CN113218243A - Shunting device for weakening energy of tail flame emitted by missile - Google Patents
Shunting device for weakening energy of tail flame emitted by missile Download PDFInfo
- Publication number
- CN113218243A CN113218243A CN202110434530.7A CN202110434530A CN113218243A CN 113218243 A CN113218243 A CN 113218243A CN 202110434530 A CN202110434530 A CN 202110434530A CN 113218243 A CN113218243 A CN 113218243A
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- turbine shaft
- turbine
- missile
- tail flame
- energy
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F41—WEAPONS
- F41F—APPARATUS FOR LAUNCHING PROJECTILES OR MISSILES FROM BARRELS, e.g. CANNONS; LAUNCHERS FOR ROCKETS OR TORPEDOES; HARPOON GUNS
- F41F1/00—Launching apparatus for projecting projectiles or missiles from barrels, e.g. cannons; Harpoon guns
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F41—WEAPONS
- F41F—APPARATUS FOR LAUNCHING PROJECTILES OR MISSILES FROM BARRELS, e.g. CANNONS; LAUNCHERS FOR ROCKETS OR TORPEDOES; HARPOON GUNS
- F41F7/00—Launching-apparatus for projecting missiles or projectiles otherwise than from barrels, e.g. using spigots
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- Engineering & Computer Science (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention discloses a flow dividing device for weakening energy of missile-launched tail flames, which is characterized in that a flow guide cone is used for guiding missile-launched tail flame jet flow, the direction of the jet flow is changed for the first time, then the jet flow guided by the flow guide cone acts on turbine blades, the turbine blades are used for guiding the missile-launched tail flame jet flow again and reducing the speed, the vector direction of the tail flame jet flow is changed, the energy field of the tail flame jet flow is reduced, the quantity of stone impurities which can be rolled up by the tail flame jet flow and the initial kinetic energy of the impurities in the tail flame jet flow are reduced, and the damage to surrounding personnel and equipment is reduced or even avoided; the flow dividing device has the advantages of simple structure, reasonable design, convenience in use and the like, and can realize the weakening of energy and reduce the damage to surrounding personnel and equipment by changing the direction and reducing the speed of tail flame jet flow.
Description
Technical Field
The invention relates to the technical field of missile accessories, in particular to a flow dividing device for weakening the energy of a tail flame emitted by a missile.
Background
Missiles, as modern defense weapon systems, have been equipped more and more in various military classes. Various technical studies on missile technology are also on the rise. In the process from ignition to takeoff, the heat-emitting missile can jet out high-speed and high-temperature tail flame fuel gas, and the tail flame fuel gas can generate strong thermal shock on an emission platform.
When the missile is launched in the past, tail flame jet flow launched by the missile rapidly extends outwards to form a vortex, a large amount of impurities such as stones can be curled up, the impurities are scattered at a high speed in a launching field, surrounding personnel and equipment are easily damaged, and huge potential safety hazards exist.
Therefore, how to develop a device to reduce the damage caused by the missile launching tail flame is a problem to be solved urgently.
Disclosure of Invention
In view of the above, the invention provides a flow dividing device for weakening the energy of the tail flame emitted by the missile, so that the energy of the tail flame emitted by the missile is weakened, the amount of impurities such as stones rolled up by the tail flame emitted by the missile is reduced, and potential safety hazards are reduced.
The technical scheme provided by the invention is that the shunting device for weakening the energy of the tail flame emitted by the missile comprises: a fixed mount and a turbine;
the fixed frame is used for integral support;
the turbine includes: a turbine shaft, a guide cone, and a plurality of turbine blades;
the turbine blades are all positioned at the upper part of the turbine shaft and are respectively arranged along the circumferential direction of the turbine shaft at intervals;
the guide cone is arranged at the upper end of the turbine shaft;
the lower part of the turbine shaft is fixedly arranged in the fixed frame.
Preferably, the flow dividing device for attenuating the energy of the tail flame emitted by the missile further comprises: a rotary damper;
the rotary damper is sleeved on the lower part of the turbine shaft and then fixedly arranged in the fixed frame.
