A solid propellant grain structure and its preparation method

Abstract

The present invention provides a solid propellant grain structure, including a side burning-limiting layer, an end face burning-limiting layer, and a propellant grain. The propellant grain is of a cylindrical structure. The side burning-limiting layer and the end face burning-limiting layer are respectively fixed to the side and the end face of the propellant grain. A plurality of round holes are formed in the end face burning-limiting layer. The design of this invention is to form a plurality of round holes in the end face burning-limiting layer. By adjusting the relationship between the radius of the round holes and the web thickness of the propellant grain, the burning surface doubling rate can reach more than 10 times, effectively solving the problem that the existing charge shape cannot achieve a high doubling of the burning surface of the grain under the condition of a small web thickness of the grain.

Description

Technical Field

[0001] The present invention belongs to the technical field of solid propellant charging, and particularly relates to a solid propellant grain structure and a preparation method thereof. Background Art

[0002] Ejection refers to "using elastic force to release or eject an object from an inclined track, a launcher or other devices". Ejection technology has many applications in the military field, and the ejected objects include missiles, shipborne aircraft, unmanned aerial vehicles, aviation rescue seats, rockets, etc. The basic principle of ejection is, taking the most commonly used gas ejection as an example, in the space behind the missile in the launch tube or in other special cylindrical devices, the gas expands and acts on the pressure-bearing surface to form an ejection force. After being directly or indirectly transmitted to the missile, it pushes the missile to move forward along the director to complete the launch process. From the perspective of energy conversion, the ejection process is essentially a process in which the chemical energy of gunpowder (propellant grain) is burned in a very short time (tens of milliseconds or hundreds of milliseconds) to be converted into the thermal energy of combustion products, and then through the expansion work of the combustion products, it is further converted into the kinetic energy of the movement of the missile, gas and launch tube. The gas generator is a device that converts the chemical energy of the propellant grain into thermal energy.

[0003] A gas generator generally consists of an igniter, a combustion chamber, a propellant grain and a nozzle. Through the combustion of the propellant grain, high-temperature and high-pressure gas is generated in the combustion chamber, and then the gas is discharged into a specified cavity through the nozzle, such as the inner cavity at the bottom of the missile launch tube. Then the gas does work in the launch tube to push the missile out of the launch tube.

[0004] At present, there are many charging structures for gas generators, such as star-shaped, finned-columnar, tube-grooved, tubular, wheel-shaped, spherical and other grain shapes. When a missile is launched using a gas generator, it is often necessary to control the launch recoil force and the missile ejection speed from the tube. The most important factor in controlling the recoil force is to control the gas pressure in the launch tube, and the most important factor determining this gas pressure is the internal pressure of the gas generator combustion chamber. The internal pressure of the gas generator combustion chamber is generated by the combustion of the propellant grain and can be calculated by formula (1). From formula (1), it can be seen that the internal pressure of the combustion chamber is mainly determined by the burning surface of the propellant grain. Therefore, it can be said that the burning surface of the grain basically determines the magnitude of the missile launch recoil force when the grain burns.

[0005]

[0006] Wherein, P c is the combustion chamber pressure;

[0007] ρ p , a, c*, n are all parameters characterizing the performance of the grain propellant;

[0008] A b is the grain burning area;

[0009] A t is the nozzle throat area.

[0010] When controlling missile launch, the launch recoil force must not exceed the required value. Since the free volume in the launch tube continuously increases as the missile moves forward, and the ejection time is extremely short, in the order of milliseconds. According to the theoretical formula of the internal ballistics of the engine, a charge structure with a small web thickness and a high burning surface doubled must be designed. However, the current charge structures are difficult to meet the requirements. The spherical charge shape has the largest burning surface doubling among common charge shapes, and its burning surface is proportional to the square of the ball radius. Assuming that the web thickness of the charge column is only 3 mm and the initial ball radius is 3 mm, the maximum burning surface doubling rate is only 4 times, and it is very difficult in terms of process forming and product installation. When the missile is launched, the burning surface doubling rate often needs to reach more than 5 times. Therefore, the conventional charge shape design method is difficult to meet the requirements. Summary of the Invention

[0011] The purpose of the present invention is to provide a solid propellant charge column structure to solve the problem that the existing charge shapes cannot achieve a high doubling of the burning surface of the charge column under the condition of a small web thickness of the charge column.

[0012] To this end, the present invention adopts the following technical solutions:

[0013] A solid propellant charge column structure includes a side flame-limiting layer, an end face flame-limiting layer, and a propellant charge column. The propellant charge column is of a cylindrical structure. The side flame-limiting layer and the end face flame-limiting layer are respectively fixed to the side and end face of the propellant charge column, and a plurality of circular holes are provided on the end face flame-limiting layer.

[0014] Further, the radius of the circular hole on the end face flame-limiting layer is r, and the length of the propellant charge column is 2e, satisfying the relationship: where e is the web thickness of the propellant charge column.

[0015] Further, the radius r of the circular hole on the end face flame-limiting layer is 3 mm, and the web thickness e of the propellant charge column is 5 mm.

[0016] Further, the shortest distance between the centers of any two circular holes on the end face flame-limiting layer is not less than 2(r + e).

