JP5359794B2 - Wings and manufacturing method thereof - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To sufficiently promote a reduction in weight of a turbine stationary blade by thinning the thickness of the turbine stationary blade while ensuring the rigidity of the turbine stationary blade made of ceramic composite material. <P>SOLUTION: A hollow inner layer blade 17 is integrally provided on the inner surface of an outer layer blade 15 and extends from the leading edge-side inner surface 15a of the outer layer blade 15 to the trailing edge-side inner surface 15t thereof. The rear outer surface 17d of the inner layer blade 17 is integrally joined to the rear inner surface 15d of the outer layer blade 15, and the belly-side outer surface 17v of the inner layer blade 17 is integrally joined to the belly-side inner surface 15v of the outer layer blade 15. The inner layer blade 17 is dividedly composed of a plurality of hollow blade divided members 19, 21 adjacently combined. <P>COPYRIGHT: (C)2011,JPO&amp;INPIT

Description

  The present invention relates to a blade such as a turbine blade used in a turbine or a compressor of a gas turbine engine such as a jet engine and a manufacturing method thereof.

  Turbine blades used in jet engine turbines are usually made of a heat-resistant alloy such as a nickel alloy. On the other hand, in recent years, composite materials such as ceramic matrix composite materials (CMC), which have better heat resistance and lower specific gravity than nickel alloys, have been attracting attention, and turbine blades made of ceramic matrix composite materials (made of ceramic matrix composite materials) Turbine blades) have also been developed (see Patent Document 1 and Patent Document 2). A turbine blade made of a ceramic matrix composite material is manufactured by the following method.

  That is, a mandrel having a surface shape corresponding to the inner shape of the turbine blade made of ceramic matrix composite material is used, and ceramic fibers (fiber bundles of ceramic fibers) are two-dimensionally and / or three-dimensionally along the mandrel surface. By weaving, a woven fabric formed of ceramic fibers is formed on the surface of the mandrel. Subsequently, the ceramic matrix is impregnated with the ceramic matrix based on an impregnation method such as a vapor phase impregnation method (CVI) or a liquid phase impregnation firing method (PIP). A turbine blade made of a ceramic matrix composite material can be manufactured by performing machining or the like on the impregnated textile molded body.

Japanese Patent Laid-Open No. 2001-206779 JP 2003-148105 A

  By the way, turbine blades made of heat-resistant alloys (turbine blades made of heat-resistant alloys) are manufactured by the lost wax precision casting method and have a high degree of freedom in shape, but turbine blades made of ceramic matrix composite materials have a degree of freedom in shape. Therefore, it is not easy to form a reinforcing rib or the like inside the turbine blade. For this reason, the thickness of the turbine blade made of the ceramic matrix composite material must be increased to ensure rigidity, and there is a problem that the weight reduction of the turbine blade cannot be promoted sufficiently.

  The above-mentioned problems occur not only with composite material turbine blades used in gas turbine engine turbines but also with composite material compressor blades used in gas turbine compressors.

  Accordingly, an object of the present invention is to provide a blade having a novel structure and a method for manufacturing the same, which can solve the above-described problems.

  A first feature of the present invention is a blade (composite blade made of a composite material) that is used in a turbine or a compressor of a gas turbine engine and includes a composite material composed of a reinforcing fiber and a matrix. A hollow outer layer wing (outer wing) made of a composite material comprising: a composite material comprising a reinforcing fiber and a matrix integrally provided on the inner surface of the outer wing, the leading edge side of the outer wing A hollow shape that extends from the inner surface to the inner surface of the trailing edge, the back outer surface is integrally joined to the back inner surface of the outer wing, and the vent outer surface is integrally joined to the inner surface of the outer wing. An inner layer blade (inner blade), and the inner layer blade is divided by a plurality of hollow blade dividing members (inner layer blade constituting members) combined so as to be adjacent to each other.

  In the specification and claims of the present application, the blade means a turbine blade and a compressor blade, and the turbine blade means a turbine stationary blade and a turbine blade, and is compressed. The machine blade is meant to include a compressor stationary blade and a compressor blade.