Further preferably, the flow dividing device for attenuating the energy of the tail flame emitted by the missile further comprises: a duct and a plurality of support rods;
mounting threaded blind holes are arranged on the peripheral side wall of the upper part of the rotary damper at intervals;
the duct is sleeved outside the turbine, and mounting holes which are in one-to-one correspondence with the blind mounting threaded holes in the side wall of the rotary damper are circumferentially arranged on the side wall of the lower part of the duct at intervals;
the supporting rods correspond to the mounting holes in the duct one to one, one end of each supporting rod penetrates through the corresponding mounting hole in the duct, threads are mounted in the threaded blind holes in the side wall of the rotary damper, the other end of each supporting rod is located outside the duct, and nuts are sleeved on the outer portions of the other ends of the supporting rods in threaded mode.
Further preferably, the duct is a cylinder with an upper opening and a lower opening.
Further preferably, the fixing frame is a U-shaped frame.
Further preferably, the turbine shaft, the guide cone and the plurality of turbine blades in the turbine are all of an integrated structure.
Preferably, the turbine shaft, the guide cone and the turbine blades in the turbine are of a split structure;
the turbine shaft comprises an upper turbine shaft and a lower turbine shaft which are integrally connected, a plurality of mounting grooves are arranged on the peripheral side wall of the upper turbine shaft at intervals, and the lower turbine shaft is fixedly mounted in the fixing frame or the rotary damper;
the turbine blades correspond to the mounting grooves one by one, and each turbine blade is mounted in the corresponding mounting groove;
the guide cone is positioned at the upper end of the upper turbine shaft in the turbine shaft and is fixedly connected with the upper turbine shaft through a bolt.
Further preferably, the diameter of the upper turbine shaft in the turbine shaft is larger than the diameter of the lower turbine shaft.
Further preferably, the turbine further comprises: a baffle plate;
the baffle plate is an annular plate, is sleeved outside the lower turbine shaft, abuts against the lower end face of the upper turbine shaft, and is fixedly connected with the upper turbine shaft through a bolt.
The flow dividing device for weakening the energy of the tail flame emitted by the guided missile provided by the invention has the advantages that the flow guide cone guides the jet flow of the tail flame emitted by the guided missile, the direction of the jet flow is changed for the first time, the jet flow guided by the flow guide cone acts on the turbine blades, the turbine blades guide the jet flow of the tail flame emitted by the guided missile again and reduce the speed, the vector direction of the jet flow of the tail flame is changed, the energy field of the jet flow of the tail flame is reduced, the quantity of stone impurities which can be rolled up by the jet flow of the tail flame is reduced, the initial kinetic energy of the impurities in the jet flow of the tail flame is weakened, and the damage to surrounding personnel and equipment is reduced or even avoided.
The shunting device for weakening the energy of the tail flame emitted by the missile has the advantages of simple structure, reasonable design, convenience in use and the like, and the energy is weakened and the damage to surrounding personnel and equipment is reduced by changing the direction and reducing the speed of the tail flame jet flow.
It is to be understood that both the foregoing general description and the following detailed description are exemplary and explanatory only and are not restrictive of the disclosure.
Drawings
The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments consistent with the invention and together with the description, serve to explain the principles of the invention.
In order to more clearly illustrate the embodiments or technical solutions in the prior art of the present invention, the drawings used in the description of the embodiments or prior art will be briefly described below, and it is obvious for those skilled in the art that other drawings can be obtained based on these drawings without creative efforts.
FIG. 1 is a schematic structural diagram of a first flow-dividing device for attenuating the energy of a missile-launching tail flame, provided in an embodiment of the present disclosure;
FIG. 2 is a schematic diagram of a second flow divider for attenuating the energy of the tail flame of a missile according to the disclosed embodiment of the invention;
FIG. 3 is a schematic diagram of the use of a third flow-splitting device for attenuating the energy of the tail flame of a missile according to the disclosed embodiment of the invention;
FIG. 4 is a schematic diagram of a fourth flow divider for attenuating the energy of the tail flame of a missile according to the disclosed embodiment of the invention;
FIG. 5 is a schematic diagram of a fifth flow divider for attenuating the energy of the tail flame of a missile according to the disclosed embodiment of the invention.
Detailed Description
Reference will now be made in detail to the exemplary embodiments, examples of which are illustrated in the accompanying drawings. When the following description refers to the accompanying drawings, like numbers in different drawings represent the same or similar elements unless otherwise indicated. The embodiments described in the following exemplary embodiments do not represent all embodiments consistent with the present invention. Rather, they are merely examples of apparatus consistent with certain aspects of the invention, as detailed in the appended claims.
