阅读说明：本技术 一种小型固体火箭发动机推力终止装置 (Thrust termination device of small solid rocket engine ) 是由 曾雪宁 郝雪杰 胡佳智 于 2020-06-05 设计创作，主要内容包括：本发明公开了一种小型固体火箭发动机推力终止装置,包括壳体、堵塞部件、点火药盒部件、拉断螺钉、活塞部件；壳体设有法兰端,并通过法兰端固接于发动机外壳上,壳体内部设置有相互垂直贯通的一级泄压通道和二级泄压通道,其中,一级泄压通道与发动机燃烧室连通,二级泄压通道与发动机喷管方向平行；点火药盒部件固定安装于壳体尾部,其排气端正对二级泄压通道；活塞部件通过拉断螺钉固定于二级泄压通道内；堵塞部件被活塞部件限位于一级泄压通道与二级泄压通道之间。本装置通过设置两级互锁的泄压通道,两级泄压通道根据指令逐级打开后实现固体火箭发动机推力终止,提高了推力终止过程的可靠性与安全性；而且结构简单紧凑、质量轻便、成本低。(The invention discloses a thrust termination device of a small solid rocket engine, which comprises a shell, a blocking part, an ignition medicine box part, a snapping screw and a piston part, wherein the blocking part is arranged on the shell; the shell is provided with a flange end and is fixedly connected to the engine shell through the flange end, and a primary pressure relief channel and a secondary pressure relief channel which are mutually vertically communicated are arranged in the shell, wherein the primary pressure relief channel is communicated with a combustion chamber of the engine, and the secondary pressure relief channel is parallel to the direction of an engine spray pipe; the ignition medicine box part is fixedly arranged at the tail part of the shell, and the exhaust end of the ignition medicine box part is aligned to the secondary pressure relief channel; the piston component is fixed in the secondary pressure relief channel through a snapping screw; the blocking component is limited between the first-stage pressure relief channel and the second-stage pressure relief channel by the piston component. The device is provided with the two-stage interlocking pressure relief channels, and the thrust termination of the solid rocket engine is realized after the two-stage pressure relief channels are opened step by step according to the instruction, so that the reliability and the safety of the thrust termination process are improved; and simple and compact structure, light weight and low cost.)

1. A thrust termination device of a small solid rocket engine is characterized in that: comprises a shell, a blocking component, an ignition medicine box component, a snapping screw and a piston component;

the shell is provided with a flange end and is fixedly connected to the engine shell through the flange end, and a primary pressure relief channel and a secondary pressure relief channel which are mutually perpendicular and communicated are arranged in the shell, wherein the primary pressure relief channel is communicated with a combustion chamber of the engine, and the secondary pressure relief channel is parallel to the direction of an engine spray pipe;

the ignition medicine box part is fixedly arranged at the tail part of the shell, and the exhaust end of the ignition medicine box part is aligned to the secondary pressure relief channel; the piston component is fixed in the secondary pressure relief channel through a snapping screw; the blocking component is limited between the first-stage pressure relief channel and the second-stage pressure relief channel by the piston component.

2. The small solid rocket engine thrust termination apparatus of claim 1 wherein: still include the insulator, press the spiral shell, protect a section of thick bamboo, the insulator is located in the one-level pressure release passageway, presses the spiral shell cover to locate the insulator afterbody, can carry on spacingly with piston part cooperation to the jam part, protects a section of thick bamboo fixed mounting in the casing export.

3. A small solid-rocket engine thrust termination apparatus according to claim 2 wherein: the plug component comprises a heat insulation layer, a metal plug and an O-shaped sealing ring I, the heat insulation layer is arranged between the metal plug and the pressing screw, the contact surface of the metal plug and the piston component is an ellipsoidal curved surface, and the O-shaped sealing ring I is arranged between the metal plug and the shell.

