CN218598267U - Charging structure of rocket engine, rocket engine and rocket - Google Patents
Charging structure of rocket engine, rocket engine and rocket Download PDFInfo
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- CN218598267U CN218598267U CN202222414503.6U CN202222414503U CN218598267U CN 218598267 U CN218598267 U CN 218598267U CN 202222414503 U CN202222414503 U CN 202222414503U CN 218598267 U CN218598267 U CN 218598267U
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Abstract
The utility model provides a loaded constitution, rocket engine and rocket of rocket engine, loaded constitution includes: coating the sleeve and the explosive column; a coating cavity with one side opened is formed in the coating sleeve; the explosive column is arranged in the coating cavity; wherein, a blind hole is formed on one side of the explosive column close to the opening of the coating cavity; the grain comprises: a first combustion section and a second combustion section; the first combustion section and the second combustion section are sequentially connected along the axial direction of the explosive column; wherein the first combustion section is the endurance stage of the rocket; the second combustion section is a booster stage of the rocket. The utility model discloses a set up the blind hole towards the uncovered one side of cladding cover at the grain, realize the single big thrust ratio characteristic requirement under the speed of burning of single room.
Description
Technical Field
The utility model relates to a rocket engine technical field especially relates to a charging structure, rocket engine and rocket of rocket engine.
Background
The rocket system can realize double thrust by adopting a mode of two-stage engine series connection and single-chamber double thrust. The two-stage engine series connection is characterized in that two thrust engines with different magnitudes are connected in series, the two-stage engine series connection has the advantages of being capable of meeting the requirement of any thrust ratio, and the two-stage engine series connection has the defects of more components of an engine system, low reliability and high cost.
The single-chamber double-thrust is to divide the combustion chamber into two working stages of a boosting stage and a cruising stage, wherein the boosting stage has short working time and large thrust and provides main thrust for rocket launching; the endurance stage has long working time and small thrust, and provides necessary thrust for the cruising flight of the rocket. Compared with the former design scheme, the mode is simple and has good realizability, and becomes the preferred scheme of the solid rocket engine with low cost, small size and large thrust ratio.
There are generally two ways for a single chamber dual thrust charge design:
firstly, a powder column end face combustion mode is adopted, two propellants with different burning rates are cast and molded in sequence, and the method has the advantages that the design scheme is simple, the manufacturability is poor, and interface component migration phenomena occur on interfaces of the propellants with different burning rates, so that the thrust transition section is abnormal;
secondly, the mode of designing different types of medicine is adopted, the thrust of different sizes is realized by controlling the size of the combustion surface, the advantages are that the single-combustion-rate propellant is used for pouring, the manufacturability is good, but the conventional inner hole, star-shaped and wing column-shaped charging structures at present can be realized with smaller propelling, the loading fraction is lower, and the overall requirement of the rocket is difficult to meet.
SUMMERY OF THE UTILITY MODEL
The utility model provides a loaded constitution, rocket engine and rocket of rocket engine for solve above-mentioned defect among the prior art, through setting up the blind hole towards the uncovered one side of cladding cover at the grain, realize the big thrust ratio characteristic requirement under the single combustion rate of single room.
According to the utility model discloses a charge structure of rocket engine that first aspect provided includes: coating the sleeve and the explosive column; a coating cavity with one open side is formed in the coating sleeve; the explosive column is arranged in the coating cavity; and a blind hole is formed on one side of the explosive column close to the opening of the coating cavity.
Optionally, the cartridge comprises: a first combustion section and a second combustion section; the first combustion section and the second combustion section are sequentially connected along the axial direction of the explosive column; wherein the first combustion section is the endurance stage of the rocket; the second combustion section is a booster stage of the rocket.
Optionally, at least a portion of the second combustion section is external to the enclosed cavity.
Optionally, the cartridge further comprises: a transition section connected to the first and second combustion sections, respectively; wherein the outer diameters of the first, transition and second combustion sections are progressively reduced.
Optionally, the blind hole extends from the end face of the second combustion section into the transition section and is partially within the cladding cavity.
Optionally, the second combustion section has an outer diameter equal to the sum of the blind bore diameter and a combustion parameter; wherein the combustion parameter is twice the product of the second combustion stage combustion speed and the second combustion stage operating time.
Optionally, the covering sleeve is a covering structure made of nitrile rubber.
Optionally, the grain is fuel made of hydroxyl-terminated butane.
According to a second aspect of the present invention, there is provided a rocket engine, comprising the charge structure of the rocket engine.