  According to the first feature, the hollow inner layer blade is integrally provided on the inner surface of the hollow outer layer blade, and is divided by a plurality of hollow blade dividing members combined so that the inner layer blades are adjacent to each other. Therefore, reinforcing ribs can be formed inside the blades by adjacent portions of the plurality of blade dividing members.

  According to a second aspect of the present invention, in the method for manufacturing a wing (composite wing) configured from the first feature, a mandrel having a surface shape corresponding to the inner surface shape of each blade dividing member is used. A wing splitting member fabric forming step (fabric splitting member fabric forming step) for forming a wing splitting member fabric molding composed of reinforcing fibers on the surface of each mandrel; and a matrix for each wing splitting member fabric molding By impregnating, each wing division member impregnation step for finishing each wing division member fabric molded body into a hollow wing division member, and during or after the impregnation step for the wing division member, each mandrel is divided into each wing division. After the completion of the removal step of removing the fabric molded body for members or each blade split member, the fabric forming step for the blade split member, the impregnation step for the blade split member, and the removal step, a plurality of the blade split members are adjacent to each other. A combination step of forming the inner layer blade composed of a plurality of the blade dividing members by combining them, and after forming the inner layer blade, forming a fabric for the outer layer blade formed of reinforcing fibers on the surface of the inner layer blade After forming the outer layer wing fabric forming step (outer layer wing fabric forming step) and the outer layer wing fabric forming step, the outer layer wing fabric molded body is impregnated with a matrix to form the outer layer wing fabric. And an impregnation step for outer layer wings for finishing the fabric molded body into the outer layer wings.

  In the specification and claims of the present application, the blade manufacturing method includes a turbine stator blade manufacturing method, a turbine rotor blade manufacturing method, a compressor stator blade manufacturing method, and a compressor rotor blade manufacturing method. It is meant to include.

  According to the second feature, the inner layer wing is formed by combining a plurality of hollow blade dividing members so as to be adjacent to each other, and the outer layer wing fabric molding is formed of reinforcing fibers on the surface of the inner layer wing. Since the body is formed, the reinforcing rib can be formed inside the wing by the adjacent portions of the plurality of wing dividing members.

  According to a third feature of the present invention, in the method for manufacturing a wing (composite wing) configured from the first feature, a mandrel having a surface shape corresponding to the inner surface shape of each wing split member is used. After the completion of the wing split member fabric forming step (the wing split member fabric forming step) for forming a wing split member fabric molded body composed of reinforcing fibers on the surface of each mandrel, and the wing split member fabric forming step A combination step of forming a plurality of the wing split member fabric molded articles by combining a plurality of the wing split member fabric molded articles, and a combination step of forming a plurality of the wing split member fabric molded articles. After completion, the outer layer wing fabric forming step (outer layer wing fabric forming step) for forming the outer layer wing fabric molded body composed of reinforcing fibers on the surface of the inner layer wing fabric molded body, and the outer layer wing fabric forming After completion of process The outer layer wing fabric molded body and the inner layer wing fabric molded body are impregnated with a matrix to finish the outer layer wing fabric molded body and the inner layer wing fabric molded body into the outer layer wing and the inner layer wing, respectively. The gist includes an impregnation step and a removal step of removing each mandrel from each divided member fabric molded body or each blade divided member during or after the impregnation step.

  According to the third feature, a plurality of hollow fabric molded bodies for blade dividing members are combined so as to be adjacent to each other, thereby forming the fabric molded body for the inner layer blade, and on the surface of the fabric molded body for the inner layer blade. Since the fabric molded body for outer layer wings composed of reinforcing fibers is formed, reinforcing ribs can be formed inside the wings by adjacent portions of the plurality of wing dividing members.

  According to the present invention, since the reinforcing rib can be formed inside the wing made of composite material by the adjacent portions of the plurality of wing dividing members, the rigidity of the wing made of composite material is ensured, and The wall thickness can be reduced to sufficiently promote the weight reduction of the wing.

