WO2022057752A1 - 适用于空间装置的低温发动机 - Google Patents

适用于空间装置的低温发动机 Download PDF

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WO2022057752A1
WO2022057752A1 PCT/CN2021/117908 CN2021117908W WO2022057752A1 WO 2022057752 A1 WO2022057752 A1 WO 2022057752A1 CN 2021117908 W CN2021117908 W CN 2021117908W WO 2022057752 A1 WO2022057752 A1 WO 2022057752A1
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Prior art keywords
nozzle
flow channel
space
injector body
low
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PCT/CN2021/117908
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English (en)
French (fr)
Inventor
程诚
周海清
田桂
熊靖宇
曾夜明
周国峰
许宏博
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上海空间推进研究所
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Priority to US18/026,601 priority Critical patent/US12025075B2/en
Publication of WO2022057752A1 publication Critical patent/WO2022057752A1/zh

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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
    • F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
    • F02K9/44—Feeding 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
    • F02K9/95—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof characterised by starting or ignition means or arrangements
    • 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
    • B64G1/244—Spacecraft control systems
    • 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
    • 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
    • F02K9/44—Feeding propellants
    • F02K9/52—Injectors
    • 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
    • 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
    • F02K9/64—Combustion or thrust chambers having cooling arrangements

Definitions

  • the present invention relates to the technical field of rocket engines, in particular, to a low-temperature engine suitable for space devices.
  • Patent document CN108321678B discloses a pre-combustion chamber ignition structure, system and working method, including a center electrode extending into the interior of the pre-combustion chamber, and a ground electrode is arranged at the end of the pre-combustion chamber.
  • the purpose of the present invention is to provide a low-temperature engine suitable for a space device.
  • a low-temperature engine suitable for a space device provided according to the present invention includes an injector body, a thrust chamber, and a spark plug;
  • An accommodating space is provided in the injector body, the spark plug is installed on one side of the injector body, and the electrode on the spark plug extends into the accommodating space;
  • the thrust chamber is installed on the other side of the injector body and communicates with the accommodating space;
  • the injector body is provided with a combustion-supporting agent flow channel and a combustible-agent flow channel, and the combustion-supporting agent flow channel and the inflammable-agent flow channel are all connected with the accommodating space.
  • it also includes a first valve and a second valve;
  • the first valve is installed on the oxidant flow passage, and the second valve is installed on the flammable agent flow passage.
  • the oxidant flow channel is connected to the accommodating space through a self-percussion primary nozzle.
  • the self-percussion type primary nozzle includes a plurality of first nozzle holes, and every two first nozzle holes forms a self-percussion pair, and the number of the self-percussion pairs is 2 to 8, and the first nozzle holes
  • the aspect ratio is 2 to 4, and the impact angle of the self-percussion pair in the self-percussion pair is 60 to 90°.
  • an inner convex cathode is arranged on the inner wall of the injector body, the inner convex cathode divides the accommodation space into an atomization vaporization chamber and a combustion chamber, and one end of the electrode along the extending direction is installed in the gap In the convex cathode, the gap forms an annular secondary nozzle, and the annular secondary nozzle communicates with the atomization vaporization chamber and the combustion chamber.
  • a first sub-flow channel and a second sub-flow channel extend from the combustible agent flow channel, the first sub-flow channel is connected to the accommodating space through an inclined core nozzle, and the second sub-flow channel is connected to the accommodating space through a vortex nozzle.
  • a vaporization baffle is provided in the direction of the self-percussion primary nozzle toward the accommodating space, and the vaporization baffle is connected to the injector body.
  • the inclined core nozzle is connected to the combustion chamber, wherein the orientation of the inclined core nozzle and the orientation of the annular secondary nozzle are arranged at an included angle, and the included angle is 0-90°.
  • the inclined core nozzle includes a plurality of second nozzle holes, the length-diameter ratio of the second nozzle holes is 2-4, and the impact angle of every two adjacent second nozzle holes is 60-90°.
  • the height of the end face of the electrode extension end and the impact point along the axis direction of the electrode is 3-5 mm.
