WO2014153712A1 - Dispositif dynamique de propulsion de fusée - Google Patents

Dispositif dynamique de propulsion de fusée Download PDF

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Publication number
WO2014153712A1
WO2014153712A1 PCT/CN2013/073154 CN2013073154W WO2014153712A1 WO 2014153712 A1 WO2014153712 A1 WO 2014153712A1 CN 2013073154 W CN2013073154 W CN 2013073154W WO 2014153712 A1 WO2014153712 A1 WO 2014153712A1
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WO
WIPO (PCT)
Prior art keywords
rocket
solid fuel
memory
dynamic
casing
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Application number
PCT/CN2013/073154
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English (en)
Chinese (zh)
Inventor
刘志伟
Original Assignee
Liu Zhiwei
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Filing date
Publication date
Application filed by Liu Zhiwei filed Critical Liu Zhiwei
Priority to PCT/CN2013/073154 priority Critical patent/WO2014153712A1/fr
Publication of WO2014153712A1 publication Critical patent/WO2014153712A1/fr

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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
    • 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
    • F02K9/22—Shape or structure of solid propellant charges of the front-burning type

Definitions

  • the invention belongs to the field of rockets, and in particular relates to a dynamic rocket propeller.
  • the rocket is a jet propulsion device that is jetted back at a high speed by a hot air flow and moves forward using the generated reaction force. It carries the combustion agent and oxidant itself, does not rely on the oxygen in the air to support combustion, and can fly in the atmosphere or in outer space. Modern rockets can be used as fast long-distance vehicles for launching satellites and delivering weapons warheads (such as warheads).
  • Traditional rockets are generally composed of rocket casings, fuels, engines, and the like.
  • the role of the rocket casing is to protect the internal components of the rocket and to support the rocket.
  • the energy generated by the rocket fuel burning powers the rocket flight.
  • the engine controls the combustion of the fuel and the injection of the flame.
  • the existing method is to make the rocket into a multi-stage rocket. Each stage of the rocket is equipped with an engine, and the empty shell of the first-class rocket is discarded every time the fuel of the first-class rocket is used.
  • the rocket should be manufactured in an infinite number of stages so that excess weight can be discarded as quickly as possible.
  • the more rocket stages the more complex the rocket is, and each stage must have its own engine, which in turn increases the weight, so the rocket stage can not be too much.
  • the current rocket series usually does not exceed three levels. Therefore, the existing rocket cannot do so in a very timely manner to discard excess weight.
  • the shells discarded by the rocket may hit people or objects on the ground, so the discarding location cannot be chosen at will, so it will also limit the flight path of the rocket to some extent.
  • Existing rocket engines use the engine to control fuel combustion and injection to power the rocket.
  • the main components of the engine are the combustion chamber and the nozzle.
  • the fuel is combusted in the combustion chamber and then ejected through the nozzle to generate power.
  • Both the combustion chamber and the nozzle are made of high temperature resistant, high strength materials. Sometimes these two parts are also equipped with cooling devices to protect the mechanism from burning at high temperatures. It has the following shortcomings: First, the engine needs to be made of high temperature resistant and high strength materials, which results in higher engine manufacturing costs. Second, the engine is an integral part of the rocket and needs to be used on each stage of the rocket.
  • Equipped with an engine since the weight of the engine itself is usually heavier, the setting of the engine is also objectively exacerbated by the weight of the entire rocket; third, due to the heat resistance of the engine material, the flame temperature in the engine cannot be too high. This is not conducive to improving the propulsion power of the entire rocket; Fourth, the structure of the engine is complex, especially liquid fuel rockets, the engine pipeline is many and easy to leak liquid, which is extremely dangerous; fifth, the engine nozzle opening usually cannot follow the external air pressure. Adjust the size of the opening to make the engine lose efficiency.