When the missile is launched in the past, tail flame jet flow launched by the missile rapidly extends outwards to form a vortex, a large amount of impurities such as stones can be curled, the impurities are scattered at a high speed in a launching field, and surrounding personnel and equipment are easily damaged. In the past, researchers have tried to reduce the damage by using a blocking method, that is, a barrier, such as a steel plate, is erected on the ground, and the barrier blocks the scattered impurities, such as stones, but the effect is general. Therefore, in the present embodiment, the damage is reduced by reducing the energy of the tail flame jet, and the amount of the foreign matter such as stones to be rolled up can be greatly reduced by reducing the energy of the tail flame jet.
The present embodiment is based on the idea of weakening the energy of the tail flame jet, and designs a flow dividing device, see fig. 1, which is mainly composed of a fixed frame 1 and a turbine 2, wherein the fixed frame 1 is used for the integral support of the flow dividing device, and usually the fixed frame 1 is directly fixed with the ground; the turbine 2 mainly comprises a turbine shaft 21, a guide cone 23 and a plurality of turbine blades 22, wherein the plurality of turbine blades 22 are located on the upper portion of the turbine shaft 21 and are respectively arranged at intervals along the circumferential direction of the turbine shaft 21, the guide cone 23 is arranged at the upper end of the turbine shaft 21, the lower portion of the turbine shaft 21 is fixedly installed in the fixed frame 1, generally, the turbine shaft 21 and the fixed frame 1 are fixedly connected through bolts 6, specifically, threaded blind holes are arranged on the side wall of the turbine shaft 21, installation holes are arranged on the fixed frame 1, and after the bolts 6 pass through the installation holes on the fixed frame 1, the bolts are screwed in the threaded blind holes of the turbine shaft 21, so that the turbine shaft and the fixed frame are fixedly connected.
The shunting device is integrally supported by a fixed frame 1, when in use, the fixed frame 1 is fixedly arranged on the ground to limit the turbine 2, the flow dividing device mainly reduces the energy of tail flame jet flow through the turbine 2, and particularly, when the missile is launched, the high-speed tail flame jet flow ejected from the tail of the missile acts on the turbine 2, firstly touches the guide cone 23, the outer surface of the diversion cone 23 is a curved surface, which has the functions of reducing the speed and guiding the high-speed tail flame jet flow, and then, the exhaust flame jet slowed down and guided by the guide cone 23 will act on the turbine blades 22, be redirected and slowed down by the turbine blades, attenuate the energy of the exhaust flame jet, and thus, when the missile is launched, after tail flame jet flows through the flow dividing device, the flow speed is reduced, the flow direction is changed, the energy is reduced, and further the damage to surrounding personnel and equipment is weakened.
In order to further reduce the energy of the high-speed tail flame jet flow, as an improvement of the technical scheme, referring to fig. 2 and fig. 3, a rotary damper 3 is further arranged in the flow dividing device, after the rotary damper 3 is additionally arranged, the rotary damper 3 is sleeved on the lower part of a turbine shaft 21 and fixedly installed in a fixed frame 1, wherein the rotary damper 3 and the fixed frame 1 can also be fixedly connected through a bolt 6, specifically, threaded blind holes are arranged on the peripheral side wall of the rotary damper 3 at intervals, an installation hole is arranged on the fixed frame 1, and the bolt 6 is installed in the threaded blind hole on the rotary damper 3 after penetrating through the installation hole on the fixed frame 1, so that the fixed connection between the rotary damper 3 and the fixed frame 1 is realized.
Through the setting of above-mentioned rotary damper 3, this diverging device's whole working process does: after the missile is launched, the high-speed tail flame jet flow is formed, firstly, the flow guide cone 23 is used for changing the direction of the first air flow of the high-speed tail flame jet flow sprayed from the tail of the missile, the tail flame jet flow acts on the turbine blades 22 along the surface of the flow guide cone 23 and pushes the turbine blades 22 to rotate, the turbine blades 22 rotate, the flow field direction of the tail flame jet flow is changed for the second time, and the flow field energy can be dispersed through the arrangement of the plurality of turbine blades 22. The turbine blades 22 rotate to drive the turbine shaft 21 and the guide cone 23 to rotate synchronously, at the moment, the turbine shaft 21 and the rotary damper 3 move relatively, the rotary damper 3 consumes the rotation energy of the turbine shaft 21 through braking, the direction of partial tail flame jet flow is further changed, the flow speed is reduced, and the energy of the tail flame jet flow is weakened.