4. A small solid-rocket engine thrust termination apparatus according to claim 3 wherein: the metal plug is made of high-strength alloy structural steel, and a hardened layer is arranged on the surface of the metal plug.

5. A small solid-rocket engine thrust termination apparatus according to claim 2 wherein: the inner diameter of the protective cylinder is larger than that of the secondary pressure relief channel, and an O-shaped sealing ring II is arranged between the protective cylinder and the shell.

6. The small solid rocket engine thrust termination apparatus of claim 1 wherein: ignition medicine box part includes medicine box body, flame igniter, ignition tablet, medicine box seat, flame igniter installs at medicine box body afterbody, is equipped with O type sealing washer III between medicine box body and the casing, and the ignition tablet is flaky firework igniter, freely loads in the space that medicine box body and medicine box seat formed.

7. The small solid rocket engine thrust termination apparatus of claim 6 wherein: fire transfer holes are formed in the medicine box body and the medicine box seat, and a sealing film I and a sealing film II are bonded to the end faces of the fire transfer holes.

8. The small solid rocket engine thrust termination apparatus of claim 1 wherein: the snapping screw is provided with a V-shaped weakening groove along the circumferential direction.

9. The small solid rocket engine thrust termination apparatus of claim 1 wherein: the piston part comprises a piston, the piston is of a cylindrical structure with an opening at one end, a threaded hole matched with the external thread of the breaking screw is formed in the center of the bottom of the piston, and a plurality of O-shaped sealing rings IV are arranged between the piston and the shell.

10. The small solid rocket engine thrust termination apparatus of claim 9 wherein: the piston is made of high-strength alloy structural steel, and a hardened layer is arranged on the surface of the piston.

Technical Field

The invention relates to the field of solid rocket engines, in particular to a thrust termination device of a small-sized solid rocket engine.

Background

Thrust termination is a mature and effective technology for realizing the interstage separation of the solid rocket engine, adjusting the missile range, controlling the flight precision and starting or terminating the flight for multiple times, and is widely applied to various large and medium-sized solid rocket engines at home and abroad. The thrust termination scheme of the existing solid rocket engine is mainly divided into three categories, namely a reverse thrust method, namely a plurality of holes are symmetrically formed in the tail part of an engine shell, a reverse jet pipe which forms a certain angle with the axial direction of the engine is installed, and when the thrust is terminated, an opening mechanism opens the reverse jet pipe to counteract the thrust of the engine; the other is a pressure-reducing extinguishing method, which mainly adopts the ways of energy-gathering cutting, bolt explosion, cap-plugging opening device and the like to rapidly open an additional gas release channel of the engine shell, so that the pressure in a combustion chamber is rapidly reduced at a certain speed, the stable combustion condition of the propellant is destroyed, and the purpose of flameout or thrust termination is achieved; and thirdly, a forced cooling method is adopted, coolant is sprayed to the combustion chamber, so that the combustion flame of the propellant is extinguished, and thrust is not generated any more.

For a small solid rocket engine with a thrust termination function, the thrust termination method has the following problems when applied:

a) the back spray pipe scheme needs to arrange a plurality of back spray pipes along the circumferential direction of the engine shell, and is structurally difficult to realize for a small-diameter engine or a propellant end face combustion engine;

b) although the energy-gathered cutting rope, the explosion bolt and the plug cover opening scheme are easy to realize in structure, the energy-gathered cutting rope, the explosion bolt and the plug cover all contain energetic materials such as explosives, instantaneous impact overload generated by explosion is up to hundreds of thousands g, splashed fragments can damage surrounding structures, reverse instantaneous thrust can be brought to a large cutting separation surface, and great thrust disturbance is generated to a small-thrust engine;

c) the forced cooling method needs to add a cooling mechanism with a complex structure in the combustion chamber, flameout is completed after the coolant is fully cooled, the thrust termination time is longer, and the forced cooling method is generally suitable for engine bench tests or occasions with low requirements on the thrust termination time.

