Composite skirt connecting structure of solid rocket engine and forming method of composite skirt connecting structure

Abstract

The invention discloses a solid rocket engine composite material skirt connecting structure and a forming method thereof.The connecting structure comprises a composite material skirt, a skirt hole reinforcing ring, a connected cabin section and a connecting piece, the composite material skirt and the skirt hole reinforcing ring are integrally formed, and the rear end of the composite material skirt and the rear end of the skirt hole reinforcing ring are connected with the connected cabin section through the connecting piece; the connected cabin section is of an L-shaped nested inner lap joint structure and has the butt joint limiting and bearing functions, and the thickness transition section at the front end of the composite skirt is connected with a fiber winding structure layer of a combustion chamber shell. Compared with a traditional metal skirt connecting structure, the weight reduction effect is obvious, compared with a traditional composite skirt hinge block connecting structure, the connecting mode is simpler, and the fiber continuity of a composite skirt connecting area is guaranteed; the problems that in-plane shear rigidity and strength of the composite material are insufficient, and orifice positions are prone to damage are solved. The structure and the forming process are simple, the maturity is high, and the cost is low.

Description

Technical Field

[0001] The invention relates to the field of solid rocket engines, and in particular to a solid rocket engine composite material skirt connection structure and a molding method thereof. Background Art

[0002] With the continuous development of material science and molding technology, fiber-reinforced composite materials have gradually replaced traditional metal materials due to their high specific strength, high specific modulus and lightweight performance advantages, and have become the main material for solid rocket engine shells. The skirt is an important component of the solid rocket engine as a carrier that connects the combustion chamber shell of the solid rocket engine to other compartments. When the engine is working, the connection position is often subjected to the combined action of complex loads such as bending moment, axial pressure, axial tension, shear, and temperature shock. The reliability of the connection structure is crucial. Generally speaking, the engine axial pressure load is much greater than the axial tension load, and the bending moment load caused by the lateral overload during large engine maneuvers is also the main influencing factor causing the engine to disintegrate in the air. Therefore, the engine load should be comprehensively considered when designing the skirt connection.

[0003] Composite skirts have a significant weight reduction effect compared to traditional metal skirts, and can effectively improve the mass ratio of the engine. However, due to the mechanical properties of composite materials such as transverse isotropy, low shear stiffness, and low shear strength, as well as the process defects of traditional composite skirt column sections and connection sections, such as discontinuous thickness and overlap of layers, the actual performance of composite skirts is greatly different from the theoretical design, and the connection design is difficult. Therefore, the solid rocket engine composite skirt connection structure and its forming method proposed in the patent of this invention comprehensively consider the above problems and effectively improve the reliability of the composite skirt connection structure. Summary of the invention

[0004] The purpose of the present invention is to provide a solid rocket engine composite skirt connection structure and a forming method thereof, so as to solve the problems of difficult forming process of composite skirt, complex connection structure, low structural strength and low reliability of connection structure. The specific technical scheme is as follows:

[0005] The present invention provides a composite skirt connection structure for a solid rocket engine, the connection structure comprising a composite skirt, a skirt hole reinforcement ring, a connected cabin section, and a connecting piece; wherein the composite skirt and the skirt hole reinforcement ring are integrally formed; the composite skirt and the rear end of the skirt hole reinforcement ring are connected to the connected cabin section through a connecting piece; the connected cabin section is an L-shaped nested inner overlap structure, which has docking limit and load-bearing functions; the front end thickness transition section of the composite skirt is connected to the fiber winding structure layer of the combustion chamber shell.

[0006] Furthermore, the connecting member is any one of a countersunk screw, a small pan head screw, a pan head screw and a bearing structure pin.

[0007] Furthermore, when the connecting member is a load-bearing structure pin, the connecting hole of the connected compartment section is an ordinary hole and 1 to 2 mm is reserved at the bottom for limiting; when the connecting member is a countersunk screw, a small pan head screw, or a pan head screw, the connecting hole of the connected compartment section is a threaded through hole.

