CN111188697A - Solid rocket engine for electromagnetic ejection - Google Patents
Solid rocket engine for electromagnetic ejection Download PDFInfo
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- CN111188697A CN111188697A CN202010005353.6A CN202010005353A CN111188697A CN 111188697 A CN111188697 A CN 111188697A CN 202010005353 A CN202010005353 A CN 202010005353A CN 111188697 A CN111188697 A CN 111188697A
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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
- F02K9/10—Shape or structure of solid propellant charges
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- C—CHEMISTRY; METALLURGY
- C06—EXPLOSIVES; MATCHES
- C06B—EXPLOSIVES OR THERMIC COMPOSITIONS; MANUFACTURE THEREOF; USE OF SINGLE SUBSTANCES AS EXPLOSIVES
- C06B33/00—Compositions containing particulate metal, alloy, boron, silicon, selenium or tellurium with at least one oxygen supplying material which is either a metal oxide or a salt, organic or inorganic, capable of yielding a metal oxide
- C06B33/06—Compositions containing particulate metal, alloy, boron, silicon, selenium or tellurium with at least one oxygen supplying material which is either a metal oxide or a salt, organic or inorganic, capable of yielding a metal oxide the material being an inorganic oxygen-halogen salt
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- C—CHEMISTRY; METALLURGY
- C06—EXPLOSIVES; MATCHES
- C06D—MEANS FOR GENERATING SMOKE OR MIST; GAS-ATTACK COMPOSITIONS; GENERATION OF GAS FOR BLASTING OR PROPULSION (CHEMICAL PART)
- C06D5/00—Generation of pressure gas, e.g. for blasting cartridges, starting cartridges, rockets
- C06D5/06—Generation of pressure gas, e.g. for blasting cartridges, starting cartridges, rockets by reaction of two or more solids
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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
- F02K9/24—Charging rocket engines with solid propellants; Methods or apparatus specially adapted for working solid propellant charges
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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
- F02K9/32—Constructional parts; Details not otherwise provided for
- F02K9/34—Casings; Combustion chambers; Liners thereof
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- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
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- Inorganic Chemistry (AREA)
- Materials Engineering (AREA)
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Abstract
The invention discloses a solid rocket engine for electromagnetic ejection, which comprises a combustion chamber shell, an outer heat-proof layer, an ignition device, a propellant grain and a spray pipe, wherein the combustion chamber shell is internally provided with a combustion chamber; the propellant grain comprises a grain body, a front umbrella, an inner hole and a rear wing, wherein the inner hole penetrating through the grain body is formed in the grain body, the front umbrella is an annular groove formed by the inner wall of the grain body along the circumferential direction in a sunken mode, and the rear wing is formed by the inner wall of the grain body in a sunken mode.
Description
Technical Field
The invention relates to the technical field of solid rocket engines, in particular to a solid rocket engine for electromagnetic ejection.
Background
The electromagnetic ejection technology is a new linear propulsion technology, is suitable for short-stroke emission of large loads, and has wide application prospect in military, civil and industrial fields. The electromagnetic ejection is to use electromagnetic energy to push an ejected object to move outwards, and is a form of an electromagnetic gun. The ejection commonly used in the past has been mechanical ejection, such as springs, rubber bands, and the like. The energy is ejected as follows: bullets (using the energy of the instantaneous explosion of gunpowder), maglev trains, and the like.
The solid rocket engine has the advantages of simple use, good safety, good storage performance, convenient maintenance and the like, and is widely used in the field of electromagnetic ejection. Generally, the diameter of the barrel of the solid rocket is determined, and the diameter of the solid rocket engine can be determined according to the diameter of the barrel, so that the length of the solid rocket engine needs to be designed to be long enough to meet the requirement of high energy of the solid rocket engine, and the length-diameter ratio of the solid rocket engine is large. The length-diameter ratio of the propellant grain of the solid rocket engine with the large length-diameter ratio is also large, unstable combustion is easy to occur when the propellant grain is combusted, and once the unstable combustion occurs, the solid rocket engine can be caused to work abnormally and cause the failure of a flight task, so that great loss can be brought.
Disclosure of Invention
Aiming at the defects in the prior art, the invention aims to provide a solid rocket engine for electromagnetic ejection, wherein the structural design of a propellant grain can inhibit unstable combustion and ensure the stable operation of the solid rocket engine.
