Composite material hollow joint structure and layering method thereof

Abstract

The invention discloses a composite material hollow joint structure and a layering method thereof, the joint is of a structure with two hollow ends, the main body is cylindrical, the main structure is the main structure of the joint and is used for connecting a machine body, the structure is gradually transited to an inclined plane, and the beveling design is used for conveniently inserting and connecting a metal lining; an internal spline is arranged on the surface of the metal lining, connecting bosses are designed at the two ends of the metal lining, the composite material and the lining are prevented from sliding, an ear-shaped protrusion is arranged on the upper portion of the middle of the connector, a through hole is formed in the ear-shaped protrusion and used for installation of a pin, and the connector can be fixed to other structures. The overall structure of the connector is optimized, the inner spline area laying layer is nested on an inner spline of the metal lining, the outer equal-thickness area laying layer is laid on a boss of the metal lining, the connecting strength of metal and a composite material is guaranteed, one end of the inner main rod first area laying layer is embedded into a connector connecting area and is directly connected with the connector connecting area, and the other end of the inner main rod first area laying layer is directly connected with the connector connecting area. And the continuity of the whole structure is ensured.

Description

Technical Field

[0001] The invention relates to the technical field of composite material structure design, in particular to a composite material hollow joint structure and a laying method thereof. Background Art

[0002] Composite materials have the advantages of high specific stiffness and specific strength, excellent fatigue performance and designability. They have been widely used in the aerospace field. The proportion of composite materials in the weight of aircraft structures has become one of the evaluation indicators of the advancement of aircraft. Compared with traditional metal joints, composite joints have significant weight reduction effects and excellent mechanical properties, and have important economic and military significance.

[0003] In recent years, composite material technology has developed rapidly. The application of composite materials has gradually transitioned from structural secondary load-bearing parts to the manufacturing of primary load-bearing parts. In large structures such as aircraft, spacecraft, and ships, composite joints are often used to connect and transfer loads to intersecting components, playing an important role in external connection and pressure-bearing sealing. A large number of experiments have shown that joints are the weakest link in the structure. Military and civilian airworthiness regulations in various countries around the world have put forward strict requirements on the load-bearing capacity of composite joints. Therefore, the overall layup technology of complex joints is of great significance to ensure the safety of aircraft structures.

[0004] A patent discloses a method for preparing carbon fiber composite joints. Although the multi-stage synchronous preforming method is used in the layup process to replace the original layer-by-layer preforming, which improves production efficiency, the integrity of the joint during the layup process cannot be guaranteed. In addition, the bonding connection method is used in the connection between the metal nesting and the composite joint, and the interface at the connection is prone to damage, which reduces the effective load-bearing capacity of the composite joint. For composite joints with drastic changes in hollow thickness at both ends, there are still problems such as poor load-bearing capacity, difficulty in forming, and long process cycle.

[0005] To this end, we propose a composite hollow joint structure and its laying method to solve the above problems. Summary of the invention

[0006] The object of the present invention is to provide a composite hollow joint structure and a laying method thereof to solve the problems raised in the above background technology.

[0007] To achieve the above object, the present invention provides the following technical solutions:

[0008] A composite hollow joint structure and a laying method thereof, wherein the joint is a hollow structure at both ends, the main body is cylindrical and is the main structure of the joint, and is used to connect the fuselage, and the structure gradually transitions to an inclined surface, and the inclined design is for the convenience of inserting and connecting a metal bushing, the surface of the metal bushing has an internal spline and connecting bosses are designed at both ends to prevent the composite material and the bushing from sliding, and an ear-shaped protrusion is provided at the upper middle part of the joint, and a through hole is provided on the ear-shaped protrusion for installing a pin, so that the joint can be fixed on other structures;

[0009] The structure is divided into three parts: the joint connection area, the outer main fuselage area and the inner fuselage area;

[0010] The joint connection area is divided into an outer equal thickness area and an inner spline area;

[0011] The outer main fuselage area is divided into two parts: the ear area and the outer fuselage area;

[0012] The inner fuselage area is divided into four parts: inner main rod area 1, area 2, area 3 and area 4; among them, the inner main rod area 1 is directly connected to the joint connection area, ensuring the continuity of the overall structure;

[0013] The specific process is as follows:

[0014] S1, according to the structural characteristics and the straightness of the fiber during processing, the complex joint structure is divided into three major areas: the joint connection area, the outer main fuselage area and the inner fuselage area. The ply shape and thickness of each area are obtained. According to the overall structural relationship, the ply order of each area is determined: the ply of the joint connection area is carried out first. Since the first area of ​​the inner main pole is directly connected to the joint connection area, the inner fuselage area is plyed layer by layer, and finally the outer main fuselage area is plyed;

[0015] Preferably, after obtaining the spatial curved surface shape of each layer, the spatial curved surface layer is converted into a two-dimensional plane layer using three-dimensional software, including most of the closed curved surface shapes. The mapping position needs to be set during the mapping process, that is, the position of the cut when the spatial overall curved surface is converted into a two-dimensional plane state. When designing the cut position, in order to avoid stress concentration caused by the cuts at the same or nearby positions, which leads to premature failure of the material, the cut positions of adjacent layers should differ by 60° when designing the cuts, and a cycle of 6 layers should be used to avoid cut concentration;

[0016] S2, joint connection area laying steps:

[0017] S2-1, first fix both ends of the metal bushing joint, and lay the inner spline area. The thickness of this area is 0.8 mm, including 4 layers of carbon fiber unidirectional cloth, and there are 8 cycles of variable thickness areas on the upper and lower sides. In the variable thickness area, the horizontal spacing between the first and second layers is 1 mm, and the horizontal spacing between the second, third and fourth layers is 2 mm. The laying order is [0° / 45° / -45° / 0°];

[0018] S2-2, the thickness of the outer equal thickness zone is 7.6 mm, which is an equal thickness zone. It adopts a ply / weaving coupling structure, including 8 woven fabric plies and 22 unidirectional fabric plies. The ply order of the innermost and outermost layers is [0° / 45° / 90° / -45° / 0°], and the ply order of the middle layers is [0° / 20° / 90° / -20° / 0°];

[0019] S3, inner fuselage zone 3 lay-up steps:

[0020] S3-1, fix the main rod core mold by a fixed bracket, and lay the inner main rod area at the front end of the core mold. This area is an equal thickness area with a thickness of 1mm. Set 5 carbon fiber unidirectional cloth layers in the order of [0° / 45° / 90° / -45° / 0°]. Heat set the inner main rod area and then inlay it with the joint connection area 1 to ensure the continuity of the overall structure;

[0021] S3-2, the second area of ​​the inner main pole is an important connection area between the joint and the main body, and local reinforcement is required under the structure, resulting in a dramatic change in thickness in this area. The thickest is 22.8mm at the bottom, 10mm at the top, and only 0.4mm at the two ends. One complete ply is retained in the innermost and outermost layers respectively. According to the structure, a semicircular arc-shaped piece is set at the top and a "M"-shaped piece is set at the bottom. The shape and size are changed according to the thickness change, and the patches are laid one by one. Among them, the top ply is all unidirectional cloth, and the plying order is [45° / 0° / 0° / -45°]. The bottom is thicker, and unidirectional cloth and woven cloth are used for plying, and the plying order is [0° / 20° / 45° / -20° / -45° / 0°].

[0022] S3-3, the three zones of the inner main pole are equal thickness transition zones with a thickness of 0.6 mm. Three complete layers are set to cover the area with sharply varying thickness of the inner main pole zone. The layering sequence is [45° / 0° / -45°];

[0023] S3-4, the front end of the inner main rod four zones is a small variable thickness zone, the bottom thickness is 6mm, the thickness at the bottom corner is 6.8mm, the thickness on both sides gradually increases from 6.8mm until the maximum thickness reaches 11mm on the top surface, and the rear end of this zone is a uniform thickness zone with a thickness of 4.8mm; therefore, when setting the ply, retain the 4.8mm unidirectional fabric complete ply, every 6 layers are a complete ply group, and the ply sequence is [0° / 45° / 90° / 90° / -45° / 0°]4. Insert the front variable thickness zone and insert it evenly between the 4 complete ply groups according to the thickness;

[0024] S4, outer main fuselage area layer laying steps:

[0025] S4-1, install the fixing bracket of the lifting ear part, first lay the layers in the lifting ear area, which is an overlap area of ​​16 mm, and use the layering / weaving coupling structure for layering, and the layering sequence is [0° / 90° / 20° / -20° / -90° / 0°];

[0026] S4-2, to ensure the integrity of the structure, on the basis of the laying of the ear area, the outer fuselage area is laid. The structure of this area is similar to that of the inner main pole area. The laying steps refer to S3-4;

[0027] Among them, in the process of laying layers in the three areas, in order to avoid stress concentration, the method shown in S1-1 is used when setting the cutouts for all closed curved surface shapes.

