress deformation of the airfoil 11 on the hub side due to the bending moment caused by bending in the blade length direction SD during a foreign object collision.

[0053] The second region 32B may be set from the boundary 34 with the first region 32A set on the hub side to the tip 11f of the airfoil portion 11. That is, the surface layer region 32 may be divided into two regions. In this case, the range in which excessive distortion of the airfoil portion 11 curved along the chord direction CD is suppressed can be extended to the tip side. Also, the composite ply 20 in the surface layer region 32 can be laid manually without using an automatic laying device.

[0054] The present disclosure is not limited to the above-described embodiments, but is defined by the claims, and includes all modifications within the meaning and scope equivalent to the claims.

Claims

1. A composite wing comprising composite plies laminated in the thickness direction, having a predetermined orientation angle with respect to the span direction, and constituting an airfoil portion of the composite wing, wherein a surface region located on the surface side of the airfoil portion is divided into a first region and a second region arranged in the span direction, wherein the composite plies include a first outer ply laminated in the first region, a second outer ply laminated in the second region, and an inner ply laminated closer to the center of the airfoil portion than the surface region, wherein the orientation angle of the first outer ply has the same value as the orientation angle of the inner ply, and the orientation angle of the second outer ply has a value larger than the orientation angle of the first outer ply.

2. The composite blade according to claim 1, wherein the orientation angle of the second outer ply increases stepwise from the center side to the surface side of the airfoil portion.

3. The composite wing as set forth in claim 2, wherein the second outer ply is disposed so as to progressively approach the tip of the airfoil portion and progressively move away from the hub of the airfoil portion as it approaches the surface layer of the airfoil portion.

4. The composite wing according to claim 2, wherein the minimum value of the orientation angle of the second outer ply is smaller than the maximum value of the orientation angle of the inner ply.

5. The composite blade according to claim 1, wherein the second outer ply includes a pair of plies that are laminated in the blade thickness direction and have positive and negative orientation angles that are equal in absolute value.

6. The composite wing of claim 1, wherein the boundary between the first region and the second region is offset in the spanwise direction for each layer.

7. The composite wing according to claim 1, wherein the first outer ply and the second outer ply are layered so as to alternate between overlapping and non-overlapping states.

8. A composite wing according to any one of claims 1 to 7, wherein the first region is set on each side of the second region in the wing span direction.

9. A composite wing as claimed in any one of claims 1 to 7, wherein the second region is set from the boundary with the first region to the tip of the airfoil portion.

10. The composite blade of any one of claims 1 to 7, wherein the composite plies include at least one third outer ply having a constant orientation angle and extending from the hub to the tip of the airfoil portion through the first and second regions.

11. A composite wing as claimed in any one of claims 1 to 7, wherein each of the composite plies is formed from parallel aligned composite prepreg tapes stretched at the orientation angle.

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