In order to achieve the above purposes, the technical scheme adopted by the invention is as follows:
a solid rocket engine for electromagnetic ejection comprises a combustion chamber shell, an ignition device and a propellant grain which are accommodated in the combustion chamber shell, and a spray pipe which is close to the propellant grain and is arranged at one end of the combustion chamber shell; the propellant grain comprises:
the powder column body is internally provided with an inner hole which penetrates through the powder column body;
the front umbrella is an annular groove formed by circumferentially sinking the inner wall of the grain body, and is close to the ignition device;
the rear wing is formed by sinking the inner wall of the grain body, and the rear wing is close to the spray pipe.
On the basis of the technical scheme, the rear wing comprises a plurality of wing grooves distributed along the circumferential direction of the inner hole at intervals.
On the basis of the technical scheme, the wing groove comprises a first wing groove and a second wing groove which are axially arranged and communicated with each other along the inner hole, the first wing groove is close to the front umbrella and forms a first step with the second wing groove, and the distance between the second wing groove and the axis of the inner hole is larger than that between the first wing groove and the axis of the inner hole.
On the basis of the technical scheme, the inner hole comprises a first inner hole and a second inner hole which are axially arranged along the inner hole and communicated with each other, the first inner hole is close to the front umbrella and forms a second step with the second inner hole, and the diameter of the second inner hole is larger than that of the first inner hole.
On the basis of the technical scheme, the powder column body is close to one end of the front umbrella and a combustion limiting layer is arranged between the front umbrella and coated on the inner wall of the powder column body.
On the basis of the technical scheme, the propellant of the propellant grain adopts a three-component butyl hydroxyl propellant.
On the basis of the technical scheme, the three-component hydroxyl butyl propellant comprises, by mass, 12.5 parts of an adhesive, 65 parts of an oxidizing agent, 17.5 parts of aluminum powder and 5 parts of a speed reducer.
In the above technical schemeOn the basis, the adhesive is hydroxyl-terminated polybutadiene, the oxidant is ammonium perchlorate, and the particle size of the aluminum powder is D50。
On the basis of the technical scheme, the speed reducing agent comprises 2.5 parts of carbonate and 2.5 parts of calcium carbonate in parts by mass.
On the basis of the technical scheme, an outer heat-proof layer is adhered to the outer surface of an ignition engine of the ignition device.
Compared with the prior art, the invention has the advantages that:
(1) the solid rocket engine for electromagnetic ejection is mainly characterized in that the front umbrella structure is designed at one end of the propellant grain close to the ignition device, the front umbrella can disturb the waveform of sound energy in the combustion chamber shell, the sound energy is dispersed, sound waves of the sound energy cannot be converged to form strong waves, and the vibration of the combustion chamber shell is prevented, so that the effect of inhibiting unstable combustion is achieved, and the stable work of the solid rocket engine is ensured.
(2) The wing groove comprises a first wing groove and a second wing groove which are axially arranged along the inner hole and are communicated, the first wing groove is close to the front umbrella and forms a first step with the second wing groove, the first step formed between the first wing groove and the second wing groove can enable a sound field in the combustion chamber shell to be reflected back and forth at the first step, and a part of sound energy of the combustion chamber shell can be consumed when the sound field passes through the first step, so that the collection of the sound energy is inhibited, the effect of inhibiting unstable combustion is achieved, and the thrust of an engine is more stable.
(3) The inner hole of the invention comprises a first inner hole and a second inner hole which are axially arranged along the inner hole and are communicated, the first inner hole is close to the front umbrella, and a second step is formed between the first inner hole and the second inner hole, so that a sound field in the combustion chamber shell can be reflected back and forth at the second step, and a part of sound energy of the combustion chamber shell can be consumed when passing through the second step, thereby inhibiting the collection of the sound energy and playing a role of inhibiting unstable combustion.
(4) A fire-limiting layer is arranged between one end of the grain body close to the front umbrella and coated on the inner wall of the grain body. The combustion limiting layer can enable the grain body in the area not to be combusted within 1-2 seconds after ignition of the ignition engine, the large length-diameter ratio is higher than the initial pressure peak of the engine, and the combustion limiting layer can play a role in reducing the initial pressure peak of the solid rocket engine and prevent excitation triggered by unstable combustion.
