A composite material winged projectile structure forming method

Abstract

The present invention discloses a composite material winged projectile structure forming method, including: S1, pre-fixing the projectile and the wing; pre-assembling the projectile and the wing through a fixing tool, preliminarily determining the relative position of the upper fixing frame and the lower fixing frame; applying structural glue to the wing root surface where the wing and the projectile are in direct contact, determining the final position of the wing, and the wing inclination α must be less than 0.2 degrees; S2, secondary fixing the projectile and the wing; determining the side structural glue application radius R according to the projectile diameter, curing the side structural glue, and secondary fixing the wing and the projectile; S3, forming the projectile and wing reinforcement area; laying fabric prepregs of different sizes and quantities in the reinforcement area in sequence, and then heating and curing. The process controllable and reliable forming method provided by the present invention improves the safety and reliability of the winged projectile structure by controlling the application radius of the structural glue on the wing root surface and the accuracy and consistency of the laying of the reinforcement area, which is conducive to exerting the ultimate performance of the aircraft.

Description

Technical Field

[0001] The present invention relates to the technical field of resin-based composite material component manufacturing, and in particular to a method for forming a composite material winged projectile structure. Background Art

[0002] Wings effectively control the trajectory and flight behavior of a missile, enhancing its maneuverability. Metal wings can be connected to the missile body in a variety of ways, including welding, riveting, and bolting. Weight reduction using composite materials is crucial for aircraft, enabling them to fly higher, farther, and achieve even greater performance. For composite materials, especially thermoset composite missile structures, ensuring a secure, reliable connection between the wings and the body, while ensuring optimal performance, is crucial.

[0003] Therefore, in response to the above technical problems, it is necessary to provide a composite material winged projectile structure forming method to solve the problem of safe, reliable and efficient connection between the composite material projectile and the projectile wings. Summary of the Invention

[0004] In order to overcome the deficiencies of the prior art, the present invention discloses a method for forming a composite material winged projectile structure.

[0005] To achieve the above object, the present invention is implemented through the following technical solutions:

[0006] The present invention discloses a method for forming a composite material winged projectile structure, comprising the following steps:

[0007] S1. Pre-fixation of missile body and wings:

[0008] Surface treatment of the outer surface of the missile and the bonding area of ​​the wing;

[0009] A fixing fixture is provided, which includes an upper fixing frame and a lower fixing frame. The missile body and the missile wings are pre-assembled through the fixing fixture, and the relative positions of the upper fixing frame and the lower fixing frame are preliminarily adjusted to determine;

[0010] After the relative positions of the upper and lower fixing frames are adjusted, remove the wing and apply structural adhesive to the wing root surface where the wing directly contacts the missile body. Determine the final position of the wing on the missile body. The wing inclination α must be less than 0.2°. After the structural adhesive cures, pre-fix the wing and missile body together.

[0011] S2. Secondary fixation of missile body and wings:

[0012] The side structural adhesive application radius R is determined according to the diameter of the projectile body. The side structural adhesive is cured and the projectile wing and the projectile body are fixed together for the second time.

[0013] S3, Forming of the missile body and wing reinforcement areas:

[0014] The body of the projectile is divided into four reinforcement areas by four wings. Multiple fabric prepregs of different sizes are laid on the reinforcement areas in order from small to large, put into vacuum bags, and placed in an oven for vacuuming, heating and curing to obtain the product.

[0015] In S2, the side structural adhesive application radius R can be determined according to the following formula:

[0016]

[0017] Where D is the diameter of the projectile, π is the circumference of the circle, To round up.

