CN113551564B - Radially-contractible ejection device conical buffering stop tail cover structure - Google Patents
Radially-contractible ejection device conical buffering stop tail cover structure Download PDFInfo
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- CN113551564B CN113551564B CN202110791692.6A CN202110791692A CN113551564B CN 113551564 B CN113551564 B CN 113551564B CN 202110791692 A CN202110791692 A CN 202110791692A CN 113551564 B CN113551564 B CN 113551564B
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- tail cover
- contraction
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F42—AMMUNITION; BLASTING
- F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
- F42B15/00—Self-propelled projectiles or missiles, e.g. rockets; Guided missiles
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F42—AMMUNITION; BLASTING
- F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
- F42B15/00—Self-propelled projectiles or missiles, e.g. rockets; Guided missiles
- F42B15/36—Means for interconnecting rocket-motor and body section; Multi-stage connectors; Disconnecting means
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- Chemical & Material Sciences (AREA)
- Aviation & Aerospace Engineering (AREA)
- Combustion & Propulsion (AREA)
- General Engineering & Computer Science (AREA)
- Aiming, Guidance, Guns With A Light Source, Armor, Camouflage, And Targets (AREA)
Abstract
The invention discloses a conical buffering and stopping tail cover of a radially-retractable ejection device, wherein the tail cover is positioned in an ejection barrel (91) and is arranged at the bottom end of a missile (92), a retraction hole (911) is formed in the top end of the ejection barrel (91), and the inner diameter of the upper end of the retraction hole (911) is smaller than that of the lower end of the retraction hole; the missile launching device is characterized in that a contraction body (1) is arranged on the circumferential direction of the tail cover, the contraction body (1) contracts radially after contacting with a contraction port (911), so that the tail cover can be buffered and stopped, after a missile is launched, the tail cover contacts with the contraction port (911), so that a speed difference is generated between the tail cover and the missile (92), and the separation of the tail cover and the missile (92) is further realized. The conical buffering stop tail cover of the radially-retractable ejection device disclosed by the invention can naturally separate the tail cover from a missile, and has the advantages of simple structure, low cost, high safety and the like.
Description
Technical Field
The invention relates to a conical buffering and stopping tail cover structure of a radially contractible ejection device, belonging to the technical field of missile launching.
Background
In the process that the cold-launched missile is pushed to pop out by the fuel gas generated by the barrel bottom power device, in order to avoid the damage of high-temperature and high-pressure fuel gas to the tail engine and other equipment of the missile during the launching, the tail cover of the missile is usually adopted to bear the huge airflow impact force generated by the fuel gas. After the missile leaves the barrel, the tail cover is separated from the missile body, and the missile engine is ignited to work.
At present, a plurality of technical modes are adopted, namely the missile is fixedly connected with a tail cover through an explosion bolt, the explosion bolt is unlocked after the missile is discharged from a barrel, the missile is separated from the tail cover under the action of the pretightening force of a spring, and the separation scheme is particularly divided into the modes of side-throwing separation, rotary separation and the like. In the side throwing mode, after the tail cover is separated from the projectile body, the tail cover engine is ignited to work, and the tail cover moves towards the side of the projectile body; in the rotation mode, the tail cover rotates for a certain angle along the rotating shaft on one side of the projectile body under the action of the pre-tightening force of the spring, and then the tail cover is separated from the projectile body.
In the separation mode, the tail cover movement after separation has uncertainty, so that the tail cover is difficult to ensure not to cause damage to ground equipment after falling off. And secondly, the tail cover is low in separation speed, and potential influence is caused on missile direction control in the separation process.
In addition, through search, the prior art also has a tail cover separating device (CN201711039142.9) based on collision braking buffering, however, the existence of a dropping part still exists in the stopping mode, the dropping part still brings potential safety hazards to missiles and other equipment and personnel, and the volume of the tail cover is obviously increased when the stopping mode is used for stopping, so that the size and the safety design of other equipment are influenced.
Therefore, there is a need for a missile tail cap designed to solve the above problems.
