CN115199435A

CN115199435A - Small solid rocket engine capable of controlling shutdown

Info

Publication number: CN115199435A
Application number: CN202210847367.1A
Authority: CN
Inventors: 吴浩东; 任翼翔; 虞国军
Original assignee: Hubei Sanjiang Aerospace Honglin Exploration and Control Co Ltd
Current assignee: Hubei Sanjiang Aerospace Honglin Exploration and Control Co Ltd
Priority date: 2022-07-19
Filing date: 2022-07-19
Publication date: 2022-10-18

Abstract

The invention discloses a small solid rocket engine with controllable shutdown, which belongs to the technical field of rocket engines and comprises a charging chamber, a sealing cover, a long tail nozzle and main charges, wherein the engine can rapidly discharge fuel gas in the charging chamber and reduce the pressure of the fuel gas in the charging chamber after opening a second channel by utilizing the corresponding design of a first channel/a second channel on the long tail nozzle and the characteristics of the main charges and matching with the arrangement of a thrust termination device, so that the pressure reduction and flameout of the main charges and the forced shutdown of the engine are completed. The small solid rocket engine capable of being shut down controllably has the advantages of compact structure, convenience in assembly and control, capability of realizing forced shutdown of the engine while ensuring normal propelling work of the engine, quick shutdown response of the engine, small impact overload and no structural damage, capability of realizing shutdown operation at any time within the design working time of the engine, full guarantee of working performance and control accuracy of the small solid rocket engine, and good application prospect and popularization value.

Description

Small solid rocket engine capable of controlling shutdown

Technical Field

The invention belongs to the technical field of rocket engines, and particularly relates to a small solid rocket engine capable of being controlled to shut down.

Background

In the driving process of aircrafts such as rockets, missiles and the like, an engine plays an extremely important role. The shutdown controllability of the solid rocket engine serving as a power device in the use process is very important for improving the maneuverability, the interstage separation, the penetration capability and the flight precision of an aircraft.

Generally, a shutdown method of a solid rocket engine is divided into a normal shutdown and a forced shutdown. Wherein, the normal shutdown refers to normal combustion of the propellant until the propellant is finished, and the forced shutdown refers to no thrust output of the engine by adopting a special means. The problem that the shutdown is not timely when the aircraft or the projectile body is normally shut down can be solved by forced shutdown, so that the flight range, the flight precision, the hit precision and the like of the aircraft or the projectile body are influenced, and the problem can be effectively avoided by forced shutdown.

At present, the forced shutdown method generally includes a reverse thrust method, a forced cooling method and a decompression flameout method, and the forced shutdown of the engine is completed by providing a reverse thrust, cooling a propellant in a combustion chamber and reducing the pressure in the combustion chamber respectively. Although the above-mentioned methods can both satisfy the forced shutdown of the engine to a certain extent, for small solid rocket engines, the counterthrust method and the forced cooling method cannot accurately satisfy the requirements in terms of structure and response time, and the pressure-reducing flameout method relying on the existing methods of energy-gathering cutting, bolt explosion and the like has the problems of overlarge impact overload, structural damage and overlarge thrust interference, and cannot accurately satisfy the forced braking requirements of the small solid rocket engines.

Disclosure of Invention

Aiming at one or more of the defects or the improvement requirements in the prior art, the invention provides the small solid rocket engine with controllable shutdown, which can accurately realize forced shutdown of the small solid rocket engine, realize controllable shutdown of the engine and ensure the reliability and the accuracy of the setting and the use of the engine.

In order to achieve the purpose, the invention provides a small solid rocket engine capable of being controlled to shut down, which comprises a charging chamber, a sealing cover and a long tail nozzle which are sequentially connected, wherein main charge is packaged in the charging chamber;

the main charge is a double-base propellant, has the characteristics of pressure reduction and flameout, and can be flameout when the pressure of the gas in the charging chamber is suddenly reduced;

one end of the long tail spray pipe is communicated with the medicine loading chamber through a sealing cover and comprises a first channel and a second channel;

the first channel is a fuel gas spraying channel when the engine works normally, and a throat liner is arranged in one end of the first channel, which is far away from the sealing cover, and is used for reducing the exhaust area of the first channel; one end of the second channel is communicated with one end of the first channel close to the sealing cover, and the other end of the second channel is provided with a sealing component and a thrust termination device which can relieve the sealing of the sealing component; and is provided with

And the exhaust area of the second channel after the blockage is removed is larger than that of the throat part of the first channel, and the second channel is used for quickly discharging the gas in the charging chamber, so that the sudden drop of the gas pressure in the charging chamber and the decompression and flameout of the main charging are realized.

