WO2021128448A1 - 基于液体二氧化碳相变的推进方法及其推进装置 - Google Patents

基于液体二氧化碳相变的推进方法及其推进装置 Download PDF

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Publication number
WO2021128448A1
WO2021128448A1 PCT/CN2020/070212 CN2020070212W WO2021128448A1 WO 2021128448 A1 WO2021128448 A1 WO 2021128448A1 CN 2020070212 W CN2020070212 W CN 2020070212W WO 2021128448 A1 WO2021128448 A1 WO 2021128448A1
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carbon dioxide
phase
propulsion
liquid
thrust
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PCT/CN2020/070212
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English (en)
French (fr)
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常正实
王聪
张冠军
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西安交通大学
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Publication of WO2021128448A1 publication Critical patent/WO2021128448A1/zh
Priority to US17/514,887 priority Critical patent/US11858666B2/en

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    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/40—Arrangements or adaptations of propulsion systems
    • B64G1/405—Ion or plasma engines
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/26—Guiding or controlling apparatus, e.g. for attitude control using jets
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/40—Arrangements or adaptations of propulsion systems
    • B64G1/401—Liquid propellant rocket engines
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/40—Arrangements or adaptations of propulsion systems
    • B64G1/402—Propellant tanks; Feeding propellants
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/40—Arrangements or adaptations of propulsion systems
    • B64G1/409—Unconventional spacecraft propulsion systems
    • 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
    • F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
    • F02K9/425—Propellants
    • 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
    • F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
    • F02K9/60—Constructional parts; Details not otherwise provided for
    • F02K9/62—Combustion or thrust chambers
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F03—MACHINES OR ENGINES FOR LIQUIDS; WIND, SPRING, OR WEIGHT MOTORS; PRODUCING MECHANICAL POWER OR A REACTIVE PROPULSIVE THRUST, NOT OTHERWISE PROVIDED FOR
    • F03H—PRODUCING A REACTIVE PROPULSIVE THRUST, NOT OTHERWISE PROVIDED FOR
    • F03H1/00—Using plasma to produce a reactive propulsive thrust
    • F03H1/0093—Electro-thermal plasma thrusters, i.e. thrusters heating the particles in a plasma

Definitions

  • the invention belongs to the technical field of aerospace and extraterrestrial exploration and propulsion, in particular to a propulsion method and propulsion device based on phase change of liquid carbon dioxide.
  • propulsion technology is the basic guarantee for the transportation of various types of aircraft.
  • the existing mature technologies include launch vehicle technology and electric propulsion technology.
  • the thrust of the launch vehicle mainly comes from the energy conversion when the propellant is burned.
  • the type of propellant it can be divided into: liquid rocket propulsion and solid rocket propulsion. While these two technologies obtain large thrust, some of the drawbacks are also exposed.
  • ultra-high temperature ablation materials environmental hazards caused by toxic and harmful or even highly toxic emissions, non-reusable and low fuel controllability (such as solid propellants).
  • Electric propulsion technology is widely used in the propulsion, deceleration and attitude adjustment of small and medium-sized products in near-Earth and deep space exploration due to the characteristics of low thrust, repeatable ignition and long endurance. However, it is widely used in obtaining large thrust and high-efficiency propulsion applications. Looks powerless.
  • the above-mentioned propulsion technologies are all chemical propulsion.
  • the electric propulsion technology is mainly the plasma generated by the ionization of the working fluid, which generates acceleration and thrust under the action of an electromagnetic field.
  • the present invention proposes a liquid carbon dioxide phase change-based propulsion method and propulsion device, which can match thrust according to different application requirements, avoid high temperature ablation and environmental pollution, and achieve high reliability and repeatability.
  • the goal of use is to meet the requirements of promoting physical advancement in different fields.
  • a propulsion method based on the phase change of liquid carbon dioxide includes the following steps:
  • the carbon dioxide is contained in a thermally insulated container in the form of a liquid phase
  • the instantaneous heating makes the carbon dioxide change from the liquid phase to the gas phase.
  • the carbon dioxide gas after the phase change is sprayed in a predetermined direction with a predetermined amount of ejection to obtain thrust.
  • carbon dioxide is contained in a thermally insulated container in a liquid phase at a predetermined temperature, and the predetermined temperature is a room temperature lower than the phase transition temperature of the carbon dioxide liquid phase to the gas phase.
  • the predetermined temperature is 10°C.
  • the instantaneous heating takes milliseconds.
  • heat is instantaneously heated through heat transfer, heat exchange or energy conversion.
  • the temperature rise of instant heating does not exceed 21°C.
  • the phase-change carbon dioxide gas injection is controlled by the relief valve to obtain thrust.
  • the thrust is a continuous force that lasts for a predetermined time.
  • a propulsion device includes,
  • Insulated container which contains carbon dioxide in liquid phase
  • the instant heating module which instantly heats the carbon dioxide in the liquid phase in the insulated container, so that the carbon dioxide is converted from the liquid phase to the gas phase,
  • the injection module which can controllably eject the phase-change carbon dioxide gas to obtain thrust.