It is an expanded sectional view along the II line in FIG. It is a side view of the turbine stator segment concerning a 1st embodiment of the present invention, and is partially sectioned. 3 (a) and 3 (b) are diagrams for explaining a blade forming member fabric forming process in the turbine vane manufacturing method according to the second and third embodiments of the present invention. 4 (a) and 4 (b) are diagrams for explaining the impregnation step for blade dividing members in the method for manufacturing a turbine stationary blade according to the second embodiment of the present invention. FIGS. 5A and 5B are views for explaining a removal step in the method for manufacturing a turbine stationary blade according to the second embodiment of the present invention. FIG. 6A is a diagram for explaining a combination process in the method for manufacturing a turbine stator blade according to the second embodiment of the present invention, and FIG. 6B is a diagram of the turbine stator blade according to the second embodiment of the present invention. It is a figure explaining the textile formation process for outer layer wings in a manufacturing method. FIG. 7 is a diagram for explaining an impregnation step for outer layer blades in the method for manufacturing a turbine stationary blade according to the second embodiment of the present invention. FIG. 8A is a diagram for explaining a combination process in the method for manufacturing a turbine vane according to the third embodiment of the present invention, and FIG. 8B is a diagram of the turbine vane according to the third embodiment of the present invention. It is a figure explaining the textile formation process for outer layer wings in a manufacturing method. FIG. 9A is a diagram for explaining an impregnation step in the method for manufacturing a turbine stator blade according to the third embodiment of the present invention, and FIG. 9B is a diagram of the turbine stator blade according to the third embodiment of the present invention. It is a figure explaining the removal process in a manufacturing method.

(First embodiment)
A first embodiment of the present invention will be described with reference to FIGS. In the drawings, “FF” indicates the forward direction (upstream direction), “FR” indicates the backward direction (downstream direction), “In” indicates the radially inner side, and “Out” indicates the radially outer side. .

  As shown in FIGS. 1 and 2, a turbine stator segment (turbine nozzle segment) 1 according to a first embodiment of the present invention is a turbine stator (turbine nozzle, not shown) used in a turbine (not shown) of a jet engine. Is divided in the circumferential direction.

  The turbine stator segment 1 includes a plurality of turbine vanes 3 (only one is shown). Each turbine vane 3 includes a ceramic matrix (an example of a matrix) and ceramic fibers (an example of reinforcing fibers). A ceramic matrix composite material (an example of a composite material) is used as a constituent material. The details of the configuration of the turbine stationary blade 3 will be described later.

  An arc-shaped outer band 5 is integrally connected to outer end portions (radially outer end portions) of the plurality of turbine vanes 3, and the outer band 5 is made of nickel alloy or the like. It is made of a heat-resistant alloy. An arc-shaped front claw 7 that can be locked to a part of a turbine case (not shown) of the turbine is formed on the front side of the outer band 5, and a part of the turbine case is formed on the rear side of the outer band 5. An arcuate rear claw 9 that can be locked to the rear is formed. The outer band 5 may be made of a composite material such as a ceramic matrix composite material or a carbon matrix composite material instead of the heat resistant alloy.

  An arc-shaped inner band 11 is provided so as to be integrally connected to inner end portions (radially inner end portions) of the plurality of turbine vanes 3, and the inner band 11 is made of a nickel alloy or the like. It is made of a heat-resistant alloy. Further, an arcuate protruding piece 13 that can be fitted into a fitting groove (not shown) of a stator support member (not shown) integrally connected to the turbine case of the turbine is formed on the inner peripheral surface of the inner band 11. Has been. The inner band 11 may be made of a composite material such as a ceramic-based composite material or a carbon-based composite material instead of being made of a heat-resistant alloy.

  Then, the detail of a structure of the turbine stationary blade 3 which concerns on 1st Embodiment is demonstrated.

