A low-overload thrust termination device for solid rocket motors

Abstract

The present invention discloses a low-overload thrust termination device for a solid rocket engine, comprising a shell component, a screw-pressing component, an isolation component, a piston component, a cover component and a gas-producing component; one axial end of the shell component is connected to the engine nozzle front chamber, and the other end is provided with a cover component; an isolation structure is provided in the middle of the shell component to separate the shell into an isolation channel and a drive cavity; a through-channel is provided on the side of the isolation channel; the piston component is arranged in the shell component, one end of which is provided with an isolation component, and the other end is locked by a locking pin perpendicular to it in the cover component; a gas-producing component is provided on the side of the shell component, and its inflation channel is connected to the drive cavity; after receiving a shutdown command, the gas-producing component generates gas to enter the drive cavity, and the locking pin is unlocked after the internal pressure reaches a certain value, and the piston component moves to the end, so that the isolation channel and the counter-channel are connected, and the high-temperature and high-pressure gas in the engine is discharged through the counter-channel, so as to realize the engine with no thrust output.

Description

Technical Field

[0001] The invention belongs to the technical field of solid rocket engines, and in particular relates to a low-overload thrust termination device for a solid rocket engine. Background Art

[0002] As a power plant, the controllable shutdown of solid rocket motors is crucial for improving missile maneuverability, stage separation, penetration capability, and flight accuracy. Solid rocket motor shutdown methods include normal shutdown and forced shutdown. Normal shutdown allows the propellant to burn normally until completion, while forced shutdown uses special means to eliminate the engine's thrust output. Normal shutdown can lead to untimely shutdown, causing problems such as missile overshoot, which can affect the missile's range, flight accuracy, and impact precision. Forced shutdown avoids this problem.

[0003] Currently, the forced shutdown methods include reverse thrust, forced cooling and pressure reduction flameout. However, for small solid rocket engines, the existing forced shutdown methods have the following problems: the reverse thrust and forced cooling methods cannot meet the requirements in terms of structure and response time; the pressure reduction flameout method is suitable for small solid rocket engines, but there are problems such as excessive impact overload, high-speed release, structural damage, excessive thrust interference, and thrust output after re-ignition after shutdown.

[0004] The applicant's previous patent, CN 111622863 A, discloses a small solid rocket engine thrust termination device. This device utilizes two interlocking pressure relief channels that open sequentially upon command to terminate the solid rocket engine's thrust. While this prior art can achieve engine thrust termination, it uses pressure to break screws, releasing the piston assembly and allowing it to rapidly move outward, connecting the solid rocket engine's combustion chamber to the outside world. The overall device still experiences significant impact overload during operation, and suffers from drawbacks such as high-speed release and residual thrust when the propellant is not extinguished. Summary of the Invention

[0005] In response to one or more of the above-mentioned defects or improvement needs in the prior art, the present invention provides a low-overload thrust termination device for a solid rocket engine. After receiving a shutdown command, the gas-producing component generates gas that enters the drive cavity. After the internal pressure reaches a certain value, the locking pin is unlocked, the piston component moves to the lower end, the isolation channel and the exhaust channel are connected, and the high-temperature and high-pressure gas in the engine is discharged through the exhaust channel, thereby achieving zero thrust output of the engine.

[0006] To achieve the above-mentioned object, the present invention provides a low-overload thrust termination device for a solid rocket engine, characterized in that it includes a shell component, a screw-pressing component, an isolation component, a piston component, a sealing component, and a gas-generating component;

[0007] One axial end of the shell component is connected to the engine nozzle front chamber, and the other axial end thereof is provided with the cover component; an isolation structure is provided in the middle of the shell component to separate the shell into an isolation channel and a drive cavity; an exhaust pipe is provided on the side of the isolation channel to form a counter-exhaust channel;