[0017] Further, the shortest distance from the center of the outermost circular hole on the end face flame-limiting layer to the inner side wall of the side flame-limiting layer is not less than (r + e).

[0018] Further, the circular holes on the two end faces of the end face flame-limiting layer are symmetrically arranged.

[0019] Further, the side flame-limiting layer and the end face flame-limiting layer are made of a heat-insulating and non-combustible non-metallic material.

[0020] Further, the propellant grain comprises an energetic non-metallic material composed of a fuel and an oxidizer mixed together.

[0021] In addition, the present invention also provides a preparation method for the above-mentioned solid propellant grain structure, which comprises the following steps:

[0022] 1) Fabricate a lateral combustion-limiting layer in the shape of a hollow cylinder, and fabricate an end combustion-limiting layer, and drill a round hole in the end combustion-limiting layer;

[0023] 2) Pour the propellant raw material of the propellant grain into the lateral combustion-limiting layer. After the propellant raw material is solidified into a grain, cut it to a designed length to form a disc shape;

[0024] 3) Bond the end combustion-limiting layer to the disc-shaped propellant grain obtained in step 2).

[0025] Compared with the prior art, the beneficial effects of the present invention are as follows:

[0026] (1) By designing and drilling a plurality of round holes in the end combustion-limiting layer of the solid propellant grain structure provided by the present invention, and by adjusting the relationship between the radius of the round hole and the web thickness of the propellant grain, the burning surface doubling rate can reach more than 10 times, effectively solving the problem that the existing charge grain shape cannot achieve a high burning surface doubling of the grain under the condition of a small web thickness of the grain.

[0027] (2) The preparation method of the solid propellant grain structure provided by the present invention is simple, effectively overcoming the problems of the existing other grain shapes that need to open holes or wings on the propellant grain, require forming molds, the forming process is relatively complex, and the quality is difficult to guarantee.

[0028] The present invention will be further described in detail below with reference to the accompanying drawings. Description of the Drawings

[0029] Figure 1 is the end view of the solid propellant grain structure of the present invention;

[0030] Figure 2 is Figure 1 the sectional view of the solid propellant grain structure along the A-A direction in

[0031] Description of the reference numerals in the drawings: 1, lateral combustion-limiting layer; 2, end combustion-limiting layer; 3, round hole; 4, propellant grain. Detailed Embodiments

[0032] Next, the technical solutions in the embodiments of the present invention will be clearly and completely described in conjunction with the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are only a part of the embodiments of the present invention, rather than all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative efforts shall fall within the protection scope of the present invention.

[0033] In the description of the present invention, it should be understood that the orientation or positional relationship indicated by the terms "center", "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc. is based on the orientation or positional relationship shown in the drawings, and is only for the convenience of describing the present invention and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore should not be construed as a limitation to the present invention.

[0034] In the description of the present invention, it should be noted that unless otherwise clearly specified and limited, the terms "installation", "connection", and "connection" should be understood in a broad sense. For example, it can be a fixed connection, a detachable connection, or a contact connection or an integral connection; for those of ordinary skill in the art, the specific meanings of the above terms in the present invention can be understood according to specific circumstances; in the description of the present invention, unless otherwise stated, the meaning of "a plurality" is two or more.

[0035] As Figure 1 and Figure 2 shown, this embodiment provides a solid propellant grain structure, including a side burning-limiting layer 1, an end face burning-limiting layer 2, and a propellant grain 4. The propellant grain 4 is a cylindrical structure. The side burning-limiting layer 1 and the end face burning-limiting layer 2 are respectively fixed on the side and end face of the propellant grain 4. A plurality of round holes 3 are opened on the end face burning-limiting layer 2. Through the round holes 3 opened on the end face burning-limiting layer 2, the propellant grain 4 is exposed, and the area of all the exposed propellant grains 4 is the initial burning surface.

[0036] Among them, the side burning-limiting layer 1 and the end face burning-limiting layer 2 are made of non-metallic materials with good heat insulation performance and non-combustibility; the propellant grain 4 is made of an energetic non-metallic material composed of a mixture of a fuel and an oxidizer and other substances, which can burn to generate a large amount of gas. The raw materials of the propellant grain 4 can adopt the propellant components of the existing grains, and the specific components are not described here in detail.

[0037] According to requirements such as the recoil force and velocity of the missile when it exits the tube, combined with characteristics such as the burning rate of the propellant, a variable burning surface that meets the requirements is designed. The initial burning surface is the sum of the areas of all the circular holes 3 on the end face limited combustion layer 2. As the gas generator igniter is ignited by the missile current, the propellant grain 4 is thus ignited. The burning surface of the grain continuously advances forward according to the parallel layer combustion theory, and a new burning surface and the corresponding internal pressure are formed until the entire propellant grain 4 is burned out. In this embodiment, circular holes 3 are provided on the end face limited combustion layers 2 at both ends of the side limited combustion layer 1 to increase the burning area and ensure that the total burning area meets the requirements. The circular holes 3 on the end face limited combustion layers 2 at both ends are symmetrically arranged. At the same time, it is set that the radius of each circular hole 3 on the end face limited combustion layer 2 is r, and the length of the propellant grain is 2e, where e is the web thickness of the propellant grain 4. Based on the geometric structure of the solid propellant grain structure designed here, the maximum doubling rate of the burning surface is calculated. Among them, r is the radius of the circular hole on the end face limited combustion layer, e is the web thickness of the propellant grain, S0 is the area of the initial burning surface, and S1 is the area of the maximum advancing burning surface formed by the burning surface of the grain according to the parallel layer combustion theory.