As an improvement of the technical solution, a duct 4 and a plurality of support rods 5 may be further added to the flow dividing device, referring to fig. 4 and 5, in order to fit the duct 4 and the support rods 5, mounting threaded blind holes are provided at intervals on the peripheral side wall of the upper portion of the rotary damper 3, the duct 4 is sleeved outside the turbine 2, mounting holes corresponding to the mounting threaded blind holes on the side wall of the rotary damper 3 are provided at intervals along the circumferential direction on the lower portion side wall of the duct 4, the support rods 5 correspond to the mounting holes on the duct 4 one by one, after the uniform end of each support rod 5 passes through the corresponding mounting hole on the duct 4, the support rod is screwed in the threaded blind hole on the side wall of the rotary damper 3, the other end of each support rod 5 is located outside the duct 4, and a nut 51 is screwed on the outer portion of the other end of each support rod 5. The support bar 5 is essentially mainly used for fixing the position of the duct 4, wherein the duct 4 is preferably a cylinder with an upper opening and a lower opening.
Set up the diverging device of duct and bracing piece, it is to the falling of the high-speed tail flame jet of guided missile can the process as follows specifically: when the missile is launched, the diversion cone 23 changes the direction of the first air flow of the high-speed tail flame jet flow ejected from the tail of the missile, the tail flame jet flow acts on the turbine blades 22 along the surface of the diversion cone 23 and pushes the turbine blades 22 to rotate, the turbine blades 22 rotate, the direction of the flow field of the tail flame jet flow is changed for the second time, the energy of the flow field can be dispersed by arranging the plurality of turbine blades 22, the dispersed flow field flows through the inside of the duct 4, the diversion is carried out again, and the energy of the flow field is further reduced. The turbine blades 22 rotate to drive the turbine shaft 21 and the guide cone 23 to rotate synchronously, at the moment, the turbine shaft 21 and the rotary damper 3 move relatively, the rotary damper 3 consumes the rotation energy of the turbine shaft 21 through braking, the direction of partial tail flame jet flow is changed, the flow speed is reduced, and the energy of the tail flame jet flow is weakened.
In the above embodiments, the structure of the fixing frame 1 may be selected from various structures as long as the supporting and fixing functions with the ground can be achieved, and the U-shaped frame is adopted as the fixing frame 1 in the above embodiments.
In the above embodiment, the turbine 2 may be of an integral or split structure, and the energy reduction effect on the tail flame jet is not affected.
With regard to the integrated structure, referring to fig. 1, 2 and 4, the turbine shaft 21, the guide cone 23 and the plurality of turbine blades 22 in the turbine 2 are all of an integrated structure, i.e., the entire turbine 2 is a single piece, preferably a centrifugal turbine structure.
For a split structure, referring to fig. 3 and 5, a turbine shaft 21, a guide cone 23 and a plurality of turbine blades 22 in a turbine 2 are split structures, wherein the turbine shaft 21 comprises an upper turbine shaft and a lower turbine shaft which are integrally connected, a plurality of mounting grooves are arranged on the outer peripheral side wall of the upper turbine shaft at intervals, the lower turbine shaft is fixedly mounted in a fixed frame 1 or a rotary damper 3, for a flow splitting device without the rotary damper 3, the lower turbine shaft is directly and fixedly connected with the fixed frame 1 through bolts 6, for the flow splitting device provided with the rotary damper 3, the lower turbine shaft is directly inserted into the rotary damper 3, then the rotary damper 3 is fixedly connected with the fixed frame 1 through the bolts 6, the turbine blades 22 correspond to the mounting grooves one to one, and each turbine blade 22 is mounted in the corresponding mounting groove, the guide cone 23 is located at the upper end of the upper turbine shaft in the turbine shaft 21 and is fixedly connected with the upper turbine shaft through a long bolt, specifically, a screw hole penetrating through the upper turbine shaft from top to bottom is formed in the upper turbine shaft, a threaded blind hole is formed in the lower end face of the guide cone 23, after the guide cone 23 is placed at the upper end of the upper turbine shaft, the upper end thread of the long bolt penetrates through the screw hole in the upper turbine shaft, and then the thread is screwed in the threaded blind hole of the guide cone 23, so that the upper turbine shaft is fixedly connected with the guide cone 23.