[0028] Compared with the prior art, the present invention has the following beneficial effects:

[0029] 1. The present invention optimizes the overall structure of the joint. The inner layer spline area ply is embedded in the inner spline of the metal bushing, and the outer layer uniform thickness area ply is laid on the metal bushing boss, thereby ensuring the connection strength between the metal and the composite material. One end of the inner layer main rod area ply is embedded in the joint connection area and directly connected thereto, thereby ensuring the continuity of the overall structure and improving the overall bearing capacity of the joint.

[0030] 2. In the process of converting the spatial overall curved surface into a two-dimensional plane state, in order to avoid stress concentration at the same or nearby positions of the cuts, which may lead to premature failure of the material, the cut positions of adjacent layers are designed to differ by 60°, thereby avoiding the reduction of joint performance due to the concentration of the cut positions;

[0031] 3. The present invention adopts a ply / weaving coupling structure in the ply design. The ply fabric includes 0.2 mm / layer unidirectional carbon fiber cloth and 0.4 mm / layer woven structure cloth with a weaving angle of 20°. The woven cloth ply is used in the high-thickness area, which can effectively resist damage and deformation, improve the stability of the overall structure and increase the service life;

[0032] 4. Compared with the prior art, the present invention discloses a composite material joint with a complex structure that is hollow at both ends, inclined and staggered, and has a drastic thickness change. This joint structure is widely used in large structures such as spacecraft and ships. The present invention provides a complete laying method for each part of the structure, provides an overall laying method for composite hollow joints, improves the production efficiency of complex joints, optimizes the process, and reduces costs. BRIEF DESCRIPTION OF THE DRAWINGS

[0033] Figure 1 It is a diagram of the overall structure of the joint of the present invention;

[0034] Figure 2 It is a schematic diagram of the structure of the metal bushing of the present invention;

[0035] Figure 3 is a schematic diagram of the joint connection region of the present invention;

[0036] Figure 4 is a schematic diagram of the outer body area of ​​the joint of the present invention;

[0037] Figure 5 It is a schematic diagram of the lifting ear area of ​​the joint of the present invention;

[0038] Figure 6 is a schematic diagram of the outer fuselage area of ​​the joint of the present invention;

[0039] Figure 7 is a schematic diagram of the inner fuselage area of ​​the joint of the present invention;

[0040] Figure 8 This is a schematic diagram of a first area of ​​the inner main rod of the joint of the present invention;

[0041] Fig. 9 It is a schematic diagram of the second zone of the inner main rod of the joint of the present invention;

[0042] Fig.10 It is a schematic diagram of three zones of the inner main rod of the joint of the present invention;

[0043] Fig.11 It is a schematic diagram of four zones of the inner main rod of the joint of the present invention;

[0044] Fig.12 It is a flow chart of the present invention;

[0045] Fig.13 It is a flow chart of layer conversion in the layer laying process of the present invention.

[0046] In the figure: 1 is the joint connection area, 2 is the outer main body area, 3 is the inner fuselage area, 11 is the outer equal thickness area, 12 is the inner spline area, 21 is the lifting ear area, 22 is the outer fuselage area, 31 is the inner main rod area 1, 32 is the inner main rod area 2, 33 is the inner main rod area 3, and 34 is the inner main rod area 4. DETAILED DESCRIPTION

[0047] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. 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 creative work are within the scope of protection of the present invention.

[0048] Embodiment 1

[0049] A composite hollow joint structure and a laying method thereof, wherein the joint is a hollow structure at both ends, the main body is cylindrical and is the main structure of the joint, and is used to connect the fuselage, and the structure gradually transitions to an inclined surface, and the inclined design is for the convenience of inserting and connecting a metal bushing, the surface of the metal bushing has an internal spline and connecting bosses are designed at both ends to prevent the composite material and the bushing from sliding, and an ear-shaped protrusion is provided at the upper middle part of the joint, and a through hole is provided on the ear-shaped protrusion for installing a pin, so that the joint can be fixed on other structures;

