A single chamber dual thrust solid rocket engine with high thrust ratio

Abstract

This invention provides a high thrust-to-weight ratio single-chamber dual-thrust solid rocket motor, comprising a propellant combustion chamber and a nozzle. The propellant combustion chamber includes a shell, an insulation layer, and a propellant grain. The insulation layer is disposed on the inner surface of the shell, and the propellant grain fills the insulation layer. The thickness of the insulation layer gradually increases from the end of the propellant combustion chamber furthest from the nozzle to the end closest to the nozzle. The propellant grain includes a first-stage propellant grain section and a second-stage propellant grain section, with the second-stage propellant grain section located on the side of the first-stage propellant grain section furthest from the nozzle. This invention, without changing the engine's maximum operating pressure, employs a gradually thickening insulation layer and propellant grain structure design to maximize the propellant load within a limited space, thereby improving the engine's mass ratio and total impulse. Furthermore, the nozzle, combined with a throat liner material with a high ablation rate, further enhances the engine's two-stage thrust ratio, achieving a two-stage thrust ratio of over 10 for the single-chamber dual-thrust solid rocket motor.

Description

Technical Field

[0001] This invention belongs to the field of solid rocket engine technology, specifically relating to a single-chamber dual-thrust solid rocket engine with a high thrust-to-weight ratio. Background Technology

[0002] Missile weapon systems employ different flight trajectories depending on their operational missions, generally consisting of launch-to-orbit (or launch tube exit), acceleration flight, constant-speed endurance, and high-speed attack phases. To meet the propulsion requirements of each flight phase, multi-stage engines are often required, namely multi-chamber multi-thrust, single-chamber multi-thrust, or multi-pulse solid rocket engines. Among these, single-chamber dual-thrust engines, with only one combustion chamber, one nozzle, and one ignition system, offer advantages such as simple and compact structure, light weight, and high reliability, making them the most widely used.

[0003] Single-chamber dual-thrust solid rocket motors, having only one nozzle, achieve dual thrust in several ways: a) using two propellant grains with different burning rates; b) adjusting the nozzle throat area; and c) using the same propellant but with two different shaped propellant grains. Method a) requires two propellant formulations and separate molding of the two propellant grains, making its manufacturing process complex and costly. Method b), while allowing both stages to operate under optimal pressure and without a clearly defined thrust ratio, enabling random adjustment, suffers from a complex nozzle adjustment structure, significant negative mass, and high material requirements. Method c), with its simple structure, lower cost, and straightforward propellant grain molding process, is therefore the most common method, employing the same propellant with two different shaped propellant grains to achieve dual thrust.

[0004] For single-chamber dual-thrust solid rocket motors, since the propellant grains of both stages are located in the same combustion chamber and use the same nozzle, the operating characteristics of the two stages inevitably affect each other. This means the two-stage thrust ratio is limited and cannot be too high; otherwise, the operating pressure difference between the first and second stage combustion chambers will be too large (too high for the first stage and too low for the second), resulting in insufficient utilization of the shell structure strength and increased negative mass. Simultaneously, too low a second-stage operating pressure leads to reduced propellant combustion efficiency, even falling below the propellant's critical operating pressure, thus preventing proper propellant combustion. Therefore, the two-stage thrust ratio of most single-chamber dual-thrust engines should not exceed 10. However, with the overall development of missiles, there is a demand for further increases in the two-stage thrust ratio of single-chamber dual-thrust engines, which has brought significant challenges to the design of solid rocket motors.

[0005] In addition, to ensure the total thrust energy of the engine, most engines currently use throat liner materials with low ablation rate (i.e., low ablation rate) for their nozzles, such as refractory metal tungsten infiltrated with copper, molybdenum, and ceramic materials. If a throat liner with a high ablation rate is used, as the engine operating time increases, the nozzle throat diameter increases, the engine operating pressure decreases, and the propellant energy loss is greater, which may fail to meet the overall tactical and technical specifications. Summary of the Invention

[0006] The purpose of this invention is to provide a single-chamber dual-thrust solid rocket motor with a high thrust-to-weight ratio, which can at least solve some of the defects existing in the prior art.

