CN113623088A - Non-straight-through type composite pipeline of small solid rocket engine - Google Patents
Non-straight-through type composite pipeline of small solid rocket engine Download PDFInfo
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- CN113623088A CN113623088A CN202111044174.4A CN202111044174A CN113623088A CN 113623088 A CN113623088 A CN 113623088A CN 202111044174 A CN202111044174 A CN 202111044174A CN 113623088 A CN113623088 A CN 113623088A
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- 229910001069 Ti alloy Inorganic materials 0.000 description 2
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/97—Rocket nozzles
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/97—Rocket nozzles
- F02K9/978—Closures for nozzles; Nozzles comprising ejectable or discardable elements
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- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Filling Or Discharging Of Gas Storage Vessels (AREA)
- Rigid Pipes And Flexible Pipes (AREA)
Abstract
The invention relates to the technical field of solid rocket engines, and discloses a non-straight-through composite pipeline of a small-sized solid rocket engine, wherein one end of a first pipeline is connected with a rocket engine, and the other end of the first pipeline is closed; the second pipeline is connected with the side wall of the first pipeline, and divides the first pipeline into a conveying section and an accommodating section by the connecting point of the second pipeline and the first pipeline; the conveying section is used for conveying the propellant gas, the accommodating section is used for accommodating solid particles in the propellant gas, the propellant gas is turned at the accommodating section and is introduced into the second pipeline, and then the separation of the propellant gas and the solid particles is realized; the third pipeline is connected with the second pipeline and is used for accelerating the expansion and ejection of the propellant gas in the second pipeline. The non-straight-through composite pipeline of the small solid rocket engine has a relatively simple integral structure, improves the anti-scouring capability of the pipeline, improves the conveying smoothness degree of propellant gas in the pipeline, and ensures the thrust stability of the propellant gas.
Description
Technical Field
The invention relates to the technical field of solid rocket engines, in particular to a non-straight-through composite pipeline of a small solid rocket engine.
Background
The solid rocket engine is a chemical rocket engine using a solid propellant, and the solid propellant is usually ignited and then combusted in a combustion chamber to convert chemical energy into heat energy and generate combustion products with high temperature and high pressure, wherein the combustion products flow through a spray pipe, are expanded and accelerated in the spray pipe, convert the heat energy into kinetic energy and are discharged from the spray pipe at high speed to generate thrust.
In the design of the solid rocket engine, the design of the spray pipe and the shell and the selection of corresponding materials are key technologies in the solid rocket engine, and the use performance of the solid rocket engine can be directly influenced. The thrust of the solid rocket is mainly generated by the propellant gas accelerating above the sonic velocity at the nozzle and then diffusing through the nozzle. And often can carry a certain amount of burning granule in the propellant gas, these granules can continuously erode the spray tube pipe wall under the promotion of high temperature high pressure propellant gas, and one side of it can cause the damage to the inside pipe wall of spray tube, also can influence the flow direction of propellant gas simultaneously, causes the unstability of propellant gas thrust.
Disclosure of Invention
In view of one or more of the above drawbacks or needs for improvement in the prior art, the present invention provides a non-straight composite pipeline for a small solid rocket engine, which is used to solve the problem of serious erosion of the pipeline by combustion particles in the existing solid rocket engine.
In order to achieve the purpose, the invention provides a non-straight-through type composite pipeline of a small solid rocket engine, which comprises a first pipeline, a second pipeline and a third pipeline;
one end of the first pipeline is connected with the rocket engine, and the other end of the first pipeline is closed;
the second pipeline is connected with the side wall of the first pipeline, and the first pipeline is divided into a conveying section and an accommodating section by a connecting point of the second pipeline and the first pipeline; wherein,
the conveying section is used for conveying propellant gas,
the accommodating section is used for accommodating solid particles in the propellant gas, and the propellant gas is turned at the accommodating section and is introduced into the second pipeline so as to separate the propellant gas from the solid particles;
the third pipeline is connected with the second pipeline and is used for expanding, accelerating and spraying out the propellant gas in the second pipeline.
As a further development of the invention, the first line is parallel or coincident with the axial direction of the third line.