Disclosure of Invention
Specifically, the present invention aims to provide the following:
in one aspect, the invention provides a conical buffering and stopping tail cover of a radially-contractible ejection device, which is positioned inside a launching tube 91) and is arranged at the bottom end of a missile 92),
the top end of the launching tube 91 is provided with a contraction port 911, and the inner diameter of the upper end of the contraction port 911 is smaller than that of the lower end;
after the missile is launched, the tail cover is in contact with the contraction port 911, so that a speed difference is generated between the tail cover and the missile 92, and the tail cover is separated from the missile 92.
Further, the tail cover is circumferentially provided with a contraction body 1, and the contraction body 1 contracts radially after contacting with the contraction opening 911, so that the tail cover can be stopped in a buffering mode.
In a preferred embodiment, the tail cap comprises an upper panel 2, a middle body 3 and a lower panel 4, the middle body 3 connects the upper panel 2 and the lower panel 4, the shrinkable body 1 is connected with the middle body 3,
before the contraction body 1 is not contacted with the contraction opening 911, the contraction body 1 circumferentially protrudes out of the lower panel 4, and the outer diameter of the contraction body 1 is the same as the inner diameter of the launching tube wall, so that the tail cover and the launching tube 91 form a sealing structure, and further gas in the launching tube can push the tail cover to ascend along the launching tube axis.
In a preferred embodiment, the contraction body 1 comprises a plurality of arc-shaped sections 11, the arc-shaped sections 11 form a ring shape, and a sleeve section 12 is arranged on the surface of the arc-shaped section 11 facing the center of the circle;
one end of the arc-shaped section 11 is a convex end 111, the other end is a concave end 112, the convex end 111 and the concave end 112 can be inserted into each other, so that the arc-shaped sections 11 can be combined into a ring shape, the insertion positions of the convex end 111 and the concave end 112 can slide relatively, so that the contraction body 1 can contract radially,
in the contraction process, a closed structure is still formed between the lower surfaces of the convex end 111 and the concave end 112, the lower panel 4 and the launching tube wall, so that the fuel gas in the launching tube is not leaked.
In a preferred embodiment, the female end 112 includes a middle groove 1121 and a bottom convex ring 1122, the middle groove 1121 is a square groove, is located at the central position of the end of the female end 112,
the bottom convex ring 1122 is located at the lower end of the middle groove 1121, protrudes from the end of the arc-shaped section 11, and the protruding length of one side far away from the axle center of the tail cover is longer than the protruding length of one side close to the axle center of the tail cover;
the protruding end 111 includes a middle protruding groove 1111 and a bottom protruding ring 1112, the structure of the middle protruding groove 1111 corresponds to the middle groove 1121, the structure of the bottom protruding ring 1112 corresponds to the bottom protruding ring 1122, so that the protruding end 111 and the recessed end 112 can be inserted into each other.
In a preferred embodiment, a plurality of guide tubes 31 are arranged in the circumferential direction of the central body 3, and the guide tubes 31 are arranged on the sleeve section 12, or the sleeve section 12 is arranged on the guide tubes 31, so that the arc-shaped section 11 can only contract radially towards the tail cover.
In a preferred embodiment, a buffer body 5 is arranged inside the guide tube 31 and/or the sleeve section 12, and the buffer body 5 is a structure capable of absorbing kinetic energy through deformation.
In a preferred embodiment, the tail cover further comprises one or more air leakage cylinders 6, the air leakage cylinders 6 penetrate the upper panel 2 and the lower panel 4 up and down, a flow guide pipe 7 is arranged in the air leakage cylinders 6, the top end of the flow guide pipe 7 is provided with an upper flange cover 71, the top end of the side part of the flow guide pipe 7 is provided with an exhaust hole 72,
before the tail cover is not separated from the guided missile, the guided missile presses an upper flange cover 71 on an upper panel 2, an exhaust hole 72 is positioned in the air leakage cylinder 6, and fuel gas below the tail cover cannot be exhausted through the exhaust hole 72;
after the tail cover is separated from the missile, the upper flange cover 71 is influenced by the gas pressure below the tail cover, rises relative to the upper panel 2, and drives the exhaust holes 72 to rise above the upper panel 2, so that the gas below the tail cover is exhausted and released.