  • the thermal insulation container includes a thermal insulation tank that is subjected to a predetermined pressure
  • the injection module includes a plurality of relief valves arranged on the thermal insulation tank with different orientations, and the relief valve adjusts the carbon dioxide gas injection in response to a predetermined thrust. ⁇ .
  • the present invention has the following advantages:
  • the present invention is a physical change that releases huge thrust, makes up for the shortcomings of the existing propulsion technology, fills the gap between the existing liquid/solid fuel propulsion and electric propulsion, and is an innovation and upgrade to the existing technology.
  • the technical temperature rise of the present invention is only about 21°C, which has almost no thermal effect on the engine and avoids ablation damage;
  • the present invention is a purely physical process, no chemical reaction, no emission of any toxic and harmful substances, and it is green and environmentally friendly;
  • the propellant used in the present invention is only carbon dioxide, which is very low in cost and easy to obtain. At the same time, the thruster based on this technology can be reused, which can greatly reduce the cost of use;
  • the phase change process can be realized in milliseconds by heating and other methods.
  • the temperature and thrust response time are rapid.
  • the carbon dioxide after the phase change is ejected in the form of gas on demand to obtain thrust for different applications, and the reliability is greatly improved.
  • Fig. 1 is a schematic diagram of the steps of a propulsion method based on a phase change of liquid carbon dioxide according to an embodiment of the present invention
  • FIG. 2 is a schematic flow chart of a propulsion method based on phase change of liquid carbon dioxide according to an embodiment of the present invention
  • FIG. 3 is a schematic structural diagram of a propulsion device based on liquid carbon dioxide phase change according to an embodiment of the present invention
  • Fig. 4 is a schematic structural diagram of a propulsion device based on a phase change of liquid carbon dioxide according to an embodiment of the present invention.
  • FIGS. 1 to 4 Although specific embodiments of the present invention are shown in the drawings, it should be understood that the present invention can be implemented in various forms and should not be limited by the embodiments set forth herein. On the contrary, these embodiments are provided to enable a more thorough understanding of the present invention and to fully convey the scope of the present invention to those skilled in the art.
  • FIG. 1 is a schematic diagram of the steps of a method according to an embodiment of the present invention.
  • the propulsion method based on the phase change of liquid carbon dioxide includes the following steps:
  • the instantaneous heating causes the carbon dioxide to change from the liquid phase to the gas phase
  • the carbon dioxide gas after the phase change is sprayed in a predetermined direction in a predetermined ejection amount to obtain thrust.
  • the liquid-phase carbon dioxide of the present invention can be stored at 10°C, the liquid-gas phase transition temperature is 31°C, and the phase transition temperature is close to room temperature; when the liquid-gas phase transition occurs, the volume or pressure instantly increases by 500-600 times , From liquid to gas, the process is purely physical, without any toxic and harmful substances; the gas after phase change can be ejected through the directional, which can produce huge thrust, which can achieve any propulsion requirements, such as ground carrier rockets one and two
  • carbon dioxide is contained in the insulating container 1 in a liquid phase at a predetermined temperature, and the predetermined temperature is a room temperature lower than the phase transition temperature of the carbon dioxide liquid phase to the gas phase.
  • the predetermined temperature is 10°C.
  • the instantaneous heating time is on the order of milliseconds.
  • instant heating is performed via heat transfer, heat exchange or energy conversion.
  • the temperature rise of the instant heating does not exceed 21°C.
  • the phase-change carbon dioxide gas injection is controlled by the relief valve 5 to obtain thrust.
  • the thrust is a continuous force that lasts for a predetermined time.
  • the heat-insulating container 1 is a high-pressure-resistant closed container.
  • the propulsion method based on liquid carbon dioxide phase change includes: liquid-gas phase transition realization technology and propulsion control technology; among them,
  • the liquid phase-gas phase change realization technology mainly realizes the transformation of carbon dioxide from the liquid phase into the gas phase through instant heating.
  • the instant heating technology includes but not limited to electric heating, heat transfer and heat exchange, and energy conversion;
  • the propulsion control technology mainly adopts automatic
  • the adaptive program commands the 5-control technology of the automatic relief valve to spray the phase-change carbon dioxide gas in the gas-phase storage part on demand to obtain the target thrust.
  • the propulsion method based on the phase change of liquid carbon dioxide includes:
  • Liquid carbon dioxide is injected into a heat-insulating container 1, such as a specific heat-insulating tank body.
  • the tank body is provided with electrodes 4, for example, the liquid carbon dioxide is heated by micro-current and high heat-sensitive materials, or the liquid carbon dioxide is heated by high-voltage discharge, or other heating methods, etc.
  • the liquid carbon dioxide heats up instantaneously and reaches the temperature phase change point to realize the phase change.
  • the end of the tank is equipped with an automatic control pressure relief valve 5 to release the gas after the phase change to obtain the thrust of different requirements.
  • a propulsion device includes,
  • Insulated container 1 which contains carbon dioxide in liquid phase
  • Instantaneous heating module 2 which instantaneously heats the carbon dioxide in the liquid phase in the thermal insulation container 1, so that the carbon dioxide is converted from the liquid phase to the gas phase,
  • the injection module 3 which can controllably eject the phase-change carbon dioxide gas to obtain thrust.