(5) The outer heat-proof layer is coated outside the ignition engine of the ignition device, so that the ignition device can be prevented from being burnt in a high-temperature and high-pressure environment in the working process of the ignition engine, and the ignition device can be used as a damping device, can be continuously disturbed at the front end of the combustion chamber shell and can absorb the sound energy of the gas reflected from the tail end of the combustion chamber shell, and plays a role in inhibiting the sound energy aggregation of the combustion chamber.
Drawings
FIG. 1 is a schematic structural diagram of a solid rocket engine for electromagnetic catapulting according to an embodiment of the present invention;
FIG. 2 is a half sectional view of FIG. 1;
FIG. 3 is a full sectional view of FIG. 1;
FIG. 4 is a schematic diagram of a propellant grain;
FIG. 5 is a full sectional view of FIG. 4;
FIG. 6 is a schematic view of the ignition device;
FIG. 7 is a full sectional view of FIG. 6;
FIG. 8 is a schematic structural view of the nozzle;
fig. 9 is a half sectional view of fig. 8.
In the figure: 1-combustion chamber shell, 10-heat insulating layer, 2-ignition device, 20-ignition engine, 21-outer heat-proof layer, 3-propellant grain, 30-grain body, 31-inner hole, 310-inner hole I, 311-inner hole II, 312-second step, 32-front umbrella, 33-rear wing, 330-wing groove, 3300-wing groove I, 3301-wing groove II, 3302-first step, 34-flame-limiting layer, 4-spray pipe, 40-shell, 41-outer heat-proof layer, 42-back lining, 43-throat lining, 44-convergence section outer heat-proof layer, 45-expansion section outer heat-proof layer and 46-blocking cover.
Detailed Description
The present invention will be described in further detail with reference to the accompanying drawings and examples.
Referring to fig. 1 to 3, the solid rocket engine for electromagnetic ejection according to the embodiment of the present invention includes a combustion chamber housing 1, an ignition device 2 and a propellant grain 3 housed in the combustion chamber housing 1, and a nozzle 4 disposed at one end of the combustion chamber housing 1 near the propellant grain 3; the inner surface of the combustion chamber housing 1 is provided with a heat insulating layer 10 which thermally protects the combustion chamber housing 1 during combustion of the propellant grains 3. The propellant grain 3 comprises a grain body 30, a front umbrella 32 and a rear wing 33, an inner hole 31 penetrating through the grain body 30 is arranged in the grain body 30, under the condition that the length-diameter ratio of an engine exceeds 10, in order to improve the stress intensity of the propellant grain 3 and reduce the initial erosion combustion effect, the inner hole 31 must not be too small, and the inner hole 31 forms the inner wall of the grain body 30. After the ignition device 2 ignites the charge body 30, the charge body 30 burns from the inner wall to the outer wall and from the front umbrella 32 along the rear wing 33; the front umbrella 32 is close to the ignition device 2, the front umbrella 32 is an annular groove formed by the inner wall of the grain body 30 along the circumferential direction in a concave mode, the annular groove is communicated with the inner hole 31, the front umbrella 32 can disturb the waveform of sound energy in the combustion chamber shell 1, the sound energy is dispersed, the sound waves of the sound energy cannot be converged to form strong waves, the combustion chamber shell 1 is prevented from vibrating, and therefore the effect of restraining unstable combustion is achieved. The rear wing 33 is close to the nozzle 4, the rear wing 33 is formed by the inner wall of the grain body 30 in a concave mode and is communicated with the inner hole 31, and the rear wing 33 is of a wing section structure similar to an airplane tail wing and can provide large thrust for an engine. The length-diameter ratio of the solid rocket engine is not less than 15, the vibration damping and the natural frequency of the combustion chamber shell 1 in a certain vibration frequency range are calculated through the optimized design of the thickness, the length and the mass of the combustion chamber shell 1, and the unstable combustion phenomenon generated by mutual coupling of the vibration of the combustion chamber shell 1 and the acoustics and the flow of the combustion chamber is avoided.