  • the swirl nozzle includes a plurality of third injection holes, the plurality of third injection holes are uniformly arranged along the circumference of the combustion chamber, and the third injection holes face the tangential direction of the circumference of the combustion chamber, the The aspect ratio of the third orifice is 2-4.
  • the present invention has the following beneficial effects:
  • the present invention adopts the structural design of the combination of the inclined core nozzle and the vortex nozzle, and independently controls the mixture ratio of the ignition core area and the total mixture ratio of the engine through staged combustion to ensure the reliability and thermal structure safety of the engine ignition work.
  • the advantage of unlimited number of times can meet the special requirements of tens of thousands of pulse ignition work and long-life use of attitude control engines in space applications.
  • the present invention adopts the structural design of the self-percussion primary nozzle and the atomization vaporization chamber to ensure that the liquid oxygen can be fully evaporated and vaporized into gas oxygen after being atomized by the primary nozzle, and then passes through the annular secondary nozzle. Entering the ignition core area, it is easy to be broken down and ionized during the ignition process to improve the ignition reliability of the engine, and a uniform oxygen film can be formed during steady-state operation to make the downstream combustion more stable.
  • the present invention adopts the structure design of annular secondary nozzle and inclined core methane nozzle, which can continuously generate multiple pure oxygen plasma streams during the ignition process, further improving the ignition reliability of the engine, and can effectively eliminate the burning of electrodes during steady-state operation. Corrosion hidden danger, is conducive to extending the life of the engine.
  • the wall of the thrust chamber where the vortex nozzle 24 is structurally designed to form a wall-mounted uniform film cooling can ensure the thermal structure safety of the engine.
  • the present invention is suitable for low-thrust liquid oxygen and methane attitude control engines, and is also suitable for small-thrust rocket engines combined with low-temperature two-component non-self-igniting propellants such as liquid oxygen/liquid hydrogen or liquid oxygen/kerosene, and has a wide range of applications , which improves the practicability of the device.
  • FIG. 1 is a schematic structural diagram of an embodiment of the present invention.
  • FIG. 2 is a working principle diagram of an embodiment of the present invention.
  • FIG. 3 is a schematic structural diagram of a self-percussion primary nozzle according to an embodiment of the present invention.
  • FIG. 4 is a schematic structural diagram of an annular secondary nozzle according to an embodiment of the present invention.
  • FIG. 5 is a schematic structural diagram of a vortex nozzle according to an embodiment of the present invention.
  • the figure shows:
  • Injector body 1 Flammable agent flow channel 12 Self-percussion primary nozzle 21
  • Thrust chamber 2 Vaporization baffle 13
  • Annular secondary nozzle 22
  • the present invention provides a low-temperature engine suitable for a space device, as shown in Figures 1 to 5, comprising an injector body 1, a thrust chamber 2, and a spark plug 3, wherein the injector body 1 is provided with a accommodating space, The spark plug 3 is installed on one side of the injector body 1 and the electrode 15 on the spark plug 3 extends into the accommodating space, and the thrust chamber 2 is installed on the other side of the injector body 1 and communicates with the accommodating space , wherein, the spark plug 3 is preferably installed on the injector body 1 through threads, and the thrust chamber 2 can be installed on the injector body 1 by welding, screwing or flange form, and the injector body 1 is provided with There are oxidant flow channels 11 and combustible agent flow channels 12, and the combustible agent flow channels 11 and flammable agent flow channels 12 are all connected with the accommodation space, wherein the combustible agent in the present invention adopts liquid oxygen, and the combustible agent can be methane, Liquid hydrogen,
  • the "low temperature engine” in the present invention means that the atomizing vaporization chamber 16 is an engine suitable for low temperature propellants, and the liquid temperature range of liquid oxygen/liquid methane is 90K-110K.
  • the present invention further includes a first valve 4 and a second valve 5 , the first valve 4 is installed on the flammable agent flow passage 11 , and the second valve 5 is installed on the flammable agent flow passage 11 12, wherein the first valve 4 and the second valve 5 are preferably installed on the injector body 1 in an embedded manner, wherein the oxidant flow channel 11 and the combustible agent flow channel 12 are provided with liquid accumulation chambers.