[0007] To achieve the above objectives, the present invention adopts the following technical solution:

[0008] A single-chamber dual-thrust solid rocket motor with a high thrust-to-weight ratio includes a propellant combustion chamber and a nozzle installed at the outlet end of the propellant combustion chamber. The propellant combustion chamber includes a shell, an insulation layer, and a propellant grain. The insulation layer is disposed on the inner surface of the shell, and the propellant grain fills the insulation layer. The thickness of the insulation layer gradually increases from the end of the propellant combustion chamber away from the nozzle to the end closer to the nozzle. The propellant grain includes a primary propellant grain segment and a secondary propellant grain segment, and the secondary propellant grain segment is located on the side of the primary propellant grain segment away from the nozzle.

[0009] Furthermore, the primary propellant column segment adopts a star-shaped or inner hole side propellant column, and the secondary propellant column segment adopts an end face propellant column.

[0010] Furthermore, the primary and secondary propellant segments are made using the same solid propellant.

[0011] Furthermore, the solid propellant used in the primary and secondary propellant grain segments has a burning rate of 4–30 mm / s, and the pressure index of the solid propellant is 0.2–0.25 in the low pressure range of 0.8–6 MPa.

[0012] Furthermore, the thickness of the insulation layer is formed by the angle α of the combustion chamber gradually expanding from the end furthest from the nozzle to the end closest to the nozzle, which is 1 to 3°.

[0013] Furthermore, the insulation layer is molded and then adhered to the inner wall of the shell.

[0014] Furthermore, the medicinal column is cast into the insulation layer inside the shell.

[0015] Furthermore, the throat liner of the nozzle is made of a material with an ablation rate ≥0.1mm / s.

[0016] Furthermore, the outlet end of the propellant combustion chamber is also provided with a rear end cap, an igniter, and a long tailpipe. One end of the rear end cap is connected to the end of the shell of the propellant combustion chamber, and the other end of the rear end cap is connected to the long tailpipe. The igniter is installed on the rear end cap, and the nozzle is installed on the long tailpipe.

[0017] Compared with the prior art, the beneficial effects of the present invention are as follows:

[0018] The single-chamber dual-thrust solid rocket motor with high thrust ratio provided by this invention, without changing the maximum operating pressure of the engine, i.e. without increasing the negative mass of the engine, adopts a gradually thickened insulation layer and propellant grain structure design to maximize the amount of propellant in a limited space, thereby improving the engine's mass ratio and total impulse. Furthermore, the use of a throat liner material with a high ablation rate in the nozzle can further improve the two-stage thrust ratio of the engine, making the two-stage thrust ratio of the single-chamber dual-thrust solid rocket motor reach more than 10.

[0019] The present invention will now be described in further detail with reference to the accompanying drawings. Attached Figure Description

[0020] Figure 1 This is a schematic diagram of the structure of the single-chamber dual-thrust solid rocket motor with high thrust ratio of the present invention;

[0021] Figure 2 This is a comparison chart of thrust-time curves for nozzles with throat liners of different ablation rates in embodiments of the present invention.

[0022] Explanation of reference numerals in the attached drawings: 1. Shell; 2. Insulation layer; 3. Propellant grain; 4. Rear end cap; 5. Igniter; 6. Long tailpipe; 7. Nozzle; 31. Primary propellant grain section; 32. Secondary propellant grain section. Detailed Implementation

[0023] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0024] In the description of this invention, it should be understood that the terms "center", "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this invention.

[0025] In the description of this invention, it should be noted that, unless otherwise explicitly specified and limited, the terms "installation", "connection", and "joining" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, an abutting connection, or an integral connection. Those skilled in the art can understand the specific meaning of the above terms in this invention based on the specific circumstances.