Disclosure of Invention

The invention provides a thrust termination device of a small solid rocket engine, which is provided with two-stage interlocking pressure relief channels, wherein the pressure generated by combustion of a pyrotechnic ignition agent is used as a power source to open the pressure relief channels step by step without detonation output.

In order to achieve the purpose, the invention adopts the following technical scheme:

a thrust terminating device of a small solid rocket engine comprises a shell, a blocking part, an ignition medicine box part, a breaking screw and a piston part;

the shell is provided with a flange end and is fixedly connected to the engine shell through the flange end, and a primary pressure relief channel and a secondary pressure relief channel which are mutually perpendicular and communicated are arranged in the shell, wherein the primary pressure relief channel is communicated with a combustion chamber of the engine, and the secondary pressure relief channel is parallel to the direction of an engine spray pipe;

the ignition medicine box part is fixedly arranged at the tail part of the shell, and the exhaust end of the ignition medicine box part is aligned to the secondary pressure relief channel; the piston component is fixed in the secondary pressure relief channel through a snapping screw; the blocking component is limited between the first-stage pressure relief channel and the second-stage pressure relief channel by the piston component.

Preferably, the pressure relief device further comprises a heat insulator, a pressing screw and a protecting cylinder, wherein the heat insulator is arranged in the first-stage pressure relief channel, the tail of the heat insulator is sleeved with the pressing screw, the pressing screw can be matched with the piston component to limit the blocking component, and the protecting cylinder is fixedly arranged at the outlet of the shell.

Preferably, the blocking component comprises a heat insulating layer, a metal plug and an O-shaped sealing ring I, the heat insulating layer is arranged between the metal plug and the pressing screw, the contact surface between the metal plug and the piston component is an ellipsoidal curved surface, and the O-shaped sealing ring I is arranged between the metal plug and the shell.

Preferably, the metal plug is made of high-strength alloy structural steel, and a hardened layer is provided on the surface of the metal plug.

Preferably, the inner diameter of the protective cylinder is larger than that of the secondary pressure relief channel, and an O-shaped sealing ring II is arranged between the protective cylinder and the shell.

Preferably, the ignition medicine box part comprises a medicine box body, a flame igniter, an ignition tablet and a medicine box seat, the flame igniter is arranged at the tail of the medicine box body, an O-shaped sealing ring III is arranged between the medicine box body and the shell, and the ignition tablet is a flaky pyrotechnic ignition agent and is freely filled in a space formed by the medicine box body and the medicine box seat.

As the preferred of above-mentioned scheme, all seted up on medicine box body and the medicine box seat and passed the fire hole, it has sealing film I and sealing film II to pass the fire hole terminal surface bonding.

Preferably, the snap screw is provided with a "V-shaped" weakening groove along the circumferential direction.

Preferably, the piston part comprises a piston, the piston is of a cylindrical structure with one open end, a threaded hole matched with the external thread of the breaking screw is formed in the center of the bottom of the piston, and a plurality of O-shaped sealing rings IV are arranged between the piston and the shell.

Preferably, the piston is made of high-strength alloy structural steel, and a hardened layer is arranged on the surface of the piston.

Due to the structure, the invention has the advantages that:

1. the two-stage interlocking pressure relief channels are arranged, the thrust termination of the solid rocket engine is realized after the two-stage pressure relief channels are opened step by step according to the instruction, and the reliability and the safety of the thrust termination process are improved.

2. A flame igniter is used for igniting a small amount of pyrotechnic ignition agent, the pressure generated by combustion of the pyrotechnic ignition agent is used as a power source, the response is rapid, detonation output is not generated, the instantaneous impact overload is small, and splashing explosive fragments are not formed.

3. The pressure relief area is moderate, reverse instantaneous thrust brought by the separation of the piston is small, and generated thrust disturbance is small.