[0008] Furthermore, the front thickness transition section of the composite skirt and the fiber winding structure layer of the combustion chamber shell are obtained by the following connection method:

[0009] S1. Apply nitrile rubber-based adhesive interface glue to the inner surface of the docking position, and then connect the thickness transition section of the front end of the composite skirt and the fiber winding structure layer of the combustion chamber shell, wherein the thickness of the adhesive layer is greater than or equal to 0.3 mm;

[0010] S2. Apply nitrile rubber-based adhesive interface glue on the upper surface of the docking position and wrap the fixed layer of the composite skirt.

[0011] Furthermore, the fixing layer is composed of hoop-wound fibers and an axial reinforcement layer, and the proportion of the axial reinforcement layer is adjusted according to the axial load requirement.

[0012] Furthermore, the composite material skirt is laid out according to a blanket curve method, and the ±45° laying ratio is not less than 50%; the skirt hole reinforcement ring (2) is staggered with double rows of holes when the load is severe.

[0013] The present invention also provides a method for forming a composite skirt connection structure of a solid rocket engine, wherein the method uses a composite skirt and a skirt hole reinforcement ring to be integrally formed, and the specific steps are: taking a positive mold and a negative mold of a mold, laying fiber layers layer by layer on the surface of the positive mold, and the transition position has a smooth profile, stacking from the 1st layer to the nth layer, where n is an integer greater than or equal to 2, and sandwiching a thickness transition layer between two layers of continuous layered parts, assembling a skirt hole reinforcement ring blank after laying, and after the assembly is completed, closing the mold with the negative mold, and high-temperature curing and compression molding to obtain an integrally formed composite skirt blank and skirt hole reinforcement ring blank.

[0014] Furthermore, the integrally formed composite material skirt blank and skirt hole reinforcement ring blank are further machined to obtain an integrally formed composite material skirt and skirt hole reinforcement ring.

[0015] Furthermore, the thermal expansion coefficients of the male and female molds match those of the composite skirt; the thermal expansion coefficient of the skirt hole reinforcement ring blank is higher than that of the composite skirt blank, and during the process of high-temperature curing and cooling to room temperature, the two achieve interference fit, which is used to improve the connection reliability between the two.

[0016] Furthermore, the skirt hole reinforcement ring blank is made of alloy steel and titanium alloy. Before assembly, the mating surface of the skirt hole reinforcement ring blank is sandblasted and coated with interface glue. The thickness of the interface glue is greater than or equal to 0.2 mm, and a 60°C insulation treatment is used during assembly. The female mold includes 3 to 6 circumferential petals to achieve uniform distribution of pressure during the curing process.

[0017] The beneficial effects of the present invention compared with the prior art are:

[0018] (1) The present invention has a significant weight reduction effect compared to the traditional metal skirt connection structure, and a simpler connection method compared to the traditional composite skirt hinge block connection structure;

[0019] (2) The present invention has high structural reliability. The design ensures the fiber continuity of the composite skirt connection area, and adopts the skirt hole reinforcement ring as a metal reinforcement to solve the problems of insufficient in-plane shear stiffness and strength of the composite material and easy damage to the hole position;

[0020] (3) The present invention has high adaptability and can be applied to engines of various projectile diameters by adopting radial connection with screws, thus solving the connection problem of composite shells with small projectile diameters;

[0021] (4) The present invention has a simple structure and molding process, is easy to install and assemble, has a high degree of process maturity, is low in cost, and has good economy. BRIEF DESCRIPTION OF THE DRAWINGS

[0022] Figure 1 It is a schematic diagram of the connection structure of the composite material skirt of the solid rocket engine of the present invention.

[0023] Figure 2 It is a schematic diagram of preparing a composite material skirt of a solid rocket engine composite material skirt equipped with a metal skirt hole reinforcement ring according to the present invention.

[0024] Figure 3 It is a schematic diagram of the integrated molding method of the composite material skirt and skirt hole reinforcement ring of the solid rocket engine of the present invention.

[0025] The accompanying drawings are marked as follows: 1. composite material skirt; 2. skirt hole reinforcement ring; 3. connected compartment section; 4. connecting piece; 5. female mold; 6. skirt hole reinforcement ring blank; 7. male mold; 8. composite material skirt blank. DETAILED DESCRIPTION

[0026] The specific embodiments of the present invention are further described in detail below with reference to the accompanying drawings.

