Composite material wing stiffened wall panel and rear beam trailing edge integrated structure and manufacturing method

Abstract

The present invention discloses an integrated structure of a composite wing stiffened wall panel and a rear beam trailing edge and a manufacturing method thereof, relating to the field of aviation technology, comprising an integrally solidified wing upper wall panel skin, a wing trailing edge skin, a wall panel long stringer, and a grooved rear beam, wherein the trailing edge of the wing upper wall panel skin is provided with the wing trailing edge skin, the wall panel long stringer is provided on the lower wing surface side of the wing upper wall panel skin, long stringer vertical ribs are provided on the wall panel long stringer, a long stringer vertical rib sandwich is provided in the middle of the long stringer vertical ribs, the grooved rear beam is provided at one end of the wall panel long stringer close to the trailing edge, the grooved rear beam is located on the inner side of the wing trailing edge skin, and the trailing edge space formed by the grooved rear beam and the wing trailing edge skin is filled with a trailing edge filler. The composite material structure of the present invention has a high degree of integration, a light weight of the wing structure, can ensure that the wing-like structure has good watertightness, and ensures the stability of the performance of the wing internal equipment and the trailing edge filler.

Description

Technical Field

[0001] The present invention relates to the field of aviation technology, and in particular to an integrated structure of a composite material wing stiffened wall panel and a rear beam trailing edge and a manufacturing method thereof. Background Art

[0002] Advanced composite materials have been widely used in the aviation field due to their advantages, including high specific strength and stiffness, corrosion resistance, fatigue resistance, and customizable performance. In particular, in the past decade or so, the use of composite materials in advanced passenger aircraft structures has exceeded 50%. The use of composite materials in aircraft structures has also expanded from secondary load-bearing structures such as ailerons, flaps, elevators, and rudders to primary load-bearing structures such as wings and fuselages.

[0003] With the rapid development of aviation technology, the use of composite materials in aircraft wing structures is increasing. Aviation products are extremely weight-sensitive, and in pursuit of higher performance, new ideas and innovative structural forms are constantly being applied to reduce weight. Furthermore, aircraft operate in a variety of atmospheric environments, exposed to the test of rain and snow, placing higher demands on watertightness while reducing structural weight. There is an urgent need to leverage the advantages of composite materials and structures being formed simultaneously, to reduce weight while continuously innovating new manufacturable structural forms to enhance the watertightness of wing structures. Summary of the Invention

[0004] The purpose of the present invention is to provide an integrated structure of a composite wing stiffened wall panel and a rear beam trailing edge and a manufacturing method to solve the problems existing in the above-mentioned prior art, which can reduce the weight of composite wing-like structures, ensure that the wing-like structures have good watertight properties, and ensure the stability of the performance of the equipment inside the wing and the filling material inside the trailing edge.

[0005] To achieve the above object, the present invention provides the following solutions:

[0006] The present invention provides an integrated structure of a composite material wing stiffened wall panel and a rear beam trailing edge, comprising an upper wing wall panel skin, a wing trailing edge skin, a wall panel long stringer and a groove-shaped rear beam that are integrally solidified and formed, the wing trailing edge skin being arranged at the trailing edge of the upper wing wall panel skin, the wall panel long stringer being arranged on one side of the lower wing surface of the upper wing wall panel skin, long stringer vertical ribs being arranged on the wall panel long stringer, a long stringer vertical rib sandwich being arranged in the middle of the long stringer vertical ribs, the groove-shaped rear beam being arranged at one end of the wall panel long stringer close to the trailing edge, the groove-shaped rear beam being located on the inner side of the wing trailing edge skin, and the trailing edge space formed by the groove-shaped rear stringer and the wing trailing edge skin being filled with a trailing edge filler.