As a further improvement of the invention, the main charge comprises nitrocotton with the content of 50-56% and nitroglycerin with the content of 27.2-35%, and the main charge has the characteristics of pressure reduction and flameout when the pressure reduction rate reaches 50 MPa/s.

As a further improvement of the present invention, the exhaust area of the second passage after unblocking is 17 times or more the exhaust area of the throat portion of the first passage.

As a further improvement of the invention, a heat insulating layer is arranged between the main charge and the charging chamber; the heat insulation layer is formed on the inner wall surface of the medicine loading chamber.

As a further improvement of the invention, one end of the main charge, which faces away from the cover, is provided with a compensation pad;

the compensation pad is arranged between the end part of the main charge and the closed end of the charge chamber, the initial compression amount of the compensation pad is 6% -9.5%, and the maximum compression amount is not less than 60%.

As a further improvement of the invention, a heat insulating material is arranged on one side of the sealing cover connected with the medicine loading chamber in a mould pressing mode, an ignition medicine box is embedded in the heat insulating material, and at least one igniter is arranged on the outer side of the sealing cover corresponding to the ignition medicine box.

As a further improvement of the invention, the first channel is coaxially arranged with the loading chamber, and the second channel is perpendicular to the first channel, so that the long tail nozzle is of a T-shaped structure.

As a further development of the invention, a diaphragm is arranged at the end of the first channel facing away from the cover for closing this end of the first channel.

As a further improvement of the present invention, heat insulators are provided on inner wall surfaces of the first passage and the second passage for heat insulation during operation and exhaust of the two passages.

As a further improvement of the invention, the heat insulator is a high silica/carbon phenolic composite material which is formed by winding and is arranged on the inner wall surface of the long tail nozzle in a mould pressing mode.

The above-described improved technical features may be combined with each other as long as they do not conflict with each other.

In general, compared with the prior art, the technical scheme conceived by the invention has the following beneficial effects:

(1) The invention relates to a small solid rocket engine capable of being controlled to shut down, which comprises a loading chamber, a sealing cover, a long tail nozzle and main charges, wherein the corresponding design of a first channel and a second channel with different exhaust areas on the long tail nozzle is utilized, the main charges are set into double-base propellants with the pressure reduction and flameout characteristics, and a controllably communicated thrust termination device is arranged corresponding to the second channel, so that the second channel can be opened when the engine has a forced shutdown requirement, the fuel gas in the loading chamber is rapidly discharged, the fuel gas pressure in the loading chamber is reduced, the main charges are rapidly flameout, the forced shutdown control of the engine is completed, the controllability of the engine is improved, the accuracy and the control accuracy of the flight control of a corresponding aircraft or a projectile body are ensured, and the working accuracy of the engine and even equipment carrying the engine is improved.

(2) According to the small solid rocket engine capable of being controlled to shut down, the pressure reduction and flameout characteristics of the engine can be accurately guaranteed by further optimizing the composition of the main charge and the ratio of the exhaust areas of the two channels, the normal propelling work of the engine is ensured, the flameout and braking process during forced shutdown is accurately finished, and the response rate and the control accuracy during forced shutdown are improved.

(3) The small solid rocket engine with controllable shutdown ensures the reliability of the engine and other equipment at the assembly part during ignition operation by arranging the heat insulating material at the inner sides of the structures such as the charging chamber, the long tail pipe, the sealing cover and the like, and avoids the overheating influence on other equipment during the operation of the engine. Simultaneously, through set up the compensation pad between the blind end at the charging chamber and the main charge, the pressure effect to this side when fully compensating main charge ignition work guarantees the reliability that sets up of charging chamber blind end, ensures the driven accuracy of igniting.

(4) The small solid rocket engine capable of being shut down controllably has the advantages of compact structure, convenience in assembly and control, capability of realizing forced shutdown of the engine while ensuring normal propelling work of the engine, quick shutdown response of the engine, small impact overload and no structural damage, capability of realizing shutdown operation of the engine at any time in designed working time, full guarantee of working performance and control accuracy of the small solid rocket engine, and good application prospect and popularization value.

Drawings

In order to more clearly illustrate the technical solutions in the embodiments of the present invention, the drawings needed to be used in the embodiments will be briefly described below, and it is obvious that the drawings in the following description are only some embodiments of the present invention, and it is obvious for those skilled in the art that other drawings can be obtained based on these drawings without creative efforts.