[0018] Preferably, in S3, sequentially laying fabric prepregs of different sizes and quantities on the reinforcement area specifically comprises the following steps:

[0019] Determine the number of fabric prepreg layers N in the reinforcement area based on the overall structural strength and stiffness, where N is ≥ 3 and is an integer;

[0020] The shapes of the prepreg layers from the first layer to the N-1 layer are all isosceles trapezoids. The shape of the first layer of prepreg layers has an upper and lower length of Its hypotenuse is parallel to the projected side E1 of the wing's side hypotenuse, and its lower base is parallel to the projection line E2 of the widest part of the two adjacent wing bodies. Subsequent layers are expanded on the basis of the first layer. The upper base, lower base and hypotenuse of the second to N-1 layers are respectively parallel to the upper base, lower base and hypotenuse of the first layer. The lengths of the upper bases of the first to N-1 layers are equal. The spacing between the upper bases of any two adjacent layers from the first to N-1 layers is equal. The spacing between the hypotenuses of any two adjacent layers from the first to N-1 layers is equal. The spacing between the lower bases of any two adjacent layers from the first to N-1 layers is equal, which are denoted as K1, H1 and M1 respectively. The Nth layer paving area is the overall reinforcement area, where K1, H1 and M1 are determined according to the following formulas:

[0021]

[0022] Among them, K is the distance between the upper bottom of the first layer and the top structural adhesive projection of the upper wing root surface, H is the distance between the oblique side of the first layer and the E1 edge, and M is the distance between the lower bottom of the first layer and the bottom structural adhesive projection of the lower wing root surface. To round down.

[0023] Preferably, in S3, a plurality of fabric prepregs of different sizes are laid in order from small to large on the reinforcement area, and then covered with an isolation film, a release cloth and a breathable felt, put into a vacuum bag, put into an oven for vacuum heating and curing, and the auxiliary materials are removed after curing to obtain the product.

[0024] Preferably, the fabric prepreg is a reinforced fiber prepreg, and the resin of the prepreg is a thermosetting resin.

[0025] Further preferably, the thermosetting resin is epoxy resin, phenolic resin, bismaleimide resin, or cyanate resin, and the reinforcing fiber is one or more of carbon fiber, glass fiber, basalt fiber, and aramid fiber.

[0026] Preferably, the long-term operating temperature of the elastic body, elastic wings and structural adhesive is greater than the curing temperature of the fabric prepreg.

[0027] Preferably, the fixing fixture further includes connecting screws for adjusting the positions of the upper fixing frame and the lower fixing frame.

[0028] Preferably, a middle first positioning hole for inserting the projectile body is respectively provided in the center of the upper fixed frame and the lower fixed frame, and a second positioning hole for inserting the projectile wing is respectively provided on the four sides of the upper fixed frame and the lower fixed frame, and the second positioning hole is connected to the first positioning hole.

[0029] Compared with the prior art, the present invention has at least the following advantages:

[0030] The composite material winged projectile structure forming method provided by the present invention has a controllable and reliable process. By controlling the application radius of the structural adhesive on the wing root surface and the accuracy and consistency of the paving in the reinforcement area, the safety and reliability of the winged projectile structure are improved, which is conducive to exerting the ultimate performance of the aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0031] In order to more clearly illustrate the technical solutions in the embodiments of the present invention, the following briefly introduces the drawings required for describing the embodiments.

[0032] Figure 1 This is a flow chart of a composite material winged projectile structure forming method described in an embodiment of the present application;

[0033] Figure 2 This is an axial view of the winged bomb structure described in an embodiment of the present application;

[0034] Figure 3 This is a diagram illustrating the pre-fixation of the missile body and missile wings described in the embodiment of the present application;

[0035] Figure 4 A side view of the winged bomb structure described in an embodiment of the present application;

[0036] Figure 5 A top view of the winged bomb structure described in an embodiment of the present application;

[0037] Figure 6This is a projection diagram of the first layer reinforcement area described in the embodiment of the present application;

[0038] Figure 7 This is a projection diagram of the second layer reinforcement area described in the embodiment of the present application;

[0039] Figure 8 This is a projection diagram of the third layer reinforcement area described in the embodiment of the present application;

[0040] Figure 9 This is a projection diagram of the fourth layer reinforcement area described in the embodiment of the present application;

[0041] Figure 10 This is a projection diagram of the fifth layer reinforcement area described in the embodiment of this application. DETAILED DESCRIPTION

[0042] The present invention will be further described in detail below with reference to the embodiments and drawings so that those skilled in the art can implement the invention with reference to the description.