[0008] A piston component is axially provided inside the shell, and the piston component passes through the isolation structure, and one end of the piston component extends into the isolation channel and is connected to the isolation component at the end of the shell, and the end of the shell is also provided with a pressing screw component, and the pressing screw component is threadedly connected to the inner side of the shell and pressed on the isolation component; the piston component passes through the other end through the drive cavity and extends into the cover component, and is locked by a locking pin in the cover component perpendicular to it;

[0009] The piston structure of the piston component is arranged in the driving cavity;

[0010] The gas-generating component is provided on the side of the shell component, which is connected to the driving cavity through the inflation channel provided on the shell. The generated gas pressure pushes the piston component to move toward one end of the cover component, so that the isolation channel and the opposite channel are connected.

[0011] As a further improvement of the present invention, the shear force F of the locking pin 剪切力 Should meet the following requirements:

[0012] F 剪切力 <P 发动机 ×S 隔离 +P 产气部件 ×S 活塞

[0013] Among them, P 发动机 is the engine gas pressure, S 隔离 is the area of ​​the isolation component, P 产气部件 is the gas production pressure of the gas production component, S 活塞 is the area of ​​the piston member; and / or,

[0014] The shear force of the locking pin is 1.2 kN to 1.6 kN.

[0015] As a further improvement of the present invention, the amount m of the gas generating agent in the gas generating component should satisfy:

[0016]

[0017] Where R is the gas constant, T is the gas temperature, M is the average molecular weight of the gas, V R2+R3 is the sum of the volume of the charging channel and the driving cavity, F 剪切力 is the shear force of the locking pin, S 隔离 is the area of ​​the isolation component, P发动机 is the engine gas pressure, S 活塞 is the area of ​​the piston member; and / or,

[0018] The amount of the gas generating agent in the gas generating component is 2.5g to 3g.

[0019] As a further improvement of the present invention, the piston component includes a piston rod and the piston structure provided on the piston rod; the piston structure separates the driving cavity into a first cavity and a second cavity, a first cavity is formed between the piston structure and the isolation structure, and a second cavity is formed between the piston structure and the covering component; the inflation channel is connected to the first cavity.

[0020] As a further improvement of the present invention, a sealing ring is provided between the piston rod and the isolation structure; and a sealing ring is provided between the piston structure and the housing.

[0021] As a further improvement of the present invention, the isolation component includes an isolation body and an isolation layer. The isolation body is fixedly connected to the end of the piston component, and the isolation layer is provided on the top surface of the isolation body.

[0022] As a further improvement of the present invention, a sealing ring is provided between the isolating body and the shell; and / or the compression amount of the isolating layer is 0.15 mm to 0.35 mm.

[0023] As a further improvement of the present invention, the compression screw component includes a compression screw insulation layer, a compression screw and a support ear. The compression screw is threadedly connected to the inner side of the shell port. Several symmetrical support ears are provided at the bottom of the inner hole of the compression screw, and a compression screw insulation layer is provided on the inner side of the compression screw.

[0024] As a further improvement of the present invention, the sealing cover component includes a sealing cover and a buffer layer. The sealing cover is threadedly connected to the outer side of the port of the shell component. A buffer material is provided on the inner side of the sealing cover to form a buffer layer. A piston rod through hole is provided in the middle of the sealing cover, and an exhaust groove is provided on the side of the piston rod through hole.

[0025] As a further improvement of the present invention, the gas producing component includes an igniter, one end of which is connected to the ignition cable and the other end is provided on the gas producing housing, and an igniter sealing ring is provided between the igniter and the gas producing housing;

[0026] The igniter is connected to the housing component through the gas generating housing, and a sealing ring is provided between the gas generating housing and the housing component;

[0027] The igniter is correspondingly provided with a gas generating agent, and the gas generating agent is filled between the gas generating shell and the gas generating cover.