  The turbine vane 3 includes a hollow outer layer blade (outer blade) 15, and the outer layer blade 15 is made of a ceramic matrix composite material including ceramic fibers and a ceramic matrix. Instead of using ceramic fibers as the constituent material of the outer wing 15, carbon fibers, glass fibers, or mixed fibers thereof (mixed fibers of at least two of ceramic fibers, carbon fibers, and glass fibers) are used. Alternatively, instead of using a ceramic matrix as a matrix as a constituent material of the outer layer blade 15, a carbon matrix or a glass matrix may be used.

  A hollow inner layer blade (inner blade) 17 is integrally provided on the inner surface of the outer layer blade 15, and the inner layer blade 17 is made of a ceramic matrix composite material composed of ceramic fibers and a ceramic matrix. . The inner layer wing 17 extends from the front edge side inner surface 15a of the outer layer wing 15 to the rear edge side inner surface 15t. The back side outer surface 17d of the inner layer wing 17 is integrally joined to the back side inner surface 15d of the outer layer wing 15. The ventral outer surface 17v of the inner wing 17 is integrally joined to the ventral inner surface 15v of the outer wing 15. In addition, instead of using a fiber that is a constituent material of the inner layer blade 17 as a ceramic fiber, carbon fiber, glass fiber, or a mixed fiber thereof, or instead of using a matrix that is a constituent material of the inner layer blade 17 as a ceramic matrix, A carbon matrix or a glass matrix may be used.

  The inner layer blade 17 is configured by being divided by a pair of hollow (two) blade dividing members (inner layer blade constituting members) 19 and 21 combined so as to be adjacent to each other. The number of blade dividing members 19 and 21 is not limited to two, and may be any number.

  A plurality of injection holes 23 for injecting cooling air introduced through a front insert (not shown) into one blade dividing member 19 are formed in the front edge 3a and the abdominal surface 3v of the turbine vane 3. Each injection hole 23 continuously penetrates the outer layer blade 15 and one blade dividing member 19. In addition, a plurality of discharge holes 25 are formed in the rear edge 3t of the turbine stationary blade 3 to discharge the cooling air introduced into the other blade dividing member 21 via a rear insert (not shown). The discharge hole 25 continuously penetrates the outer layer blade 15 and the other blade dividing member 21. The cooling air is compressed air extracted from a jet engine compressor (not shown).

  Then, the effect | action and effect of 1st Embodiment of this invention are demonstrated.

  A hollow inner layer blade 17 is integrally provided on the inner surface of the hollow outer layer blade 15 and is divided by a pair of hollow blade dividing members 19 and 21 combined so that the inner layer blades 17 are adjacent to each other. Therefore, the reinforcing rib R can be easily formed inside the turbine stationary blade 3 made of the ceramic matrix composite material by the adjacent portions of the pair of blade dividing members 19 and 21.

  Therefore, according to the first embodiment of the present invention, the rigidity of the turbine vane 3 made of the ceramic matrix composite material is secured, and the thickness of the turbine vane 3 is reduced to reduce the weight of the turbine vane 3. Can be promoted sufficiently.

(Second Embodiment)
A method for manufacturing a turbine vane according to a second embodiment of the present invention will be described with reference to FIGS. 3 (a) and 3 (b) to FIGS. 7 (a) and 7 (b).

  The method for manufacturing a turbine vane according to the second embodiment of the present invention is a method for manufacturing the turbine vane 3 according to the first embodiment of the present invention, and includes a blade forming member fabric forming step (for the blade dividing member). A woven fabric forming step), a blade impregnating member impregnation step, a removal step, a combining step, an outer layer wing fabric forming step (outer layer wing fabric forming step), an outer layer wing impregnation step, and a machining step. And the concrete content of each process in the manufacturing method of the turbine vane concerning a 2nd embodiment of the present invention is as follows.