Preferably, the diameter of the upper turbine shaft is greater than the diameter of the lower turbine shaft in the turbine shaft 21, and in this case, when the turbine shaft 21 is mounted with the rotary damper 3, the upper turbine shaft is caught outside the upper end of the rotary damper 3, and the lower turbine shaft is entirely inserted inside the rotary damper 3.
In order to improve the attenuation effect of the device on the energy of the tail flame jet. As the improvement of the technical proposal, the plurality of mounting grooves on the upper turbine shaft are oblique grooves which are parallel to each other.
As a modification of the proposal, the turbine 2 is also provided with a baffle plate 24, the baffle plate 24 is an annular plate, is sleeved outside the lower turbine shaft, is abutted against the lower end surface of the upper turbine shaft, and is fixedly connected with the upper turbine shaft through a long bolt so as to avoid the abrasion of the upper turbine shaft.
In the turbine scheme provided with the blocking piece, the connection modes among the blocking piece 24, the turbine shaft 21 and the guide cone 23 can be selected from multiple modes, as long as the fixed connection among the three can be realized, a structural form with a relatively simple structure and convenience in assembly is provided below, specifically, refer to fig. 3 and 5, the blocking piece 24 is provided with countersunk holes at intervals along the circumferential direction, the upper turbine shaft is provided with through holes at intervals along the circumferential direction, the lower end face of the guide cone 23 is provided with threaded blind holes at intervals along the circumferential direction, the threaded blind holes are in one-to-one correspondence with the through holes, the long bolt 7 sequentially penetrates through the countersunk holes on the blocking piece 24 and the through holes on the upper turbine shaft, the long bolt is installed in the threaded blind holes on the guide cone 23 in a threaded manner, and the blocking piece 24, the upper turbine shaft and the guide cone 23 are fixedly connected.
Other embodiments of the invention will be apparent to those skilled in the art from consideration of the specification and practice of the invention disclosed herein. This application is intended to cover any variations, uses, or adaptations of the invention following, in general, the principles of the invention and including such departures from the present disclosure as come within known or customary practice within the art to which the invention pertains. It is intended that the specification and examples be considered as exemplary only, with a true scope and spirit of the invention being indicated by the following claims.
It is to be understood that the present invention is not limited to what has been described above, and that various modifications and changes may be made without departing from the scope thereof. The scope of the invention is limited only by the appended claims.
Claims (9)
1. A flow divider for attenuating missile launching tail flame energy, comprising: a fixed frame (1) and a turbine (2);
the fixed frame (1) is used for integral support;
the turbine (2) comprises: a turbine shaft (21), a guide cone (23) and a plurality of turbine blades (22);
a plurality of turbine blades (22) are positioned on the upper part of the turbine shaft (21) and are respectively arranged along the circumferential direction of the turbine shaft (21) at intervals;
the guide cone (23) is arranged at the upper end of the turbine shaft (21);
the lower part of the turbine shaft (21) is fixedly arranged in the fixed frame (1).
2. The shunt device for attenuating the energy of a missile launch tail flame as defined in claim 1, further comprising: a rotary damper (3);
the rotary damper (3) is sleeved on the lower part of the turbine shaft (21) and then fixedly installed in the fixed frame (1).
3. The shunt device for attenuating the energy of a missile launch tail flame as defined in claim 2, further comprising: a duct (4) and a plurality of support rods (5);
blind mounting thread holes are formed in the peripheral side wall of the upper part of the rotary damper (3) at intervals;
the duct (4) is sleeved outside the turbine (2), and mounting holes which are in one-to-one correspondence with the blind mounting thread holes in the side wall of the rotary damper (3) are formed in the side wall of the lower part of the duct (4) at intervals along the circumferential direction;
the utility model discloses a duct, including the duct, bracing piece (5) with mounting hole one-to-one on the duct (4), every bracing piece (5) uniform end passes behind the mounting hole that corresponds on duct (4), the screw thread is installed in the screw thread blind hole on rotary damper (3) lateral wall, every the other end of bracing piece (5) all is located the outside of duct (4), and every the equal screw thread cover in the other end outside of bracing piece (5) is equipped with nut (51).