[0028] In general, the above technical solutions conceived by the present invention have the following beneficial effects compared with the prior art:

[0029] (1) The low-overload thrust termination device of the solid rocket engine of the present invention locks the piston component and the isolation component through the locking pin before receiving the shutdown command, so as to ensure that the isolation component isolates the combustion gas in the engine; after receiving the shutdown command, the gas-generating component generates gas to enter the driving cavity, and the locking pin is unlocked after the internal pressure reaches a certain value, and the piston component moves to the end, the isolation channel and the exhaust channel are connected, and the high-temperature and high-pressure gas in the engine is discharged through the exhaust channel, so that the engine has no thrust output.

[0030] (2) The low-overload thrust termination device for the solid rocket engine of the present invention, through the provision of a sealing ring between the gas-producing shell and the shell component, can seal the gas generated by the gas-producing component in the inflation channel and the drive cavity to form a closed chamber, so that the piston component is always pressed against the end of the shell to achieve a locking effect.

[0031] (3) The low-overload thrust termination device for the solid rocket engine of the present invention can form a double sealing structure between the compression screw component and the isolation component. The first sealing structure is formed by the compression screw component squeezing the isolation layer and deforming it, and the second sealing structure is formed by the second sealing ring and the inner wall of the shell component, which can play a better role in isolating and sealing the engine nozzle gas.

[0032] (4) The low-overload thrust termination device for a solid rocket engine of the present invention has the characteristics of compact structure, rapid response, no residual thrust, and no release material. It can effectively reduce the impact interference when the engine thrust is terminated, and can achieve a zero-thrust output effect regardless of whether the engine charge is extinguished. BRIEF DESCRIPTION OF THE DRAWINGS

[0033] Figure 1 This is a schematic diagram of the isolation state of the low-overload thrust termination device according to an embodiment of the present invention;

[0034] Figure 2 Schematic diagram of the low-overload thrust termination device in the conductive state according to an embodiment of the present invention;

[0035] Figure 3 Schematic diagram of the planar structure of the screw-type compression component involved in the low-overload thrust termination device according to an embodiment of the present invention;

[0036] Figure 4 Schematic diagram of the cross-sectional three-dimensional structure of the screw-type compression component involved in the low-overload thrust termination device according to an embodiment of the present invention;

[0037] Figure 5 Schematic diagram of the cover exhaust hole structure involved in the low-overload thrust termination device according to an embodiment of the present invention;

[0038] Figure 6 Schematic diagram of the structure of the gas-generating components involved in the low-overload thrust termination device according to an embodiment of the present invention;

[0039] Figure 7 This is a schematic diagram of the connection between the low-overload thrust termination device and the solid rocket motor according to an embodiment of the present invention;

[0040] Figure 8 A comparison chart of thrust results between a low-overload thrust termination device according to an embodiment of the present invention and an earlier thrust termination device;

[0041] Figure 9 For the embodiment of the present invention Figure 8 Partial detail drawing.

[0042] In all the drawings, the same reference numerals represent the same technical features, specifically: 1-shell component, 2-pressing screw component, 3-isolating component, 4-piston component, 5-cover component, 6-locking pin, 7-gas generating component;

[0043] 11-housing, 12-exhaust pipe, 13-isolating structure, 14-first sealing ring; 21-pressed screw insulation layer, 22-pressed screw, 23-support ear; 31-isolating layer, 32-second sealing ring, 33-isolating body; 41-piston rod, 42-piston structure, 43-third sealing ring; 51-sealing cover, 52-buffer layer, 53-piston rod through hole, 54-exhaust groove; 71-ignition cable, 72-igniter, 73-igniter sealing ring, 74-gas generating housing, 75-sealing diaphragm, 76-gas generating cover, 77-gas generating charge, 78-fourth sealing ring, 79-fifth sealing ring;

[0044] 100-thrust termination device, 200-solid rocket motor;

[0045] R1-isolation channel, R2-inflation channel, R3-driving cavity, R4-opposite channel. DETAILED DESCRIPTION

[0046] In order to make the objectives, technical solutions and advantages of the present invention more clearly understood, the present invention is further described in detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only intended to illustrate the present invention and are not intended to limit the present invention. In addition, the technical features involved in the various embodiments of the present invention described below may be combined with each other as long as they do not conflict with each other.