  • the oxidant flow channel 11 is connected to the accommodating space through a self-percussion type primary nozzle 21 , wherein the self-percussion type primary nozzle 21 includes a plurality of first injection holes, every two One nozzle hole forms a self-percussion pair, the number of the self-percussion pairs is 2-8, the aspect ratio of the first nozzle hole is 2-4, and the impact angle of the self-percussion pair in the self-percussion pair is 60 ⁇ 90°, the impact angle is the angle between the injection directions of the two first nozzle holes in the self-percussion pair.
  • a plurality of self-percussion primary nozzles 21 are arranged along the circumferential direction of the atomization vaporization chamber 16, which greatly improves the vaporization efficiency of liquid oxygen.
  • an inner convex cathode 14 is provided on the inner wall of the injector body 1 , and the inner convex cathode 14 divides the accommodation space into an atomization vaporization chamber 16 and a combustion chamber 25 .
  • One end of the electrode 15 along the extending direction is installed in the convex cathode 14 with a gap, wherein the gap forms an annular secondary nozzle 22 , and the annular secondary nozzle 22 communicates with the atomization vaporization chamber 16 and the combustion chamber 25 .
  • a first sub-flow channel and a second sub-flow channel extend from the combustible agent flow channel 12 , and the first sub-flow channel is connected with the accommodating space through the inclined core nozzle 23 , and the The second sub-flow channel is connected to the accommodating space through the swirl nozzle 24 .
  • the inclined core nozzle 23 is connected to the combustion chamber 25 , wherein the orientation of the inclined core nozzle 23 and the orientation of the annular secondary nozzle 22 are arranged at an included angle, and the included angle is 0-90°.
  • the inclined core nozzles 23 and the swirl nozzles 24 are evenly arranged in the circumference of the combustion chamber 25, as shown in FIG.
  • a schematic diagram of the arrangement of the first sub-flow channel and the second sub-flow channel on the combustible agent flow channel 12 , and the inclined core nozzle 23 and the vortex nozzle 24 are correspondingly matched.
  • the inclined core nozzle 23 includes a plurality of second nozzle holes, and the length-diameter ratio of the second nozzle holes is 2 ⁇ 4.
  • the impact angle A of the orifice is 60-90°.
  • the height of the end face of the extended end of the electrode 15 and the impact point along the axis of the electrode 15 is 3-5 mm, and the height of the impact point is represented by Lc. The height arrangement makes the end of the electrode 15 away from the core high-temperature combustion zone, which can effectively reduce the heat transfer of the high-temperature gas to the electrode 15 .
  • the swirl nozzle 24 includes a plurality of third injection holes, preferably 4 to 8, the plurality of the third injection holes are uniformly arranged along the circumference of the combustion chamber 25 and the third injection holes face the combustion chamber In the tangential direction of the 25 circumferential direction, the aspect ratio of the third nozzle hole is 2 to 4, and the swirl nozzle 24 is arranged so that more than 80% of the total flow of the engine fuel is injected tangentially along the inner wall of the thrust chamber 2 to form a sticking wall.
  • the uniform film cooling ensures the thermal structure safety of the thrust chamber 2.
  • a plurality of vortex nozzles 24 rotate clockwise or counterclockwise toward the combustion chamber 25 to form a vortex airflow arrangement, so that the combustion agent It can be fully mixed with the combustion accelerant.
  • the inclined core nozzle 23 is located upstream of the flow of the oxidant in the combustion chamber 25
  • the swirl nozzle 24 is located downstream of the flow of the oxidant in the combustion chamber 25 compared to the inclined core nozzle 23 .
  • the self-percussion primary nozzle 21 is provided with a vaporization baffle 13 in the direction toward the accommodating space, and the vaporization baffle 13 is connected to the injector body 1 .