  • the thermally insulated container 1 includes a thermally insulated tank that is subjected to a predetermined pressure
  • the injection module 3 includes a plurality of relief valves 5 arranged on the thermally insulated tank with different orientations, and the relief valve 5 The carbon dioxide gas ejection amount is adjusted in response to the predetermined thrust.
  • the inner diameter of the insulated container 1 such as the thruster tank is adjustable from 5cm to 50cm, and the height, that is, the plane of the jet module 3 with the pressure relief valve 5 at its tail is located to the smallest radius of curvature of the thruster head Adjustable from 10cm-150cm; heat-insulating container 1 uses heat-insulating materials, such as heat-insulating carbon steel, alloys, etc., to maintain the temperature of liquid carbon dioxide in the tank always below the phase transition temperature point, when triggered by micro-current or high-voltage discharge plasma When heated, it can quickly rise to the phase transition point.
  • heat-insulating container 1 uses heat-insulating materials, such as heat-insulating carbon steel, alloys, etc.
  • the head of the heat-insulating container 1 is ellipsoidal, the length of the short axis of the ellipse is adjustable from 5cm to 50cm, and the length of the semi-major axis is adjustable from 6cm to 100cm; it reduces wind resistance and facilitates stable flow field formation and heat loss during the propulsion process. .
  • the instant heating module 2 includes an electrode 4, which is a good conductor, with a diameter of 0.1mm-2mm adjustable.
  • the material is not limited to copper and stainless steel. It is introduced into the thruster tank through an insulator sleeve at both ends of the electrode 4. A certain voltage is applied, for example, the voltage amplitude is adjustable from a few volts to a few hundred volts, so as to generate a micro current to heat the liquid carbon dioxide.
  • the insulator sleeve is an arc umbrella skirt structure to increase the creepage distance and ensure insulation safety; the electrode 4 is placed in the middle to achieve insulation from the inner and outer walls of the thruster tank.
  • the high-temperature-sensitive material is in the shape of " ⁇ ", spiral, etc.; the two electrodes 4 inside the thruster tank are connected to the high-temperature-sensitive material.
  • the high heat-sensitive material generates heat instantaneously, with a response time ranging from several hundred milliseconds to several milliseconds, and instantly releases huge heat, which is used to heat liquid carbon dioxide to achieve phase change.
  • the strength of the instant heating module 2 is greater than the pressure generated by the phase change.
  • the instantaneous heating module 2 is arranged 1 cm-4 cm away from the wall of the insulating container 1.
  • the thermally insulated container 1 is provided with a protective structure for protecting the instantaneous heating module 2.
  • the strength of the protective structure is greater than the pressure generated by the phase change, and further, the pressure difference is 10 MPa-500 MPa, which can be adjusted according to different needs.
  • the protection structure and the thermal insulation container 1 are integrally formed.
  • the protection structure is an arc-shaped shielding structure.
  • the injection module 3 includes a relief valve 5 arranged at the tail of the thermal insulation container 1 and a nozzle 6 connected to the relief valve 5, and the pressure relief threshold is adjustable from 15MPa to 100MPa.
  • an automatic pressure relief valve 5 is provided at the rear of the thruster, and the pressure release threshold is automatically controlled according to the pressure in the phase change process inside the thruster tank and the required thrust.
  • the inner diameter of the nozzle 6 is adjustable from 1 cm to 15 cm, which is a cone, bell, plug, expansion-divergent flow and other geometric shapes of the nozzle 6 structure, and an adjustable angle nozzle 6 is provided at the tail of the thruster. , Used to guide the gas phase carbon dioxide released from the relief valve 5 to generate directional thrust.
  • the inner diameter of the insulated container 1 such as the thruster tank is adjustable from 5cm to 50cm, and the height, that is, the plane of the jet module 3 with the pressure relief valve 5 at its tail is located to the smallest radius of curvature of the thruster head Adjustable from 10cm-150cm; heat-insulating container 1 uses heat-insulating materials, such as heat-insulating carbon steel, alloys, etc., to maintain the temperature of liquid carbon dioxide in the tank always below the phase transition temperature point, when triggered by micro-current or high-voltage discharge plasma When heated, it can quickly rise to the phase transition point.
  • heat-insulating container 1 uses heat-insulating materials, such as heat-insulating carbon steel, alloys, etc.
  • the head of the heat-insulating container 1 is ellipsoidal, that is, the head of the thruster is ellipsoidal, the length of the short axis of the ellipse is adjustable from 5cm to 50cm, and the length of the semi-major axis is adjustable from 6cm to 100cm; it reduces wind resistance and facilitates stability. The heat lost during the formation and propulsion of the flow field.
  • the instant heating module 2 includes an electrode 4, which is a good conductor, with a diameter of 0.1mm-2mm adjustable.
  • the material is not limited to copper and stainless steel. It is introduced into the thruster tank through an insulator sleeve at both ends of the electrode 4. A certain voltage is applied, for example, the voltage amplitude is adjustable from several hundreds of volts to several tens of thousands of volts, based on the discharge plasma generated in the gap of the high thermal sensitive electrode 4 inside the tank.