[0027] The technical solutions in the embodiments of the present invention will be described clearly and completely below in conjunction with the drawings 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. Therefore, the present invention is not limited to the specific embodiments disclosed below.

[0028] Figure 1 It is a schematic diagram of a connection structure of a composite skirt of a solid rocket engine. The present invention provides a connection structure of a composite skirt of a solid rocket engine, the connection structure comprising a composite skirt 1, a skirt hole reinforcement ring 2, a connected compartment section 3, and a connector 4; wherein the composite skirt 1 and the skirt hole reinforcement ring 2 are integrally formed; the rear ends of the composite skirt 1 and the skirt hole reinforcement ring 2 are connected to the connected compartment section 3 through the connector 4; the connected compartment section 3 is an L-shaped nested inner overlap structure, which has docking limit and load-bearing functions; the front end thickness transition section of the composite skirt 1 is connected to the fiber winding structure layer of the combustion chamber shell.

[0029] The thickness of the composite skirt 1 is H≥k*(2π*R / (2π*Rn*d))*T / (2π*R*σ C ), where k is the design safety factor, k≥2; T is the circumferential tension or axial tension; R is the outer diameter of the composite skirt; n is the number of circumferential openings; d is the aperture, d is 0.5H~1.5H, σ C is the axial tensile strength of the composite skirt σ φ Or hoop tensile strength σ θ ; The thickness h of the skirt hole reinforcement ring is 0.15H~0.35H; The distance between the hole and the end face is e≥T / (2*(Hh)*τ c ); the arc length between the openings w≥d+[T / (σ θ *(Hh))]; the skirt hole reinforcement ring overlap length l≥2e; the axial stiffness E φ , hoop stiffness E θ , in-plane shear stiffness G, axial tensile strength σ φ , hoop tensile strengthσ θ , in-plane shear strength τ c It is calculated based on the equivalent stiffness theory of composite materials; the skirt hole reinforcement ring is staggered with double rows of holes when the load is severe, wherein the row spacing B≥2d.

[0030] When the connecting member 4 is a bearing structure pin, the connecting hole of the connected compartment section 3 is a common hole and a 1-2 mm hole is reserved at the bottom for position limiting; when the connecting member 4 is a countersunk screw, a small pan head screw, or a pan head screw, the connecting hole of the connected compartment section 3 is a threaded through hole. In this embodiment, the connecting structure uses a countersunk screw, and the corresponding connecting hole of the connecting compartment section is a threaded M8 through hole.

[0031] The method for connecting the front thickness transition section of the composite skirt 1 and the fiber winding structure layer of the combustion chamber shell is as follows: S1, brushing the nitrile rubber-based adhesive interface glue on the inner surface of the docking position, and then connecting the front thickness transition section of the composite skirt 1 and the fiber winding structure layer of the combustion chamber shell, wherein the thickness of the adhesive layer is greater than or equal to 0.3mm; S2, brushing the nitrile rubber-based adhesive interface glue on the upper surface of the docking position, and winding the fixed layer of the composite skirt 1. The fixed layer is composed of hoop-wound fibers and an axial reinforcement layer, and the proportion of the axial reinforcement layer is adjusted according to the axial load requirements.

[0032] In this embodiment, the thickness of the adhesive layer is 0.3 mm, the skirt fixing layer is composed of circumferentially wound fibers and an axial reinforcement layer, the thickness of the circumferential single layer is 0.2 mm, and 0.4 mm thick carbon fiber axial reinforcement cloth is sandwiched between every two circumferentially wound layers, with a total thickness of 2.4 mm.

[0033] like Figure 2 The schematic diagram of preparing a composite skirt of a solid rocket engine composite skirt with a metal skirt hole reinforcement ring is shown in the figure. In this embodiment, the specific design method of the ply, thickness and opening of the composite skirt 1 is as follows:

[0034] The skirt material is T800 grade carbon fiber prepreg with a tensile strength of 2600MPa in the fiber direction. The ply is designed according to the blanket curve method. The ±45° ply ratio is less than 55%, the 0° ply ratio is less than 30%, and the 90° ply ratio is less than 15%. The ply sequence is [90 / ±45 / 0 / ±45 / 0]n.