  • the injector body 1 is provided with a vaporization baffle 13 at the outlet of the self-percussion type primary nozzle 21, and the liquid oxygen enters the atomization and vaporization chamber 16 after passing through the self-percussion type primary nozzle 21.
  • the vaporization area after the liquid oxygen atomization increases significantly.
  • the initial low pressure environment of the atomization vaporization chamber 16 makes the liquid oxygen atomized.
  • the vaporized oxygen gas is more easily broken down and ionized than liquid oxygen when it passes through the annular secondary nozzle 22, which effectively improves the ignition reliability of the engine; during the steady operation of the engine, in the described In the atomization vaporization chamber 16, the hot end of the electrode 15 is used to regenerate the atomized liquid oxygen to vaporize, and the vaporized oxygen gas forms a uniform oxygen film through the annular secondary nozzle 22, so that the downstream combustion is more uniform and stable , which improves the combustion performance of the engine.
  • both sides of the annular secondary nozzle 22 form a high-voltage breakdown circuit of several thousand volts through the electrode 15 and the convex cathode 14, and the annular breakdown gas is controlled by the convex cathode 14.
  • the size of the gap Le during the ignition process of the engine, a high-frequency breakdown voltage (typical frequency value 100-200 Hz, typical voltage value 2000V) is applied through the electrode 15, and the continuous breakdown flow through the annular secondary oxygen nozzle 22 is continuous.
  • the gas and oxygen form a pure oxygen plasma flow and enter the combustion chamber 25, which is the mixing area of the core combustible agent and the accelerant, so as to achieve reliable ignition of the engine.
  • this annular breakdown air gap structure is conducive to the formation of random multi-point breakthroughs, which almost simultaneously generate multi-point breakthroughs.
  • a pure oxygen plasma flow can be used to further improve the ignition reliability; during the steady-state operation of the engine, when the gas and oxygen are accelerated through the annular secondary nozzle 22, the side surface of the electrode 15 can be cooled, and the side of the electrode 15 can be cooled.