[0047] In the description of the present invention, it should be understood that the terms "center", "longitudinal", "lateral", "length", "width", "thickness", "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inside", "outside", "clockwise", "counterclockwise", "axial", "radial", "circumferential" and the like to indicate orientations or positional relationships based on the orientations or positional relationships shown in the accompanying drawings, and are 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 understood as limiting the present invention.

[0048] Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of the technical features being referred to. Thus, a feature defined as "first" or "second" may explicitly or implicitly include at least one such feature. In the description of the present invention, "plurality" means at least two, such as two, three, etc., unless otherwise specifically defined.

[0049] In the present invention, unless otherwise specified or limited, the terms "installed," "connected," "connect," "fixed," etc. should be understood in a broad sense. For example, they can refer to fixed connection, detachable connection, or integration; mechanical connection, electrical connection; direct connection, or indirect connection through an intermediate medium; internal communication between two components, or interaction between two components, unless otherwise specified. Those skilled in the art will understand the specific meanings of the above terms in the present invention based on specific circumstances.

[0050] In the present invention, unless otherwise expressly specified or limited, when a first feature is "above" or "below" a second feature, it may mean that the first and second features are in direct contact, or that the first and second features are in indirect contact through an intermediary. Furthermore, when a first feature is "above," "above," or "above" a second feature, it may mean that the first feature is directly above or diagonally above the second feature, or simply means that the first feature is at a higher level than the second feature. When a first feature is "below," "below," or "below" a second feature, it may mean that the first feature is directly below or diagonally below the second feature, or simply means that the first feature is at a lower level than the second feature.

[0051] like Figures 1 to 7 As shown, the low-overload thrust termination device of the solid rocket engine of an embodiment of the present invention is connected to the engine nozzle front chamber of the solid rocket, and includes a shell component 1, a compression screw component 2, an isolation component 3, a piston component 4, a sealing component 5, a locking pin 6 and a gas producing component 7.

[0052] One axial end of the shell component 1 is connected to the engine nozzle front chamber, and the other end is provided with a cover component 5. The isolation structure 13 in the middle of the shell component 1 divides the shell into an isolation channel R1 and a drive cavity R3. An exhaust pipe 12 is provided on the side of the isolation channel R1 of the shell component 1 to form an exhaust channel R4. A through hole is provided in the middle of the isolation structure 13 to match the piston rod of the piston component 4. The piston component 4 passes through the isolation structure 13, and one end extends into the isolation channel R1 and is connected to the isolation component 3. The isolation component 3 is provided at the end of the shell, and the end of the shell is provided with a screw-pressing component 2. The screw-pressing component 2 is threadedly connected to the inner side of the shell and pressed on the isolation component 3; the other end of the piston component 4 even extends into the cover component 5 and is locked by a locking pin 6 perpendicular to it in the cover component 5. A gas-generating component 7 is provided on the side of the shell component 1, which is connected to the driving cavity R3 through the inflation channel R2 provided on the shell. The gas pressure generated pushes the piston component 4 to move toward one end of the cover component 5, so that the isolation channel R1 and the opposite channel R4 are connected.

[0053] Specifically, the housing component 1 comprises a housing 11, one axial end of which is flanged for connection to the engine nozzle, and the other end is provided with a capping member 5. An axial through-hole is provided within the housing 11, and a partitioning structure 13 is located in its center, separating the housing component 1 into an isolation channel R1 and a drive chamber R3. Exhaust pipes 12 extend through the sides of isolation channel R1, forming a counter-exhaust channel R4. Exhaust pipes 12 are connected to the housing 11 at one end and have exhaust ports symmetrically arranged at right angles on either side of the exhaust pipe, ensuring that the output thrust can offset each other.

[0054] In a preferred embodiment, the effective diameters of the isolation channel R1 and the counter-drain channel R4 are not less than Φ28 mm.