4. The small-sized solid rocket engine has the advantages of simple and compact structure, light weight and low cost, and can adapt to small-sized solid rocket engines with different sizes by adjusting the area of the pressure relief channel.

Drawings

In order to more clearly illustrate the technical solutions in the embodiments of the present invention, the drawings used in the description of the embodiments will be briefly introduced below.

FIG. 1 is a schematic view of the overall structure of the present invention;

FIG. 2 is an enlarged partial schematic view of the present invention;

fig. 3 is a schematic view of the plugging member of the present invention.

In the drawings, the correspondence between each component and the reference numeral is:

1-ignition medicine box component, 11-flame igniter, 12-ignition tablet, 13-medicine box body, 14-medicine box seat, 15-sealing film I, 16-sealing film II, 2-O type sealing ring III, 3-breaking screw, 31- 'V type' weakening groove, 4-piston component, 41-piston, 42-O type sealing ring IV, 5-blocking component, 51-heat insulating layer, 52-metal block, 53-O type sealing ring I, 6-pressing screw, 7-heat insulating body, 8-shell, 81-flange end, 9-O type sealing ring II, 10-protective cylinder, R1-primary pressure relief channel and R2-secondary pressure relief channel.

Detailed Description

The technical solution of the present invention will be clearly and completely described below with reference to the accompanying drawings of the present invention. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.

As shown in fig. 1 to 3, the present embodiment provides a thrust terminating device for a small solid rocket engine, which includes an ignition cartridge unit 1, a snap screw 3, a piston unit 4, a blocking unit 5, a press screw 6, a heat insulator 7, a housing 8, a casing 10, a primary pressure relief passage R1, and a secondary pressure relief passage R2.

The shell 8 is provided with a flange end 81 and is fixedly connected to the engine shell through the flange end 81, and a primary pressure relief channel R1 and a secondary pressure relief channel R2 which are mutually perpendicular and communicated are arranged in the shell 8; the first-stage pressure relief channel R1 is communicated with an engine combustion chamber, and the second-stage pressure relief channel R2 is parallel to the direction of an engine spray pipe.

The ignition medicine box component 1 is arranged at the tail part of the shell 8, the exhaust end of the ignition medicine box component is aligned to the secondary pressure relief channel R2 and is communicated through a plurality of exhaust holes on the shell 8; the piston component 4 is fixed in the secondary pressure relief channel R2 by a breaking screw 3; the blocking component 5 is limited between the first-stage pressure relief channel R1 and the second-stage pressure relief channel R2 by the piston component 4 and the pressing screw 6, and plays a role in isolating the two-stage pressure relief channels; a shroud 10 is mounted to the housing 8 outlet.

In this embodiment, the ignition cartridge component 1 includes a cartridge body 13, a flame igniter 11, an ignition tablet 12, and a cartridge seat 14, the cartridge body 13 is sealed with the housing 8 by an O-ring seal iii 2, the flame igniter 11 is installed at the tail of the cartridge body, and the ignition tablet 12 is a pre-made sheet-like pyrotechnic igniter and is freely filled in the cartridge body 13 and the cartridge seat 14. After the flame igniter ignites the ignition tablet, the gas pressure is quickly generated, so that detonation output can be avoided, and overlarge impact overload is caused.

In this embodiment, the medicine box body 13 and the medicine box seat 14 are both provided with a plurality of fire transmission holes, and the end faces of the fire transmission holes of the medicine box body 13 and the medicine box seat 14 are bonded with a sealing film i 15 and a sealing film ii 16 for ensuring the sealing performance of the ignition medicine box component 1 in the long-term storage process and preventing the ignition tablet from being affected with damp.

In this embodiment, the cylindrical section of the snap screw 3 is provided with a "V-shaped" weakening groove 31 in the circumferential direction, and the snap screw 3 breaks from the "V-shaped" weakening groove 31 when the pressure to which the piston member is subjected reaches a breaking threshold of the "V-shaped" weakening groove.