In a preferred embodiment, the bottom end of the draft tube 7 is provided with a lower flange ring 73, the outer diameter of the lower flange ring 73 is larger than the bore diameter of the gas leakage cylinder 6, and the length of the draft tube 7 is larger than the length of the gas leakage cylinder 6, so that the draft tube 7 cannot fly out of the tail cover after the tail cover is separated from the missile.
On the other hand, the invention also provides a conical buffering and stopping method for radial contraction of the tail cover, which is preferably realized by the conical buffering and stopping of the tail cover of the radially-contractible ejection device, and particularly, after the missile is launched, a contraction port arranged on the launch canister is contacted with a tail cover contraction body, so that the contraction body contracts radially, the kinetic energy of the tail cover along the axial direction of the launch canister is converted into the radial kinetic energy of the tail cover, and the velocity difference is generated between the tail cover and the missile through absorption of the buffering body, so that the natural separation of the tail cover and the missile is realized.
The conical buffering and stopping tail cover structure of the radially contractible ejection device provided by the invention has the beneficial effects that:
(1) when the tail cover passes through the contraction port of the launching barrel opening, the contraction body extrudes and deforms the buffer body under the action of the side wall of the contraction port, so that the tail cover is stopped, and the tail cover can be naturally separated from the guided missile;
(2) the buffer body structure adopts a metal multi-cell thin-wall structure, and has the characteristics of simple manufacturing process, light weight, high axial strength, good bearing efficiency and the like;
(3) the tail cover structure is retained in the launching barrel after the missile leaves the barrel, so that damage to surrounding equipment caused by falling of the tail cover is avoided;
(4) after the guided missile is separated from the tail cover, residual fuel gas in the launching tube can still be discharged through the guide pipe, and the safety is high.
Drawings
Fig. 1 is a schematic structural view of the whole structure of a cone-shaped buffering stop tail cover of a radially contractible ejection device according to a preferred embodiment of the invention;
FIG. 2 is a schematic view of a radially collapsible ejector cone buffer stop boot configuration launch canister in accordance with a preferred embodiment of the present invention;
fig. 3 is a schematic view showing the overall cross-sectional structure of a tail cap in the structure of a cone-shaped buffering and stopping tail cap of a radially contractible catapult according to a preferred embodiment of the invention;
figure 4 is a schematic view showing the structure of a contractible body portion of a cone-shaped buffering stop tail cap structure of a radially contractible catapult according to a preferred embodiment of the present invention;
figure 5 is a schematic view showing the construction of the contractible body portion of the conical buffering stop tail cap structure of the radially contractible catapult according to a preferred embodiment of the present invention;
figure 6 is a schematic view showing the structure of a contractible body portion of a cone-shaped buffering stop tail cap structure of a radially contractible catapult according to a preferred embodiment of the present invention;
fig. 7 is a schematic view showing a buffer structure in a conical buffer stop tail cap structure of a radially contractible catapult according to a preferred embodiment of the invention;
figure 8 is a schematic view showing the configuration of a blow-out cylinder in the conical buffer stop tail cap structure of a radially collapsible ejector according to a preferred embodiment of the present invention;
fig. 9 is a schematic view showing a flow guide tube structure in a conical buffering stopping tail cover structure of a radially contractible ejection device according to a preferred embodiment of the invention;
figure 10 is a schematic cross-sectional view of a tail cap in a tapered buffer stop tail cap configuration for a radially collapsible ejector according to a preferred embodiment of the present invention;
figure 11 is a schematic diagram showing the construction of the central body of a radially collapsible ejector cone bumper stop tail cap construction in accordance with a preferred embodiment of the present invention;
figure 12 is a schematic structural view of a radially collapsible ejector cone bumper stop tail cap arrangement according to a preferred embodiment of the present invention prior to collapse;
fig. 13 is a schematic structural view showing a contracted conical buffer stop tail cap structure of a radially contractible ejection device according to a preferred embodiment of the invention.