FIG. 1 is a sectional view of a small solid rocket engine with controlled shutdown in an embodiment of the present invention;

FIG. 2 is a schematic end view of a small solid rocket engine with controlled shutdown according to an embodiment of the present invention;

FIG. 3 is a schematic view of a small solid rocket engine with controlled shutdown in various states according to an embodiment of the present invention;

FIG. 4 is data obtained from a controlled shutdown small solid rocket engine test in an embodiment of the present invention;

throughout the drawings, like reference numerals designate like features, and in particular:

1. an igniter; 2. sealing the cover; 3. an ignition cartridge; 4. a compensation pad; 5. a main charge; 6. a medicine loading chamber; 7. a long tail nozzle; 8. a heat insulator; 9. a throat liner; 10. a membrane; 11. pressing the snail; 12. an isolation gasket; 13. and a thrust termination device.

Detailed Description

In order to make the objects, technical solutions and advantages of the present invention more apparent, 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 merely illustrative of the invention and are not intended to limit the invention. In addition, the technical features involved in the respective embodiments of the present invention described below may be combined with each other as long as they do not conflict with each other.

In the description of the present invention, it is to be understood that the terms "central," "longitudinal," "transverse," "length," "width," "thickness," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," "outer," "clockwise," "counterclockwise," "axial," "radial," "circumferential," and the like are used in the orientations and positional relationships indicated in the drawings for convenience in describing the invention and to simplify the description, but are not intended to indicate or imply that the device or element so referred to must have a particular orientation, be constructed and operated in a particular orientation, and are not to be construed as limiting the invention.

Furthermore, the terms "first", "second" and "first" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance or implicitly indicating the number of technical features indicated. 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, "a plurality" means at least two, e.g., two, three, etc., unless explicitly specified otherwise.

In the present invention, unless otherwise expressly stated or limited, the terms "mounted," "connected," "secured," and the like are to be construed broadly and can, for example, be fixedly connected, detachably connected, or integrally formed; can be mechanically or electrically connected; they may be directly connected or indirectly connected through intervening media, or they may be interconnected within two elements or in a relationship where two elements interact with each other unless otherwise specifically limited. The specific meanings of the above terms in the present invention can be understood by those skilled in the art according to specific situations.

In the present invention, unless otherwise expressly stated or limited, the first feature "on" or "under" the second feature may be directly contacting the first and second features or indirectly contacting the first and second features through an intermediate. Also, a first feature "on," "over," and "above" a second feature may be directly or diagonally above the second feature, or may simply indicate that the first feature is at a higher level than the second feature. A first feature being "under," "below," and "beneath" a second feature may be directly under or obliquely under the first feature, or may simply mean that the first feature is at a lesser elevation than the second feature.

Example (b):

referring to fig. 1-2, a controlled shutdown small-sized solid rocket engine according to a preferred embodiment of the present invention comprises a charging chamber 6, a cover 2, an ignition module and a long tail nozzle 7. Wherein, the main charge 5 is arranged in the charge chamber 6 in a filling way, so that the main charge can be ignited under the action of the ignition assembly, and high-pressure gas is sprayed out from the long tail nozzle 7, thereby completing the driving process of the engine.

Specifically, the loading chamber 6 in the preferred embodiment is a tubular structure with one end closed and one end open, and the cover 2 is connected to the opening of the loading chamber 6 for forming a closed accommodation cavity in the loading chamber 6. Accordingly, the main charge 5 is freely loaded in the charge chamber 6 and is provided with a compensation pad 4 on its side facing away from the closure 2 (i.e. on the left side as viewed in fig. 1). Meanwhile, corresponding to the filling of the main charge 5, an insulating layer is laid on the inner wall of the charge chamber 6, and the insulating layer is further preferably formed in a high-temperature and pressurization vulcanization mode, so that the influence of the charge chamber 6 on other equipment when the main charge 5 is ignited for work is avoided.

In more detail, in the preferred embodiment, the main charge 5 is of end-fired cylindrical structure, is a double-base propellant, and is formulated to have a pressure-reducing and flameout characteristic, so that flameout shutdown can be completed when the gas pressure is suddenly reduced. In a specific design, the main charge 5 comprises a first base material and a second base material, and has decompression and flameout characteristics when the decompression rate reaches 50 MPa/s. Wherein, the first base material is preferably nitrocotton, and the content thereof is preferably 50 to 56 percent; the second base material is preferably nitroglycerin, and the content of the nitroglycerin is preferably 27.2-35%.

It will of course be appreciated that in actual implementation, in addition to the two materials mentioned above, the main charge 5 will contain minor amounts of other admixtures such as dinitrotoluene, petrolatum, medium size II, etc.