  • High-voltage insulator bushing Designed as a curved umbrella skirt structure to increase the creepage distance and improve the insulation level; the electrode 4 is placed in the middle to achieve insulation from the inner and outer walls of the thruster tank.
  • the insulating container 1 is provided with a protective structure for protecting the instantaneous heating module 2, and the strength of the protective structure is greater than the pressure generated by the phase change. Further, the protection structure and the thermal insulation container 1 are integrally formed.
  • the protection structure is an arc-shaped shielding structure.
  • the arrangement between every two horizontally high thermal sensitive electrodes 4 is designed as a needle-needle, rod-rod electrode 4 structure, and the inside of the tank can be arranged according to thrust requirements. 1-20 pairs of electrode 4 structure; so that when a high voltage is applied to the electrode 4, a discharge plasma can be generated in the gap; while the discharge plasma channel is generated, a higher current is formed, which heats the high heat sensitive electrode 4 material.
  • the ends of the two electrodes 4 inside the thruster tank are connected to a high-temperature-sensitive material.
  • a discharge plasma is generated, a large current flows through the high-temperature-sensitive material, and the high-temperature-sensitive material generates instantaneous heat, and the response time is Ranging from a few microseconds to a few milliseconds, it releases huge heat instantly, which is used to heat the liquid phase carbon dioxide and realize the phase change.
  • the injection module 3 includes a relief valve 5 provided at the end of the insulated container 1 and a nozzle 6 connected to the relief valve 5.
  • an automatic pressure relief valve 5 is provided at the end of the thruster, The pressure relief threshold is adjustable from 15MPa-100MPa, and the pressure relief threshold is automatically controlled according to the pressure in the phase change process of the thruster tank and the required thrust.
  • the inner diameter of the nozzle 6 is adjustable from 1cm to 15cm; the nozzle 6 is designed as a conical, bell-shaped, plug-type, expansion-divergent flow and other geometric shapes, and an adjustable-angle nozzle 6 is provided at the tail of the thruster. It is used to guide the gas-phase carbon dioxide released from the relief valve 5 to generate a directional thrust.

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Abstract