[0055] A piston assembly 4 is axially disposed within the housing 11 of the housing component 1. This assembly includes a piston rod 41, preferably made of 30CrMnSi structural alloy steel. A piston structure 42 is mounted on the piston rod 41. This piston structure 42 is located within the drive cavity R3, aligning with the inner wall of the housing 11 and capable of axial movement. The piston structure 42 divides the drive cavity R3 into a first cavity and a second cavity. The first cavity is formed between the piston structure 42 and the isolation structure 13, while the second cavity is formed between the piston structure 42 and the cover component 5.

[0056] The piston rod 41 passes through a central through-hole in the isolation structure 13. One end of the piston rod 41 extends into the isolation passage R1 and is connected to the isolation component 3, located at the end of the shell. This component is used to isolate the high-temperature, high-pressure combustion gases generated by the rocket engine. The isolation component 3 comprises an isolation layer 31 and an isolation body 33. The isolation body 33 is fixedly connected to the end of the piston rod 41 and has the isolation layer 31 disposed on its top surface, providing rigid support for the isolation layer 31.

[0057] Preferably, a second sealing ring 32 is provided between the isolation body 33 and the housing to improve the sealing performance between the isolation component 3 and the housing.

[0058] Preferably, the isolation layer 31 is made of 824 flexible ablation-resistant material, the isolation body 33 is made of alloy structural steel 30CrMnSi material, and the second sealing ring 32 is made of perfluororubber material.

[0059] Preferably, the compression amount of the isolation layer 31 is 0.15 mm to 0.35 mm.

[0060] Furthermore, the corresponding isolation component 3 is also provided with a screw pressing component 2, which is threadedly connected to the inner side of the shell 1 port so that the screw pressing component 2 is pressed on the inner step surface of the shell 11 and the top of the isolation layer 31 at the same time. The screw pressing component 2 includes a screw pressing insulation layer 21, a screw pressing 22 and a lug 23, combined with the screw pressing component 2. Figure 3 and Figure 4 As shown, the pressing screw 22 is threadedly connected to the inner side of the port of the shell 11, and a number of symmetrically arranged support ears 23 are provided at the bottom of the inner hole of the pressing screw 22, and the inner side of the pressing screw 22 is molded with insulation material to form a pressing screw insulation layer 21. The molded insulation material is preferably a high-silicon composite material.

[0061] In the specific embodiment shown in the drawings of the present invention, four lugs 23 are symmetrically arranged along the inner side of the pressure screw 22, which can ensure the stability of the pressure connection with the isolation component while ensuring a large ventilation area, which is conducive to the discharge of gas. Preferably, the ventilation area of ​​the pressure screw component 2 is not less than 650mm 2 .

[0062] The thrust termination device of the present invention can form a double sealing structure between the compression screw component 2 and the isolation component 3. The first sealing structure is formed by the compression screw component squeezing the isolation layer 31 and deforming it. The second sealing ring 32 forms a second seal with the inner wall of the shell component 1, which can play a better isolating and sealing role for the engine nozzle gas.

[0063] Furthermore, the other end of the piston rod 41 extends out of the driving cavity R3 and is connected to the end cover assembly 5 at the end of the shell, and is locked by a transverse locking pin 6 provided in the cover component 5.

[0064] Specifically, see again Figure 1 The cover component 5 includes a cover 51, which is threadedly connected to the outer side of the shell 11 port. A buffer material is bonded to the inside of the cover 51 to form a buffer layer 52, which can buffer and decelerate the piston component 4 when it moves to the end of the shell. Figure 5As shown, a piston rod through hole 53 matching the piston rod 41 is provided in the middle of the cover 51, and an exhaust groove 54 is provided on the side of the piston rod through hole 53, which is used to discharge the air in the second cavity of the driving cavity R3 when the piston structure moves toward the cover component 5. Through the design of the exhaust groove, the initial air in the second cavity can be quickly discharged, thereby improving the response speed.