4. The flow divider for attenuating the energy of the tail flame of a missile according to claim 1, characterized in that the duct (4) is a cylinder with an upper and a lower opening.
5. The flow divider for attenuating the energy of a missile-launching tail flame according to claim 1, characterized in that the fixed mount (1) is a U-shaped mount.
6. The flow divider device for attenuating missile-launching tail flame energy according to any one of claims 1-5, characterized in that the turbine shaft (21), the guide cone (23) and the plurality of turbine blades (22) in the turbine (2) are all of an integrated structure.
7. The flow divider device for attenuating missile-launching tail flame energy according to any one of claims 1-5, characterized in that a turbine shaft (21), a guide cone (23) and a plurality of turbine blades (22) in the turbine (2) are of a split structure;
the turbine shaft (21) comprises an upper turbine shaft and a lower turbine shaft which are integrally connected, a plurality of mounting grooves are arranged on the peripheral side wall of the upper turbine shaft at intervals, and the lower turbine shaft is fixedly mounted in the fixed frame (1) or the rotary damper (3);
the turbine blades (22) correspond to the mounting grooves one by one, and each turbine blade (22) is mounted in the corresponding mounting groove;
the guide cone (23) is positioned at the upper end of an upper turbine shaft in the turbine shaft (21) and is fixedly connected with the upper turbine shaft through a bolt.
8. The splitter device for attenuating the energy of the tail flames of a missile according to claim 7, wherein the diameter of the upper turbine shaft in the turbine shaft (21) is greater than the diameter of the lower turbine shaft.
9. The splitter device for attenuating the energy of a missile launch tail flame according to claim 8, wherein the turbine (2) further comprises: a baffle plate (24);
the baffle plate (24) is an annular plate, is sleeved outside the lower turbine shaft, abuts against the lower end face of the upper turbine shaft, and is fixedly connected with the upper turbine shaft through bolts.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110434530.7A CN113218243B (en) | 2021-04-22 | 2021-04-22 | Shunting device for weakening energy of tail flame emitted by missile |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110434530.7A CN113218243B (en) | 2021-04-22 | 2021-04-22 | Shunting device for weakening energy of tail flame emitted by missile |
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| CN113218243A true CN113218243A (en) | 2021-08-06 |
| CN113218243B CN113218243B (en) | 2022-10-25 |
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| Application Number | Title | Priority Date | Filing Date |
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| CN202110434530.7A Active CN113218243B (en) | 2021-04-22 | 2021-04-22 | Shunting device for weakening energy of tail flame emitted by missile |
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Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB2124741A (en) * | 1982-07-15 | 1984-02-22 | British Aerospace | Missile launcher |
| CN207229173U (en) * | 2017-08-03 | 2018-04-13 | 东莞理工学院 | A kind of device for realizing small sized turbine jet engine energy rapid conversion |
| CN108007272A (en) * | 2018-01-24 | 2018-05-08 | 安徽工业大学 | A kind of concentric launching tube of helix flow-guiding type |
| CN111238305A (en) * | 2020-03-16 | 2020-06-05 | 上海机电工程研究所 | Anti-impact tail cover suitable for missile launcher and missile launcher |
| CN112461045A (en) * | 2020-11-26 | 2021-03-09 | 北京星途探索科技有限公司 | Detachable simple launching platform |
-
2021
- 2021-04-22 CN CN202110434530.7A patent/CN113218243B/en active Active
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB2124741A (en) * | 1982-07-15 | 1984-02-22 | British Aerospace | Missile launcher |
| CN207229173U (en) * | 2017-08-03 | 2018-04-13 | 东莞理工学院 | A kind of device for realizing small sized turbine jet engine energy rapid conversion |
| CN108007272A (en) * | 2018-01-24 | 2018-05-08 | 安徽工业大学 | A kind of concentric launching tube of helix flow-guiding type |
| CN111238305A (en) * | 2020-03-16 | 2020-06-05 | 上海机电工程研究所 | Anti-impact tail cover suitable for missile launcher and missile launcher |
| CN112461045A (en) * | 2020-11-26 | 2021-03-09 | 北京星途探索科技有限公司 | Detachable simple launching platform |
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| CN113218243B (en) | 2022-10-25 |
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