  • the technical problem to be solved by the present invention is to provide a dynamic rocket thruster, thereby reducing the following problems:
  • the existing rockets produce discards that are too large in size and cannot discard the excess weight of the rocket in time;
  • the dynamic rocket thruster provided by the present invention is implemented as follows:
  • a dynamic rocket propeller comprising an arrow body filled with solid fuel and having a control element on the arrow body that dynamically generates a Laval nozzle structure suitable for rocket flight when the solid fuel is burned with the solid fuel
  • the control element includes an outer casing encasing the solid fuel and having a lower edge rim that can be consumed in conjunction with combustion dynamic combustion of the solid fuel when the solid fuel is combusted.
  • the arrow body is provided with a plurality of electrodes and a discharge control system for controlling the discharge state of each of the electrodes and/or at least one cooling passage for injecting the coolant.
  • the outer casing is a memory casing made of a shape memory material, the memory casing having a cross-sectional area that is smaller than a cross-sectional area of the arrow body.
  • the memory housing surface is provided with lateral scratches and/or the interior of the memory housing material is provided with a transverse fibrous material.
  • control element further includes a tubular member disposed within the body and made of a memory material, the tubular member memorizing a cross-sectional area greater than a cross-sectional area of the tubular member.
  • tubular member surface is provided with lateral scratches and/or its material is internally provided with a transverse fibrous material.
  • control element further includes a nozzle generating member disposed under the solid fuel inside the arrow body; and the first portion of the nozzle generating member penetrates from the top to the bottom with at least one shape such as a Laval nozzle
  • the passage and/or the outer wall of the nozzle generating member together with the inner wall of the outer casing form a second passage shaped like a Laval nozzle.
  • the dynamic rocket propeller is internally provided with a traction climbing mechanism that connects the nozzle generating members.
  • a cooling device is disposed in the nozzle generating member, and a surface of the nozzle generating member is provided with a cooling material outlet that communicates with the cooling device.
  • the invention has the beneficial effects that the dynamic rocket propeller provided by the invention adopts solid fuel and is provided with a control element, so that the bottom of the arrow body generates a Laval nozzle structure suitable for rocket flight.
  • the control element includes an arrow housing that can be dynamically ablated with the combustion surface of the solid fuel as the solid fuel burns. In this way, during the flight of the rocket, the bottom of the arrow body can dynamically generate a Laval nozzle structure suitable for rocket flight, and the generated discards are substantially used as the working medium for propelling the rocket to fly by the Laval spray.
  • the excess weight of the rocket can be discarded in time, and no large discards are generated, thereby improving the freedom of the rocket flight path design.
  • the pusher since the pusher has a simple structure, the manufacturing cost can be reduced. Since the propeller acts as a part of the nozzle during operation, the propeller is easy to generate a larger Laval nozzle structure, and the larger Laval nozzle structure is advantageous for improving the energy utilization efficiency of the fuel.
  • FIG. 1 is a schematic structural view of a first embodiment of a dynamic rocket propeller provided by the present invention
  • FIG. 2 is a schematic structural view of a second embodiment of a dynamic rocket propeller provided by the present invention.
  • FIG 3 is a schematic structural view of a third embodiment of a dynamic rocket propeller provided by the present invention.
  • a dynamic rocket propeller includes an arrow body filled with a solid fuel 1 a and the arrow body is provided with a dynamic force for burning the bottom of the arrow body to facilitate rocket flight when the solid fuel 1 a is burned.
  • the dynamic rocket thruster provided by the present invention adopts an arrow body internally filled with the solid fuel 1a, and a control element 2a is disposed on the solid fuel 1a, so that the bottom of the arrow body generates a nozzle structure 3a suitable for rocket flight.
  • the surface of the solid fuel 1a is provided with a casing 21a which encloses the solid fuel 1a and which can be ablated along with the arrow body when the solid fuel 1a is burned.
  • the bottom of the solid fuel 1a can dynamically generate a nozzle structure 3a suitable for rocket flight, and the generated discard substantially serves as a working medium for propelling the rocket to fly from the nozzle structure.
  • 3a is ejected, the excess weight of the rocket can be discarded in time, and no large discards are generated, thereby improving the freedom of design of the rocket flight path.