[0026] like Figure 1 As shown, this embodiment provides a single-chamber dual-thrust solid rocket motor with a high thrust-to-weight ratio, including a propellant combustion chamber and a nozzle 7 installed at the outlet end of the propellant combustion chamber. The propellant combustion chamber includes a shell 1, an insulation layer 2, and a propellant grain 3. The insulation layer 2 is disposed on the inner surface of the shell 1, and the propellant grain 3 is disposed on the inner surface of the insulation layer 2. The insulation layer 2 adopts a gradually increasing thickness design, that is, the thickness of the insulation layer 2 gradually increases from the end of the propellant combustion chamber away from the nozzle 7 to the end closer to the nozzle 7, forming a certain taper. Preferably, the thickness of the insulation layer 2 gradually increases from the end of the propellant combustion chamber away from the nozzle 7. The included angle α formed by the gradual expansion towards the nozzle 7 is 1 to 3°; the propellant grain 3 includes a primary propellant grain segment 31 and a secondary propellant grain segment 32. The secondary propellant grain segment 32 is located on the side of the primary propellant grain segment 31 away from the nozzle 7. The primary propellant grain segment 31 and the secondary propellant grain segment 32 are filled in the insulation layer 2 and coaxially connected in sequence to form an integral propellant grain 3. The primary propellant grain segment 31 is the booster stage of the rocket, and the secondary propellant grain segment 32 is the endurance stage of the rocket. The outer diameter of the propellant grain 3 is designed to gradually increase from the end of the combustion chamber away from the nozzle 7 to the end closer to the nozzle 7 to match the inner surface of the insulation layer 2 with a gradually changing thickness. In this embodiment, the combustion chamber is designed with a gradually thickened insulation layer 2, and the outer diameter of the propellant grain 3 is designed to match the inner surface of the gradually thickened insulation layer 2. That is, the outer diameters of the first-stage propellant grain segment 31 and the second-stage propellant grain segment 32 are also gradually thickened. The insulation layer 2 corresponding to the first-stage propellant grain segment 31 is thicker, and the insulation layer 2 corresponding to the second-stage propellant grain segment 32 is thinner. The insulation layer 2 is thinnest at the front end cap near the inlet end of the combustion chamber. In this way, the amount of propellant in the combustion chamber is maximized without changing the volume of the combustion chamber, thereby improving the mass ratio of the solid rocket engine and thus increasing the total impulse of the solid rocket engine.

[0027] The insulation layer 2 is molded and then bonded to the inner wall of the shell 1, and then machined to form the required taper. Since the insulation layer 2 in this embodiment adopts a gradually changing thickness structure, its machining and shaping are convenient and the required taper of the insulation layer can be precisely controlled. The drug cartridge 3 is cast into the inner surface of the insulation layer 2 inside the shell 1. This process is simple to operate and easy to implement, and the forming quality of the insulation layer 2 and the drug cartridge 3 is controllable.

[0028] To further increase the two-stage thrust ratio, the first-stage propellant section 31 adopts a star-shaped or inner-hole side propellant form with a larger burning surface to generate a larger first-stage thrust, while the second-stage propellant section 32 adopts an end-face propellant form with a smaller burning surface to generate a smaller second-stage thrust. By controlling the size of the burning surface, different sizes of thrust can be achieved, thereby realizing a large two-stage thrust ratio.

[0029] To optimize the above technical solution, the first-stage propellant segment 31 and the second-stage propellant segment 32 are made of the same solid propellant. In order to meet the requirement of a large thrust ratio between the two stages, a composite propellant with a wide range of adjustable burning rate, a wide pressure range (0.8MPa~20MPa), a small pressure index at low pressure, good mechanical properties, and a low critical working pressure is selected. Specifically, in this embodiment, the burning rate of the solid propellant used in the first-stage propellant segment 31 and the second-stage propellant segment 32 is 4~30mm / s, and the pressure index of the solid propellant is 0.2~0.25 in the low pressure range of 0.8~6MPa.