[0038] When the missile is launched, a burning surface doubling rate of more than 5 times is required. Using the solid propellant grain structure of this embodiment, the following relationship is satisfied: That's all. For the ejection of small-mass missiles, since the working time of the gas generator is extremely short, in the millisecond level, and the amount of propellant is small, only about 10 g, the web thickness e of the propellant grain 4 is generally 2 mm to 5 mm. Therefore, by designing an appropriate radius of the circular hole 3 and the web thickness of the propellant grain 4, a grain shape with a small web thickness and a high doubling rate of the burning surface can be achieved to meet the engineering requirements. For example, when the radius r of the circular hole 3 on the end face limited combustion layer 2 is designed to be 3 mm and the web thickness e of the propellant grain 4 is 5 mm, the calculated maximum doubling rate n of the burning surface can reach 16 times.

[0039] To optimize the above technical solution, since the burning area of a single circular hole is small and cannot meet the actual needs, multiple circular holes 3 are provided on the end face limited combustion layer 2 to increase the burning area. For the distribution of multiple circular holes 3 on the end face limited combustion layer 2, it is designed that the shortest distance between the centers of any two circular holes 3 on the end face limited combustion layer 2 is not less than 2(r + e), so as to ensure that during the combustion process of the grains of any two circular holes 3, their burning surfaces will not intersect in advance, thus ensuring the required burning area and burning surface doubling rate. And it is designed that the shortest distance from the center of the outermost circular hole 3 on the end face limited combustion layer 2 to the inner wall of the side limited combustion layer 1 is not less than (r + e), so as to ensure that during the combustion process of the grain of the outermost circular hole 3, its burning surface will not burn to the side limited combustion layer 1 (the side limited combustion layer is non-combustible) in advance, and the burning surface will not shrink in advance, thus ensuring the required burning area and burning surface doubling rate.

[0040] When manufacturing the solid propellant grain structure of this embodiment, first fabricate the lateral restricted combustion layer 1 in the shape of a hollow cylinder, and then pour the propellant raw material of the propellant grain 4 into the hollow cylinder. After the propellant raw material solidifies into a grain, it is truncated into a disc shape according to the designed length; drill a round hole 3 in the end restricted combustion layer 2, and then bond the end restricted combustion layer 2 to the disc-shaped propellant grain 4, so that it can be formed, and the forming process is simple. For other existing grain shapes, it is necessary to open holes or wings on the propellant grain, and forming molds are required, and its forming process is complex and the quality is difficult to guarantee.

[0041] The above examples are only illustrative of the present invention and do not constitute a limitation on the protection scope of the present invention. Any design identical or similar to the present invention falls within the protection scope of the present invention.

Claims

1. A solid propellant grain structure, characterized in that: It includes a side flame-limiting layer, an end face flame-limiting layer and a propellant grain. The propellant grain is of a cylindrical structure. The side flame-limiting layer and the end face flame-limiting layer are respectively fixed to the side and the end face of the propellant grain. A plurality of circular holes are formed in the end face flame-limiting layer. The radius of the circular hole on the end face restricted burning layer is r, and the length of the propellant grain is 2e, satisfying the relationship: ; where e is the web thickness of the propellant grain; the shortest distance between the centers of any two circular holes on the end face restricted burning layer is not less than 2(r + e); the shortest distance from the center of the outermost circular hole on the end face restricted burning layer to the inner wall of the side face restricted burning layer is not less than (r + e).

2. The solid propellant grain structure according to claim 1, wherein: The radius r of the circular holes in the end face flame-limiting layer is 3 mm, and the web thickness e of the propellant grain is 5 mm.

3. The solid propellant grain structure according to claim 1, characterized in that: The circular holes on the two end faces of the end face flame-limiting layer are symmetrically arranged.

4. The solid propellant grain structure according to claim 1, characterized in that: The side flame-limiting layer and the end face flame-limiting layer are made of a heat-insulating and non-combustible non-metallic material.

5. The solid propellant grain structure according to claim 1, characterized in that: The propellant grain comprises an energetic non-metallic material composed of a fuel binder and an oxidizer.

6. The preparation method of the solid propellant grain structure according to any one of claims 1 to 5, characterized in that, It includes the following steps: 1) Fabricate a hollow cylindrical side flame-limiting layer, and fabricate an end face flame-limiting layer, and form circular holes in the end face flame-limiting layer; 2) Pour the propellant raw material of the propellant grain into the side flame-limiting layer. After the propellant raw material is solidified into a grain, cut it into a disc shape according to the designed length; 3) Bond the end face flame-limiting layer to the disc-shaped propellant grain obtained in step 2).

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