[0035] The axial stiffness E φ , hoop stiffness E θ , in-plane shear stiffness G, axial tensile strength σ φ , hoop tensile strengthσ θ , in-plane shear strength τ c Calculated based on composite material equivalent stiffness theory;

[0036] Skirt thickness H ≥ k*(2π*R / (2π*Rn*d))*T / (2π*R*σ C), n = 24, d = 8mm, k = 2, R = 250mm, T = 500kN, according to the calculation results of the axial tensile load and combined with the process conditions, the skirt thickness H is 4mm; the thickness of the skirt hole reinforcement ring 2 h = 1mm; the distance between the hole and the end face e ≥ T / (2n*(Hh)*τ c ) is 16 mm, and when n is 24, the arc length between the openings w≥d+[T / (n*σ θ *(Hh))], the lap length of the skirt hole reinforcement ring is l = 30mm.

[0037] like Figure 3 As shown in the schematic diagram of the integrated molding method of the composite material skirt and the skirt hole reinforcement ring of a solid rocket engine, the present invention also provides a method for molding a composite material skirt connection structure of a solid rocket engine, the method adopts a composite material skirt 1 and a skirt hole reinforcement ring 2 to be integrally molded, and the specific steps are:

[0038] Take the male mold 7 and the female mold 5, and lay the fiber layers on the surface of the male mold 7 layer by layer. The transition position has a smooth profile, and the layers are stacked from the first layer (lay1) to the nth layer (layn). A thickness transition layer is interposed between two layers of continuous layers. After laying, the skirt hole reinforcement ring blank 6 is assembled. After the assembly is completed, the female mold 5 is matched to close the mold, and the composite material skirt blank 8 and the skirt hole reinforcement ring blank 6 are obtained by high temperature curing and compression molding. In this embodiment, the mold material is a glass fiber / epoxy resin composite material.

[0039] The integrally formed composite skirt blank 8 and skirt hole reinforcement ring blank 6 are further machined to obtain an integrally formed composite skirt 1 and skirt hole reinforcement ring 2 .

[0040] The thermal expansion coefficients of the male mold 7 and the female mold 5 match the thermal expansion coefficient of the composite skirt 1; the thermal expansion coefficient of the skirt hole reinforcement ring blank 6 is higher than the thermal expansion coefficient of the composite skirt blank 8, and the two achieve interference fit during the process of cooling to room temperature after high-temperature curing, which is used to improve the connection reliability between the two.

[0041] The skirt hole reinforcement ring blank 6 is made of alloy steel and titanium alloy. Before assembly, the matching surface of the skirt hole reinforcement ring blank 6 is sandblasted and coated with interface glue. The thickness of the interface glue is greater than or equal to 0.2 mm. During assembly, a 60° C. insulation treatment is used. In this embodiment, the skirt hole reinforcement ring uses a TC4 titanium forging ring, and the thickness of the interface glue is 0.2 mm.

[0042] The female mold 5 comprises 3 to 6 circumferential petals, which are used to achieve uniform distribution of pressure during the curing process. In this embodiment, the female mold consists of 4 circumferential petals.

[0043] The contents not described in detail in this specification belong to the prior art known to those skilled in the art. It is obvious to those skilled in the art that the present invention is not limited to the details of the exemplary embodiments described above, and that the present invention can be implemented in other specific forms without departing from the spirit or essential features of the present invention. Therefore, no matter from which point of view, the embodiments should be regarded as exemplary and non-restrictive, and the scope of the present invention is defined by the appended claims rather than the above description, and it is intended that all changes falling within the meaning and scope of the equivalent elements of the claims are included in the present invention.