[0007] The present invention also provides a method for manufacturing the above-mentioned composite material wing stiffened wall panel and rear beam trailing edge integrated structure, comprising the following steps:

[0008] S1: using wet prepreg to lay the wing upper panel skin and the wing trailing edge upper wing skin on the surface of the first mold, and reserving prepreg required for laying the wing trailing edge lower wing skin;

[0009] S2: Laying the front long stringer cap and the long stringer rib of the panel long stringer using wet prepreg on the inner side of the wing upper panel skin laid in S1, and placing a second mold on the inner side of the front long stringer cap to assist in laying the front long stringer cap and the long stringer rib;

[0010] S3: placing the long girder reinforcement interlayer between the long girder reinforcements laid in S2;

[0011] S4: On the inner side of the wing upper panel skin laid in S1 and behind the long stringer ribs laid in S2, continue laying the rear long stringer cap of the panel long stringer and the upper wing panel of the channel-shaped rear beam, and reserve prepreg required for laying the web of the channel-shaped rear beam and the lower wing panel. Place a third mold inside the rear long stringer cap to assist in laying the rear long stringer cap;

[0012] S5: placing a grooved rear beam mold inside the upper wing panel of the grooved rear beam, and laying the reserved prepreg required for laying the web of the grooved rear beam and the lower wing panel in reverse order above the grooved rear beam mold;

[0013] S6: placing the trailing edge inner filler on the inner side of the wing trailing edge upper wing skin laid in S1 and behind the channel-shaped rear beam laid in S5, and reversely laying the reserved prepreg required for laying the wing trailing edge lower wing skin on the trailing edge inner filler and the lower wing panel of the channel-shaped rear beam;

[0014] S7: solidifying the prepreg laid out in S6, the first mold, the second mold, the third mold, the channel-shaped rear beam mold, the long stringer reinforcement sandwich layer, and the rear edge inner filler as a whole to form an integrated structure;

[0015] S8: After removing the second mold, the third mold and the grooved rear beam mold, removing the solidified part from the first mold;

[0016] S9: Grinding the edge of the solidified workpiece, applying sealant, and performing non-destructive testing to complete the preparation of the integrated structure of the composite wing stiffened wall panel and the rear beam trailing edge.

[0017] Preferably, the thickness of the prepreg laid on the wing upper panel skin and the wing trailing edge skin is 0.5 to 3 mm.

[0018] Preferably, the thickness of the prepreg laid on the front long stringer edge strip, the long stringer vertical rib, the rear long stringer edge strip and the channel-shaped rear beam is 0.5-4 mm.

[0019] Preferably, the trailing edge inner filler is made of foam or honeycomb material.

[0020] Preferably, the long stringer reinforcement interlayer is made of composite material prepreg, foam or honeycomb material, and has a laying thickness of 1 to 3 mm.

[0021] Preferably, before curing, the prepreg laid out in S6, the first mold, the second mold, the third mold, the grooved rear beam mold, the long truss vertical reinforcement sandwich and the rear edge inner filler are placed in a vacuum bag as a whole, the vacuum bag is evacuated, and then cured together with the vacuum bag.

[0022] Preferably, in step S7, curing is carried out in an autoclave, the curing temperature is 120-180° C., the curing holding time is 120-180 min, and the pressure in the autoclave is 0.6 MPa.

[0023] Compared with the prior art, the present invention has achieved the following technical effects:

[0024] The present invention provides an integrated structure and manufacturing method of a composite material wing reinforced wall panel and a rear beam trailing edge. The wing upper wall panel skin, the wing trailing edge skin, the wall panel long stringer and the grooved rear beam are manufactured in one piece without a structural separation surface, thereby reducing the mechanical connection between the grooved rear beam, the wing trailing edge and the upper wall panel structure, reducing the fasteners or adhesives used for the connection, effectively reducing the weight of the structure, and at the same time ensuring that the wing-like structure has good watertight properties. In rainy and snowy weather, the watertight properties of the entire wing structure are guaranteed, thereby ensuring the stability of the performance of the equipment inside the wing and the filler inside the trailing edge. BRIEF DESCRIPTION OF THE DRAWINGS

[0025] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the embodiments. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying any creative work.

[0026] Figure 1 The present invention provides a schematic diagram of the principle of a manufacturing method for an integrated structure of a composite wing stiffened wall panel and a rear beam trailing edge;

[0027] Figure 2 A schematic diagram of the structure of the integrated composite wing stiffened wall panel and the rear beam trailing edge structure provided by the present invention;

[0028] In the figure: 1-wing upper panel skin, 2-wing trailing edge skin, 3-panel stringer, 4-trough rear beam, 5-stringer ribs, 6-stringer rib sandwich, 7-trailing edge filler, 8-first mold, 9-wing trailing edge upper wing skin, 10-wing trailing edge lower wing skin, 11-front stringer rib, 12-second mold, 13-rear stringer rib, 14-upper wing panel, 15-web, 16-lower wing panel, 17-third mold, 18-trough rear beam mold DETAILED DESCRIPTION

[0029] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative efforts are within the scope of protection of the present invention.