As a further improvement of the invention, a turning part for changing the gas flow direction in the second pipeline is arranged at the joint of the second pipeline and the third pipeline.
As a further improvement of the present invention, the third pipeline includes a compression section and an injection section, the compression section is connected to the second pipeline, and the inner diameter of the connection end of the compression section and the second pipeline is larger than the inner diameter of the connection end of the compression section and the injection section.
As a further improvement of the invention, a sealing part for sealing the nozzle of the injection section, which is away from the compression section, is arranged at the nozzle of the injection section.
As a further improvement of the invention, the sealing part is provided with a weakening groove for the sealing part to fall off integrally under the impact of propellant gas.
As a further improvement of the present invention, heat insulation layers are disposed on inner walls of the first pipeline, the second pipeline and the third pipeline.
As a further improvement of the present invention, the heat insulation layer is arranged in sections, and includes a first heat insulation layer, a second heat insulation layer, and a third heat insulation layer, which are respectively attached to the first pipeline, the second pipeline, and the third pipeline.
As a further improvement of the present invention, a first fixing portion for fixing the first pipe and the first heat insulating layer is provided at a connection end of the first pipe and the rocket engine.
As a further improvement of the present invention, the third pipeline is provided with a second fixing portion for fixing the third pipeline and the third heat insulating layer at the propellant gas ejection end.
The above-described improved technical features may be combined with each other as long as they do not conflict with each other.
Generally, compared with the prior art, the above technical solution conceived by the present invention has the following beneficial effects:
(1) according to the non-straight-through type composite pipeline of the small solid rocket engine, the spray pipe of the rocket engine is set to be of a multi-section non-straight-through type, and solid particles in propellant gas are intercepted and stored through shunting and intercepting of the containing section, so that the amount of the solid particles on the spray pipe path is reduced, the abrasion of the solid particles to the spray pipe is reduced, and the thrust stability of the rocket engine is ensured.
(2) According to the non-straight-through composite pipeline of the small solid rocket engine, the steering part and the compression part are arranged on the third pipeline aiming at the propellant gas, the power damage caused by collision of the propellant gas and the pipe wall is reduced by utilizing the steering part, the gas is compressed by utilizing the compression part, the injection power of the propellant gas is increased, and the rocket engine has higher driving force.
(3) According to the non-straight-through type composite pipeline of the small solid rocket engine, the first fixing part is arranged between the first pipeline and the first heat insulation layer, and the third fixing part is arranged between the third pipeline and the third heat insulation layer, so that the pipelines and the heat insulation layers are tightly connected, the good heat insulation effect of the heat insulation layers is ensured, the manufacturing and installation difficulty of the heat insulation layers is reduced, and the stable operation of the heat insulation layers under the impact of high-temperature and high-pressure propellant gas is ensured.
(4) According to the non-straight-through type composite pipeline of the small solid rocket engine, the sealing part is arranged at the outlet end of the third pipeline, so that external water vapor and the like of the rocket engine are prevented from entering the spray pipe under the unused state; set up the weakening groove in the sealing simultaneously for rocket engine starts the back, propellant gas when breaching the sealing, the sealing can wholly drop along weakening the groove, avoids the sealing part to remain in the pipe mouth department of third pipeline, influences the injection efficiency of propellant gas.
Drawings
FIG. 1 is a schematic view of the overall structure of a non-straight composite pipeline of a small solid rocket engine according to an embodiment of the present invention;
FIG. 2 is a schematic cross-sectional view of the interior of a non-straight composite piping structure of a small solid rocket engine according to an embodiment of the present invention;
FIG. 3 is an enlarged schematic view at A in FIG. 2;
fig. 4 is an enlarged schematic view at B in fig. 2.
In all the figures, the same reference numerals denote the same features, in particular:
1. a first pipeline; 2. a second pipeline; 3. a third pipeline; 4. a rocket motor; 5. a first insulating layer; 6. a second thermal insulation layer; 7. a third thermal insulation layer; 8. a first fixed part; 9. a second fixed part; 10. a termination mechanism;
101. a conveying section; 102. an accommodating section;
301. a steering section; 302. a compression section; 303. a spraying section; 304. a sealing part; 305. a weakening groove;
901. pressing the snail; 902. and (5) sealing rings.