(2-1) Fabric formation process for blade dividing members As shown in FIGS. 3 (a) and 3 (b), mandrels 27 and 29 having surface shapes corresponding to the inner surface shapes of the blade dividing members 19 and 21 are used. Ceramic fibers (fiber bundles of ceramic fibers) are woven two-dimensionally and / or three-dimensionally along the surface of each mandrel 27, 29 by blade weaving or plain weaving. As a result, the wing split member fabric molded bodies 19F and 21F formed of ceramic reinforcing fibers are formed on the surfaces of the mandrels 27 and 29, respectively. The method of weaving the ceramic fibers can be changed as appropriate. Instead of weaving the ceramic fibers, the wing splitting member is formed by winding a fabric (not shown) made of ceramic fibers around the surfaces of the mandrels 27 and 29. The woven fabric molded bodies 19F and 21F may be molded.

(2-2) Impregnation step for blade split member As shown in FIGS. 4 (a) and 4 (b), based on the vapor phase impregnation method (CVI method), the liquid phase impregnation firing method (PIP method), and the solid phase impregnation method Then, the woven fabric molded bodies 19F and 21F for the blade dividing members are impregnated with a ceramic matrix. Thereby, each woven fabric molded body 19F, 21F for blade dividing members can be finished into the blade dividing members 19, 21. The ceramic matrix impregnation method can be changed as appropriate.

(2-3) Removal step During or after the impregnation step for blade dividing members, each mandrel 27, 29 is moved in the left-right direction in FIG. The mandrels 27 and 29 are removed from the wing split member fabric molded articles 19F and 21F or the wing split members 19 and 21, respectively.

(2-4) Combination Step After the completion of the wing split member fabric forming step, the wing split member impregnation step, and the removal step, as shown in FIG. By setting the combination jig 31 at a predetermined position, the pair of blade dividing members 19 and 21 are combined so as to be adjacent to each other. Thereby, the inner layer blade | wing 17 which consists of a pair of blade | wing division | segmentation members 19 and 21 can be formed.

(2-5) Fabrication process for outer layer wings After the completion of the combination process, as shown in FIG. 6B, ceramic fibers (fiber bundles of ceramic fibers) are laid along the surface of the inner layer wings 17 by blade weaving or plain weaving. 2D and / or 3D. Thereby, the outer-layer wing fabric molded body 15 </ b> F composed of the reinforcing fibers can be formed on the surface of the inner-layer wing 17. The method of weaving the ceramic fibers can be changed as appropriate. Instead of weaving the ceramic fibers, the fabric of the outer layer blades is formed by winding a fabric made of ceramic fibers (not shown) around the surface of the inner layer blades 17. The body 15F may be molded.

(2-6) Outer wing impregnation step After the outer wing fabric forming step, as shown in FIG. 7 (a), a gas phase impregnation method (CVI method), a liquid phase impregnation firing method (PIP method), and Based on the solid phase impregnation method, the outer wing fabric molded body 15F is impregnated with a ceramic matrix. Thereby, the outer-layer wing fabric molded body 15 </ b> F can be finished into the outer-layer wing 15. The ceramic matrix impregnation method can be changed as appropriate.

(2-7) Machining Step After the outer blade impregnation step, a plurality of injection holes 23 and a plurality of discharge holes 25 are formed by machining. In addition, after forming the several injection hole 23 and the discharge hole 25, it is desirable to perform the coating process on the surface of the outer-layer blade 15 and the inner-layer blade 17 by an appropriate impregnation method.

  As described above, the turbine stationary blade 3 made of a ceramic matrix composite material can be manufactured.

  Then, the effect | action and effect of 2nd Embodiment are demonstrated.

  An inner layer blade 17 is formed by combining a pair of hollow blade dividing members 19 and 21 so as to be adjacent to each other, and a fabric formed body 15F for outer layer blades composed of ceramic reinforcing fibers is formed on the surface of the inner layer blade 17. Therefore, the reinforcing rib R (see FIG. 1) can be formed inside the turbine stationary blade 3 made of the ceramic matrix composite material by the adjacent portions of the pair of blade dividing members 19 and 21.

  Therefore, according to 2nd Embodiment of this invention, there exists an effect similar to 1st Embodiment of this invention.

(Third embodiment)
A third embodiment of the present invention will be described with reference to FIGS. 3A and 3B, FIGS. 8A and 8B, and FIGS. 9A and 9B.