[0030] The purpose of the present invention is to provide an integrated structure of a composite wing stiffened wall panel and a rear beam trailing edge and a manufacturing method to solve the problems existing in the prior art, reduce the weight of composite wing-like structures, ensure that the wing-like structures have good watertight properties, and ensure the stability of the performance of the equipment inside the wing and the filling material inside the trailing edge.

[0031] In order to make the above-mentioned objects, features and advantages of the present invention more obvious and easy to understand, the present invention is further described in detail below with reference to the accompanying drawings and specific embodiments.

[0032] like Figure 1-Figure 2 As shown, this embodiment provides an integrated structure of a composite material wing stiffened wall panel and a rear beam trailing edge, comprising an integrally solidified wing upper wall panel skin 1, a wing trailing edge skin 2, a wall panel long girder 3 and a grooved rear beam 4, the wing trailing edge skin 2 being arranged at the trailing edge of the wing upper wall panel skin 1, a wall panel long girder 3 being arranged on the lower wing surface side of the wing upper wall panel skin 1, long girder vertical ribs 5 being arranged on the wall panel long girder 3, a long girder vertical rib sandwich 6 being arranged between the two long girder vertical ribs 5, a grooved rear beam 4 being arranged at one end of the wall panel long girder 3 close to the trailing edge, the grooved rear beam 4 being located on the inner side of the wing trailing edge skin 2, and a trailing edge filler 7 being filled in the trailing edge space formed by the grooved rear beam 4 and the wing trailing edge skin 2.

[0033] The wing upper panel skin 1, wing trailing edge skin 2, panel stringers 3, and channel-shaped rear spar 4 are integrally manufactured through curing. This eliminates structural separations, reduces the mechanical connections between the channel-shaped rear spar, wing trailing edge, and upper panel structure, and reduces the number of fasteners or adhesives used for these connections, effectively reducing structural weight. This ensures excellent watertightness for the wing structure, ensuring the overall wing is watertight in rainy and snowy conditions, and further ensuring the stability of the wing's internal equipment and trailing edge filler 7. The wing trailing edge skin 2 comprises an integrally curing upper wing trailing edge skin 9 and a lower wing trailing edge skin 10.

[0034] The method for manufacturing the above-mentioned integrated structure of composite wing stiffened wall panel and rear beam trailing edge comprises the following steps:

[0035] S1: Using wet prepreg, the wing upper panel skin 1 and the wing trailing edge upper wing skin 9 are laid on the surface of the first mold 8, and the wing trailing edge lower wing skin 10 is reserved for laying the required prepreg;

[0036] S2: The front long stringer flange 11 and the long stringer rib 5 of the panel long stringer 3 are laid out on the inner side of the wing upper panel skin 1 laid out in S1 using wet prepreg, and a second mold 12 is placed inside the front long stringer flange 11 to assist in the laying of the front long stringer flange 11 and the long stringer rib 5;

[0037] S3: a long girder reinforcement interlayer 6 is placed between the long girder reinforcement 5 laid in S2;

[0038] S4: On the inner side of the wing upper panel skin 1 laid in S1 and behind the long stringer ribs 5 laid in S2, continue to lay the rear long stringer cap 13 of the panel long stringer 3 and the upper wing panel 14 of the channel rear beam 4. Reserve space for the web 15 of the channel rear beam 4 and the lower wing panel 16 for laying the required prepreg. Place a third mold 17 inside the rear long stringer cap 13 to assist in laying the rear long stringer cap 13.

[0039] S5: Place a grooved rear beam mold 18 inside the upper wing panel 14 of the grooved rear beam 4, and lay the reserved prepreg required for laying the web 15 and the lower wing panel 16 of the grooved rear beam 4 in reverse on top of the grooved rear beam mold 18;

[0040] S6: Place the trailing edge inner filler 7 inside the wing trailing edge upper wing skin 9 laid in S1 and behind the channel-shaped rear beam 4 laid in S5, and lay the reserved prepreg required for laying the wing trailing edge lower wing skin 10 in reverse on the trailing edge inner filler 7 and the lower wing panel 16 of the channel-shaped rear beam 4;

[0041] S7: The prepreg material laid out in S6, the first mold 8, the second mold 12, the third mold 17, the channel-shaped rear beam mold 18, the long stringer vertical reinforcement sandwich 6 and the rear edge inner filler 7 are solidified as a whole to form an integrated structure;

[0042] S8: After removing the second mold 12, the third mold 17 and the grooved rear beam mold 18, the solidified part is removed from the first mold 8;

[0043] S9: The edges of the cured parts are polished, sealant is applied, and after non-destructive testing, the preparation of the integrated structure of the composite wing stiffener panel and the trailing edge of the rear beam is completed.