The reference numbers indicate:
1-a shrinkable body;
11-arc segment;
111-male ends;
112-a female end;
1111-middle convex groove;
1112-bottom concave ring;
1121-middle groove;
1122-bottom convex ring;
12-a sleeve;
2-upper panel;
3-an intermediate;
31-a guide tube;
4-a lower panel;
5-a buffer;
6-air release cylinder;
7-a flow guide pipe;
71-upper flange cover;
72-vent hole;
73-lower flange ring;
74-a cover plate;
91-a launch canister;
911-constriction;
92-missile.
Detailed Description
The invention is explained in more detail below with reference to the figures and examples. The features and advantages of the present invention will become more apparent from the description.
The word "exemplary" is used exclusively herein to mean "serving as an example, embodiment, or illustration. Any embodiment described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other embodiments. While the various aspects of the embodiments are presented in drawings, the drawings are not necessarily drawn to scale unless specifically indicated.
In one aspect, the present invention provides a radially retractable projectile launching device tapered buffer stop tail cap that is located inside the launch canister 91 and placed at the bottom end of the projectile 92, as shown in FIG. 1.
Further, the top end of the launcher 91 is provided with a constriction 911, and the inner diameter of the upper end of the constriction 911 is smaller than that of the lower end, as shown in fig. 2.
In a preferred embodiment, the inner diameter of the lower end of the constriction 911 is the same as the inner diameter of the launch canister 91.
More preferably, the inner diameter of the constriction 911 is tapered, and the inner diameter decreases uniformly from bottom to top, i.e. the inner side of the constriction 911 is provided with a slope.
In a preferred embodiment, the gradient of the change of the inner diameter is less than 1 so as to reduce the lateral force of the side wall of the contraction opening, which is applied to the wall surface along the central axis of the launching tube, and increase the lateral force of the tail cover, which is applied to the launching tube along the central axis of the launching tube.
More preferably, the included angle alpha between the inner side inclined plane of the contraction opening 911 and the axis of the launching tube is 25-35 degrees.
According to the invention, after the missile 92 is launched, the tail cover is contacted with the contraction port 911, so that a speed difference is generated between the tail cover and the missile 92, and the separation of the tail cover and the missile 92 is further realized.
In the invention, the missile 92 does not need to be fixed with the tail cover by adopting a fixing part, and the tail cover can only move along the axis of the launching tube when the missile is launched, so that the problems of tail cover deviation and the like can be avoided.
Further, according to the present invention, the tail cover is circumferentially provided with the contracting body 1, the contracting body 1 contracts radially after contacting with the contracting opening 911, and absorbs kinetic energy through deformation during contraction, so that the tail cover can buffer and stop, as shown in fig. 3.
In a preferred embodiment, the height H of the inner slope of the constricting orifice 911 is related to the constriction length L of the constriction 1 by: h is more than or equal to L/tan alpha, wherein alpha is an included angle between an inner side inclined plane of the contraction opening 911 and the axis of the launching tube.
In a preferred embodiment, the tail cap comprises an upper face plate 2, a middle body 3 and a lower face plate 4, the middle body 3 connects the upper face plate 2 and the lower face plate 4, and the shrinkable body 1 is connected with the middle body 3.
Preferably, the upper panel 2 and the lower panel 4 are circular flat plates, and the diameters of the upper panel 2 and the lower panel 4 are smaller than the inner diameter of the launch canister.
Further, the intermediate body 3 is preferably located at a central position of the upper panel 2 and the lower panel.
Further, the lower surface of the contraction body 1 is in contact with the lower panel 4, so that no gap exists between the contraction body 1 and the lower panel 4, and a large amount of gas is prevented from leaking from the contact position of the contraction body 1 and the lower panel 4.