[0050] The structure is divided into three parts: the joint connection area 1, the outer main body area 2 and the inner body area 3; Figure 3 As shown, the joint connection area 1 is divided into an outer layer uniform thickness area 11 and an inner layer spline area 12; Figure 4-6 As shown, the outer main body fuselage area 2 is divided into two parts: the ear area 21 and the outer fuselage area 22; Figure 7-11 As shown, the inner fuselage area 3 is divided into four parts: inner main rod area 1 31, area 2 32, area 33 and area 4 34; wherein, the inner main rod area 1 31 is directly connected to the joint connection area 1, thereby ensuring the continuity of the overall structure;

[0051] Preferably, the structure of the present embodiment is mostly a complex variable thickness irregular structure, and a ply / weaving coupling structure is adopted in the ply design. The ply fabric includes a 0.2mm / layer unidirectional carbon fiber cloth and a 0.4mm / layer woven structure cloth with a weaving angle of 20°. The innermost and outermost layers are carbon fiber unidirectional cloth, and the middle part is a carbon fiber unidirectional cloth and a woven fabric coupling structure. The ply angles of the present embodiment are mainly 0°, 90°, +45°, and -45°, interspersed with +20° and -20° woven fabric plies, wherein 0° provides axial strength and stiffness, 90° provides lateral strength and stiffness, and 45° improves toughness and shear resistance. The ply angle is selected according to the stress conditions of each area. The ply steps are as follows: Fig.12 As shown, the specific process is as follows:

[0052] S1, according to the structural characteristics and the straightness of the fiber during processing, the complex joint structure is divided into three major areas: the joint connection area 1, the outer main fuselage area 2 and the inner fuselage area 3, and the ply shape and thickness of each area are obtained. According to the overall structural relationship, the ply order of each area is determined: first, the ply of the joint connection area 1 is carried out. Since the inner main rod area 1 31 is directly connected to the joint connection area 1, the inner fuselage area 3 is plyed layer by layer, and finally the outer main fuselage area 2 is plyed;

[0053] Preferably, after obtaining the spatial curved surface shape of each layer, the spatial curved surface layer is converted into a two-dimensional plane layer using three-dimensional software, including most of the closed curved surface shapes. The mapping position needs to be set during the mapping process, that is, the position of the cut when the spatial overall curved surface is converted into a two-dimensional plane state. When designing the cut position, in order to avoid stress concentration caused by the cuts at the same or nearby positions, which leads to premature failure of the material, the cut positions of adjacent layers should differ by 60° when designing the cuts, and a cycle of 6 layers should be used to avoid cut concentration;

[0054] S2, joint connection area 1 layer laying steps:

[0055] S2-1, first fix both ends of the metal bushing joint, and lay the inner spline area 12. The thickness of this area is 0.8 mm, including 4 layers of carbon fiber unidirectional cloth, and there are 8 cycles of variable thickness areas on the upper and lower sides. In the variable thickness area, the horizontal spacing between the first layer and the second layer is 1 mm, and the horizontal spacing between the second layer, the third layer and the fourth layer is 2 mm. The laying order is [0° / 45° / -45° / 0°];

[0056] S2-2, the thickness of the outer equal thickness zone 11 is 7.6 mm, which is an equal thickness zone. It adopts a ply / weaving coupling structure, including 8 woven fabric plies and 22 unidirectional fabric plies. The ply order of the innermost and outermost layers is [0° / 45° / 90° / -45° / 0°], and the ply order of the middle layers is [0° / 20° / 90° / -20° / 0°];

[0057] S3, inner fuselage zone 3 lay-up steps:

[0058] S3-1, fix the main rod core mold by a fixing bracket, and lay the inner main rod area 1 31 at the front end of the core mold. This area is an equal thickness area with a thickness of 1mm. Set 5 carbon fiber unidirectional cloth layers in the order of [0° / 45° / 90° / -45° / 0°]. Heat-set the inner main rod area 31 and then inlay it with the joint connection area 1 to ensure the continuity of the overall structure;

[0059] S3-2, the second zone 32 of the inner main pole is an important connection area between the joint and the main body, and local reinforcement is required under the structure, resulting in a dramatic change in thickness in this area. The thickest layer is 22.8mm at the bottom, 10mm at the top, and only 0.4mm at the two ends. One complete ply is retained in the innermost layer and the outermost layer respectively. According to the structure, a semicircular arc-shaped piece is set at the top and a "M"-shaped piece is set at the bottom. The shape and size are changed according to the thickness change, and the patches are laid one by one. Among them, the top ply is all unidirectional cloth, and the plying order is [45° / 0° / 0° / -45°]. The bottom is thicker, and unidirectional cloth and woven cloth are used for plying, and the plying order is [0° / 20° / 45° / -20° / -45° / 0°].