[0044] The thickness of the prepreg laid on the wing upper panel skin 1 and the wing trailing edge skin 2 is 0.5 to 3 mm, preferably 1 mm.

[0045] The thickness of the prepreg laid on the front long stringer cap 11, the long stringer vertical rib 5, the rear long stringer cap 13 and the channel-shaped rear beam 4 is 0.5-4 mm, preferably 2 mm.

[0046] The rear edge inner filler 7 is made of foam or honeycomb material.

[0047] The truss reinforcement interlayer 6 is made of composite material prepreg, foam or honeycomb material, and has a laying thickness of 1 to 3 mm, preferably 1 mm.

[0048] Before curing, the prepreg laid out in S6, the first mold 8, the second mold 12, the third mold 17, the channel-shaped rear beam mold 18, the long stringer vertical reinforcement sandwich 6 and the rear edge inner filler 7 are placed in a vacuum bag as a whole, the vacuum bag is evacuated, and then cured together with the vacuum bag.

[0049] In step S7, curing is performed in an autoclave at a curing temperature of 120 to 180° C., a curing holding time of 120 to 180 min, and a pressure of 0.6 MPa in the autoclave.

[0050] The technical solution of the present invention is described in detail below with reference to specific examples.

[0051] The method for manufacturing the integrated structure of the composite material wing stiffened wall panel and the rear beam trailing edge of the present invention specifically comprises the following steps:

[0052] Step 1: Clean the surface of the first mold 8, apply a release agent, and then use carbon fiber resin-based composite material prepreg to lay the wing upper panel skin 1 and the wing trailing edge upper wing surface skin 9. The thickness of the wing upper panel skin 1 and the wing trailing edge upper wing surface skin 9 is preferably 0.5-3 mm, more preferably 1 mm, and the ply layup order is preferably [45 / -45 / 0 / 90] s , and reserve space for laying the prepreg required for the lower wing skin 10 at the trailing edge of the wing.

[0053] Step 2: On the inner side of the wing upper panel skin 1 prepreg laid in step 1, the front long stringer cap strip 11 and the long stringer rib 5 of the panel long stringer 3 are laid using carbon fiber resin-based composite material prepreg. After the surface of the second mold 12 is coated with a release agent, the second mold 12 is placed to assist in the laying of the long stringer rib 5 and the front long stringer cap strip 11. The thickness of the front long stringer cap strip 11 is preferably 0.5-4 mm, more preferably 2 mm. The height of the long stringer rib 5 is preferably 20-40 mm, more preferably 26 mm. The layup order is preferably [45 / -45 / 0 / 90]. 2s .

[0054] Step 3: Place the long stringer reinforcement sandwich 6 between the long stringer reinforcements 5 laid in step 2. The long stringer reinforcement sandwich 6 is preferably made of composite material prepreg, foam or honeycomb material. In this embodiment, composite material prepreg is selected, and the thickness is preferably 1-3 mm, more preferably 1 mm. The preferred laying order is [45 / -45 / 0 / 90] s .

[0055] Step 4: Continue laying the rear stringer edge strips 13 of the panel stringer 3 and the upper wing panel 14 of the channel rear beam 4, inside the prepreg material of the wing upper panel skin 1 laid in step 1 and behind the stringer ribs 5 laid in step 2. Reserve the web 15 and lower wing panel 16 of the channel rear beam 4 for laying the required prepreg. Place a third mold 17 coated with a release agent behind the stringer ribs 5. Apply release agent to the surface of the channel rear beam mold 18. After placing the channel rear beam mold 18, lay the reserved prepreg materials of the web 15 and lower wing panel 16 of the channel rear beam 4 in the opposite direction on top of the channel rear beam mold 18. The layer thickness is preferably 0.5-4mm, more preferably 2mm, and the layering order is preferably [45 / -45 / 0 / 90]. 2s The distance between the long stringer reinforcement 5 and the channel-shaped rear beam 4 is preferably 100 to 200 mm, more preferably 115 mm.