According to the invention, when the missile is launched, the gas in the launching tube pushes the tail cover to rise along the launching tube, before the contraction body 1 is not contacted with the contraction opening 911, the contraction body 1 protrudes out of the lower panel 4 in the circumferential direction, the outer diameter of the contraction body 1 is the same as the inner diameter of the launching tube wall, so that the tail cover and the launching tube 91 form a sealing structure, and the gas in the launching tube can push the tail cover to rise along the launching tube axial direction.
In a preferred embodiment, the contraction body 1 comprises a plurality of arc-shaped sections 11, the arc-shaped sections 11 form a ring shape, and a sleeve 12 is arranged on the surface of the arc-shaped section 11 facing the center of the circle, as shown in fig. 4 to 6;
in a preferred embodiment, the number of arcuate segments 11 is not less than 6.
In a preferred embodiment, the edges of the arcuate segments 11 have a beveled slope that guides the arcuate segments 11 as they contact the constrictions 911, as shown in FIG. 1.
Furthermore, one end of the arc-shaped section 11 is a convex end 111, the other end is a concave end 112, the convex end 111 and the concave end 112 can be inserted into each other, so that the plurality of arc-shaped sections 11 can be combined into a ring shape, the insertion positions of the convex end 111 and the concave end 112 can slide relatively, so that the contraction body 1 can contract radially,
according to the invention, as the tail cover ascends, the contraction body 1 is contacted with the contraction opening 911, the contraction opening 911 circumferentially presses the contraction body 1, so that the convex end 111 and the concave end 112 slide relatively, and the contraction body 1 contracts radially,
further, in the relative sliding and shrinking process, a closed structure is still formed between the lower surfaces of the convex end 111 and the concave end 112, the lower panel 4 and the launching tube wall, so that the gas in the launching tube is not leaked, the missile is not completely separated from the tail cover or the distance between the missile and the tail cover is short after the missile and the tail cover are separated, and the missile is prevented from being damaged by gas leakage pressure.
In a preferred embodiment, the female end 112 includes a middle groove 1121 and a bottom convex ring 1122, the middle groove 1121 is a square groove, is located at the central position of the end of the female end 112,
the bottom convex ring 1122 is located at the lower end of the middle groove 1121, protrudes out of the end of the arc-shaped section 11, and the protruding length of one side far away from the axle center of the tail cover is longer than the protruding length of one side close to the axle center of the tail cover, so that in the process that the convex end 111 and the concave end 112 are inserted and slide, the lower surface of the contraction body, the lower panel and the wall of the launching cylinder still form a closed structure;
the protruding end 111 includes a middle protruding groove 1111 and a bottom protruding ring 1112, the structure of the middle protruding groove 1111 corresponds to the middle groove 1121, the structure of the bottom protruding ring 1112 corresponds to the bottom protruding ring 1122, so that the protruding end 111 and the recessed end 112 can be inserted into each other.
In a preferred embodiment, a plurality of guide tubes 31 are arranged in the circumferential direction of the middle body 3, as shown in fig. 11, the guide tubes 31 are sleeved on the sleeve 12, or the sleeve 12 is sleeved on the guide tubes 31, so that the arc-shaped section 11 can only contract radially towards the tail cover, and the shrink body is ensured not to rotate during the contraction process.
In the present invention, the guide tube 31 and the sleeve 12 may be circular tubes or square tubes, preferably square tubes, which have better guiding and anti-rotation effects.
Further, according to the present invention, a buffer body 5 is provided inside the guide tube 31 and/or the sleeve 12, and the buffer body 5 is a structure or a material capable of absorbing kinetic energy by deformation.
According to the invention, the convex end 111 and the concave end 112 are inserted with each other, so that radial contraction can be realized under the condition that the arc-shaped section in the contraction body does not generate structural deformation, the states before and after contraction are shown in figures 12-13, and the energy of the radial contraction is transmitted to the buffer body 5.
The arc-shaped section in the contraction body does not generate structural deformation, so that the tail cover is integrally stable in the process of separating from the missile, and the tail cover can be repeatedly used.