In this embodiment, the piston part 4 includes a piston 41, a threaded hole matched with the external thread of the snap screw 3 is formed in the center of the bottom of the piston 41 for installation and fixation, after the snap screw 3 is broken, the piston part 4 is released from constraint in the secondary pressure relief channel R2, and a plurality of O-ring seals iv 42 are arranged on the periphery of the piston 41 for ensuring sealing with the housing.

In the present embodiment, the plug member 5 includes a heat insulating layer 51, a metal plug 52, and an O-ring i 53; the heat insulation layer 51 is arranged between the metal plug 52 and the pressing screw 6, is vulcanized and molded on the basis of the metal plug 52 by adopting a flexible ablation-resistant material and is used for isolating high-temperature fuel gas, so that the O-shaped sealing ring I53 is ensured to work at normal temperature, and the high-temperature and high-pressure fuel gas of the engine is isolated in the primary pressure relief channel; the contact surface of the metal plug 52 and the piston member 4 is an ellipsoidal curved surface structure, and the sliding friction resistance between the two can be reduced.

In this embodiment, the inner diameter of the casing 10 is larger than the inner diameter of the secondary pressure relief passage R2, and the casing 10 is sealed with the housing 8 by the O-ring ii 9, so as to protect the surrounding structure of the engine from being damaged or affected by the piston flying out at high speed and high-temperature and high-pressure combustion gas.

In this embodiment, the piston 41 and the metal plug 52 are both made of high-strength alloy structural steel, and subjected to surface hardening treatment such as ternary co-carburization, so as to form a hardened layer, specifically, a carburized layer or a nitrided layer, on the surface, thereby avoiding piston member motion stagnation caused by contact surface deformation.

The working process principle of the structure is as follows: when the solid rocket engine needs thrust termination, a command is sent to a thrust termination device, a flame igniter 11 ignites an ignition tablet 12 in an ignition medicine box component 1, and the generated high-pressure gas quickly builds pressure in a sealing area between the ignition medicine box component 1 and a piston component 4; when the pressure reaches a damage threshold value, the snapping screw 3 is broken from the V-shaped weakening groove 31, and the piston part 4 is released from the constraint in the secondary pressure relief channel R2 and moves outwards rapidly; when the piston part 4 completely passes over the arc surface of the blocking part 5, the blocking part 5 loses support, and high-temperature and high-pressure fuel gas in a primary pressure relief channel R1 communicated with a combustion chamber of the solid rocket engine pushes the blocking part out. At the moment, the secondary pressure relief channel R2 and the primary pressure relief channel R1 are sequentially opened step by step, the solid rocket engine combustion chamber is communicated with the outside and starts to be quickly decompressed, the stable combustion condition of the propellant is destroyed, and the purpose of flameout or thrust termination is realized.

According to the thrust termination device of the small solid rocket engine, the two-stage interlocking pressure relief channels are arranged, and the thrust termination of the solid rocket engine is realized after the two-stage pressure relief channels are opened step by step according to the instruction, so that the reliability and the safety of the thrust termination process are improved; a flame igniter is used for igniting a small amount of pyrotechnic ignition agent, the pressure generated by the combustion of the pyrotechnic ignition agent is used as a power source, the response is rapid, the detonation output is not generated, the instantaneous impact overload is small, and the splashed explosive fragments are not formed; the pressure relief area is moderate, reverse instantaneous thrust is small when the piston is separated, and generated thrust disturbance is small; the small-sized solid rocket engine has the advantages of simple and compact structure, light weight and low cost, and can adapt to small-sized solid rocket engines with different sizes by adjusting the area of the pressure relief channel.

The above is only a preferred embodiment of the present invention, and is not intended to limit the present invention, and various modifications and changes will occur to those skilled in the art. Any modification, equivalent replacement, or improvement made within the spirit and principle of the present invention should be included in the protection scope of the present invention.