[0065] A locking pin 6 is provided in the cover member 5 and is perpendicular to the piston rod 41. The locking pin 6 is threadedly provided in the cover member 5 and passes through the piston rod 41 in the piston rod through hole to lock the piston member. The locking pin 6 is preferably made of aluminum alloy.

[0066] Preferably, the shear force F of the locking pin 6 剪切力 (kN) should meet the following conditions:

[0067] F 剪切力 <P 发动机 ×S 隔离 +P 产气部件 ×S 活塞

[0068] Among them, P 发动机 is the engine gas pressure, S 隔离 is the area of ​​the isolation component, P 产气部件 is the gas production pressure of the gas production component, S 活塞 is the area of ​​the piston component. When the above formula is satisfied, the locking pin 6 is ensured to break under the pressure of the engine gas and the gas-producing components, releasing the constraint on the piston component 4 and allowing the piston component 4 to move toward the cover component 5 within the drive chamber R3, thereby connecting the isolation channel R1 and the counter-discharge channel R4.

[0069] In a specific embodiment of the present invention, the shear force of the locking pin 6 is preferably 1.2 kN to 1.6 kN.

[0070] In addition, a first sealing ring 14 is preferably provided between the piston rod 41 of the piston rod 4 and the isolation structure 13, and a double sealing ring structure is further preferably adopted; a third sealing ring 43 is provided between the piston structure 42 of the piston rod 4 and the housing, and a double sealing ring structure is further preferably adopted.

[0071] Furthermore, a gas-producing component 7 is provided on the side of the shell 11 of the shell component 1, which is connected to the driving cavity R3 through the inflation channel R2, and the inflation channel R2 is connected to the first cavity of the driving cavity R3. After receiving the ignition signal, the gas-producing component 7 generates high-temperature and high-pressure gas, and fills it into the first cavity through the inflation channel R2, pushing the piston component 4 to move toward the covering component 5.

[0072] like Figure 6As shown, the gas-generating component 7 of the present invention includes an ignition cable 71, an igniter 72, an igniter sealing ring 73, a gas-generating housing 74, a sealing diaphragm 75, a gas-generating cover 76, a gas-generating agent 77, and a fourth sealing ring 78. One end of the igniter 72 is connected to the ignition cable 71, and the other end is mounted on the gas-generating housing 74. The igniter sealing ring 73 is positioned between the igniter 72 and the gas-generating housing 74. The igniter 72 is connected to the housing 11 via the gas-generating housing 74. A fifth sealing ring 79 is positioned between the gas-generating housing 74 and the housing 11 to ensure a tight seal between the gas-generating component 7 and the drive chamber.

[0073] A gas generating agent 77 is provided corresponding to the igniter 72. Gas generating agent 77 is loaded between the gas generating housing 74 and the gas generating cover 76. Gas generating agent 77 is preferably a sheet-like pyrotechnic agent. A sealing membrane 75 is bonded to the end surface of the gas generating agent 77 to prevent moisture during storage. Preferably, a fourth sealing ring 78 is provided between the gas generating cover 76 and the housing 11 to ensure a tight seal.

[0074] After the gas generating component 7 receives the ignition signal, the igniter 72 ignites the gas generating agent 77 to generate high-temperature and high-pressure gas, which enters the charging channel R2 and is filled into the first cavity, pushing the piston component 4 to move toward the covering component 5.

[0075] In the specific embodiment shown in the drawings of the present invention, two gas generating components 7 are provided, and the angle between the two is preferably less than 90°, more preferably 60°, so as to avoid the impact effect of the igniter.

[0076] Preferably, the amount m (g) of the gas generating agent in the gas generating component 7 should meet the following conditions:

[0077]

[0078] Where R is the gas constant, T is the gas temperature, M is the average molecular weight of the gas, V R2+R3 is the sum of the cavity volumes of the inflation channel R2 and the driving cavity R3, F 剪切力 is the shear force of the locking pin, S 隔离 is the area of ​​the isolation component, P 发动机 is the engine gas pressure, S 活塞 When the amount of gas generating agent m meets the above conditions, it can ensure that the gas generating pressure pushes the piston member to the piston end.