[0060] S3-3, the third area 33 of the inner main rod is a transition area of ​​equal thickness with a thickness of 0.6 mm. Three complete layers are set to cover the area of ​​the inner main rod area 32 with a sharp change in thickness. The layering sequence is [45° / 0° / -45°];

[0061] S3-4, the front end of the inner main rod four zones 34 is a small variable thickness zone, the bottom thickness is 6mm, the thickness at the bottom corner is 6.8mm, the thickness on both sides gradually increases from 6.8mm until the maximum thickness reaches 11mm at the top surface, and the rear end of this zone is a uniform thickness zone with a thickness of 4.8mm; therefore, when setting the ply, retain the 4.8mm unidirectional fabric complete ply, every 6 layers are a complete ply group, and the ply sequence is [0° / 45° / 90° / 90° / -45° / 0°]4. Insert the front variable thickness zone and insert it evenly between the 4 complete ply groups according to the thickness;

[0062] S4, outer layer main fuselage area 2 laying steps:

[0063] S4-1, install the fixing bracket of the lifting ear part, first lay the layers in the lifting ear area 21, which is an overlap area of ​​16 mm, and use the layering / weaving coupling structure for layering, and the layering sequence is [0° / 90° / 20° / -20° / -90° / 0°];

[0064] S4-2, to ensure the integrity of the structure, on the basis of the laying of the ear area, the outer fuselage area 22 is laid. The structure of this area is similar to that of the inner main pole area 34. The laying steps refer to S3-4;

[0065] Among them, in the process of laying layers in the three areas, in order to avoid stress concentration, the method shown in S1-1 is used when setting the cutouts for all closed curved surface shapes.

[0066] It will be apparent to those skilled in the art that the invention is not limited to the details of the exemplary embodiments described above and that the invention can be implemented in other specific forms without departing from the spirit or essential features of the invention. Therefore, the embodiments should be considered exemplary and non-limiting in all respects, and the scope of the invention is defined by the appended claims rather than the foregoing description, and it is intended that all variations falling within the meaning and scope of the equivalent elements of the claims be included in the invention. Any reference numeral in a claim should not be considered as limiting the claim to which it relates.

[0067] In addition, it should be understood that although the present specification is described according to implementation modes, not every implementation mode contains only one independent technical solution. This description of the specification is only for the sake of clarity. Those skilled in the art should regard the specification as a whole. The technical solutions in each embodiment may also be appropriately combined to form other implementation modes that can be understood by those skilled in the art.

Claims

1. A composite hollow joint structure, characterized in that: The joint is a hollow structure at both ends. The main body is cylindrical and is the main structure of the joint. It is used to connect the fuselage. The structure gradually transitions to an inclined surface. This bevel design is to facilitate the insertion and connection of the metal bushing. The surface of the metal bushing has internal splines and connecting bosses are designed at both ends. There is an ear-shaped protrusion in the upper middle part of the joint, which has a through hole for the installation of pins, so that the joint can be fixed on other structures. The structure is divided into three parts: the joint connection area (1), the outer main body area (2) and the inner body area (3). The joint connection area (1) is divided into an outer uniform thickness area (11) and an inner spline area (12). The outer main body region (2) is divided into two parts: a lifting ear region (21) and an outer body region (22); The inner fuselage area (3) is divided into four parts: inner main rod area 1 (31), area 2 (32), area 3 (33) and area 4 (34); wherein the inner main rod area 1 (31) is directly connected to the joint connection area (1).

2. A method for laying layers of a composite hollow joint structure as claimed in claim 1, characterized in that: The steps include: S1, according to the structural characteristics and the straightness of the fibers during processing, the complex joint structure is divided into three major areas: the joint connection area (1), the outer main fuselage area (2) and the inner fuselage area (3), and the ply shape and thickness of each area are obtained. According to the overall structural relationship, the ply laying order of each area is determined: first, the ply laying of the joint connection area (1) is carried out. Since the inner main rod area 1 (31) is directly connected to the joint connection area (1), the inner fuselage area 3 is ply laid layer by layer, and finally the outer main fuselage area (2) is ply laid; S2, joint connection area (1) ply; S3, inner fuselage zone (3) layup; S4, outer main fuselage area (2) lay-up.