[0030] Generally, when selecting the throat liner material for the nozzle 7, it is desirable that the throat liner of the nozzle 7 is resistant to ablation and that the throat diameter remains constant, thereby achieving the thrust-time curve of the design state. However, in this invention, in order to further improve the two-stage thrust ratio, i.e., the first-stage thrust is greater and the second-stage thrust is smaller, since the first-stage operating time is short, its ablation on the nozzle 7 throat liner is negligible. Experiments show that when the first stage operates alone, there is basically no ablation on the nozzle 7 throat liner, and the throat diameter can remain constant. However, if the second stage operates, as the engine operates, the nozzle 7 throat diameter increases, and the second-stage pressure decreases, thus obtaining a smaller second-stage thrust. Therefore, the nozzle 7 throat liner of this invention is made of a material with an ablation rate ≥0.1mm / s, such as graphite or C / C material. This allows the engine thrust ratio to be greater than 10, from less than 10 when the throat diameter remains constant, while ensuring that the total engine stroke does not decrease or even increases.

[0031] Optionally, a rear end cap 4, an igniter 5, and a long tailpipe 6 are also provided at the outlet end of the propellant combustion chamber. One end of the rear end cap 4 is connected to the end of the shell 1 of the propellant combustion chamber by a retaining ring, and the other end of the rear end cap 4 is connected to the long tailpipe 6. The igniter 5 is mounted on the rear end cap 4, and the nozzle 7 is mounted on the long tailpipe 6. The propellant combustion chamber, rear end cap 4, long tailpipe 6, and nozzle 7 are all coaxially arranged. This embodiment uses one nozzle 7, one type of propellant, and different propellant grain combustion surfaces to achieve two-stage thrust. The engine structure is simple and reliable, and the production process is easy to implement and economical.

[0032] The following specific examples illustrate the impact of propellant grain structure and nozzle throat liner material on the total impulse and two-stage thrust ratio of a solid rocket motor.

[0033] Example 1:

[0034] In this embodiment, a comparative experiment was conducted between an experimental group and a control group. The experimental group adopted the technical solution of this invention, where the insulation layer 2 of the combustion chamber had a gradually varying thickness, and correspondingly, the outer diameter of the two-stage propellant grains 3 within the combustion chamber also had a gradually varying shape. Simultaneously, its nozzle throat liner was made of graphite material with an ablation rate of 0.12 mm / s. The control group adopted a conventional technical solution, where the insulation layer 2 of the combustion chamber had a uniform thickness structure, and its nozzle throat liner was made of tungsten-copper infiltrated material with a non-ablation rate (i.e., an ablation rate of 0). Both the experimental and control groups used a star-shaped propellant grain 3. The main structural parameters of the propellant grain 3 are shown in Table 1. The thrust-time curves of the solid rocket motors measured in the two experiments are compared as follows: Figure 2 As shown in Table 2, the two-stage thrust ratio and total impulse of the solid rocket motor were obtained by processing the experimental data.

[0035] Table 1:

[0036]

[0037]

[0038] Table 2:

[0039] project control group experimental group Thrust ratio 8.9 11 Total impulse (N·s) 9.6 10.3

[0040] As shown in Table 2, with the propellant combustion chamber space remaining unchanged, the total thrust of the engine using this invention can be increased by about 7%, raising the two-stage thrust ratio from 8.9 to 11, an increase of 24%. If a further increase in the two-stage thrust ratio is desired, a throat liner material with a higher ablation rate can be selected.

[0041] Example 2:

[0042] This embodiment conducts tests on the insulation layer 2 of the propellant combustion chamber with different tapers α and on the throat liner of the nozzle 7 with different ablation rate materials. Specifically, when the insulation layer 2 uses the same taper α (α = 1.0°), tests are conducted on the throat liner of the nozzle 7 using different ablation rate materials, and the performance results of the solid rocket motor are shown in Table 3. When the throat liner of the nozzle 7 uses the same ablation rate material (i.e., an ablation rate of 0.1 mm / s), tests are conducted on the insulation layer 2 with different tapers α, and the performance results of the solid rocket motor are shown in Table 4.

[0043] Table 3:

[0044]

[0045] Table 4:

[0046]

[0047] As shown in Table 3, under the same taper α, the thrust ratio increases with the increase of the ablation rate of the nozzle throat liner material. When the ablation rate increases from 0 to 0.4 mm / s, the thrust ratio can increase from 8.90 to 16.52, but the total impulse of the solid rocket motor decreases.