According to a third aspect of the present invention, there is provided a rocket including the charge structure of the rocket motor, or the rocket motor.
The utility model provides an above-mentioned one or more technical scheme has one of following technological effect at least: the utility model provides a pair of loaded constitution, rocket engine and rocket of rocket engine sets up the blind hole through the uncovered one side of grain orientation cladding cover, realizes the single room big thrust ratio characteristic requirement under the speed of burning alone.
Drawings
In order to clearly illustrate the technical solutions of the present invention or the prior art, the drawings used in the embodiments or the prior art descriptions will be briefly introduced below, and it is obvious that the drawings in the following description are some embodiments of the present invention, and other drawings can be obtained by those skilled in the art without creative efforts.
FIG. 1 is one of the schematic views of the assembled relationship of the charge structure of a rocket engine provided by the present invention;
FIG. 2 is a second schematic view showing the assembly of the charging structure of the rocket engine according to the present invention;
FIG. 3 is a schematic view of a sheath in a rocket motor charging configuration;
FIG. 4 is a schematic structural diagram of a charge column in a charging structure of a rocket engine provided by the present invention;
fig. 5 is a schematic view of a combustion surface-thickness curve generated when a propellant is combusted in a charge structure of a rocket engine provided by the present invention.
Reference numerals:
10. coating a sleeve; 20. carrying out grain treatment; 30. blind holes; 40. a first combustion section; 50. a second combustion stage; 60. a transition section.
Detailed Description
To make the purpose, technical solution and advantages of the embodiments of the present invention clearer, the technical solution in the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings in the embodiments of the present invention, and obviously, the described embodiments are some, but not all, embodiments of the present invention. Based on the embodiments in the present invention, all other embodiments obtained by a person skilled in the art without creative efforts belong to the protection scope of the present invention.
In the description of the embodiments of the present invention, it should be noted that the terms "center", "longitudinal", "lateral", "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", and the like indicate orientations or positional relationships based on the orientations or positional relationships shown in the drawings, and are only for convenience of describing the embodiments of the present invention and simplifying the description, but do not indicate or imply that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and thus should not be construed as limiting the embodiments of the present invention. Furthermore, the terms "first," "second," and "third" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance.
In some embodiments of the present invention, as shown in fig. 1 to 5, the present invention provides a charge structure of a rocket engine, comprising: the coating sleeve 10 and the explosive column 20; a coating cavity with one side opened is formed inside the coating sleeve 10; the grain 20 is arranged in the coating cavity; wherein, a blind hole 30 is formed at one side of the grain 20 close to the opening of the coating cavity.
It should be noted that, by providing the blind hole 30 at the end of the grain 20, the characteristic requirement of large thrust ratio of the small solid rocket engine is met based on the mode of adjusting the combustion surface of the propellant, i.e. the combustion surface of the grain 20, and the characteristic requirement of large thrust ratio at single-chamber single-combustion speed is met.
In a possible embodiment, the covering sleeve 10 and the charge 20 are both cylindrical in structure.
In some possible embodiments of the present invention, the cartridge 20 comprises: a first combustion section 40 and a second combustion section 50; the first combustion section 40 and the second combustion section 50 are sequentially connected along the axial direction of the grain 20; wherein, the first combustion section 40 is the endurance of the rocket; the second combustion stage 50 is the booster stage of the rocket.
Specifically, the present embodiment provides an embodiment of the first and second combustion sections 40 and 50 that satisfies the dual thrust requirements for rocket boost and endurance stages at a single chamber and a single firing rate by providing the first and second combustion sections 40 and 50 and providing the blind bore 30 in the second combustion section 50.
The single-chamber double-thrust is to divide the combustion chamber into two working stages, namely a boosting stage and a cruising stage, wherein the boosting stage has short working time and large thrust and provides main thrust for rocket launching; the endurance stage has long working time and small thrust, provides necessary thrust for the cruising flight of the rocket, and realizes the adjustment of the size of the combustion surface of the propellant by arranging the first combustion section 40 and the second combustion section 50 and arranging the blind hole 30 in the second combustion section 50.
In some possible embodiments of the present invention, at least a portion of the second combustion section 50 is external to the enclosed cavity.
Specifically, the present embodiment provides an embodiment of the second combustion section 50 and the coating cavity, the combustion chamber charge is composed of the coating sleeve 10 and the grain 20, and the second combustion section 50 is arranged in the air, so that the free filling mode is satisfied for assembling with the combustion chamber heat insulation shell.