  • the end face of the electrode 15 forms a pure oxygen vortex area for cooling, so the end of the electrode 15 is always surrounded by pure oxygen during the steady state operation of the engine.
  • the electrode can be effectively prevented from 15's tip overheated and ablated.
  • the invention solves the problems existing in the prior art, such as poor ignition reliability and thermal structure safety of the low-thrust low-temperature engine.
  • the invention is suitable for low-thrust liquid oxygen and methane attitude control engines, and is also suitable for low-thrust rocket engines combined with low-temperature dual-component non-self-ignition propellants such as liquid oxygen/liquid hydrogen or liquid oxygen/kerosene.

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Abstract

一种适用于空间装置的低温发动机，包括喷注器本体(1)、推力室(2)、火花塞(3)，喷注器本体(1)中设置有容纳空间，火花塞(3)安装在喷注器本体(1)的一侧且火花塞(3)上所具有的电极(15)延伸到容纳空间中，推力室(2)安装在喷注器本体(1)的另一侧且与容纳空间连通，喷注器本体(1)上设置有助燃剂流道(11)以及可燃剂流道(12)，助燃剂流道(11)、可燃剂流道(12)都与容纳空间连接。

Description

适用于空间装置的低温发动机 技术领域
本发明涉及火箭发动机技术领域，具体地，涉及一种适用于空间装置的低温发动机。
背景技术
高性能、无毒的低温化学推进技术已经成为液体火箭发动机的主流发展方向，基于液氧/甲烷、液氧/液氢等低温推进剂组合的火箭发动机及其推进系统，因具有较高的综合性能(比冲性能、可重复使用性能、操作维护性和空间长期贮存等)，在运载火箭、空间飞行器、星表基地建设等领域存在广泛的应用前景。然而，不同于采用四氧化二氮/肼类燃料组合的常规发动机，液氧/甲烷等低温双组元非自燃推进剂组合的发动机需要专门的点火结构才能实现发动机的工作，因此点火技术成为液氧/甲烷发动机可靠工作的核心关键技术之一。尤其是空间应用的小推力低温姿态控制发动机，几万次脉冲点火工作以及长寿命使用等特殊需求对发动机的点火可靠性和热结构安全性等提出了更高的要求。
专利文献CN108321678B公开了一种预燃室点火结构、系统及其工作方法，包括伸入预燃室内部的中心电极，在其端部配置接地电极，所述中心电极及接地电极之间形成位于预燃室内部的火花塞间隙，但该设计中电极容易被烧蚀，降低了设备的可靠性。
发明内容
针对现有技术中的缺陷，本发明的目的是提供一种适用于空间装置的低温发动机。
根据本发明提供的一种适用于空间装置的低温发动机，包括喷注器本体、推力室、火花塞；
所述喷注器本体中设置有容纳空间，所述火花塞安装在喷注器本体的一侧且火花塞上所具有的电极延伸到容纳空间中；
所述推力室安装在喷注器本体的另一侧且与容纳空间连通；
所述喷注器本体上设置有助燃剂流道以及可燃剂流道，所述助燃剂流道、可燃剂流道都与容纳空间连接。
优选地，还包括第一阀门以及第二阀门；
所述第一阀门安装在助燃剂流道上，所述第二阀门安装在可燃剂流道上。
优选地，所述助燃剂流道通过自击式一次喷嘴与容纳空间连接。
优选地，所述自击式一次喷嘴包括多个第一喷孔，每两个第一喷孔组成一个自击对，所述自击对的数量为2～8个，所述第一喷孔的长径比为2～4，在自击对中的自击对撞击角为60～90°。
优选地，所述喷注器本体的内壁上设置有内凸阴极，所述内凸阴极将所述容纳空间分割为雾化汽化室以及燃烧室，所述电极沿延伸方向的一端间隙安装在内凸阴极中，其中所述间隙形成环形二次喷嘴，所述环形二次喷嘴连通雾化汽化室和燃烧室。
优选地，所述可燃剂流道上延伸出第一子流道以及第二子流道，所述第一子流道通过倾斜核心喷嘴与容纳空间连接，所述第二子流道通过涡流喷嘴与容纳空间连接。