Meanwhile, as shown in fig. 1, the compensation pad 4 in the preferred embodiment is arranged between the main charge 5 and the closed end of the charge chamber 6, the material of the compensation pad is aviation sponge rubber, the initial compression amount of the compensation pad is 6% -9.5%, the maximum compression amount is not less than 60%, and the buffer compensation on one side of the closed end of the charge chamber 6 after the main charge 5 is ignited can be effectively realized.

Further, the closure 2 in the preferred embodiment is attached to the opening of the loading chamber 6, preferably by a screw connection, and sealed by an O-ring. Meanwhile, the inside of the cover 2 in the preferred embodiment is molded with a heat insulating material, so that after the cover 2 is connected with the medicine loading chamber 6, the side of the main medicine 5, which faces away from the compensation pad 4, can be limited by the heat insulating material.

More specifically, the ignition cartridge 3 is embedded "deeply" in the heat insulating material of the cover 2, and preferably is threadably engaged with the cover of the cover 2 to ensure reliable placement of the ignition cartridge 3. Correspondingly, an igniter 1 is arranged on the outer side of the cover 2 corresponding to the ignition medicine box 3, so that the main charge 5 can complete the ignition process through the ignition medicine box 3 and the igniter 1. In the preferred embodiment, the igniter 1 is preferably symmetrically disposed two as shown in FIG. 2, and the igniter 1 is preferably threaded with the cap 2 and sealed with a copper gasket.

Further, the long tail pipe 7 in the preferred embodiment is connected with the cover 2 through a flange, and comprises a first passage for exhausting gas during normal operation of the engine and a second passage communicated with the first passage, so that one inlet and two outlets of a gas passage are realized, and the engine can be shut down and flameout when the second passage is opened.

More specifically, the first passage in the preferred embodiment is arranged coaxially with the loading chamber 6, and the second passage communicates with the front end of the first passage and is further perpendicular to its axis, forming a "T-shaped" configuration as shown in FIG. 1. Correspondingly, a throat insert 9 is arranged in the tail end (the side departing from the sealing cover 2) of the first channel and is used for reducing the inner diameter of the fuel gas spraying channel and completing the pressurized spraying of the fuel gas. At the same time, a membrane 10 is enclosed at the end of the first channel, which is preferably limited by a press screw 11.

As shown in fig. 1, the second channel in the preferred embodiment communicates vertically with the front end of the first channel and is closed at its end by a closure assembly and in correspondence with which a thrust stop 13 is provided. In actual arrangement, the thrust termination device 13 is in sealing connection with the second channel through the isolation sealing gasket 12, so that the sealing performance of the second channel in the non-working state is ensured, and the normal use requirement of the engine is met.

The arrangement of the thrust termination device 13 can be used for reliably limiting the plugging component, so that the second channel can be accurately plugged when the engine works normally; meanwhile, when the engine needs to be braked and shut down, the limit of the plugging component can be quickly released through the control of the thrust termination device 13, and then the plugging of the second channel is released, so that the fuel gas in the charging chamber 6 can be quickly discharged through the second channel, and then the pressure reduction and flameout of the main charge 5 are realized, and the forced shutdown of the engine is completed.

It should be noted that, in the preferred embodiment of the present invention, the core concerned is not in the thrust terminating device 13, so the specific structural configuration is not particularly limited as long as the blockage control of the second passage can be realized. Accordingly, the plugging assembly matching the thrust termination device 13 and the second channel end will not be described in detail.

Preferably, the heat insulator 8 is arranged on the inner peripheral wall surface of the channel of the long tail nozzle 7, and is used for realizing heat insulation of the long tail nozzle 7 during working and avoiding the influence of the long tail nozzle 7 during air injection on surrounding equipment. In more detail, the insulator 8 in the preferred embodiment is a high silica/carbon phenolic composite material, which is further preferably formed by winding and is provided by being molded on the inner wall surface of the long nozzle 7.

Further preferably, the diaphragm 10 is made of aluminum material and is preferably provided with a preformed annular groove, which is held down by a holding-down screw 11. Correspondingly, the isolation sealing gasket 12 arranged between the thrust terminating device 13 and the long tail nozzle 7 is preferably made of reinforced flexible graphite material, the compression amount of the isolation sealing gasket is 30% -50%, and the isolation sealing gasket is further arranged in annular grooves in flanges of the sealing cover 2 and the thrust terminating device 13 so as to seal and isolate high-temperature and high-pressure fuel gas.