[0056] Step 5: Place the trailing edge inner filler 7 on the inner side of the wing surface skin 9 prepreg laid out in step 1 and behind the grooved rear beam 4 laid out in step 4. The trailing edge inner filler is preferably made of foam or honeycomb material, more preferably made of foam.

[0057] Step 6: The prepreg required for laying the wing trailing edge lower wing surface skin 10 reserved in step 1 is reversely laid on the trailing edge inner filler 7 completed in step 5 and the lower wing panel 16 of the wing channel rear beam 4 completed in step 4.

[0058] Step 7: Cover the finished product with a vacuum bag, adhere the bag to the first mold 8 around the first mold 8 with a sealant, and evacuate the vacuum bag.

[0059] Step 8: Place the product and each mold in an autoclave for curing. The curing temperature is preferably 120-180°C, more preferably 180°C; the holding time is preferably 120-180 minutes, more preferably 180 minutes. The pressure in the autoclave is preferably 0.6 MPa.

[0060] Step 9: Take out the cured product, open the vacuum bag, take out the second mold 12 , the third mold 17 and the grooved rear beam mold 18 in sequence, and remove the product from the first mold 8 .

[0061] Step 10: Grind the edge of the workpiece, apply sealant, and complete the preparation of the integrated structure of the composite wing stiffener panel and the rear beam trailing edge after non-destructive testing. Figure 2 .

[0062] The integrated structure of the composite wing stiffened wall panel and the rear beam trailing edge provided by the present invention reduces the amount of butt-jointing or overlapping materials and the number of fasteners or adhesives used for connection, thereby effectively reducing the structural weight. Compared with the traditional wing configuration in which the rear beam and upper and lower wing panels are manufactured separately and connected using fasteners, and the upper wing panel and trailing edge are manufactured separately and then assembled, the present invention can reduce the structural weight by 1.06 kg per meter along the wingspan direction, a weight reduction of 10.8% compared with the traditional configuration, thereby improving the overall performance of the aircraft.

[0063] The wing upper panel skin 1, the wing trailing edge skin 2, the panel long stringer 3 and the grooved rear beam 4 of the present invention are manufactured in one piece without a structural separation surface, thereby ensuring that the wing-like structure has good watertightness. In rainy and snowy weather, the watertightness of the entire wing structure is guaranteed, thereby ensuring the stability of the performance of the equipment inside the wing and the filling material inside the trailing edge.

[0064] The present invention manufactures the wing upper panel skin 1, the wing trailing edge skin 2, the panel long stringer 3 and the channel-shaped rear beam 4 by solidifying and molding them in one piece, and only requires the preparation of one set of molds, thereby reducing the multiple sets of molds required for the separate manufacture of each component and the occupation of the autoclave by multiple curing processes, reducing the additional auxiliary tooling and connection workload required for component assembly, greatly reducing labor consumption and processing and manufacturing costs, and improving manufacturing efficiency.

[0065] The present invention uses specific examples to illustrate the principles and implementation methods of the present invention. The above examples are only intended to help understand the method and core concept of the present invention. At the same time, those skilled in the art will find that the specific implementation methods and application scopes may vary based on the concept of the present invention. In summary, the contents of this specification should not be construed as limiting the present invention.

Claims

1. A composite wing stiffened panel and rear beam trailing edge integrated structure, characterized by: The invention comprises an integrally solidified wing upper panel skin, a wing trailing edge skin, a panel long stringer and a grooved rear beam, wherein the wing trailing edge skin is provided at the trailing edge of the wing upper panel skin, the panel long stringer is provided on the lower wing surface side of the wing upper panel skin, long stringer vertical ribs are provided on the panel long stringer, a long stringer vertical rib sandwich is provided in the middle of the long stringer vertical ribs, the grooved rear beam is provided at one end of the panel long stringer close to the trailing edge, the grooved rear beam is located on the inner side of the wing trailing edge skin, and the trailing edge space formed by the grooved rear beam and the wing trailing edge skin is filled with a trailing edge filler; The wall panel long girder includes a front long girder edge strip and a rear long girder edge strip, and the long girder vertical ribs are provided at the junction of the front long girder edge strip and the rear long girder edge strip, and the long girder vertical rib sandwich is provided between the two long girder vertical ribs. The groove-type rear beam includes an upper wing panel, a web and a lower wing panel that are integrally solidified and formed, and the rear long girder edge strip and the upper wing panel of the groove-type rear beam are an integrally solidified and formed structure; the wing trailing edge skin includes an upper wing surface skin of the wing trailing edge and a lower wing surface skin of the wing trailing edge that are integrally solidified and formed, and the lower wing surface skin of the wing trailing edge is folded downward and overlapped on the lower surface of the lower wing panel of the groove-type rear beam and is integrally solidified and formed with the lower wing panel of the groove-type rear beam.