Detailed Description
In order to make the objects, technical solutions and advantages of the present invention more apparent, the present invention is described in further 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 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," "lateral," "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, and are not intended to indicate or imply that the referenced devices or elements must have a particular orientation, be constructed and operated in a particular orientation, and are therefore not to be considered limiting of 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 specifically limited 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 connected internally or in any other suitable relationship, unless expressly stated otherwise. 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):
the non-straight-through type composite pipeline of the small solid rocket engine in the preferred embodiment of the invention is shown in figures 1-4 and comprises a first pipeline 1, a second pipeline 2 and a third pipeline 3; wherein one end of the first pipeline 1 is connected with the rocket engine 4, and the other end is closed; the second pipeline 2 is connected with the side wall of the first pipeline 1, and divides the first pipeline 1 into a conveying section 101 and a containing section 102 at the connecting point of the first pipeline 1; the conveying section 101 is used for conveying the propellant gas, the containing section 102 is used for containing solid particles in the propellant gas, the propellant gas is enabled to turn at the containing section 102 and is led into the second pipeline 2, and then separation of the propellant gas and the solid particles is achieved; the third pipeline 3 is connected with the second pipeline 2 and is used for compressing and spraying out the propellant gas in the second pipeline 2.
The pipeline of the conventional solid rocket engine 4 is of a straight-through type, although the combustion and injection factors of the solid fuel are fully considered, the combustion process of the solid fuel in the spray pipe still lasts for a period of time, and the injected fuel cannot be combusted completely instantly, so that when gas in the pipeline circulates, a large amount of solid particles are carried to wash the pipe wall and the inner wall structure, the pipeline structure is damaged, the injection rate of the gas is influenced due to the fact that the gas contains the solid particles, and the kinetic energy loss of the solid rocket engine 4 is improved. Set up to non-combined type through 4 pipelines with small-size solid rocket engine in this application, utilize multichannel pipeline cooperation form to it holds the region to remain the residue on the pipeline, carries out the gas-solid separation with the propellant gas, and saves solid particle, in order to avoid solid particle to the washing away of follow-up pipeline, overall structure design is simple, and it is high to solid particle's entrapment rate, has improved solid rocket engine 4's life and propellant gas's thrust stability greatly.
Preferably, the first pipeline 1, the second pipeline 2 and the third pipeline 3 in the invention are made of a titanium alloy material TC11, and the first pipeline 1, the second pipeline 2 and the third pipeline 3 are only named as pipelines at different positions, and the pipelines can be considered to be integrally arranged, of course, for convenience of manufacturing the pipelines and placing internal parts of subsequent pipelines, the pipelines are separately arranged, and are welded by electron beams, the penetration is 3mm, so that the airtightness, nondestructive testing, strength testing and the like of the welded pipelines meet the launching requirements of solid rockets.
Further, as shown in fig. 1, as a preferred embodiment of the present invention, the second pipeline 2 is disposed at a certain inclination angle with respect to the first pipeline 1 and the third pipeline 3, and the axial directions of the first pipeline 1 and the third pipeline 3 are parallel or consistent. Because a part of the first pipeline 1 is used for conveying the propellant gas and a part of the first pipeline is used for accommodating solid particles, when the propellant gas turns to be introduced into the second pipeline 2 at the accommodating section 102 and the solid particles are reserved in the accommodating section 102, if the accommodating section 102 is longer, the end sealing mode of the accommodating section 102 can cause the propellant gas to form hedging at the accommodating section 102, and the transportation kinetic energy of the propellant gas is reduced. Therefore, the length of the housing section 102 is strictly calculated, and the kinetic energy loss of the propellant gas at the housing section 102 needs to be considered while the solid particles are retained. Theoretically, when the first pipeline 1 and the second pipeline 2 are arranged in a straight line, the kinetic energy loss of the propellant gas is minimum, namely the propellant gas directly enters the second pipeline 2 without passing through the accommodating section 102, but the solid particles in the propellant gas cannot be intercepted at the moment. When the first pipeline 1 and the second pipeline 2 are arranged at a certain angle, the second pipeline 2 is arranged on the side wall of the first pipeline 1, the accommodating section 102 is formed at the end part of the first pipeline 1, solid particles impact the end part of the accommodating section 102 and stay, and propellant gas is transposed at the accommodating section 102, flows into the second pipeline 2 and flows through the third pipeline 3 to be sprayed out due to the gaseous fluidity of the propellant gas. At this time, the larger the angle between the first pipeline 1 and the second pipeline 2 is, the smaller the kinetic energy loss of the propellant gas is, and the poorer the retention capacity of the accommodating section 102 on the solid particles is; when the angle between first pipeline 1 and the second pipeline 2 is the acute angle, can seriously lose the kinetic energy of propellant gas to propellant gas itself high temperature, highly compressed characteristic can be strikeed first pipeline 1 and second pipeline 2 junction, cause pipeline structure's unstability, consequently the contained angle between first pipeline 1 and the second pipeline 2 of this place is 90 ~ 179, preferably 90, first pipeline 1 and second pipeline 2's connected mode is comparatively simple this moment, is higher to solid particle's entrapment rate under the less condition of loss kinetic energy.