  • the solid fuel 1a is provided with a plurality of electrodes 5a and a discharge control system for controlling the discharge state of each of the electrodes 5a and/or at least one cooling passage 6a for injecting the coolant 7a.
  • the burning state of the solid fuel 1a is controlled by injecting the discharge state of the coolant 7a and the control electrode 5a through the cooling passage 6a, thereby achieving the purpose of adjusting the shape of the combustion surface 4a.
  • the combustion surface 4a can be cooled or even extinguished by increasing the injection amount of the coolant 7a.
  • the electrode 5a may radiate an arc toward the lower end of the combustion face 4a, thereby accelerating the combustion of the solid fuel on the combustion face 4a, or accelerating the burning rate at certain specific positions, or igniting the solid fuel.
  • the gas atmosphere required for the discharge of the electrode 5a can be provided by vaporization of the coolant 7a injected into the cooling passage 6a.
  • the cooling passage 6a and the electrode 5a employed in the present invention are both dynamically ablated with the advancement of the combustion surface 4a on the solid fuel 1a, and the cooling passage 6a, the electrode 5a, and the injection passage 6a are injected.
  • the gas generated after the gasification of the coolant 7a can be ejected from the nozzle structure 3a along with the gas generated by the combustion of the solid fuel 1a, that is, the cooling passage 6a, the electrode 5a, and the cooling passage.
  • the gas produced by the gasification of the injected coolant 7a of 6a should be considered as part of the working medium that propels the rocket.
  • the electrode 5a and the cooling passage 6a are both disposed in the axial direction of the solid fuel 1a and in parallel to the axial direction of the solid fuel 1a.
  • the mechanical strength of the solid fuel 1a can also be enhanced by appropriately enhancing the mechanical strength of each of the electrodes 5a.
  • the length of the electrode can be known, and the remaining amount of the solid fuel 1a can be converted.
  • the resistance detecting device for detecting the electric resistance of the electrode 5a is used to know the temperature of the flame combustion, thereby converting the speed at which the solid fuel 1a burns and whether the rocket thruster has been turned off.
  • the area size of the combustion surface 4a is controlled by controlling the distance of the liquid surface of the coolant 7a on the cooling passage 6a from the combustion surface 4a. Specifically, when the coolant 7a is far from the combustion surface 4a, the combustion flame on the combustion surface 4a will ignite the portion of the cooling passage 6a where there is no coolant 7a, in which case The portion of the cooling passage 6a where the coolant 7a is not present also becomes a part of the combustion surface 4a, so that the total area of the combustion surface 4a is relatively large. Therefore, the area of the combustion surface 4a can be adjusted by adjusting the distance of the coolant 7a from the combustion surface 4a.
  • the outer casing 21a is a memory casing made of a shape memory material, the memory casing having a memory cross-sectional area smaller than a cross-sectional area of the solid fuel 1a.
  • the memory casing may be specifically manufactured by first making the memory casing remember a shape having a smaller cross section than the cross section of the solid fuel 1a, and expanding the cross section of the memory casing to be wrapable. The structure of the solid fuel 1a, thereby completing the process of providing a memory casing on the solid fuel 1a.
  • the shape memory material used in the above described memory casing may be a memory alloy or a polymer heat shrinkable material.
  • the outer casing of the solid fuel 1a is made of a shape memory material, wherein the material of the arrow body provides mechanical support for the arrow body, and the material also dynamically generates a Laval nozzle structure, which is also a consumable material, and is also a kind of propulsion. Working quality.
  • the memory alloy may be the following formulas Au-Cd, Ag-Cd, Cu-Zn, Cu-Zn-Al, Cu-Zn-Sn, Cu-Zn-Si, Cu-Sn, Cu-Zn-Ga, One of In-Ti, Au-Cu-Zn, NiAl, Fe-Pt, Ti-Ni, Ti-Ni-Pd, Ti-Nb, U-Nb, and Fe-Mn-Si.
  • the polymer heat-shrinkable material is also called a polymer shape memory material, and is an intelligent material which is a combination of a polymer material and a radiation processing technology. Ordinary polymer materials such as polyethylene, polyvinyl chloride, etc. are usually linear structures.