The solid rocket engine for electromagnetic ejection in the embodiment of the invention is mainly characterized in that the structure of the front umbrella 32 is designed at one end of the propellant grain 3 close to the ignition device 2, the front umbrella 32 can disturb the waveform of sound energy in the combustion chamber shell 1, disperse the sound energy, prevent the sound waves of the sound energy from converging to form strong waves and prevent the combustion chamber shell 1 from vibrating, thereby playing the role of inhibiting unstable combustion and ensuring the stable operation of the solid rocket engine.
Alternatively, as shown in FIGS. 4-5, the aft wing 33 includes a plurality of wing slots 330 spaced circumferentially about the inner bore 31. The wing slots 330 are designed with six, reducing the initial pressure peak and attenuating the triggering excitation by reducing the length, depth and number of the wing slots 330.
Preferably, as shown in fig. 5, the wing groove 330 includes a first wing groove 3300 and a second wing groove 3301 arranged along the axial direction of the inner hole 31 and communicating with each other, the first wing groove 3300 is close to the front umbrella 32 and forms a first step 3302 with the second wing groove 3301, and the distance between the second wing groove 3301 and the axis of the inner hole 31 is greater than the distance between the first wing groove 3300 and the axis of the inner hole 31. The distance between the first wing groove 3300 of the wing groove 330 and the axis of the inner hole 31 is 68mm, the distance between the second wing groove 3301 and the axis of the inner hole 31 is 80mm, and the draft angle is not less than 3%, the first step 3302 formed between the first wing groove 3300 and the second wing groove 3301 can make the sound field in the combustion chamber shell 1 reflect back and forth at the first step 3302, and can consume a part of sound energy of the combustion chamber shell 1 when passing through the first step 3302, thereby inhibiting the collection of sound energy, playing a role of inhibiting unstable combustion, and making the thrust of the engine more stable.
Further, referring to fig. 5, the inner hole 31 includes a first inner hole 310 and a second inner hole 311 which are arranged along the axial direction of the inner hole 31 and are communicated with each other, the first inner hole 310 is close to the front umbrella 32 and forms a second step 312 with the second inner hole 311, and the diameter of the second inner hole 311 is larger than that of the first inner hole 310. The second step 312 formed between the first inner hole 310 and the second inner hole 311 enables a sound field in the combustion chamber shell 1 to be reflected back and forth at the second step 312, and a part of sound energy of the combustion chamber shell 1 can be consumed when the sound field passes through the second step 312, so that the aggregation of the sound energy is inhibited, and the effect of inhibiting unstable combustion is achieved.
Further, referring to fig. 5, a fire-limiting layer 34 is disposed between one end of the grain body 30 close to the front umbrella 32 and the front umbrella 32, and the fire-limiting layer 34 is coated on the inner wall of the grain body 30. The powder column body 30 of this patent is close to and is equipped with limit burning layer 34 between the one end of preceding umbrella 32 and preceding umbrella 32, and limit burning layer 34 coats on the inner wall of powder column body 30. The combustion limiting layer 34 can prevent the grain body 30 in the area from being combusted within 1-2 seconds after the ignition engine 20 is ignited, the large length-diameter ratio is higher than the initial pressure peak of the engine, and the combustion limiting layer 34 can play a role in reducing the initial pressure peak of the solid rocket engine and prevent excitation triggered by unstable combustion.
Preferably, the propellant of the propellant grain 3 is a hydroxyl-terminated tri-component propellant. The three-component propellant comprises, by mass, 12.5 parts of an adhesive, 65 parts of an oxidizing agent, 17.5 parts of aluminum powder and 5 parts of a speed reducer. The adhesive is hydroxyl-terminated polybutadiene, the oxidant is ammonium perchlorate, and the particle diameter of the aluminum powder is D50. The speed reducer comprises, by mass, 2.5 parts of carbonate and 2.5 parts of calcium carbonate. The propellant formula of the propellant grain 3 of the solid rocket engine of the embodiment of the invention adopts a low-combustion formula, and the particle size of an oxidant, aluminum powder and a speed reducing agent is changed to increase the damping of propellant particles; the formulation adopts thinner aluminum powder, and solves the problem of oscillatory combustion by reducing the granularity of the aluminum powder.
The propellant provided by the embodiment of the invention adopts the improved butylated hydroxytoluene three-component propellant, and has the characteristics of low burning rate, small energy loss and high stability. On the basis of increasing the low-combustion and high-stability of the propellant, the energy is still not lower than 237s, the mechanical property is good, the combustion stability is good, and the specific performance comparison condition is shown in table 1.