Here, the length of the housing section 102 itself needs to be obtained according to the charge and injection rate inside the solid rocket engine 4, and in order to avoid the impact of the propellant gas injection on the closed end of the housing section 102, the closed section of the housing section 102 needs to be thickened.
Preferably, in the case of increasing the kinetic energy loss of the propellant gas, it is considered that the second pipeline 2 is formed in an "Contraband" shape, that is, the second pipeline 2 is respectively connected to the side wall between the first pipeline 1 and the third pipeline 3, so that a composite pipeline with the first pipeline 1 and the third pipeline 3 coaxially arranged can be formed.
Preferably, in order to ensure effective control of the injection speed and the launching direction of the solid rocket during launching, it is necessary to ensure that the overall installation direction of the solid rocket is consistent with the injection direction of the propellant gas injected from the third pipeline 3, that is, the installation direction of the first pipeline 1 is consistent with that of the third pipeline 3, so that the joint of the second pipeline 2 and the third pipeline 3 is also obliquely arranged, so that the gas conveying directions of the first pipeline 1 and the third pipeline 3 are the same.
Further, as shown in fig. 2, as a preferred embodiment of the present invention, a turning portion 301 for changing the flow direction of the gas in the second pipeline 2 is provided at the connection of the second pipeline 2 and the third pipeline 3. Since solid particles in the propellant gas are trapped at the accommodating section 102, the connection between the second pipeline 2 and the third pipeline 3 does not need to take on this function, and only the flow direction of the propellant gas needs to be redirected, which also means that a structure which is helpful for the air flow diversion can be arranged at the connection between the second pipeline 2 and the third pipeline 3, so as to reduce the kinetic energy loss of the propellant gas at the corner. Therefore, a turning part 301 is arranged at the joint of the second pipeline 2 and the third pipeline 3, and the turning part 301 adopts a high-strength graphite layer with a hemispherical characteristic, so that propellant gas is transited to the third pipeline 3 along an arc spherical surface, and the loss caused by collision of the propellant gas on the pipe wall of the third pipeline 3 is greatly reduced. The inner diameter of the hemispherical surface of the turning part 301 is determined according to the inner diameters of the second pipeline 2 and the third pipeline 3, so that the turning part 301 can be perfectly attached to the corners of the two pipelines, the high-strength graphite layer can directly bear the impact of propellant gas and the scouring of a small amount of residual solid particles, and the damage to the pipe wall is reduced.