一种基于液体二氧化碳相变的推进方法及推进装置，二氧化碳以液相形态容纳于绝热容器(1)中，瞬间加热模块（2）瞬间加热使得二氧化碳由液相转化为气相，相变后的二氧化碳气体在预定方向上喷射预定的喷出量以获得推力。

Description

基于液体二氧化碳相变的推进方法及其推进装置 技术领域
本发明属于航空航天和地外探索推进技术领域，特别是一种基于液体二氧化碳相变的推进方法及其推进装置。
背景技术
近年来，随着空天和地外探测技术的迅速发展，人类对地外资源的探索和开发、利用的需求愈发旺盛。而在地外探索中，推进技术是运输各类飞行器的基本保障。目前已有的成熟技术有运载火箭技术和电推进技术。其中运载火箭的推力主要来自于推进剂燃烧时的能量转化，根据推进剂的类型可分为：液体火箭推进和固体火箭推进，这两种技术在获得大推力的同时，一些弊端也暴露无遗，如超高温烧蚀材料、有毒有害甚至剧毒排出物带来的环境危害、不可重复使用和燃料可控性低(如固体推进剂)等。电推进技术由于推力小、可重复点火、续航时间长等特点，在近地以及深空探测中小型型号产品的推进、减速和姿态调整方面应用广泛，但是在获取大推力和高效率推进应用领域显得力不从心。上述推进技术均属于化学推进，其中电推进技术主要是通过工质电离产生的等离子体，在电磁场的作用下产生加速、推力。
在背景技术部分中公开的上述信息仅仅用于增强对本发明背景的理解，因此可能包含不构成在本国中本领域普通技术人员公知的现有技术的信息。
发明内容
针对现有技术中存在的问题，本发明提出一种基于液体二氧化碳相变的推进方法及其推进装置，能够根据不同应用需求匹配推力，避免高温烧蚀和环境污染，实现高可靠性、可重复使用的目标，达到为不同领域推广物理推进的要求。
本发明的目的是通过以下技术方案予以实现，一种基于液体二氧化碳相变的推进方法包括以下步骤：
第一步骤中，二氧化碳以液相形态容纳于绝热容器中，
第二步骤中，瞬间加热使得二氧化碳由液相转化为气相，
第三步骤中，相变后的二氧化碳气体在预定方向上以喷射预定喷出量以获得推力。
所述的方法中，第一步骤中，二氧化碳在预定温度下以液相形态容纳于绝热容器中，所述预定温度为低于二氧化碳液相气相相变温度的室温。
所述的方法中，第一步骤中，所述预定温度为10℃。
所述的方法中，第二步骤中，瞬间加热耗时为毫秒级。
所述的方法中，，第二步骤中，经由热传递、热交换或能量转化瞬间加热。
所述的方法中，第二步骤中，瞬间加热的温升不超过21℃。
所述的方法中，第三步骤中，通过泄放阀控制相变后的二氧化碳气体喷射以获得推力。
所述的方法中，第三步骤中，所述推力为持续预定时间的持续作用力。
根据本发明的另一方面，一种推进装置包括，
绝热容器，其容纳液相形态的二氧化碳，
瞬间加热模块，其瞬间加热所述绝热容器内液相形态的二氧化碳，使得二氧化碳由液相转化为气相，
喷射模块，其可控地喷出相变后的二氧化碳气体以获得推力。
所述的推进装置中，绝热容器包括承受预定压力的绝热罐，所述喷射模块包括设在绝热罐上的多个朝向不同的泄放阀，所述泄放阀响应于预定推力调节二氧化碳气体喷出量。
和现有技术相比，本发明具有以下优点：
1、先进性方面：本发明属于物理变化释放巨大推力，弥补现有推进技术的缺点，填补了现有液体/固体燃料推进、电推进之间的空白，是对现有技术的革新升级。
2、温度方面：本发明的技术温升只有大约21℃，对发动机而言几乎无热效应影响、避免了烧蚀损伤；
3、环保方面：本发明属于纯物理过程，无化学反应，无任何有毒有害物质排放，绿色环保；
4、经济方面：本发明使用的推进剂只有二氧化碳，成本非常低廉，获取容易，同时，基于该技术的推力器可重复使用，能够大幅降低使用成本；
5、可靠性方面：通过加热等方法在毫秒级即可实现相变过程，温度和推力响应时间迅速，相变后的二氧化碳以气体形式按需喷出，获取不同应用推力，可靠程度大幅提高。
附图说明
通过阅读下文优选的具体实施方式中的详细描述，本发明各种其他的优点和益处对于本领域普通技术人员将变得清楚明了。说明书附图仅用于示出优选实施方式的目的，而并不认为是对本发明的限制。显而易见地，下面描述的附图仅仅是本发明的一些实施例，对于本领域普通技术人员来讲，在不付出创造性劳动的前提下，还可以根据这些附图获得其他的附图。而且在整个附图中，用相同的附图标记表示相同的部件。
在附图中：
图1是根据本发明一个实施例的基于液体二氧化碳相变的推进方法的步骤示意图；
图2是根据本发明一个实施例的基于液体二氧化碳相变的推进方法的流程示意图；
图3是根据本发明一个实施例的基于液体二氧化碳相变的推进装置的结构示意图；
图4是根据本发明一个实施例的基于液体二氧化碳相变的推进装置的结构示意图。
以下结合附图和实施例对本发明作进一步的解释。
具体实施方式
下面将参照附图1至图4更详细地描述本发明的具体实施例。虽然附图中显示了本发明的具体实施例，然而应当理解，可以以各种形式实现本发明而不应被这里阐述的实施例所限制。相反，提供这些实施例是为了能够更透彻地理解本发明，并且能够将本发明的范围完整的传达给本领域的技术人员。
需要说明的是，在说明书及权利要求当中使用了某些词汇来指称特定组件。本领域技术人员应可以理解，技术人员可能会用不同名词来称呼同一个组件。本说明书及权利要求并不以名词的差异来作为区分组件的方式，而是以组件在功能 上的差异来作为区分的准则。如在通篇说明书及权利要求当中所提及的“包含”或“包括”为一开放式用语，故应解释成“包含但不限定于”。说明书后续描述为实施本发明的较佳实施方式，然所述描述乃以说明书的一般原则为目的，并非用以限定本发明的范围。本发明的保护范围当视所附权利要求所界定者为准。
为便于对本发明实施例的理解，下面将结合附图以具体实施例为例做进一步的解释说明，且各个附图并不构成对本发明实施例的限定。
为了更好地理解，图1是根据本发明一个实施例的方法的步骤示意图，如图1所示，基于液体二氧化碳相变的推进方法包括以下步骤：
第一步骤S1中，二氧化碳以液相形态容纳于绝热容器1中，
第二步骤S2中，瞬间加热使得二氧化碳由液相转化为气相，