2. A method for manufacturing the integrated structure of composite wing stiffened wall panel and rear beam trailing edge as claimed in claim 1, characterized in that: The following steps are involved: S1: using wet prepreg to lay the wing upper panel skin and the wing trailing edge upper wing skin on the surface of the first mold, and reserving prepreg required for laying the wing trailing edge lower wing skin; S2: Laying the front long stringer cap and the long stringer rib of the panel long stringer using wet prepreg on the inner side of the wing upper panel skin laid in S1, and placing a second mold on the inner side of the front long stringer cap to assist in laying the front long stringer cap and the long stringer rib; S3: placing the long girder reinforcement interlayer between the long girder reinforcements laid in S2; S4: On the inner side of the wing upper panel skin laid in S1 and behind the long stringer ribs laid in S2, continue laying the rear long stringer cap of the panel long stringer and the upper wing panel of the channel-shaped rear beam, and reserve prepreg required for laying the web of the channel-shaped rear beam and the lower wing panel. Place a third mold inside the rear long stringer cap to assist in laying the rear long stringer cap; S5: placing a grooved rear beam mold inside the upper wing panel of the grooved rear beam, and laying the reserved prepreg required for laying the web of the grooved rear beam and the lower wing panel in reverse order above the grooved rear beam mold; S6: placing the trailing edge inner filler on the inner side of the wing trailing edge upper wing skin laid in S1 and behind the channel-shaped rear beam laid in S5, and reversely laying the reserved prepreg required for laying the wing trailing edge lower wing skin on the trailing edge inner filler and the lower wing panel of the channel-shaped rear beam; S7: solidifying the prepreg laid out in S6, the first mold, the second mold, the third mold, the channel-shaped rear beam mold, the long stringer reinforcement sandwich layer, and the rear edge inner filler as a whole to form an integrated structure; S8: After removing the second mold, the third mold and the grooved rear beam mold, removing the solidified part from the first mold; S9: Grinding the edge of the solidified workpiece, applying sealant, and performing non-destructive testing to complete the preparation of the integrated structure of the composite wing stiffened wall panel and the rear beam trailing edge.

3. The method for manufacturing the integrated structure of composite wing stiffened wall panel and rear beam trailing edge according to claim 2, characterized in that: The thickness of the prepreg laid on the wing upper panel skin and the wing trailing edge skin is 0.5-3 mm.

4. The method for manufacturing the integrated structure of composite wing stiffened wall panel and rear beam trailing edge according to claim 2, characterized in that: The thickness of the prepreg laid on the front long stringer edge strip, the long stringer vertical rib, the rear long stringer edge strip and the channel-shaped rear beam is 0.5-4 mm.

5. The method for manufacturing the integrated structure of composite wing stiffened wall panel and rear beam trailing edge according to claim 2, characterized in that: The rear edge inner filler is made of foam or honeycomb material.

6. The method for manufacturing the integrated structure of composite wing stiffened wall panel and rear beam trailing edge according to claim 2, characterized in that: The long stringer reinforcement interlayer is made of composite material prepreg, foam or honeycomb material, and has a laying thickness of 1 to 3 mm.

7. The method for manufacturing the integrated structure of composite wing stiffened wall panel and rear beam trailing edge according to claim 2, characterized in that: Before curing, the prepreg laid out in S6, the first mold, the second mold, the third mold, the grooved rear beam mold, the long truss vertical reinforcement sandwich and the rear edge inner filler are placed in a vacuum bag as a whole, the vacuum bag is evacuated, and then cured together with the vacuum bag.

8. The method for manufacturing the integrated structure of composite wing stiffened wall panel and rear beam trailing edge according to claim 2, characterized in that: In step S7, curing is performed in an autoclave at a curing temperature of 120-180° C., a curing holding time of 120-180 min, and a pressure of 0.6 MPa in the autoclave.

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