Further, as a preferred embodiment of the present invention, as shown in fig. 3, the third pipeline 3 includes a compression section 302 and an injection section 303, wherein the compression section 302 is connected to the second pipeline 2, and the inner diameter of the connection end of the compression section 302 to the second pipeline 2 is larger than the inner diameter of the connection end of the compression section 302 to the injection section 303. In order to increase the injection efficiency of the propellant gas, a compression section 302 for compressing the propellant gas is provided on the third pipeline 3, so that the propellant gas which is originally transported at subsonic speed is accelerated to supersonic speed. Specifically, the compression section 302 compresses the propellant gas by reducing the inner diameter of the compression section 302, and the compression section 302 can reduce the inner diameter in the middle area to achieve the purpose of compressing the compressed gas, but a step structure is formed in the compression section 302, so that the propellant gas collides with the inner diameter dip, and the kinetic energy of the compression section 302 is reduced. Further, the inner diameter of the compression section 302 is set to be gradually reduced, that is, the compression space is set to be a conical structure, so as to gradually compress the compressed gas and reduce the loss of kinetic energy of the propellant gas. Preferably, a diffuser section, i.e. a conical diffuser space arranged opposite to the compression section 302, is further arranged at one end of the compression section 302 facing away from the second pipeline 2, so that the propellant gas sprayed from the compression section 302 expands and accelerates at the diffuser section to improve the propelling force of the propellant gas.
Preferably, in order to facilitate the arrangement of the internal structure of the compression section 302, the compression section 302 and the injection section 303 are arranged in a segmented manner, the compression section 302 and the injection section 303 are connected by means of threads and the like, and a proper amount of high-temperature sealant is coated at the threaded connection to ensure the tightness of the connection between the compression section 302 and the injection section 303. The compression section 302 and the injection section 303 cannot be directly connected by adopting a thread structure in a conventional straight-through pipeline, propellant gas in the straight-through pipeline contains a certain amount of solid particles, when the propellant gas is compressed by the compression section 302, collision of the solid particles on a pipe wall is also aggravated, the compression section 302 and the injection section 303 are easily separated, the whole pipeline of the straight-through pipeline has serious heat transfer phenomenon, the temperature of a connection part between the compression section 302 and the injection section 303 is high, the conventional thread is easy to appear or deform, and the like, so that the pipeline structure is damaged. This application is through setting up the pipeline into the multistage form to with solid particle interception storage, the very big degree has reduced the solid particle in the propellant gas, make compression section 302 when compression propellant gas, the impact force that receives is less, and the heat transfer problem is more straight formula lighter between compression section 302 and the injection section 303, can guarantee the stable connection of the two through the helicitic texture between compression section 302 and the injection section 303.
Further, as a preferred embodiment of the present invention, as shown in fig. 4, a sealing portion 304 for sealing the nozzle of the injection section 303 facing away from the compression section 302 is provided. The solid rocket itself requires stable environmental conditions to ensure its performance. The pipeline of the solid rocket motor 4 is used as a connecting port of the solid rocket and the outside, and better sealing performance is also required. That is, moisture or residual humid air is not generated inside the pipeline due to the change of the external environment, so that the sealing part 304 is arranged at the injection pipe orifice of the injection section 303 to seal the pipe orifice and ensure the dryness inside the pipe orifice. In the application, the sealing diaphragm of the sealing part 304 is usually made of 1060-O aluminum alloy, and the connection between the sealing diaphragm and the pipe wall is good, so that the sealing effect of the sealing part 304 is ensured.
Preferably, a weakening groove 305 is further provided in the sealing portion 304, and the weakening groove 305 is used for enabling the sealing portion 304 to fall off along the weakening groove 305 under the impact of propellant gas. When the solid rocket engine 4 performs combustion injection, the propellant gas needs to break the sealing portion 304, and the sealing portion 304 is made of metal, so that when the propellant gas breaks the sealing portion 304, the sealing portion 304 does not integrally separate from the pipe orifice of the third pipeline 3, but is broken by a part of the sealing portion, and the part of the sealing portion 304 remains at the pipe orifice of the third pipeline 3, thereby affecting the injection rate of the propellant gas, and even affecting the launching direction of the solid rocket. Therefore, the weakening groove 305 is formed in the sealing portion 304, the weakening groove 305 weakens the joint of the sealing portion 304 and the third pipeline 3, the weakening groove 305 becomes a weak part of the whole sealing portion 304, and when propellant gas impacts the sealing portion 304, the sealing portion 304 is entirely separated along the weakening groove 305, so that the injection efficiency of the propellant gas is ensured. Preferably, the thickness of the bottom of the weakening groove 305 is between 0.5mm and 0.6mm, preferably 0.55mm, and the weakening groove 305 of this thickness facilitates the injection of the propellant gas while ensuring the stability of the connection of the seal 304 to the third pipe 3. Preferably, the weakening groove 305 may be provided in the form of a ring, i.e. a ring of weakening ring area is formed on the sealing portion 304, facilitating the propellant gas to impact the sealing portion 304 around the weakening ring area to be entirely removed. Meanwhile, the weakening groove 305 may be provided as a plurality of weakening grooves 305 arranged along a ring shape, that is, a plurality of weakening points arranged in a ring shape, so as to facilitate the whole impact falling of the surrounding area of the weakening grooves 305.