第三步骤S3中，相变后的二氧化碳气体在预定方向上以喷射预定喷出量以获得推力。
本发明的液相二氧化碳正可储存在10℃，其液相-气相相变温度为31℃，相变温度接近室温；在发生液相-气相相变时，体积或压力瞬间增加500-600倍，由液体变成气体，该过程纯属物理过程，无任何有毒有害物质产生；将相变后的气体通过定向喷出，可产生推力巨大，可实现任何推进需求，如地面运载火箭一、二级阶段的发射，火箭分离后飞行器的继续推进，飞行器在月球、火星、金星甚至一些小行星表面着陆阶段的高效减速、再发射，火星表面基于母舰的远距离表面勘察飞行器起飞、着陆推进等。
所述的方法的优选实施方式中，第一步骤S1中，二氧化碳在预定温度下以液相形态容纳于绝热容器1中，所述预定温度为低于二氧化碳液相气相相变温度的室温。
所述的方法的优选实施方式中，第一步骤S1中，所述预定温度为10℃。
所述的方法的优选实施方式中，第二步骤S2中，瞬间加热耗时为毫秒级。
所述的方法的优选实施方式中，第二步骤S2中，经由热传递、热交换或能量转化瞬间加热。
所述的方法的优选实施方式中，第二步骤S2中，瞬间加热的温升不超过21℃。
所述的方法的优选实施方式中，第三步骤S3中，通过泄放阀5控制相变后 的二氧化碳气体喷射以获得推力。
所述的方法的优选实施方式中，第三步骤S3中，所述推力为持续预定时间的持续作用力。
所述的方法的优选实施方式中，所述绝热容器1为耐高压的密闭容器。
为了进一步理解本发明，在一个实施方式中，基于液体二氧化碳相变的推进方法包括：液相-气相相变实现技术和推进控制技术；其中，
所述液相-气相相变实现技术主要通过瞬间加热实现二氧化碳有液相变为气相，瞬间加热技术包括但不限于电加热、热传递和热交换、能量转化；所述推进控制技术主要通过自适应程序命令自动泄放阀5控技术将相变后的二氧化碳气体在气相储存部分按需喷出，获得目标推力。
如图2所示，在一个实施方式中，基于液体二氧化碳相变的推进方法包括，
液体二氧化碳注入如特定的绝热罐体的绝热容器1内，罐体设置电极4，通过例如微电流和高热敏材料加热液体二氧化碳，或者高电压放电加热液体二氧化碳，或其他加热方式等，罐内液体二氧化碳瞬时升温，达到温度相变点实现相变，罐体尾部设有自动控制泄压阀5释放相变后的气体，获得不同需求的推力。
一种推进装置包括，
绝热容器1，其容纳液相形态的二氧化碳，
瞬间加热模块2，其瞬间加热所述绝热容器1内液相形态的二氧化碳，使得二氧化碳由液相转化为气相，
喷射模块3，其可控地喷出相变后的二氧化碳气体以获得推力。
所述的推进装置的优选实施例中，绝热容器1包括承受预定压力的绝热罐，所述喷射模块3包括设在绝热罐上的多个朝向不同的泄放阀5，所述泄放阀5响应于预定推力调节二氧化碳气体喷出量。
如图3所示的实施例中，如推力器罐体的绝热容器1的内径5cm-50cm可调，高度，即尾部包括泄压阀5的喷射模块3所在平面至推力器头部曲率半径最小处10cm-150cm可调；绝热容器1采用绝热材料，如绝热碳钢、合金等，维持罐内液相二氧化碳的温度始终在相变温度点以下，当利用如微电流或高电压放电等离子体触发加热时，能够迅速上升至相变点。
在一个实施例中，绝热容器1头部为椭球状，椭圆的短轴长度5cm-50cm可 调，半长轴长度6cm-100cm可调；减少风阻，利于稳定流场形成和推进过程的热量散失。
在一个实施例中，瞬间加热模块2包括电极4，其为良导体，直径0.1mm-2mm可调，材料不限于铜、不锈钢，经由绝缘子套管引入推力器罐体内部，在电极4两端施加一定电压，例如电压幅值为几伏特至几百伏特可调，以产生微电流以便加热液体二氧化碳为准。绝缘子套管为弧形伞裙结构，增加爬电距离，确保绝缘安全；中间安置电极4，实现与推力器罐体内、外壁绝缘。
在一个实施例中，高热敏材料为“Π”、螺旋形等形状；在推力器罐体内部的两个电极4端头连接高热敏材料，当微电流流过高热敏材料时，高热敏材料瞬时发热，响应时间为几百毫秒至几毫秒不等，瞬间释放巨大热量，用于加热液相二氧化碳，实现相变。
在一个实施例中，所述瞬间加热模块2的强度大于相变产生的压强。
在一个实施例中，所述瞬间加热模块2距离绝热容器1壁1cm-4cm布置。
在一个实施例中，绝热容器1设在保护瞬间加热模块2的保护结构，所述保护结构强度大于所述相变产生的压强，进一步，压强差为10MPa-500MPa，根据不同需求调整。进一步，所述保护结构与绝热容器1一体成型。
在一个实施例中，所述保护结构为弧形遮挡结构。
在一个实施例中，喷射模块3包括设在绝热容器1尾部的泄放阀5和连接泄放阀5的喷嘴6，泄压阈值为15MPa-100MPa可调。
在一个实施例中，在推力器尾部设有自动压力泄放阀5，根据推力器罐体内部相变过程的压力和需求推力大小，自动控制压力释放阈值。
在一个实施例中，喷嘴6的内径1cm-15cm可调，其为锥形、钟形、塞式、膨胀-偏流等几何形状的喷嘴6结构，在推力器尾部设有可调节角度的喷嘴6，用于引导来自泄放阀5释放的气相二氧化碳，产生定向推力。
如图4所示的实施例中，如推力器罐体的绝热容器1的内径5cm-50cm可调，高度，即尾部包括泄压阀5的喷射模块3所在平面至推力器头部曲率半径最小处10cm-150cm可调；绝热容器1采用绝热材料，如绝热碳钢、合金等，维持罐内液相二氧化碳的温度始终在相变温度点以下，当利用如微电流或高电压放电等离子体触发加热时，能够迅速上升至相变点。
在一个实施例中，绝热容器1头部为椭球状，即推力器头部为椭球状，椭圆的短轴长度5cm-50cm可调，半长轴长度6cm-100cm可调；减少风阻，利于稳定流场形成和推进过程的热量散失。
在一个实施例中，瞬间加热模块2包括电极4，其为良导体，直径0.1mm-2mm可调，材料不限于铜、不锈钢，经由绝缘子套管引入推力器罐体内部，在电极4两端施加一定的电压，例如电压幅值为几百伏特至几万伏特可调，以在罐体内部高热敏电极4间隙产生放电等离子体为准。高压绝缘子套管：设计为弧形伞裙结构，增加爬电距离，提高绝缘等级；中间安置电极4，实现与推力器罐体内、外壁绝缘。