[0043] refer to Figure 1 As shown, the present application provides a method for forming a composite material winged missile structure, comprising:

[0044] S1. Pre-fixation of missile body and wings:

[0045] Surface treatment of the outer surface of the missile and the bonding area of ​​the wing, such as Figure 2 The winged projectile structure shown mainly includes a projectile body (1), four projectile wings (2) and four reinforcement areas (6). First, the surfaces of the projectile body (1) and the projectile wings (2) involved in the bonding area are polished and wiped clean;

[0046] The missile body (1) and the missile wing (2) are pre-assembled by means of a fixing fixture, and the relative positions of the upper fixing frame and the lower fixing frame are preliminarily determined. Figure 3 This is a diagram showing the pre-fixation of the projectile body (1) and the projectile wing (2). First, the projectile wing (2) and the projectile body (1) are pre-assembled together, and the relative positions of the upper fixing frame (4) and the lower fixing frame (5) are preliminarily determined and adjusted by connecting screws.

[0047] After the position is adjusted, the wing (2) is removed, and the structural glue is applied to the wing root surface where the wing (2) contacts the missile body (1). The final position of the wing, the tilt angle α and the thickness of the structural glue (3) on the wing root surface are determined. The tilt angle α is less than 0.2°, and the thickness of the structural glue (3) on the wing root surface is preferably 0.1mm to 0.3mm. If the tilt angle α is too large, it is easy to produce an excessive rolling rate, which is not conducive to the stability of the aircraft. The whole is placed in an oven for rotation curing, and the missile body (1) and the wing (2) are pre-fixed together;

[0048] S2. Secondary fixation of missile body and wings:

[0049] The pre-fixation of the projectile body (1) and the projectile wing (2) plays a preliminary bonding role, and the side structural glue is also needed for secondary fixing to prepare for the final reinforcement area molding. The smearing radius R of the side structural glue cannot be too small or too large. If R is too small, the bonding strength between the projectile wing (2) and the projectile body (1) cannot be guaranteed, which is not conducive to the reinforcement area molding operation. At the same time, if the radius R is too small, it is easy to make it difficult to apply the structural glue, and if the smearing radius R is too large, it will significantly increase the resistance and reduce the efficiency of the aircraft.

[0050] The side structural adhesive application radius R can be determined according to the following formula:

[0051]

[0052] Where D is the diameter of the projectile, π is the circumference of the circle, To round up;

[0053] The embodiment has a winged bomb structure such as Figure 5 , the diameter D of the projectile is 200 mm, and the radius R of the structural adhesive coating is calculated as follows:

[0054]

[0055] That is, the R value range is 10 to 20, preferably, the radius R is 15 mm, and after the side structural glue is applied, it is placed in an oven for rotation and curing, and the secondary fixation of the elastic body (1) and the elastic wing (2) is completed;

[0056] S3, Forming of the missile body and wing reinforcement areas:

[0057] The molding of the reinforced area of ​​the missile body (1) and the missile wing (2) needs to ensure that there is sufficient bonding strength between the missile wing (1) and the missile body (1) and the deformation amount that meets the design requirements, meet the force requirements of high-speed flight, and at the same time keep the ply positions and angles of the four areas accurate and consistent. The number N of fabric prepreg plies in the reinforced area is an integer greater than or equal to 3, that is, the number of plies in the reinforced area is at least 3. According to the strength and stiffness verification results, the reinforced area plies in this example are 5 layers, all of which are glass fiber plain weave fabric reinforced epoxy resin prepregs. The shapes of the first to fourth layers of fabric prepreg plies are all isosceles trapezoids, such as Figure 6 As shown, the first layer is laid at an angle of 45° and its upper bottom length is That is, the outer projection width of the side structural adhesive is rounded down. The length is 153mm, the oblique side is parallel to the projection side E1 of the oblique side of the wing, and the lower bottom is parallel to the projection line E2 of the same widest part of the two adjacent wing bodies. The shape and size of the first layer of prepreg are determined. When laying, align with the starting laying point (7) of the first layer and align with the outer side of the side structural glue. The first layer needs to lay 4 pieces;