When the contraction body 1 contracts radially, the buffer body is extruded to deform and absorb kinetic energy, so that the buffer stop effect is achieved.
In a preferred embodiment, the buffer body 5 is a multi-cell thin-wall structure, which is a structure similar to a honeycomb, and has a plurality of small holes inside the structure, so that the structure has excellent energy absorption characteristics, and also has the characteristics of light weight and low cost, and further, compared with buffer bodies of other structures, such as springs and the like, the multi-cell thin-wall structure also has the characteristics of high energy absorption efficiency, quick response and small stroke, and can realize high-energy and high-efficiency energy absorption in a limited space of the tail cover.
Preferably, the buffer body 5 is rod-shaped and made of metal multi-cell thin-wall structure, and the multi-cell pores are penetrated from one end of the rod body to the other end of the rod body along the axis, as shown in fig. 7, so that the buffer body has higher axial strength and good bearing efficiency.
Further, both ends of the rod-shaped buffer 5 abut against the guide tube 31 and the sleeve 12, respectively.
According to the invention, the tail cover also comprises one or more air leakage cylinders 6, the air leakage cylinders 6 penetrate through the upper panel 2 and the lower panel 4 up and down, as shown in figure 8, a flow guide pipe 7 is arranged in the air leakage cylinders 6, as shown in figure 9, the top end of the flow guide pipe 7 is provided with an upper flange cover 71, the top end of the side part of the flow guide pipe 7 is provided with an exhaust hole 72,
before the tail cover is not separated from the guided missile, the guided missile presses an upper flange cover 71 on the upper panel 2, the exhaust hole 72 is positioned in the air leakage cylinder 6, and the fuel gas below the tail cover cannot be exhausted through the exhaust hole 72;
after the tail cover is separated from the missile, the upper flange cover 71 is influenced by the gas pressure below the tail cover, rises relative to the upper panel 2, and drives the exhaust holes 72 to rise above the upper panel 2, so that the gas below the tail cover is exhausted and released.
Further, during pressure relief, gas pressure is sprayed out towards the radial direction of the launching barrel through the exhaust holes 72 instead of being sprayed out towards the direction of the missile, and the influence of the gas pressure in the launching barrel on the missile is further avoided.
In a preferred embodiment, the bottom end of the draft tube 7 is provided with a lower flange ring 73, the outer diameter of the lower flange ring 73 is larger than the bore diameter of the gas leakage cylinder 6, and the length of the draft tube 7 is larger than the length of the gas leakage cylinder 6, so that the draft tube 7 cannot fly out of the tail cover after the tail cover is separated from the missile.
Further, the arrangement of the lower flange ring 73 also increases the acting force area of the fuel gas on the draft tube.
In a preferred embodiment, when the plurality of air release cylinders 6 are provided, the plurality of air release cylinders 6 are uniformly arranged on the upper panel 2, so that the missile is uniformly stressed.
More preferably, when there are a plurality of air release cylinders 6, the upper flange covers 71 at the upper ends of different flow guide pipes 7 in the plurality of air release cylinders 6 are the same, or a cover plate 74 is arranged at the upper end of the upper flange cover 71, as shown in fig. 10, so that the force bearing area between the missile and the tail cover is increased, and the missile is stressed more uniformly.
In another aspect, the invention further provides a conical buffering and stopping method for radial contraction of the tail cover, which is preferably realized by the conical buffering and stopping tail cover of the radially contractible ejection device.
Specifically, after the guided missile is launched, a contraction opening formed in the launching tube is in contact with a tail cover contraction body, so that the contraction body contracts radially, the kinetic energy of the tail cover along the axial direction of the launching tube is converted into the radial kinetic energy of the tail cover, the kinetic energy is absorbed by the buffer body, the tail cover and the guided missile generate a speed difference, and the natural separation of the tail cover and the guided missile is realized.