In a possible embodiment, there is a bond between the first combustion section 40 and the sheathing 10.
In some possible embodiments of the present invention, the grain 20 further comprises: a transition section 60, the transition section 60 being connected to the first combustion section 40 and the second combustion section 50, respectively; wherein the outer diameters of the first combustion section 40, the transition section 60, and the second combustion section 50 are gradually reduced.
Specifically, the present embodiment provides an implementation of the transition section 60, which makes the connection between the drug column 20 and the covering sleeve 10 more compact by arranging the transition section 60, and also changes the shape of the covering sleeve 10 to be closed and installed.
It should be noted that the axial lengths of the first combustion section 40, the transition section 60, and the second combustion section 50 depend on the thrust-boosting requirements for the combustion face.
In a possible embodiment, when the propellant is ignited, the combustion surface consists of the surface of the blind hole 30, the outer circular surface of the second combustion section 50 and the rear end surface of the second combustion section 50, the combustion of the boosting-stage combustion surface in an equal surface mode can be realized by adjusting the axial lengths of the blind hole 30 and the outer circular surface of the second combustion section 50, and after the characteristics of the cylindrical section of the blind hole 30 disappear, the whole combustion surface is in an end surface combustion characteristic, so that the combustion of the continuation-stage combustion surface in an equal surface mode is realized.
In some possible embodiments of the present invention, the blind hole 30 extends from the end surface of the second combustion section 50 to the transition section 60 and is partially located in the cladding cavity.
Specifically, the present embodiment provides an implementation of the blind hole 30 that facilitates dual power conversion between a boost stage and a cruise stage by disposing the blind hole 30 within the transition section 60.
In a possible embodiment, the diameter of the blind hole 30 ends between 10 mm and 30 mm.
In some possible embodiments of the present invention, the outer diameter of the second combustion section 50 is equal to the sum of the diameter of the blind hole 30 and the combustion parameter; wherein the combustion parameter is twice the product of the second combustion stage 50 firing rate and the second combustion stage 50 operating time.
Specifically, the present embodiment provides an embodiment of the outer diameter of the second combustion section 50, which satisfies the requirements of different boost and endurance classes by providing a corresponding relationship between the diameter of the second combustion section 50 and the blind hole 30 and the combustion parameters.
In some possible embodiments of the present invention, the covering sheath 10 is a covering structure made of nitrile rubber.
Specifically, the present embodiment provides an embodiment of the encasement 10.
In some possible embodiments of the present invention, the grains 20 are fuel made of butylated hydroxyanisole.
Specifically, the present example provides an embodiment of a charge 20.
In one application scenario, the outer diameter of the covering sleeve 10 is 170 mm, the diameter of the blind hole 30 is 20 mm, the length of the cylindrical section of the blind hole 30 is 90 mm, the bottom of the blind hole is of a hemispherical structure with the diameter of 2 mm, the outer cylindrical surface of the suspended second combustion section 50 is an outer circular surface, the outer diameter of the suspended second combustion section is 150 mm, the length of the outer circular surface of the suspended second combustion section 50 is 80 mm, and the root of the suspended second combustion section is an arc with the radius of 5 mm.
In one application scenario, as shown in fig. 5, a combustion surface-thickness curve generated during propellant combustion is shown in fig. 5, and a single-chamber double-thrust solid rocket engine with a thrust ratio of 4.8 is predicted to be achieved, wherein the combustion surface ratio of a boosting stage and a cruising stage is about 2.8.
Further, the combustion surface of the boosting stage can be further increased by increasing the lengths of the outer circular surfaces of the blind holes 30 and the second combustion section 50, and meanwhile, the combustion surface of the endurance stage is kept unchanged, so that the combustion surface ratio of two stages can be increased, and a charge structure with a larger thrust ratio is realized.
In some embodiments of the present invention, the present invention provides a rocket engine, including the charge structure of the rocket engine.
It is noted that the design of a single-chamber single-burning-rate charging structure is adopted, and the requirement of the small solid rocket engine on the characteristic of high thrust ratio is met. The rocket system has the characteristics of large initial combustion surface, high filling ratio and stable performance, improves the comprehensive performance of the rocket system, and can ensure the stable and reliable performance of the rocket system.
Furthermore, the improved charging structure has a simple medicine shape structure and good process implementation, reduces the production cost of the engine, is beneficial to mass production of engine charging, and improves the market competitiveness of the small solid rocket engine.
In some embodiments of the present invention, the present invention provides a rocket including the charging structure of the rocket engine described above, or the rocket engine described above.