[0058] The sizes of the subsequent second, third and fourth layers of fabric need to be determined based on the values ​​of K, H and M. K is the projected distance between the upper base of the first layer of fabric and the top structural adhesive of the upper wing root surface, H is the projected distance between the oblique edge of the first layer of ply and the E1 edge, and M is the projected distance between the lower base of the first layer of ply and the bottom structural adhesive of the lower wing root surface. Figure 7 As shown, the second layer is laid at an angle of -45°, and its size is enlarged on the basis of the first layer. Its upper base is consistent with the upper base of the first layer in length of 153mm. Its upper base edge, hypotenuse and base edge are parallel to the upper base edge, hypotenuse and base edge of the first layer respectively. The distances in the embodiment are K1, H1 and M1, which are calculated as follows:

[0059]

[0060] That is, the spacing between the upper bases of the first to fourth layers, the spacing between the oblique sides, and the spacing between the lower bases are 66mm, 22mm, and 37mm respectively, and the shape and size of the 2-layer, 3-layer, and 4-layer fabric prepreg layups are determined;

[0061] Figure 7 The second layer of -45° cloth pieces has the same length of 153mm as the upper bottom of the first layer. Its upper bottom edge, bevel edge and bottom edge are parallel to the upper bottom edge, bevel edge and bottom edge of the first layer respectively, with spacing of 66mm, 22mm and 37mm respectively. When laying, align with the starting point (8) of the second layer and align with the outer side of the side structural glue. The second layer needs to lay 4 pieces.

[0062] Figure 8 The third layer of 0° cloth pieces has the same length of 153mm as the upper bottom of the first layer, and its upper bottom edge, oblique side and bottom edge are parallel to the upper bottom edge, oblique side and bottom edge of the second layer respectively, with spacing of 66mm, 22mm and 37mm respectively. When laying, align with the starting point (9) of the third layer and align with the outer side of the side structural glue. The third layer needs to lay 4 pieces;

[0063] Figure 9 The fourth layer of -45° cloth pieces has the same length of 153mm as the upper bottom of the first layer, and its upper bottom edge, oblique side and bottom edge are parallel to the upper bottom edge, oblique side and bottom edge of the third layer respectively, with spacing of 66mm, 22mm and 37mm respectively. When laying, align with the starting point (10) of the fourth layer and align with the outer side of the side structural glue. The fourth layer needs to lay 4 pieces;

[0064] Figure 10 The fifth layer of 45° fabric is laid, with the paving area covering the overall reinforcement area (6) and aligned with the starting paving point (11) of the fifth layer;

[0065] The long-term operating temperature of the projectile body, projectile wings, and the structural adhesive used in the bonding process must be higher than the curing temperature of the fabric prepreg to ensure that the performance of the projectile wings (2) and the projectile body (1) does not significantly decrease and the relative positions do not shift during the curing of the fabric prepreg. In all the layers from the first to the fifth layer, the paving area must not be larger than the reinforcement area (6), and the excess prepreg needs to be cut off during paving.

[0066] After all the layers of the four reinforcement areas (6) are laid, an isolation film, a release cloth and a breathable felt are laid in sequence, the bag is placed in a vacuum bag, vacuumed and placed in an oven, and after curing, the auxiliary materials are removed to obtain the product.

[0067] Through the above technical scheme, the composite material winged projectile structure forming method provided by the present invention has a controllable and reliable process. By controlling the application radius of the structural glue on the wing root surface and the accuracy and consistency of the paving in the reinforcement area, the weight uniformity of the winged projectile structure at the axis is improved while meeting the connection strength between the wing and the projectile, thereby improving the overall safety and reliability of the winged projectile structure, which is conducive to exerting the ultimate performance of the aircraft.

[0068] Various modifications to these embodiments will be readily apparent to those skilled in the art, and the general principles defined herein may be implemented in other embodiments without departing from the spirit or scope of the invention. Therefore, the present invention is not limited to the embodiments shown herein, but is to be construed in the widest manner consistent with the principles and novel features disclosed herein.