  • these materials After being irradiated by a radiation source such as an electron accelerator into a network structure, these materials have a unique "memory effect", expansion, cooling and shaping.
  • the material can re-shrink to restore its original shape when heated.
  • the memory properties of the heat-shrinkable material can be used to make heat-shrinkable pipes, membranes and profiled materials.
  • the main characteristic is that the heat shrinkage is coated on the outer surface of the object, which can function as insulation, moisture, seal, protection and connection, and the diameter of the shrinkage material.
  • the shrinkage rate can reach 50% ⁇ 80%.
  • the Laval nozzle is composed of three parts, which are respectively a shrink tube 31a, a expansion tube 33a, and a throat 32a connecting the shrink tube 31a and the expansion tube 33a.
  • the process of forming the various sections of the Laval nozzle structure by the memory casing is described below. Due to the difference in heating time of the respective portions of the memory casing, the higher the temperature of the portion of the memory casing which is farther from the burning surface 4a, the lower the temperature of the portion of the memory casing which is closer to the burning surface 4a.
  • the portion of the memory casing that is closer to the burning surface is contracted by heat to form the shrinking tube 31a; the portion of the memory casing that is farther from the burning surface and that is connected to the shrinking tube 31a is heated more time than the shrinking tube 31a.
  • the portion of the memory casing that is further away from the combustion surface and connected to the throat 32a has been softened due to the longer heating time, and is subjected to the combustion medium on the combustion surface 4a.
  • the memory casing of this portion will expand outward to form the expansion tube 33a. Therefore, in the present embodiment, during the combustion of the solid fuel 1a, the memory casing follows the advancing process of the combustion surface 4a of the solid fuel 1a, and the lower portion thereof will exhibit the Laval commonly used in the rocket nozzle structure. Nozzle structure.
  • the shrinkage tube 31a will gradually become the throat 32a, which will gradually become the softening expansion tube 33a, and the expansion tube 33a will increase in temperature as the temperature thereof increases. It will begin to melt or vaporize away from the memory casing.
  • the memory casing thereon will dynamically form a Laval nozzle structure, and the memory casing made of the memory material may follow the solid fuel 1a.
  • the memory casing surface is provided. Lateral scratches and/or transverse fibrous material are provided within the memory shell material. In this way, the memory casing can be ablated more evenly.
  • the center of the solid fuel 1a is provided with a rod member or a tubular member 22a. Since the cross-sectional area of the throat 32a is small, and the rod or tubular member 22a occupies a large cross-sectional area of the throat 32a, the ventilation area passing through the throat 32a can be reduced, so that the throat is passed through the throat. The air flow of the tube 32a is faster. In this way, the shrinking ability required to fabricate the memory material of the memory casing can be effectively reduced, thereby reducing the difficulty in manufacturing the memory casing.
  • dynamically generating and burning the Laval nozzle structure has the following advantages:
  • the memory case made of memory material has lower requirements on the heat resistance of the memory material.
  • the dynamic rocket propeller can dynamically generate a combustion chamber during flight.
  • the memory casing When the temperature in the combustion chamber is high and the pressure is high, the memory casing will be melted or vaporized by heat in a short time, at which time the Laval nozzle structure is generally short and the throat 32a is relatively thick and short.
  • the Laval nozzle structure when the temperature in the combustion chamber is low and the pressure is low, the Laval nozzle structure will be long overall and the throat 32a will be elongated.
  • the shape of the Laval nozzle structure also has an influence on the temperature in the combustion chamber 4a. The longer Laval nozzle structure and the elongated throat structure have a large gas discharge resistance, thereby contributing to an increase in the combustion chamber temperature.
  • the short Laval nozzle structure and the short and short throat structure have low gas discharge resistance, which is beneficial to lower the combustion chamber temperature, and the low combustion chamber temperature is beneficial to protect the Laval nozzle structure, thereby reducing the memory shell forming the Laval nozzle structure. Heat resistance requirements.