TABLE 1 comparison of the Performance of the low-flammability butylated hydroxytoluene three-component propellants of the examples of the invention with conventional propellants
As can be seen from the above table, the propellant grain 3 of the engine in the embodiment of the invention adopts the design of the low-combustion hydroxyl-terminated tri-component propellant formula, on one hand, the specific impulse of the engine can be improved due to low combustion speed of the propellant and higher energy level of the propellant, and on the other hand, the propellant has better mechanical property, can adapt to the use conditions of wide temperature range from-40 ℃ to +50 ℃, and has large initial axial overload impact without damaging the integrity of the propellant grain 3.
Referring to fig. 6 to 7, the outer surface of the ignition engine 20 of the ignition device 2 is pasted with an outer heat-proof layer 21. The ignition engine 20 has a charge of 0.25kg and an ignition flow coefficient of 0.192kg/s.cm2And the ignition use requirement is met. The outer heat-proof layer 21 is coated outside the ignition engine 20 of the ignition device 2, so that the ignition device 2 can be prevented from being burnt in a high-temperature and high-pressure environment in the working process of the solid rocket engine, and the ignition device 2 can be used as a damping device, and can continuously disturb and absorb gas sound energy reflected from the tail end of the combustion chamber shell 1 at the front end of the combustion chamber shell 1, so that the effect of inhibiting the combustion chamber sound energy from being gathered is achieved.
Referring to fig. 8-9, the nozzle 4 comprises a convergent-section outer heat protection layer 44, a throat insert 43, a backing 42, an expansion-section outer heat protection layer 45, a blanking cover 46, a shell 40 and an outer heat protection layer 41, and is adapted to the working condition that the average working pressure of an engine is 7MPa (20 ℃), the nozzle throat diameter is 90mm, the initial expansion ratio is 8.5 in the embodiment, and the outlet inner diameter is 233 mm.
The design of the aerodynamic profile of the nozzle 4 in the embodiment of the invention is characterized in that due to the limitation on the size of the spatial structure of the nozzle 4, the constraint relation between the aerodynamic structure and the performance of the nozzle 4 is comprehensively considered in the design of the aerodynamic profile, the nozzle profile with higher efficiency is realized, the convergent section profile is designed into a conical profile, a convergent half angle β s can be selected between 30 degrees and 60 degrees, the convergent angle is not too large in consideration of factors influencing aerodynamic efficiency and ablation, otherwise, the convergent section profile is seriously ablated, the convergent half angle β s is 52 degrees, the nozzle throat diameter (minimum sectional area) is 90mm, the nozzle throat diameter is determined, the critical section is mainly designed by selecting an appropriate upstream curvature radius R1, the length of a column section and a downstream curvature radius R2 to reduce ablation, the flow characteristic of fuel gas is better, the nozzle has the advantages of R1 (0.2-1.5) dt, R2 (1.2-2.0) dt, the upstream curvature radius R29 is selected to be (1.2.0.0) dt, the pressure intensity/dt 2), the nozzle flow characteristic of the nozzle has obvious two-dimensional linear expansion of the nozzle profile, the nozzle profile has the shortest effective linear expansion of the three-dimensional linear gradient, and the aerodynamic profile has the shortest linear gradient of the three-dimensional linear gradient, and the linear gradient of the nozzle, and the linear gradient of the linear gradient.
The engine of the embodiment of the invention can work at a wide temperature range of-40 ℃ to +55 ℃, can bear 35g of axial large overload impact, has a specific impact of 241s, and is designed according to a specific design rule, wherein the specific design rule specifically comprises the following steps: the working pressure of the engine is one of the most important parameters of the engine, the higher the working pressure of the engine is, the higher the specific impulse of the engine is, but the negative quality is improved, the combustion stability of the engine is poor, the working pressure is selected to be 7 MPa-9 MPa in a comprehensive consideration, and the working pressure of the engine is designed to be 7 MPa.
The present invention is not limited to the above-described embodiments, and it will be apparent to those skilled in the art that various modifications and improvements can be made without departing from the principle of the present invention, and such modifications and improvements are also considered to be within the scope of the present invention. Those not described in detail in this specification are within the skill of the art.