Claims

1. A solid rocket engine composite skirt connection structure, characterized in that: The connection structure comprises a composite material skirt (1), a skirt hole reinforcement ring (2), a connected cabin section (3), and a connection piece (4); wherein: The composite material skirt (1) and the skirt hole reinforcement ring (2) are integrally formed; The composite material skirt (1) and the rear end of the skirt hole reinforcement ring (2) are connected to the connected cabin section (3) via a connecting piece (4); The connected compartment sections (3) are L-shaped nested inner overlapping structures, which have the functions of docking limit and load bearing; The front thickness transition section of the composite material skirt (1) is connected to the fiber winding structure layer of the combustion chamber shell.

2. A solid rocket engine composite skirt connection structure according to claim 1, characterized in that: The connecting member (4) is any one of a countersunk screw, a small pan head screw, a pan head screw and a bearing structure pin.

3. A solid rocket engine composite material skirt connection structure according to claim 2, characterized in that: When the connecting member (4) is a bearing structure pin, the connecting hole of the connected compartment section is a common hole and a 1 to 2 mm hole is reserved at the bottom for limiting; when the connecting member (4) is a countersunk screw, a small pan head screw, or a pan head screw, the connecting hole of the connected compartment section is a threaded through hole.

4. A solid rocket engine composite skirt connection structure according to claim 1, characterized in that: The front end thickness transition section of the composite material skirt (1) and the fiber winding structure layer of the combustion chamber shell are obtained by the following connection method: S1, applying a nitrile rubber-based adhesive interface glue to the inner surface of the docking position, and then connecting the front end thickness transition section of the composite skirt (1) and the fiber winding structure layer of the combustion chamber shell, wherein the thickness of the adhesive layer is greater than or equal to 0.3 mm; S2. Apply nitrile rubber-based adhesive interface glue to the upper surface of the docking position, and wrap the fixed layer of the composite material skirt (1).

5. A solid rocket engine composite material skirt connection structure according to claim 4, characterized in that: The fixing layer is composed of hoop-wound fibers and an axial reinforcement layer, and the proportion of the axial reinforcement layer is adaptively adjusted according to the axial load requirements.

6. A solid rocket engine composite material skirt connection structure according to claim 1, characterized in that: The composite material skirt (1) is laid out according to the blanket curve method, with the ±45° laying ratio being not less than 50%; the skirt hole reinforcement ring (2) is staggered with double rows of holes when the load is severe.

7. The method for forming a solid rocket engine composite material skirt connection structure according to claim 1, characterized in that: The method adopts a composite material skirt (1) and a skirt hole reinforcement ring (2) to be integrally formed, and comprises the following steps: Take a male mold (7) and a female mold (5), and lay fiber layers on the surface of the male mold (7) layer by layer. The transition position has a smooth profile, and the layers are stacked from the first layer to the nth layer, where n is an integer greater than or equal to 2. A thickness transition layer is interposed between two layers of continuous layers. After laying, a skirt hole reinforcement ring blank (6) is assembled. After assembly, the female mold (5) is matched to close the mold, and high-temperature curing and compression molding are performed to obtain an integrally formed composite material skirt blank (8) and skirt hole reinforcement ring blank (6).

8. The method for forming a solid rocket engine composite material skirt connection structure according to claim 5, characterized in that: The integrally formed composite material skirt blank (8) and skirt hole reinforcement ring blank (6) are further machined to obtain an integrally formed composite material skirt (1) and skirt hole reinforcement ring (2).

9. A method for forming a solid rocket engine composite material skirt connection structure according to claim 8, characterized in that: The thermal expansion coefficients of the male mold (7) and the female mold (5) match the thermal expansion coefficient of the composite material skirt (1); the thermal expansion coefficient of the skirt hole reinforcement ring blank (6) is higher than the thermal expansion coefficient of the composite material skirt blank (8), and the two achieve interference fit during the process of cooling from high temperature curing to room temperature.

10. The method for forming a solid rocket engine composite material skirt connection structure according to claim 7, characterized in that: The skirt hole reinforcement ring blank (6) is made of alloy steel or titanium alloy. Before assembly, the mating surface of the skirt hole reinforcement ring blank (6) is sandblasted and coated with interface glue. The thickness of the interface glue is greater than or equal to 0.2 mm. During assembly, a 60° C. insulation treatment is used. The female mold (5) includes 3 to 6 circumferential petals for achieving uniform distribution of pressure during the curing process.

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