[0079] In a specific embodiment of the present invention, the amount of the gas generating agent in the gas generating component 7 is 2.5 g to 3 g.

[0080] In addition, the first sealing ring, second sealing ring, third sealing ring, fourth sealing ring, fifth sealing ring, igniter sealing ring, etc. of the present invention are preferably O-rings of the prior art, which can be selected according to specific sizes.

[0081] Furthermore, Figure 7 This is a schematic diagram of the connection between a low-overload thrust termination device and a solid rocket engine according to an embodiment of the present invention. The thrust termination device 100 of the present invention is connected to the nozzle antechamber of a solid rocket engine 200. The working process and principle of the thrust termination device of the present invention are as follows:

[0082] When the solid rocket engine is working, it generates high-temperature and high-pressure gas that is discharged through the nozzle. Before receiving the shutdown command, the thrust termination device locks the piston component and isolation component through the locking pin to ensure that the isolation component isolates the gas in the engine.

[0083] When the solid rocket engine needs to terminate thrust, a termination command is sent to the thrust termination device. After the gas-producing component receives the shutdown signal, the ignition cable is connected, and the igniter ignites the gas-producing powder. The high-temperature and high-pressure gas generated enters the drive cavity from the inflation channel. After the internal pressure reaches a certain value, the locking pin breaks under the thrust of the solid rocket engine gas and the gas-producing component, and the piston component is pushed to the end of the shell under pressure so that the piston structure contacts the buffer layer of the cover component. Because the inflation channel R2 and the drive cavity R3 form a closed chamber under the action of the sealing ring, the piston component is always pressed against the end of the shell, achieving a locking effect. The isolation channel R1 and the exhaust channel R4 are connected, and the high-temperature and high-pressure gas in the engine nozzle is discharged through the exhaust channel R4, realizing that the engine has no thrust output.

[0084] like Figure 8 and Figure 9 As shown in the figure, there is a comparison of the thrust results of the low-overload thrust termination device of the present invention and the previous thrust termination device (CN111622863A). It can be seen that the thrust of the previous thrust termination device is relatively large, up to 3800N, while the present invention can achieve zero thrust output of the engine.

[0085] The low-overload thrust termination device for a solid rocket engine of the present invention locks the piston assembly and isolation assembly via a locking pin before receiving a shutdown command, ensuring that the isolation assembly isolates the combustion gas within the engine. Upon receiving the shutdown command, the gas-generating assembly generates gas that enters the drive chamber. When the internal pressure reaches a certain value, the locking pin unlocks, the piston assembly moves to the end, connecting the isolation passage and the exhaust passage, allowing the high-temperature and high-pressure gas within the engine to be discharged through the exhaust passage, achieving zero thrust output. Regardless of whether the propellant grain is extinguished or slow-burning, the engine will not produce any residual thrust output or work.

[0086] The thrust termination device of the present invention has the characteristics of compact structure, rapid response, no residual thrust, and no released objects. It can effectively reduce the impact interference during engine thrust termination and provide a new implementation solution for thrust termination of small solid rocket engines.

[0087] It will be easily understood by those skilled in the art that the above description is merely a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, and improvements made within the spirit and principles of the present invention should be included in the scope of protection of the present invention.