Claims (10)
1. A solid rocket engine for electromagnetic ejection comprises a combustion chamber shell (1), an ignition device (2) and a propellant grain (3) which are accommodated in the combustion chamber shell (1), and a spray pipe (4) which is arranged at one end of the combustion chamber shell (1) and close to the propellant grain (3); characterized in that the propellant grains (3) comprise:
a grain body (30) which is internally provided with an inner hole (31) penetrating through the grain body (30);
the front umbrella (32) is an annular groove formed by circumferentially sinking the inner wall of the grain body (30), and the front umbrella (32) is close to the ignition device (2);
the rear wing (33), the rear wing (33) by the inner wall of powder column body (30) is sunken to form, rear wing (33) are close to spray tube (4).
2. The solid-rocket engine for electromagnetic catapulting according to claim 1, wherein said rear wing (33) comprises a plurality of wing grooves (330) spaced circumferentially along said inner bore (31).
3. The solid rocket engine for electromagnetic ejection according to claim 2, wherein said wing slot (330) comprises a first wing slot (3300) and a second wing slot (3301) which are arranged along the axial direction of said inner bore (31) and are communicated with each other, said first wing slot (3300) is close to said front umbrella (32) and forms a first step (3302) with said second wing slot (3301), and the distance between said second wing slot (3301) and the axis of said inner bore (31) is greater than the distance between said first wing slot (3300) and the axis of said inner bore (31).
4. The solid rocket engine for electromagnetic ejection according to claim 1, wherein said inner bore (31) comprises a first inner bore (310) and a second inner bore (311) which are arranged along the axial direction of said inner bore (31) and are communicated with each other, said first inner bore (310) is close to said front umbrella (32) and forms a second step (312) with said second inner bore (311), and the diameter of said second inner bore (311) is larger than the diameter of said first inner bore (310).
5. The solid rocket engine for electromagnetic ejection according to claim 1, wherein a fire-limiting layer (34) is provided between one end of the grain body (30) near the front umbrella (32) and the front umbrella (32), and the fire-limiting layer (34) is coated on the inner wall of the grain body (30).
6. The solid rocket engine for electromagnetic catapulting according to claim 1, wherein the propellant of said propellant grains (3) is a butylated hydroxytricopolymer.
7. The solid-rocket engine for electromagnetic catapults according to claim 6 wherein said butylated hydroxytristol propellant comprises, in parts by mass, 12.5 parts binder, 65 parts oxidizer, 17.5 parts aluminum powder and 5 parts decelerator.
8. The solid-rocket engine for electromagnetic catapult according to claim 7 wherein said binder is hydroxyl-terminated polybutadiene, said oxidizer is ammonium perchlorate, and said aluminum powder has a particle size of D50。
9. The solid-rocket engine for electromagnetic catapulting of claim 7 wherein said deceleration agent comprises, in parts by mass, 2.5 parts carbonate and 2.5 parts calcium carbonate.
10. The solid rocket engine for electromagnetic catapult according to claim 1, wherein an outer heat-proof layer (21) is adhered to the outer surface of the ignition engine (20) of said ignition device (2).
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Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112177797A (en) * | 2020-09-16 | 2021-01-05 | 西北工业大学 | Solid rocket engine without spray pipe |
| CN113217228A (en) * | 2021-06-18 | 2021-08-06 | 西北工业大学 | Magnetic control type thrust vector control device for solid rocket engine |
| CN115653791A (en) * | 2022-10-31 | 2023-01-31 | 北京中科宇航技术有限公司 | Solid rocket engine and cabin-penetrating type ignition device thereof |
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| 胡大宁等: "翼柱型药柱固体火箭发动机不稳定燃烧研究", 《固体火箭技术》 * |
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112177797A (en) * | 2020-09-16 | 2021-01-05 | 西北工业大学 | Solid rocket engine without spray pipe |
| CN113217228A (en) * | 2021-06-18 | 2021-08-06 | 西北工业大学 | Magnetic control type thrust vector control device for solid rocket engine |
| CN115653791A (en) * | 2022-10-31 | 2023-01-31 | 北京中科宇航技术有限公司 | Solid rocket engine and cabin-penetrating type ignition device thereof |
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