Furthermore, in the actual arrangement, the thrust stop 13 is connected to the nozzle antechamber channel, i.e. the connection point is located on the side of the throat insert 9 which is adjacent to the cover 2. Meanwhile, the exhaust area of the second channel in the preferred embodiment is more than 17 times of the throat area of the throat liner 9, so that the thrust termination device 13 can open the gas channel of the spray pipe front chamber after receiving a shutdown signal, the high-temperature gas is exhausted from the second channel, the gas pressure in the charging chamber 6 is rapidly reduced, the pressure reduction and flameout of the main charge 5 are realized, and the forced shutdown process is completed.

For the small solid rocket engine in one embodiment, after the assembly is completed, a ground hot test is performed, and the test condition is shown in fig. 3, wherein the left side of fig. 3 is the normal working condition of the engine, the middle part is the condition after the engine is shut down, and the right side is the main charge 5 left after shutdown and flameout, which indicates that the engine can be forcibly shut down under the action of the thrust termination device 13, and a certain amount of the main charge 5 remains in the charge chamber 6. Accordingly, the data obtained for the above test is shown in fig. 4, and it can be seen from the test image in fig. 3 and the test data in fig. 4 that: the engine can be shut down reliably, and each index can satisfy the operation requirement. In addition, the small solid rocket engine in the preferred embodiment participates in flight tests along with the system, which proves that the engine can normally complete the propulsion process when the thrust termination device 13 does not work, and all indexes meet the use requirements.

The small solid rocket engine capable of being shut down controllably has the advantages of compact structure, convenience in assembly and control, capability of realizing forced shutdown of the engine while ensuring normal propelling work of the engine, quick shutdown response of the engine, small impact overload and no structural damage, capability of realizing shutdown operation of the engine at any time in designed working time, full guarantee of working performance and control accuracy of the small solid rocket engine, and good application prospect and popularization value.

It will be understood by those skilled in the art that the foregoing is only an exemplary embodiment of the present invention, and is not intended to limit the invention to the particular forms disclosed, since various modifications, substitutions and improvements within the spirit and scope of the invention are possible and within the scope of the appended claims.

Claims

1. A small solid rocket engine capable of controlling shutdown comprises a loading chamber, a sealing cover and a long tail nozzle which are sequentially connected, and main charge is sealed in the loading chamber, which is characterized in that,

one end of the long tail nozzle is communicated with the medicine loading chamber through a sealing cover and comprises a first channel and a second channel;

2. A small controlled-shutdown solid-rocket engine according to claim 1 wherein said main charge comprises nitrocotton in an amount of 50-56% and nitroglycerin in an amount of 27.2-35%, and has decompression quenching properties at a decompression rate of 50 MPa/s.

3. A small controlled-shutdown solid-rocket engine according to claim 2 wherein the exhaust area of said second passageway after unblocking is more than 17 times the exhaust area of the throat of said first passageway.

4. A small solid rocket engine with controlled shutdown according to any one of claims 1-3 wherein an insulating layer is further provided between the main charge and the charging chamber; the heat insulation layer is formed on the inner wall surface of the medicine loading chamber.

5. A small solid rocket engine with controlled shutdown according to any one of claims 1 to 4 wherein the end of the main charge facing away from the cover is provided with a compensation pad;

the compensation pad is arranged between the end part of the main charge and the closed end of the charge chamber, the initial compression amount is 6-9.5%, and the maximum compression amount is not less than 60%.

6. A small-sized solid rocket engine with controllable shutdown according to any one of claims 1-5 wherein the side of the cover connected to the loading chamber is molded with a heat insulating material, an ignition cartridge is embedded in the heat insulating material, and at least one igniter is arranged outside the cover corresponding to the ignition cartridge.

7. A small solid rocket engine with controlled shutdown according to any of claims 1-6 wherein said first passageway is coaxially disposed with said loading chamber and said second passageway is perpendicular to said first passageway such that said long nozzle is of "T-shaped" configuration.

8. A small solid rocket engine with controlled shutdown according to any one of claims 1 to 7 wherein the end of the first passage facing away from the cover is provided with a membrane for closing off the end of the first passage.

9. A small-sized solid rocket engine with controllable shutdown according to any one of claims 1 to 8 wherein the inner wall surfaces of said first and second passages are provided with thermal insulation for thermal insulation during operation and exhaust of both passages.

10. A small solid rocket engine with controlled shutdown according to claim 9 wherein said thermal insulator is a high silica/carbon phenolic composite material formed by winding and molded on the inner wall of the long jet nozzle.