  The turbine stator blade manufacturing method according to the third embodiment of the present invention is similar to the turbine stator blade manufacturing method according to the second embodiment of the present invention. A method of manufacturing, a wing split member fabric forming step (wing split member fabric forming step), a combination step, an outer layer wing fabric forming step (outer layer wing fabric forming step), an impregnation step, a removal step, and a machining process. It has a process. And the concrete content of each process in the manufacturing method of the turbine vane concerning a 3rd embodiment of the present invention is as follows.

(3-1) Fabric Formation Step for Blade Splitting Member As shown in FIGS. 3 (a) and 3 (b), (2-1) Fabric formation step for blade segmentation member in the method of manufacturing a turbine vane according to the second embodiment. By executing the same process as the above, the wing split member fabric molded bodies 19F and 21F made of ceramic reinforcing fibers are formed on the surfaces of the mandrels 27 and 29, respectively.

(3-2) Combining Step After the blade forming member fabric forming step, as shown in FIG. 8 (a), by bringing the pair of mandrels 27 and 29 closer, a pair of blade forming member fabric molded body 19F , 21F are adjacently combined. Thereby, the fabric forming body 17F for inner layer blades which consists of a pair of fabric forming bodies 19F and 21F for blade division members can be formed.

(3-3) Outer wing fabric formation process After the combination step is completed, as shown in FIG. 7 (b), (2-5) Outer wing fabric formation in the turbine vane manufacturing method according to the second embodiment By performing the same process as the process, the outer layer wing fabric molded body 15F composed of reinforcing fibers is formed on the surface of the inner layer wing fabric molded body 17F.

(3-4) Impregnation step After the outer wing fabric forming step is completed, as shown in FIG. 9 (a), the vapor phase impregnation method (CVI method), the liquid phase impregnation firing method (PIP method), and the solid phase impregnation. Based on the method, a ceramic matrix is impregnated in the outer layer wing fabric molded body 15F and the inner layer wing fabric molded body 17F. Accordingly, the outer layer wing fabric molded body 15F and the inner layer wing fabric molded body 17F can be finished into the outer layer wing 15 and the inner layer wing 17, respectively.

(3-5) Removal Step After the impregnation step or after completion, the same process as the (2-3) removal step in the method for manufacturing a turbine vane according to the second embodiment is performed, so that FIG. As shown in FIG. 4, the mandrels 27 and 29 are removed from the blade divided member molded bodies 19F and 21F or the blade divided members 19 and 21, respectively.

(3-6) Machining process After the removal process is completed, a plurality of injection holes 23 are obtained by executing the same process as the (2-7) machining process in the method for manufacturing a turbine vane according to the second embodiment. And a plurality of discharge holes 25 are formed.

  As described above, the turbine stationary blade 3 made of a ceramic matrix composite material can be manufactured.

  Then, the effect | action and effect of 3rd Embodiment of this invention are demonstrated.

  A pair of hollow wing split member fabric molded bodies 19F, 21F are combined so as to be adjacent to each other to form an inner wing fabric molded body 17F, and the surface of the inner wing fabric molded body 17F is composed of ceramic fibers. Since the outer wing fabric molded body 15F is formed, a reinforcing rib R (see FIG. 1) is formed inside the turbine stationary blade 3 made of a ceramic matrix composite material by adjacent portions of the plurality of blade dividing members 19 and 21. can do.

  Therefore, according to the third embodiment of the present invention, the same effects as those of the first embodiment of the present invention are achieved.

  The present invention is not limited to the description of the above-described embodiment. For example, the configuration of the turbine stationary blade 3 according to the first embodiment of the present invention is the same as that of the turbine rotor blade (not shown), the compressor stationary blade ( (Not shown), or applied to a compressor rotor blade, or the configuration of a turbine stator blade manufacturing method according to the second and third embodiments of the present invention is a turbine rotor blade manufacturing method, compressor stator blade manufacturing method, or The present invention can be implemented in various other modes such as applying to a method of manufacturing a compressor blade. Further, the scope of rights encompassed by the present invention is not limited to these embodiments.