在一个实施例中，绝热容器1设在保护瞬间加热模块2的保护结构，所述保护结构强度大于所述相变产生的压强。进一步，所述保护结构与绝热容器1一体成型。
在一个实施例中，所述保护结构为弧形遮挡结构。
在一个实施例中，电极4结构中，每两个水平方向的高热敏电极4之间布置为设计为针-针、棒-棒等电极4结构，在罐体内部根据推力需求，可以布置1-20对电极4结构；以便于在电极4上施加高电压时，能够在间隙中产生放电等离子体；在放电等离子体通道产生的同时，形成较高的电流，加热了高热敏电极4材料。
在一个实施例中，在推力器罐体内部的两个电极4端头连接高热敏材料，当产生放电等离子体时，大电流流过高热敏材料，高热敏材料瞬时发热，响应时间为几微秒至几毫秒不等，瞬间释放巨大热量，用于加热液相二氧化碳，实现相变。
在一个实施例中，喷射模块3包括设在绝热容器1尾部的泄放阀5和连接泄放阀5的喷嘴6，在一个实施例中，在推力器尾部设有自动压力泄放阀5，泄压阈值为15MPa-100MPa可调，根据推力器罐体内部相变过程的压力和需求推力大小，自动控制压力释放阈值。
在一个实施例中，喷嘴6内径1cm-15cm可调；设计为锥形、钟形、塞式、膨胀-偏流等几何形状的喷嘴6结构，在推力器尾部设有可调节角度的喷嘴6，用于引导来自泄放阀5释放的气相二氧化碳，产生定向推力。
尽管以上结合附图对本发明的实施方案进行了描述，但本发明并不局限于上 述的具体实施方案和应用领域，上述的具体实施方案仅仅是示意性的、指导性的，而不是限制性的。本领域的普通技术人员在本说明书的启示下和在不脱离本发明权利要求所保护的范围的情况下，还可以做出很多种的形式，这些均属于本发明保护之列。

Claims (10)

  1. 一种基于液体二氧化碳相变的推进方法，所述方法包括以下步骤：
    第一步骤(S1)中，二氧化碳以液相形态容纳于绝热容器中，
    第二步骤(S2)中，瞬间加热使得二氧化碳由液相转化为气相，
    第三步骤(S3)中，相变后的二氧化碳气体在预定方向上以喷射预定喷出量以获得推力。
  2. 根据权利要求1所述的方法，其中，第一步骤(S1)中，二氧化碳在预定温度下以液相形态容纳于绝热容器中，所述预定温度为低于二氧化碳液相气相相变温度的室温。
  3. 根据权利要求2所述的方法，其中，第一步骤(S1)中，所述预定温度为10℃。
  4. 根据权利要求1所述的方法，其中，第二步骤(S2)中，瞬间加热耗时为毫秒级。
  5. 根据权利要求1所述的方法，其中，第二步骤(S2)中，经由热传递、热交换或其他能量转化瞬间加热。
  6. 根据权利要求1所述的方法，其中，第二步骤(S2)中，瞬间加热的温升不超过21℃。
  7. 根据权利要求1所述的方法，其中，第三步骤(S3)中，通过所述绝热容器设有的泄放阀控制相变后的二氧化碳气体喷射以获得推力。
  8. 根据权利要求1所述的方法，其中，第三步骤(S3)中，所述推力为持续预定时间的持续作用力。
  9. 一种推进装置，其包括，
    绝热容器，其容纳液相形态的二氧化碳，
    瞬间加热模块，其瞬间加热所述绝热容器内液相形态的二氧化碳，使得二氧化碳由液相转化为气相，
    喷射模块，其可控地喷出相变后的二氧化碳气体以获得推力。
  10. 根据权利要求9所述的推进装置，其中，绝热容器包括承受预定压力的绝热罐，所述喷射模块包括设在绝热罐上的多个朝向不同的泄放阀，所述泄放阀响应于预定推力调节二氧化碳气体喷出量。
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Citations (16)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6047541A (en) * 1998-08-26 2000-04-11 The United States Of America As Represented By The Secretary Of The Air Force HAN TEAN (xm-46) mixing gas generator propellant tank pressurizer for launch vehicles and spacecraft
CN202063261U (zh) * 2011-04-06 2011-12-07 丛洋 气箭及气箭系统
CN102735114A (zh) * 2011-04-06 2012-10-17 丛洋 气箭及气箭系统
CN107487459A (zh) * 2016-06-12 2017-12-19 梅法明 电热储能蒸汽火箭
CN206980098U (zh) * 2017-02-09 2018-02-09 合肥四叶草光电科技有限公司 一种科普用火箭模型
CN109238062A (zh) * 2018-10-19 2019-01-18 中南大学 一种动能弹装置及利用该装置疏通堵塞溜井的方法
CN109654948A (zh) * 2019-01-26 2019-04-19 中国人民解放军国防科技大学 一种二氧化碳相变膨胀发射装置及使用方法
CN208975018U (zh) * 2018-10-26 2019-06-14 山西恒赫贸易有限责任公司 一种用于消防无人机的破窗装置
CN110053789A (zh) * 2019-06-05 2019-07-26 北京宇航天驰空间科技有限公司 可变推力推进系统及推力改变方法
CN110207551A (zh) * 2019-06-13 2019-09-06 湖南瑞安安防科技有限公司 一种二氧化碳相变膨胀发射灭火装置及其使用方法
CN209639615U (zh) * 2019-01-26 2019-11-15 中国人民解放军国防科技大学 一种二氧化碳相变膨胀发射装置
CN110979715A (zh) * 2019-12-25 2020-04-10 西安交通大学 基于液相二氧化碳相变弹射方法、装置及舰、船或舰载机