Further, as a preferred embodiment of the present invention, heat insulation layers are further provided at the inner walls of the first pipeline 1, the second pipeline 2 and the third pipeline 3. The propellant gas temperature of a general small solid rocket engine 4 is above 2000 ℃, the working time is about 37.5S, the conventional titanium alloy pipeline can not completely meet the requirement, and the pipeline is easy to damage. Therefore, the heat protection design is needed to be carried out inside the pipeline, heat insulation materials are laid in the pipeline with the internal runner turning for many times, heat insulation layers are formed in the internal parts of the pipelines, and the stable use of the pipelines is guaranteed. Preferably, the heat insulation layer is formed by winding a carbon fiber phenolic layer and a high silica fiber phenolic layer, and has the characteristics of ablation resistance, scouring resistance, good heat insulation property and the like.
In particular, as a preferred embodiment of the invention, the insulation is provided in sections and comprises a first insulation 5, a second insulation 6 and a third insulation 7, provided in correspondence with the first 1, second 2 and third 3 pipes, respectively. In order to facilitate the molding of each pipeline and the heat insulation layer, each pipeline is arranged in sections, and the heat insulation layer is arranged aiming at each pipeline, so that the attachment between the heat insulation layer and the management is facilitated, and the heat insulation effect of the heat insulation layer is ensured.
Further, as a preferred embodiment of the present invention, a first fixing portion 8 for fixing the first pipe line 1 and the first heat insulating layer 5 is provided at a connection end of the first pipe line 1 and the rocket motor 4. Specifically, an annular step is arranged at a port where the first heat insulation layer 5 and the first pipeline 1 are attached, so that a pipe orifice of the first pipeline 1 protrudes out of the first heat insulation layer 5, an annular patch, namely a first fixing part 8, is arranged at the port of the heat insulation layer and abuts against the port of the first heat insulation layer 5, the annular patch is used for fixing the port of the first heat insulation layer 5, and the annular patch is fixedly connected with the pipe orifice of the first pipeline 1, so that the first pipeline 1 and the first heat insulation layer 5 are fixed relatively. Further, here, the annular patch can be partially embedded into the first pipeline 1 and the first heat insulation layer 5 through a threaded connection mode so as to realize fixation of the two.
Further, as a preferred embodiment of the present invention, in order to ensure the relative stability of the third pipeline 3 and the third thermal insulation layer 7, a second fixing portion 9 is disposed between the third pipeline 3 and the third thermal insulation layer 7 to fix the two. Specifically, due to the arrangement of the sealing portion 304, the sealing portion 304 seals the third heat insulating layer 7 and then connects the third heat insulating layer to the third pipeline 3, and therefore, the sealing portion 304 and the third heat insulating layer 7 are fixed, and then the sealing portion 304 and the third pipeline 3 are fixed. This application sets up the ring channel through the mouth of pipe department at third pipeline 3, sealing 304 sets up to the end cover structure to the end cover card is located in the ring channel, and set up sealing washer 902 in the two laminating department, in order to guarantee the two leakproofness, fixes sealing 304 and third pipeline 3 through pressing spiral shell 901 again, guarantees the relative stability of third pipeline 3 and third insulating layer 7 and the stable sealing of sealing 304 to the 3 mouths of pipe of third pipeline.