  • the dynamically generated Laval nozzle structure has an opening size of the throat 32a, which is favorable for improving the kinetic energy conversion rate.
  • the entire Laval nozzle structure is dynamically generated.
  • the outside air pressure is an important factor affecting the size of the opening of the nozzle structure 3a.
  • the opening of the nozzle structure 3a will be made smaller due to the large outside air pressure, which will cause the nozzle structure 3a to eject a gas having a higher pressure.
  • the outside air pressure is small, the nozzle structure 3a is subjected to a small outside air pressure, and its opening is large, so that the nozzle structure 3a will eject a gas having a lower pressure.
  • the opening size of the nozzle structure 3a on the dynamically generated Laval nozzle structure in the embodiment can automatically adjust the output state of the kinetic energy according to the external air pressure, which is beneficial to improving the conversion rate of the kinetic energy.
  • the gasification of the memory shell generates kinetic energy and improves the kinetic energy conversion rate of the combustion medium.
  • the structure of the entire rocket propeller is simple, low in cost, safe and reliable.
  • each part of the dynamic rocket thruster of the present embodiment undertakes a plurality of functions as much as possible, the structure thereof becomes simple, and the simple structure is advantageous in reducing the manufacturing cost, and the manufacturing error between the respective parts affects the probability of the entire performance. Decline, which increases its security.
  • Rocket vibration mainly comes from the unstable combustion of fuel in the combustion chamber. Excessive vibration easily affects the safety of the rocket and the safety of people or objects that affect the transportation.
  • the Laval nozzle structure formed is in a soft state or a highly elastic rubber state, which can effectively buffer the vibration.
  • some solid fuels are themselves rubbery, highly elastic materials, which themselves have the ability to cushion vibration.
  • the existing rocket cannot produce a nozzle with a large size, so that the conversion rate of the kinetic energy of the rocket is not high.
  • the rocket nozzle can be generated by the outer casing 21a, and a larger volume nozzle can be generated, which is beneficial for converting more exhaust heat energy into kinetic energy.
  • the embodiment is different from the first embodiment in that the control elements are arranged in different manners.
  • the control element 2b includes a casing 21b encasing the solid fuel and when the solid fuel 1b is burned, the lower end edge thereof may be consumed along with the combustion dynamic combustion of the solid fuel, and further includes A tubular member 22b in the body of the arrow and made of a memory material, the tubular member 22b memorizing a cross-sectional area greater than the cross-sectional area of the tubular member 22b.
  • the tubular member 22b will gradually return to the original memory section as the combustion surface 4b is moved, and the lower portion thereof will form an expanded state of a larger sectional area.
  • the tubular member 22b will form a Laval nozzle structure with the outer casing 21b.
  • the outer casing 21b may also be made of a memory material to facilitate the formation of the Laval nozzle structure.
  • the tubular member 22b is provided with a lateral scratch on the surface and/or a material of a transverse fiber material is disposed inside the material. In this way, the tubular member 22b can be ablated more uniformly.
  • the embodiment is different from the first embodiment or the second embodiment in that the control elements are arranged in different manners.
  • the control element 2c includes a casing 21c that encloses the solid fuel and that consumes the lower end edge of the solid fuel 1c as it burns, and further includes a nozzle generating member 22c below the solid fuel 1c inside the arrow; a first passage 221c and/or the nozzle of the nozzle forming member 22c extending from top to bottom through at least one shape such as a Laval nozzle
  • the outer wall of the generating member 22c and the inner wall of the outer casing 21c together form a second passage 222c shaped like a Laval nozzle.
  • the gas generated during the combustion of the solid fuel 1c can be ejected through the first passage 221c and/or the second passage 222c shaped as a Laval nozzle.
  • the dynamic rocket thruster is internally provided with a traction climbing mechanism 23c that connects the nozzle generating member 22c.
  • a traction climbing mechanism 23c that connects the nozzle generating member 22c.
  • a cooling device is disposed in the nozzle generating member 22c, and a surface of the nozzle generating member 22c is provided with a coolant outlet connected to the cooling device.