3. The method for laying layers of a composite hollow joint structure according to claim 2, characterized in that: In the step S1, after obtaining the spatial surface shape of each layer, the spatial surface layer is converted into a two-dimensional plane layer using three-dimensional software, which includes most of the closed surface shapes. The mapping position needs to be set during the mapping process, that is, the cutting position of the spatial overall surface converted into a two-dimensional plane state. When designing the cutting position, in order to avoid stress concentration caused by cutting at the same position or nearby positions, which leads to premature failure of the material, the cutting positions of adjacent layers should differ by 60° when designing the cutting, and a cycle of 6 layers is used to avoid cutting concentration.

4. The method for laying layers of a composite hollow joint structure according to claim 2, characterized in that: The steps of laying the joint connection area (1) in step S2 are as follows: S2-1, first fix both ends of the metal bushing joint, and lay the inner spline area (12). The thickness of this area is 0.8 mm, including 4 layers of carbon fiber unidirectional cloth, and there are 8 cycles of variable thickness areas on the upper and lower sides. In the variable thickness area, the horizontal spacing between the first layer and the second layer is 1 mm, and the horizontal spacing between the second layer, the third layer and the fourth layer is 2 mm. The laying order is [0° / 45° / -45° / 0°]; S2-2, the thickness of the outer equal thickness zone 11 is 7.6 mm, which is an equal thickness area. It adopts a ply / weaving coupling structure, including 8 woven fabric plies and 22 unidirectional fabric plies. The ply sequence of the innermost and outermost layers is [0° / 45° / 90° / -45° / 0°], and the middle ply sequence is [0° / 20° / 90° / -20° / 0°].

5. The method for laying layers of a composite hollow joint structure according to claim 2, characterized in that: The steps of laying the inner fuselage area (3) in step S3 are as follows: S3-1, fix the main rod core mold by a fixing bracket, and lay the inner main rod area 1 31 at the front end of the core mold. This area is an equal thickness area with a thickness of 1mm. Set 5 carbon fiber unidirectional cloth layers in the order of [0° / 45° / 90° / -45° / 0°]. Heat-set the inner main rod area 31 and then inlay it with the joint connection area 1 to ensure the continuity of the overall structure; S3-2, the second area (32) of the inner main pole is an important connection area between the joint and the main body. It is necessary to locally strengthen the structure below, resulting in a dramatic change in thickness in this area. The thickest thickness at the bottom is 22.8 mm, the top is 10 mm, and the thinnest thickness at both ends is only 0.4 mm. A complete ply is retained in the innermost layer and the outermost layer respectively. According to the structure, a semicircular arc-shaped piece is set at the top and a "M"-shaped piece is set at the bottom. The shape and size are changed according to the thickness change, and the patch ply is laid one by one. Among them, the top ply is all unidirectional fabric, and the plying order is [45° / 0° / 0° / -45°]. The bottom thickness is relatively large, and unidirectional fabric and woven fabric are used for plying, and the plying order is [0° / 20° / 45° / -20° / -45° / 0°]. S3-3, the third area (33) of the inner main rod is a transition area of ​​equal thickness with a thickness of 0.6 mm. Three complete layers are set to cover the area with a sharp change in thickness of the inner main rod area (32). The layering sequence is [45° / 0° / -45°] S3-4, the front end of the inner main rod four zones (34) is a slightly variable thickness zone, the bottom thickness is 6mm, the thickness at the bottom corner is 6.8mm, the thickness on both sides gradually increases from 6.8mm until the maximum thickness reaches 11mm at the top surface, and the rear end of this zone is a uniform thickness zone with a thickness of 4.8mm; therefore, when setting the ply, retain the 4.8mm unidirectional fabric complete ply, every 6 layers are a complete ply group, and the ply sequence is [0° / 45° / 90° / 90° / -45° / 0°]4. The front variable thickness zone is intercalated and evenly intercalated between the 4 complete ply groups according to the thickness.

6. The method for laying layers of a composite hollow joint structure according to claim 2, characterized in that: The steps of laying the outer layer of the main body region (2) in step S4 are as follows: S4-1, installing the fixing bracket of the lifting ear part, firstly laying the lifting ear area (21), which is an overlap area of ​​16 mm, using a laying / weaving coupling structure for laying, and the laying sequence is [0° / 90° / 20° / -20° / -90° / 0°]; S4-2, in order to ensure the integrity of the structure, on the basis of the laying of the ear area, the outer fuselage area (22) is laid. The structure of this area is similar to that of the inner main rod area (34). The laying steps refer to step S3.

Images