In a preferred embodiment, when the tail cover contraction body contracts, a closed structure is still formed among the lower surface of the contraction body, the lower panel and the wall of the launching barrel, so that gas in the launching barrel is not leaked.
In a preferred embodiment, the kinetic energy of the tail cover shrinkage is absorbed through the deformation of the buffer body, so that the buffer stop effect is achieved.
In a preferred embodiment, after the tail cover is separated from the missile, the pressure of the fuel gas in the launching barrel is relieved through the guide pipe.
Further, before the tail cover is separated from the missile, the missile presses the upper flange cover on the upper panel, so that gas cannot be discharged from the gas release cylinder.
After the tail cover is separated from the missile, the upper flange cover is influenced by the gas pressure below the tail cover, rises relative to the upper panel, and drives the exhaust holes to rise above the upper panel, so that the gas below the tail cover is exhausted and released.
In the description of the present invention, it should be noted that the terms "upper", "lower", "inner", "outer", "front", "rear", etc. indicate orientations or positional relationships based on operational states of the present invention, and are only for convenience of description and simplification of 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 present invention. Furthermore, the terms "first," "second," "third," and "fourth" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance.
In the description of the present invention, it should be noted that, unless otherwise specifically stated or limited, the terms "mounted," "connected," and "connected" are to be construed broadly, e.g., as meaning either a fixed connection, a removable connection, or an integral connection; can be mechanically or electrically connected; the connection may be direct or indirect via an intermediate medium, and may be a communication between the two elements. The specific meanings of the above terms in the present invention can be understood in specific cases to those skilled in the art.
The present invention has been described above in connection with preferred embodiments, but these embodiments are merely exemplary and merely illustrative. On the basis of the above, the invention can be subjected to various substitutions and modifications, and the substitutions and the modifications are all within the protection scope of the invention.
Claims (7)
1. A conical buffering stop tail cover of a radially contractible ejection device, which is positioned inside a launch barrel (91) and is arranged at the bottom end of a missile (92),
the top end of the launching tube (91) is provided with a contraction port (911), and the inner diameter of the upper end of the contraction port (911) is smaller than that of the lower end;
after the missile is launched, the tail cover is contacted with the contraction opening (911), so that the tail cover and the missile (92) generate a speed difference, and the tail cover and the missile (92) are separated;
the tail cover is circumferentially provided with a contraction body (1), and the contraction body (1) contracts radially after contacting with the contraction opening (911) so that the tail cover can be buffered and stopped;
the tail cover comprises an upper panel (2), a middle body (3) and a lower panel (4), the middle body (3) is connected with the upper panel (2) and the lower panel (4), the contraction body (1) is connected with the middle body (3),
before the contraction body (1) is not contacted with the contraction opening (911), the contraction body (1) circumferentially protrudes out of the lower panel (4), the outer diameter of the contraction body (1) is the same as the inner diameter of the wall of the emission tube, so that the tail cover and the emission tube (91) form a sealing structure, and further, gas in the emission tube can push the tail cover to ascend axially along the emission tube;
the contraction body (1) comprises a plurality of arc sections (11), the arc sections (11) form a ring, and a sleeve section (12) is arranged on the surface of the arc section (11) facing the circle center;
one end of each arc-shaped section (11) is a convex end (111), the other end of each arc-shaped section is a concave end (112), the convex ends (111) and the concave ends (112) can be mutually inserted, so that the arc-shaped sections (11) can be combined into a ring shape, the inserting positions of the convex ends (111) and the concave ends (112) can relatively slide, so that the contraction body (1) can radially contract,
in the contraction process, a closed structure is still formed among the lower surfaces of the convex end (111) and the concave end (112), the lower panel (4) and the emission barrel wall, so that the fuel gas in the emission barrel is not leaked.