CN110979748A (zh) * 2019-12-25 2020-04-10 西安交通大学 基于液相二氧化碳相变的缓冲方法、缓冲装置及飞行器
CN111003219A (zh) * 2019-12-25 2020-04-14 西安交通大学 基于高压放电的液相二氧化碳相变的推进方法及推进装置
CN111003216A (zh) * 2019-12-25 2020-04-14 西安交通大学 基于电加热的液相二氧化碳相变的推进方法及其推进装置
CN111003217A (zh) * 2019-12-25 2020-04-14 西安交通大学 基于液相二氧化碳相变的可持续推进方法、装置及飞行器

Family Cites Families (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5533331A (en) * 1994-05-25 1996-07-09 Kaiser Marquardt, Inc. Safe propulsion system for missile divert thrusters and attitude control thrusters and method for use of same
US6513602B1 (en) * 2000-09-13 2003-02-04 Universal Propolsion Company Gas generating device
US8487222B2 (en) * 2007-07-10 2013-07-16 Toyo Seikan Kaisha, Ltd. Heating electrode and method for heating material-to-be-heated by using the heating electrode
US9617069B2 (en) * 2013-03-11 2017-04-11 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Thermal insulation system for non-vacuum applications including a multilayer composite
CN104930933B (zh) * 2015-06-25 2016-08-24 永州市鑫东森机械装备有限公司 二氧化碳起爆器
CN205876318U (zh) * 2016-08-11 2017-01-11 湖北照东机械有限公司 一种智能二氧化碳致裂器
CN109374679A (zh) * 2018-10-24 2019-02-22 南京理工大学 一种混合气体爆炸压力和速度的测量装置

Patent Citations (16)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6047541A (en) * 1998-08-26 2000-04-11 The United States Of America As Represented By The Secretary Of The Air Force HAN TEAN (xm-46) mixing gas generator propellant tank pressurizer for launch vehicles and spacecraft
CN202063261U (zh) * 2011-04-06 2011-12-07 丛洋 气箭及气箭系统
CN102735114A (zh) * 2011-04-06 2012-10-17 丛洋 气箭及气箭系统
CN107487459A (zh) * 2016-06-12 2017-12-19 梅法明 电热储能蒸汽火箭
CN206980098U (zh) * 2017-02-09 2018-02-09 合肥四叶草光电科技有限公司 一种科普用火箭模型
CN109238062A (zh) * 2018-10-19 2019-01-18 中南大学 一种动能弹装置及利用该装置疏通堵塞溜井的方法
CN208975018U (zh) * 2018-10-26 2019-06-14 山西恒赫贸易有限责任公司 一种用于消防无人机的破窗装置
CN209639615U (zh) * 2019-01-26 2019-11-15 中国人民解放军国防科技大学 一种二氧化碳相变膨胀发射装置
CN109654948A (zh) * 2019-01-26 2019-04-19 中国人民解放军国防科技大学 一种二氧化碳相变膨胀发射装置及使用方法
CN110053789A (zh) * 2019-06-05 2019-07-26 北京宇航天驰空间科技有限公司 可变推力推进系统及推力改变方法
CN110207551A (zh) * 2019-06-13 2019-09-06 湖南瑞安安防科技有限公司 一种二氧化碳相变膨胀发射灭火装置及其使用方法
CN110979715A (zh) * 2019-12-25 2020-04-10 西安交通大学 基于液相二氧化碳相变弹射方法、装置及舰、船或舰载机
CN110979748A (zh) * 2019-12-25 2020-04-10 西安交通大学 基于液相二氧化碳相变的缓冲方法、缓冲装置及飞行器
CN111003219A (zh) * 2019-12-25 2020-04-14 西安交通大学 基于高压放电的液相二氧化碳相变的推进方法及推进装置
CN111003216A (zh) * 2019-12-25 2020-04-14 西安交通大学 基于电加热的液相二氧化碳相变的推进方法及其推进装置
CN111003217A (zh) * 2019-12-25 2020-04-14 西安交通大学 基于液相二氧化碳相变的可持续推进方法、装置及飞行器

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