Claims

1. A low-overload thrust termination device for a solid rocket engine, characterized in that: Including shell parts, screw-pressing parts, isolation parts, piston parts, cover parts and gas-generating parts; One axial end of the shell component is connected to the engine nozzle front chamber, and the other axial end thereof is provided with the cover component; an isolation structure is provided in the middle of the shell component to separate the shell into an isolation channel and a drive cavity; and a through-row channel is provided on the side of the isolation channel; A piston component is axially provided inside the shell, and the piston component passes through the isolation structure, and one end of the piston component extends into the isolation channel and is connected to the isolation component at the end of the shell, and the end of the shell is also provided with a pressing screw component, and the pressing screw component is threadedly connected to the inner side of the shell and pressed on the isolation component; the piston component passes through the other end through the drive cavity and extends into the cover component, and is locked by a locking pin in the cover component perpendicular to it; The piston structure of the piston component is arranged in the driving cavity; The gas generating component is provided on the side of the shell component, which is connected to the driving cavity through the inflation channel provided on the shell, and the generated gas pressure pushes the piston component to move toward one end of the cover component, so that the isolation channel and the opposite channel are connected; The shear force F of the locking pin 剪切力 Should meet the following requirements: Among them, P 发动机 is the engine gas pressure, S 隔离 is the area of ​​the isolation component, P 产气部件 is the gas production pressure of the gas production component, S 活塞 is the area of ​​the piston member; and / or, The shear force of the locking pin is 1.2 kN to 1.6 kN.

2. The low-overload thrust termination device for a solid rocket motor according to claim 1, characterized in that: The amount m of the gas generating agent in the gas generating component should meet the following requirements: Where R is the gas constant, T is the gas temperature, M is the average molecular weight of the gas, V R2+R3 is the sum of the volume of the charging channel and the driving cavity, F 剪切力 is the shear force of the locking pin, S 隔离 is the area of ​​the isolation component, P 发动机 is the engine gas pressure, S 活塞 is the area of ​​the piston member; and / or, The amount of the gas generating agent in the gas generating component is 2.5g to 3g.

3. The low-overload thrust termination device for a solid rocket motor according to claim 1, characterized in that: The piston component includes a piston rod and the piston structure provided on the piston rod; the piston structure divides the driving cavity into a first cavity and a second cavity, a first cavity is formed between the piston structure and the isolation structure, and a second cavity is formed between the piston structure and the covering component; the inflation channel is connected to the first cavity.

4. The low-overload thrust termination device for a solid rocket motor according to claim 3, characterized in that: A sealing ring is provided between the piston rod and the isolation structure; and a sealing ring is provided between the piston structure and the housing.

5. The low-overload thrust termination device for a solid rocket engine according to any one of claims 1 to 4, characterized in that: The isolation component includes an isolation body and an isolation layer, and the isolation body is fixedly connected to the end of the piston component. An isolation layer is provided on the top surface thereof.

6. The low-overload thrust termination device for a solid rocket motor according to claim 5, characterized in that: A sealing ring is provided between the isolating body and the shell; and / or the compression amount of the isolating layer is 0.15 mm to 0.35 mm.

7. The low-overload thrust termination device for a solid rocket motor according to any one of claims 1 to 4, characterized in that: The compression screw component includes a compression screw insulation layer, a compression screw and a lug. The compression screw is threadedly connected to the inner side of the shell port. Several symmetrical lugs are provided at the bottom of the inner hole of the compression screw, and a compression screw insulation layer is provided on the inner side of the compression screw.

8. The low-overload thrust termination device for a solid rocket motor according to any one of claims 1 to 4, characterized in that: The sealing cover component includes a sealing cover and a buffer layer. The sealing cover is threadedly connected to the outer side of the shell component port. A buffer material is provided on the inner side of the sealing cover to form a buffer layer. A piston rod through hole is provided in the middle of the sealing cover, and an exhaust groove is provided on the side of the piston rod through hole.

9. The low-overload thrust termination device for a solid rocket motor according to any one of claims 1 to 4, characterized in that: The gas producing component includes an igniter, one end of which is connected to the ignition cable and the other end is arranged on the gas producing housing, and an igniter sealing ring is provided between the igniter and the gas producing housing; The igniter is connected to the housing component through the gas generating housing, and a sealing ring is provided between the gas generating housing and the housing component; The igniter is correspondingly provided with a gas generating agent, and the gas generating agent is filled between the gas generating shell and the gas generating cover.

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