2. The tapered buffering stop boot for a radially collapsible ejector as in claim 1,
the concave end (112) comprises a middle groove (1121) and a bottom convex ring (1122), the middle groove (1121) is a square groove and is positioned in the center of the end part of the concave end (112),
the bottom convex ring (1122) is positioned at the lower end of the middle groove (1121), protrudes out of the end part of the arc-shaped section (11), and the protruding length of one side far away from the axle center of the tail cover is longer than that of one side close to the axle center of the tail cover;
the convex end (111) comprises a middle convex groove (1111) and a bottom concave ring (1112), the structure of the middle convex groove (1111) corresponds to the middle groove (1121), and the structure of the bottom concave ring (1112) corresponds to the bottom convex ring (1122), so that the convex end (111) and the concave end (112) can be mutually inserted.
3. The tapered buffering stop boot for a radially collapsible ejector as in claim 1,
a plurality of guide pipes (31) are arranged in the circumferential direction of the intermediate body (3), the guide pipes (31) are sleeved on the sleeve section (12), or the sleeve section (12) is sleeved on the guide pipes (31), so that the arc-shaped section (11) can only contract towards the radial direction of the tail cover.
4. The tapered buffering stop boot for a radially collapsible ejector as in claim 3,
the inner sides of the guide pipe (31) and/or the sleeve section (12) are/is provided with a buffer body (5), and the buffer body (5) is of a structure capable of absorbing kinetic energy through deformation.
5. The tapered buffering stop boot for a radially collapsible ejector as in claim 1,
the tail cover also comprises one or more air leakage cylinders (6), the air leakage cylinders (6) vertically penetrate through the upper panel (2) and the lower panel (4), a guide pipe (7) is arranged in the air leakage cylinders (6), the top end of the guide pipe (7) is provided with an upper flange cover (71), the top end of the side part of the guide pipe (7) is provided with an exhaust hole (72),
before the tail cover is not separated from the guided missile, the guided missile presses an upper flange cover (71) on an upper panel (2), an exhaust hole (72) is positioned in an air release cylinder (6), and gas below the tail cover cannot be exhausted through the exhaust hole (72);
after the tail cover is separated from the missile, the upper flange cover (71) is influenced by gas pressure below the tail cover, rises relative to the upper panel (2), and drives the exhaust holes (72) to rise above the upper panel (2), so that the gas below the tail cover is exhausted and released.
6. The tapered buffering stop boot for a radially collapsible ejector as in claim 5,
the bottom end of the draft tube (7) is provided with a lower flange ring (73), the outer diameter of the lower flange ring (73) is larger than the aperture of the air leakage cylinder (6), and the length of the draft tube (7) is larger than that of the air leakage cylinder (6), so that the draft tube (7) cannot fly out of the tail cover after the tail cover is separated from the guided missile.
7. A conical buffering and stopping method for radial contraction of a tail cover, which is realized by the conical buffering and stopping tail cover of the radially contractible ejection device according to any one of claims 1 to 6,
after the guided missile is launched, a contraction opening formed in the launching barrel is in contact with a tail cover contraction body, so that the contraction body contracts radially, the kinetic energy of the tail cover along the axial direction of the launching barrel is converted into the radial kinetic energy of the tail cover, the kinetic energy is absorbed by the buffer body, the tail cover and the guided missile generate a speed difference, and the tail cover and the guided missile are separated naturally.
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| US4276901A (en) * | 1979-09-27 | 1981-07-07 | Essex Cryogenics Of Missouri Inc. | High flow pressure relief valve |
| DE202005001117U1 (en) * | 2005-01-25 | 2005-03-31 | Festo Ag & Co | Piston and thus equipped working cylinder |
| US7398721B1 (en) * | 2005-03-28 | 2008-07-15 | Lockheed Martin Corporation | Cold-gas munitions launch system |
| US8919256B1 (en) * | 2013-01-09 | 2014-12-30 | Raytheon Company | Obturator ring with interlocking segments |
| CN107816914B (en) * | 2017-10-31 | 2019-07-02 | 北京理工大学 | A kind of detention hood separator based on big momentum collision brake buffering |
| CN112361895B (en) * | 2020-10-22 | 2022-11-25 | 湖北航天技术研究院总体设计所 | Retention type missile tail cover |
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