[0048] As shown in Table 4, when the ablation rate of the nozzle throat liner material is the same, the thrust ratio decreases as the taper α of the insulation layer increases. When the taper α increases from 0° to 3.0°, the thrust ratio decreases from 11.91 to 7.92. At the same time, as the taper α increases, the amount of propellant in the combustion chamber and the total thrust of the solid rocket motor increase.

[0049] In addition, this embodiment also tested the insulation layer 2 of the propellant combustion chamber with different taper α, and the throat liner of the nozzle 7 with different ablation rate materials. The performance results of the solid rocket motor are shown in Table 5.

[0050] Table 5:

[0051] Insulating layer taper α 0° 0.5° 1.0° 1.5° 2.0° 2.5° 3.0° Ablation rate (mm / s) 0.00 0.08 0.15 0.22 0.30 0.38 0.45 Thrust ratio 8.90 9.87 10.74 11.54 12.46 13.29 13.81 Charge weight (kg) 4.40 4.52 4.66 4.82 5.00 5.20 5.45 Total impulse (N·s) 9.63 10.00 10.25 10.48 10.70 10.89 11.05

[0052] As shown in Table 5, with the increase of the taper α of the insulating layer 2 and the ablation rate of the nozzle throat liner material, the two-stage thrust ratio of the solid rocket motor increases. The taper α increases from 0° to 3.0°, and the two-stage thrust ratio increases from 8.90 to 13.81, an increase of 55%. At the same time, with the increase of the taper of the insulating layer 2, the amount of propellant in the combustion chamber and the total thrust of the solid rocket motor also increase.

[0053] The above examples are merely illustrative of the present invention and do not constitute a limitation on the scope of protection of the present invention. All designs that are the same as or similar to the present invention are within the scope of protection of the present invention.

Claims

1. A single-chamber, dual-thrust solid rocket motor with a high thrust-to-weight ratio, comprising a propellant combustion chamber and a nozzle installed at the outlet end of the propellant combustion chamber, characterized in that: The combustion chamber includes a shell, an insulation layer, and a propellant grain. The insulation layer is disposed on the inner surface of the shell, and the propellant grain is filled within the insulation layer. The thickness of the insulation layer gradually increases from the end of the combustion chamber furthest from the nozzle to the end closest to the nozzle. The propellant grain includes a primary propellant grain segment and a secondary propellant grain segment, with the secondary propellant grain segment located on the side of the primary propellant grain segment furthest from the nozzle. The included angle α formed by the gradual expansion of the insulation layer from the end of the combustion chamber furthest from the nozzle to the end closest to the nozzle is 1–3°. The throat liner of the nozzle is made of a material with an ablation rate ≥0.1 mm / s.

2. The high thrust-ratio single-chamber dual-thrust solid rocket motor as described in claim 1, characterized in that: The primary propellant column segment uses a star-shaped or inner hole side propellant column, and the secondary propellant column segment uses an end face propellant column.

3. The high thrust-ratio single-chamber dual-thrust solid rocket motor as described in claim 1 or 2, characterized in that: The primary and secondary propellant segments are made using the same solid propellant.

4. The high thrust-ratio single-chamber dual-thrust solid rocket motor as described in claim 3, characterized in that: The solid propellant used in the primary and secondary propellant grain sections has a burning rate of 4–30 mm / s, and the pressure index of the solid propellant is 0.2–0.25 in the low pressure range of 0.8–6 MPa.

5. The high thrust-ratio single-chamber dual-thrust solid rocket motor as described in claim 1, characterized in that: The insulation layer is molded and then adhered to the inner wall of the shell.

6. The high thrust-ratio single-chamber dual-thrust solid rocket motor as described in claim 1, characterized in that: The medicinal powder is cast into the insulation layer inside the shell.

7. The high thrust-ratio single-chamber dual-thrust solid rocket motor as described in claim 1, characterized in that: The outlet end of the combustion chamber is also provided with a rear end cap, an igniter, and a long tailpipe. One end of the rear end cap is connected to the end of the shell of the combustion chamber, and the other end of the rear end cap is connected to the long tailpipe. The igniter is installed on the rear end cap, and the nozzle is installed on the long tailpipe.

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