DESCRIPTION OF SYMBOLS 1 Turbine stator segment 3 Turbine stationary blade 3a Front edge of turbine stationary blade 3t Rear edge of turbine stationary blade 3v Abdominal surface of turbine stationary blade 5 Outer band 11 Inner band 15 Outer blade 15F Fabric molded body 15a for outer layer blade Front edge side of outer layer blade Inner surface 15d Outer layer blade back side inner surface 15t Outer layer blade trailing edge side inner surface 15v Outer layer blade abdomen inner surface 17 Inner layer blade 17F Inner layer wing fabric molded body 17d Inner layer blade back side outer surface 17v Inner layer blade abdomen outer surface 19 Blade splitting member 19F Fabric formed body 21 for blade dividing member 21 Blade formed member 21F Fabric molded body for blade divided member 23 Injection hole 25 Discharge hole 27 Mandrel 29 Mandrel 31 Combination jig R Reinforcing rib

Claims (5)

In a blade used in a turbine or compressor of a gas turbine engine and made of a composite material composed of reinforcing fibers and a matrix,
A hollow outer wing composed of a composite material composed of reinforcing fibers and a matrix;
The outer wing is provided integrally with the inner surface of the outer wing, is composed of a composite material composed of a reinforcing fiber and a matrix, extends from the inner surface of the outer edge of the outer wing to the inner surface of the rear edge, and the back outer surface of the outer wing of the outer wing. A hollow inner layer wing integrally joined to the back side inner surface, and a ventral outer surface integrally joined to the abdominal side inner surface of the outer layer wing,
The inner layer blade is divided by a plurality of hollow blade dividing members combined so as to be adjacent to each other.   The wing according to claim 1, wherein the composite material is a ceramic matrix composite material or a carbon-based composite material. The method of manufacturing a wing according to claim 1,
Using a mandrel having a surface shape corresponding to the inner surface shape of each wing split member, forming a woven fabric for a wing split member formed of reinforcing fibers on the surface of each mandrel;
Impregnation step for wing splitting member to finish each wing splitting member fabric molding into a hollow wing splitting member by impregnating each wing splitting member fabric molding with a matrix;
In the middle or after the impregnation step for the blade split member, a removal step of removing each mandrel from each blade split member molded body or each blade split member;
After the wing split member fabric forming step, the wing split member impregnation step, and the removal step are finished, the plurality of wing split members are combined so as to be adjacent to each other, thereby forming the wing split member. A combined process of forming inner wings;
After the combination step, the outer layer wing fabric forming step for forming the outer layer wing fabric molded body composed of reinforcing fibers on the surface of the inner layer wing,
An outer layer wing impregnation step of finishing the outer layer wing fabric molded body into the outer layer wing by impregnating the outer layer wing fabric molded body with a matrix after completion of the outer layer wing fabric formation step. A method of manufacturing a wing characterized by the above. The method of manufacturing a wing according to claim 1,
Using a mandrel having a surface shape corresponding to the inner surface shape of each wing split member, forming a woven fabric for a wing split member formed of reinforcing fibers on the surface of each mandrel;
After completion of the wing split member fabric forming step, a plurality of the wing split member fabric molded bodies are combined so as to be adjacent to each other, thereby forming a plurality of the wing split member fabric molded bodies. A combined process to form
After the combination step, an outer layer wing fabric forming step of forming an outer layer wing fabric formed of reinforcing fibers on the surface of the inner layer wing fabric molded body,
After the outer layer wing fabric forming step is completed, the outer layer wing fabric molded body and the inner layer wing fabric molded body are impregnated with a matrix to obtain the outer layer wing fabric molded body and the inner layer wing fabric molded body. An impregnation step for finishing the outer layer blade and the inner layer blade, respectively;
A wing manufacturing method comprising: a step of removing each mandrel from each wing split member molded body or each wing split member during or after the impregnation step.   The wing manufacturing method according to claim 3 or 4, wherein the composite material is a ceramic matrix composite material or a carbon matrix composite material.