  • the velocity of the gas will rise sharply and exceed the speed of sound because the area through which the gas passes is small.
  • the gas exceeding the speed of sound continues to accelerate toward the portion of the outer end of the outer casing 21c which is expanded to generate thrust.
  • the combustion surface 4c moves upward, and the traction climbing mechanism 23c pulls the nozzle generating member 22c upward.
  • the lower end of the outer casing 21c is also gradually consumed by the exhaust ablation.
  • the cooling liquid of the cooling device may be stored in the nozzle generating member 22c or conveyed from the front portion of the rocket through a conveying pipe.
  • the gas generated by the evaporation of the cooling liquid by the heat may be discharged from the outlet of the cooling material, and may also provide part of the power of the rocket flight.
  • control method of the electrode and the cooling channel described in the first embodiment may be provided on the solid fuel 1c; and the design method described in the first embodiment may also be used.
  • a lateral scratch is provided on the outer casing and/or a transverse fiber material is disposed inside the outer casing material, which is not described herein.
  • the outer casing 21c may also be made of a memory material, and the inner diameter of the memory material is greater than the outer diameter of the solid fuel 1c.
  • the lower end of the outer casing 21c will automatically expand, thereby forming a diverging pipe section that facilitates the formation of the spout.

Abstract

L'invention concerne un dispositif dynamique de propulsion de fusée utilisant un carburant solide (1a) et pourvu d'un élément de commande (2a), faisant en sorte que la partie inférieure du corps de fusée forme une structure de tuyère de Laval (3a) appropriée pour un vol de fusée. L'élément de commande (2a) comprend un boîtier de corps de fusée (21a), et lorsque le carburant solide subit une combustion, un bord inférieur du boîtier (21a) peut être dynamiquement enlevé en même temps que le front de combustion du carburant solide. De cette façon, durant le vol de la fusée, la partie inférieure du corps de fusée peut être amenée dynamiquement à former la structure de tuyère (3a) appropriée pour un vol de fusée, et les déchets produits servent essentiellement de milieu de travail pour propulser le vol de la fusée et sont éjectés par la structure de tuyère, c'est-à-dire le poids en excès de la fusée peut être éliminé d'une manière opportune, et des déchets d'un volume relativement important ne sont pas produits, ce qui améliore les degrés de liberté pour la conception d'une trajectoire de vol de fusée.
PCT/CN2013/073154 2013-03-25 2013-03-25 Dispositif dynamique de propulsion de fusée WO2014153712A1 (fr)

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Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3285262A (en) * 1962-08-07 1966-11-15 Snecma Aerodynamic or hydrodynamic servovalve, especially for use for the guidance and stabilisation of rockets
US3724216A (en) * 1957-06-13 1973-04-03 Us Navy Combined rocket-ram-jet aircraft
US4630539A (en) * 1985-06-24 1986-12-23 The United States Of America As Represented By The Secretary Of The Army Device for flash suppression of a rocket motor
CN1119522C (zh) * 1998-08-19 2003-08-27 阿利安特技术系统公司 由作为母体的短纤维素纤维制成的火箭组件的烧蚀材料及采用这种材料绝热和热保护火箭组件的方法
CN201165916Y (zh) * 2008-03-24 2008-12-17 沈阳理工大学 固体火箭发动机

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3724216A (en) * 1957-06-13 1973-04-03 Us Navy Combined rocket-ram-jet aircraft
US3285262A (en) * 1962-08-07 1966-11-15 Snecma Aerodynamic or hydrodynamic servovalve, especially for use for the guidance and stabilisation of rockets
US4630539A (en) * 1985-06-24 1986-12-23 The United States Of America As Represented By The Secretary Of The Army Device for flash suppression of a rocket motor
CN1119522C (zh) * 1998-08-19 2003-08-27 阿利安特技术系统公司 由作为母体的短纤维素纤维制成的火箭组件的烧蚀材料及采用这种材料绝热和热保护火箭组件的方法
CN201165916Y (zh) * 2008-03-24 2008-12-17 沈阳理工大学 固体火箭发动机

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