Claims

1. A method for forming a composite material winged projectile structure, characterized by: The following steps are involved: S1. Pre-fixation of missile body and wings: Surface treatment of the outer surface of the missile and the bonding area of ​​the wing; A fixing fixture is provided, which includes an upper fixing frame and a lower fixing frame. The missile body and the missile wings are pre-assembled through the fixing fixture, and the relative positions of the upper fixing frame and the lower fixing frame are preliminarily adjusted to determine; After the relative positions of the upper and lower fixing frames are adjusted, remove the wing and apply structural adhesive to the wing root surface where the wing directly contacts the missile body. Determine the final position of the wing on the missile body. The wing inclination α must be less than 0.2°. After the structural adhesive cures, pre-fix the wing and missile body together. S2. Secondary fixation of missile body and wings: The side structural adhesive application radius R is determined according to the diameter of the projectile body. The side structural adhesive is cured and the projectile wing and the projectile body are fixed together for the second time. S3, Forming of the missile body and wing reinforcement areas: The body of the projectile is divided into four reinforcement areas by four wings. Multiple fabric prepregs of different sizes are laid on the reinforcement areas in order from small to large, put into vacuum bags, and placed in an oven for vacuuming, heating and curing to obtain the product. In S2, the side structural adhesive application radius R can be determined according to the following formula: ； Where D is the diameter of the projectile, π is the circumference of the circle, To round up; In S3, laying the fabric prepregs of different sizes and quantities in the reinforcement area in sequence specifically includes the following steps: Determine the number of fabric prepreg layers N in the reinforcement area based on the overall structural strength and stiffness, where N is ≥ 3 and is an integer; The shapes of the prepreg layers from the first layer to the N-1 layer are all isosceles trapezoids. The shape of the first layer of prepreg layers has an upper and lower length of , its hypotenuse is parallel to the projected side E1 of the wing side hypotenuse, and its lower base is parallel to the projection line E2 of the widest part of the two adjacent wing bodies. Subsequent layers are expanded on the basis of the first layer. The upper base, lower base and hypotenuse of the second to N-1 layers are respectively parallel to the upper base, lower base and hypotenuse of the first layer. The lengths of the upper bases of the first to N-1 layers are equal. The distances between the upper bases of any two adjacent layers from the first to N-1 layers are equal. The distances between the hypotenuses of any two adjacent layers from the first to N-1 layers are equal. The distances between the lower bases of any two adjacent layers from the first to N-1 layers are equal, which are denoted as K1, H1 and M1 respectively. The Nth layer paving area is the overall reinforcement area, where K1, H1 and M1 are determined according to the following formulas: ； Among them, K is the distance between the upper bottom of the first layer and the top structural adhesive projection of the upper wing root surface, H is the distance between the oblique side of the first layer and the E1 edge, and M is the distance between the lower bottom of the first layer and the bottom structural adhesive projection of the lower wing root surface. To round down.

2. The method for forming a composite material winged projectile structure according to claim 1, characterized in that: In S3, a plurality of fabric prepregs of different sizes are laid in order from small to large on the reinforcement area, and then covered with an isolation film, a release cloth and a breathable felt, placed in a vacuum bag, placed in an oven for vacuum heating and curing, and the auxiliary materials are removed after curing to obtain a product.

3. The method for forming a composite material winged projectile structure according to claim 1, characterized in that: The fabric prepreg is a reinforced fiber prepreg, and the resin of the prepreg is a thermosetting resin.

4. The method for forming a composite material winged projectile structure according to claim 3, characterized in that: The thermosetting resin is epoxy resin, phenolic resin, bismaleimide resin, or cyanate resin, and the reinforcing fiber is one or more of carbon fiber, glass fiber, basalt fiber, and aramid fiber.

5. The method for forming a composite material winged projectile structure according to claim 1, characterized in that: The long-term use temperature of the elastic body, the elastic wing and the structural adhesive is greater than the curing temperature of the fabric prepreg.

6. The method for forming a composite material winged projectile structure according to claim 1, characterized in that: The fixing fixture also includes a connecting screw for adjusting the positions of the upper fixing frame and the lower fixing frame.

7. The method for forming a composite material winged projectile structure according to claim 6, characterized in that: A middle first positioning hole for inserting the projectile body is respectively provided in the center of the upper fixing frame and the lower fixing frame, and a second positioning hole for inserting the projectile wing is respectively provided on the four sides of the upper fixing frame and the lower fixing frame, and the second positioning hole is connected to the first positioning hole.

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