Preferably, a terminating device 10 is also provided on the second line 2, the terminating device 10 being used primarily for diverting the propellant gas, the terminating device 10 being used only as a structural element for the rapid stopping of the solid rocket and not being shown in fig. 2 in cross section. When a shutdown signal is received in the solid rocket launching process, the propellant gas is quickly shunted through the termination mechanism 10, so that the pressure in the combustion chamber of the rocket engine 4 is quickly reduced, and the propellant is prompted to extinguish and shutdown.
The utility model provides a compound pipeline of small-size solid rocket engine non-through formula in this application aims at setting up the pipeline into non-through formula to set up solid particle detention structure in pipeline corner, separate the solid particle in the propellant gas, avoid the solid particle to the continuous washing away of pipeline structure, cause the damage of pipeline, reduce the solid particle in the propellant gas simultaneously and can improve the smooth and easy nature of propellant gas circulation in the pipeline, improve the velocity of flow of propellant gas, improve the injection efficiency of solid rocket. Simultaneously, only through twice structure of turning round in this application, once be used for solid particle to be detained, once be used for propellant gas flow to correct, when not changing solid rocket transmission kinetic energy direction, reduce the kinetic energy loss of propellant gas when turning round as far as possible, overall structure is simple relatively, has improved the scour resistance of pipeline, improves the smooth degree of transport of propellant gas in the pipeline, has guaranteed the thrust stability of propellant gas.
It will be understood by those skilled in the art that the foregoing is only a preferred embodiment of the present invention, and is not intended to limit the invention, and that any modification, equivalent replacement, or improvement made within the spirit and principle of the present invention should be included in the scope of the present invention.
Claims (10)
1. A non-straight-through type composite pipeline of a small solid rocket engine is characterized by comprising a first pipeline, a second pipeline and a third pipeline;
one end of the first pipeline is connected with the rocket engine, and the other end of the first pipeline is closed;
the second pipeline is connected with the side wall of the first pipeline, and the first pipeline is divided into a conveying section and an accommodating section by a connecting point of the second pipeline and the first pipeline; wherein,
the conveying section is used for conveying propellant gas,
the accommodating section is used for accommodating solid particles in the propellant gas, and the propellant gas is turned at the accommodating section and is introduced into the second pipeline so as to separate the propellant gas from the solid particles;
the third pipeline is connected with the second pipeline and is used for expanding, accelerating and spraying out the propellant gas in the second pipeline.
2. A small solid rocket engine non-flow-through composite conduit according to claim 1 wherein the first conduit is axially parallel to or coincident with the third conduit.
3. The small solid rocket engine non-flow-through composite conduit according to claim 2 wherein the junction of said second conduit and said third conduit is provided with a turning portion for changing the direction of gas flow in said second conduit.
4. The small solid rocket engine non-flow-through composite conduit according to claim 1 wherein said third conduit comprises a compression section and an injection section, said compression section is connected to said second conduit, and the inner diameter of the end of said compression section connected to said second conduit is greater than the inner diameter of the end of said compression section connected to said injection section.
5. A small solid rocket engine non-flow-through composite conduit according to claim 4 wherein the nozzle of the injection section facing away from the compression section is provided with a sealing portion for sealing the same.
6. A small solid rocket engine non-straight composite pipeline according to claim 5 wherein the end face of said sealing portion is provided with a plurality of weakening grooves along the circumferential direction for separating the impact-breaking unit on the sealing portion so that the impact-breaking unit can be separated from the sealing portion under the impact of gas.
7. The small solid rocket engine non-straight composite pipeline according to any one of claims 1-6, wherein the inner walls of the first pipeline, the second pipeline and the third pipeline are all provided with heat insulation layers.
8. The small solid rocket engine non-flow-through composite conduit according to claim 7 wherein said insulation is provided in sections comprising a first insulation, a second insulation and a third insulation attached to said first conduit, said second conduit and said third conduit, respectively.
9. The small solid rocket engine non-flow-through composite conduit according to claim 8 wherein the connection end of said first conduit to the rocket engine is provided with a first fixing portion for fixing said first conduit to said first insulation layer.
10. The small solid rocket engine non-flow-through composite conduit according to claim 8 wherein the third conduit end for propellant gas injection is provided with a second securing portion for securing the third conduit to the third insulation layer.
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| CN113623088B (en) | 2023-01-13 |
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