In the description of the embodiments of the present invention, it should be noted that, unless explicitly stated or limited otherwise, the terms "connected" and "connected" should be interpreted broadly, and may be, for example, fixedly connected, detachably connected, or integrally connected; can be mechanically or electrically connected; may be directly connected or indirectly connected through an intermediate. The specific meaning of the above terms in the embodiments of the present invention can be understood by those of ordinary skill in the art according to specific situations.
In the description herein, references to the description of the term "one embodiment," "some embodiments," "a manner," "a particular manner," or "some manner" or the like are intended to mean that a particular feature, structure, material, or characteristic described in connection with the embodiment or manner is included in at least one embodiment or manner of an embodiment of the invention. In this specification, the schematic representations of the terms used above are not necessarily intended to refer to the same embodiment or mode. Furthermore, the particular features, structures, materials, or characteristics described may be combined in any suitable manner in any one or more embodiments or modes. Furthermore, various embodiments or modes described in this specification, as well as features of various embodiments or modes, may be combined and combined by those skilled in the art without contradiction.
Finally, it should be noted that: the above embodiments are merely illustrative, and not restrictive, of the present invention. Although the present invention has been described in detail with reference to the embodiments, it should be understood by those skilled in the art that various combinations, modifications or equivalent substitutions may be made to the technical solution of the present invention without departing from the spirit and scope of the technical solution of the present invention, and all of the technical solutions should be covered by the scope of the claims of the present invention.
Claims (10)
1. A charge configuration for a rocket engine, comprising: a coating sleeve (10) and a medicine column (20);
a coating cavity with one side opened is formed in the coating sleeve (10);
the explosive column (20) is arranged in the coating cavity;
wherein, a blind hole (30) is formed on one side of the grain (20) close to the opening of the coating cavity.
2. A rocket engine charge construction according to claim 1, characterized in that said charge (20) comprises: a first combustion section (40) and a second combustion section (50);
the first combustion section (40) and the second combustion section (50) are sequentially connected along the axial direction of the explosive column (20);
wherein the first combustion section (40) is a endurance stage of the rocket;
the second combustion section (50) is a booster stage of the rocket.
3. A rocket engine charge construction according to claim 2, wherein at least part of said second combustion section (50) is outside said cladding cavity.
4. A rocket engine charge configuration according to claim 2, wherein said charge (20) further comprises: a transition section (60), the transition section (60) being connected with the first combustion section (40) and the second combustion section (50), respectively;
wherein the outer diameters of the first combustion section (40), the transition section (60) and the second combustion section (50) are gradually reduced.
5. A rocket engine charge structure according to claim 4, wherein said blind hole (30) extends from the end face of said second combustion section (50) into said transition section (60) and is partly inside said cladding cavity.
6. -a rocket engine charge structure according to claim 2, wherein the outer diameter of said second combustion section (50) is equal to the sum of the diameter of said blind hole (30) and a combustion parameter;
wherein the combustion parameter is twice the product of the second combustion stage (50) firing rate and the second combustion stage (50) operating time.
7. Charge construction of a rocket engine according to any one of claims 1 to 6, wherein said sheath (10) is a covering made of nitrile rubber.
8. A rocket engine charge configuration according to any one of claims 1 to 6, wherein said charge (20) is fuel made of butylated hydroxytoluene.
9. A rocket engine comprising a charge configuration for a rocket engine as recited in any one of claims 1 to 8.
10. A rocket comprising a charge for a rocket motor according to any one of claims 1 to 8 or a rocket motor according to claim 9.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202222414503.6U CN218598267U (en) | 2022-09-09 | 2022-09-09 | Charging structure of rocket engine, rocket engine and rocket |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202222414503.6U CN218598267U (en) | 2022-09-09 | 2022-09-09 | Charging structure of rocket engine, rocket engine and rocket |
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| CN218598267U true CN218598267U (en) | 2023-03-10 |
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| CN202222414503.6U Active CN218598267U (en) | 2022-09-09 | 2022-09-09 | Charging structure of rocket engine, rocket engine and rocket |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN116220944A (en) * | 2023-05-09 | 2023-06-06 | 北京星河动力装备科技有限公司 | Solid engine and rocket |
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Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN116220944A (en) * | 2023-05-09 | 2023-06-06 | 北京星河动力装备科技有限公司 | Solid engine and rocket |
| CN116220944B (en) * | 2023-05-09 | 2023-09-05 | 北京星河动力装备科技有限公司 | Solid engine and rocket |
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