优选地，所述自击式一次喷嘴朝向容纳空间的方向上设置有汽化隔板，所述汽化隔板连接所述喷注器本体。
优选地，所述倾斜核心喷嘴连接燃烧室，其中所述倾斜核心喷嘴的朝向与环形二次喷嘴的朝向呈夹角布置，所述夹角为0～90°。
优选地，所述倾斜核心喷嘴包括多个第二喷孔，第二喷孔的长径比为2～4，每相邻的两个第二喷孔的撞击角为60～90°，所述电极延伸端的端面与撞击点沿电极的轴线方向的高度为3～5mm。
优选地，所述涡流喷嘴包括多个第三喷孔，多个所述第三喷孔沿燃烧室沿周向的均匀布置且所述第三喷孔朝向燃烧室周向的切线方向，所述第三喷孔长径比为2～4。
与现有技术相比，本发明具有如下的有益效果：
1、本发明采用倾斜核心喷嘴和涡流喷嘴相结合的结构设计，通过分级燃烧独立控制点火核心区混合比和发动机总混合比，确保发动机点火工作的可靠性和热结构安全性，结合电火花点火次数不受限制的优势，能够满足空间应用的姿态控制发动 机几万次脉冲点火工作以及长寿命使用等特殊要求。
2、本发明采用所述的自击式一次喷嘴和雾化汽化室的结构设计，保障液氧经一次喷嘴雾化后能充分蒸发汽化完全变成气氧，然后通过所述的环形二次喷嘴进入点火核心区，点火过程中易被击穿电离进而提高发动机点火可靠性，稳态工作过程中能形成均匀氧气膜使得下游燃烧更加稳定。
3、本发明采用环形二次喷嘴和倾斜核心甲烷喷嘴的结构设计，点火过程中能够持续产生多条纯氧等离子体流进一步提高了发动机点火可靠性，稳态工作过程中能够有效消除电极的烧蚀隐患，有利于延长发动机的寿命。
4、本发明采用所述的涡流喷嘴24结构设计在的推力室内壁形成贴壁的均匀膜冷却，可以保障发动机的热结构安全。
5、本发明适用于小推力的液氧与甲烷姿态控制发动机，同样也适用于液氧/液氢或者液氧/煤油等低温双组元非自燃推进剂组合的小推力火箭发动机，应用范围广泛，提高了设备的实用性。
附图说明
通过阅读参照以下附图对非限制性实施例所作的详细描述，本发明的其它特征、目的和优点将会变得更明显：
图1为本发明实施例的结构示意图。
图2为本发明实施例的工作原理图。
图3为本发明实施例的自击式一次喷嘴结构示意图。
图4为本发明实施例的环形二次喷嘴结构示意图。
图5为本发明实施例的涡流喷嘴结构示意图。
图中示出：
喷注器本体1         可燃剂流道12        自击式一次喷嘴21
推力室2             汽化隔板13          环形二次喷嘴22
火花塞3             内凸阴极14          倾斜核心喷嘴23
第一阀门4           电极15              涡流喷嘴24
第二阀门5           雾化汽化室16        燃烧室25
助燃剂流道11
具体实施方式
下面结合具体实施例对本发明进行详细说明。以下实施例将有助于本领域的技术人员进一步理解本发明，但不以任何形式限制本发明。应当指出的是，对本领域的普通技术人员来说，在不脱离本发明构思的前提下，还可以做出若干变化和改进。这些都属于本发明的保护范围。
本发明提供了一种适用于空间装置的低温发动机，如图1～图5所示，包括喷注器本体1、推力室2、火花塞3，所述喷注器本体1中设置有容纳空间，所述火花塞3安装在喷注器本体1的一侧且火花塞3上所具有的电极15延伸到容纳空间中，所述推力室2安装在喷注器本体1的另一侧且与容纳空间连通，其中，火花塞3优选通过螺纹安装在喷注器本体1上，所述推力室2可以通过焊接、螺接或者法兰形式安装在喷注器本体1上，所述喷注器本体1上设置有助燃剂流道11以及可燃剂流道12，所述助燃剂流道11、可燃剂流道12都与容纳空间连接，其中，本发明中的助燃剂采用液氧，可燃剂可以采用甲烷、液氢、煤油等。
需要说明的是，本发明中的“低温发动机”是指雾化汽化室16为适用于低温推进剂的发动机，液氧/液甲烷的液态温区为90K～110K。
具体地，如图1所示，本发明还包括第一阀门4以及第二阀门5，所述第一阀门4安装在助燃剂流道11上，所述第二阀门5安装在可燃剂流道12上，其中，第一阀门4、第二阀门5优选采用嵌入的方式安装在喷注器本体1上，其中助燃剂流道11、可燃剂流道12上都设置有积液腔。
具体地，如图1所示，所述助燃剂流道11通过自击式一次喷嘴21与容纳空间连接，其中，所述自击式一次喷嘴21包括多个第一喷孔，每两个第一喷孔组成一个自击对，所述自击对的数量为2～8个，所述第一喷孔的长径比为2～4，在自击对中的自击对撞击角为60～90°，撞击角即为自击对中两个第一喷孔喷射方向的夹角，在一个优选例中，所述自击式一次喷嘴21在助燃剂流道11上设置的示意图，通过多个自击式一次喷嘴21沿雾化汽化室16的周向布置，大大提高了液氧的气化效率。
具体地，如图1所示，所述喷注器本体1的内壁上设置有内凸阴极14，所述内凸阴极14将所述容纳空间分割为雾化汽化室16以及燃烧室25，所述电极15沿延伸方向的一端间隙安装在内凸阴极14中，其中所述间隙形成环形二次喷嘴22，所述环形二次喷嘴22连通雾化汽化室16和燃烧室25。
具体地，如图1所示，所述可燃剂流道12上延伸出第一子流道以及第二子流道， 所述第一子流道通过倾斜核心喷嘴23与容纳空间连接，所述第二子流道通过涡流喷嘴24与容纳空间连接。所述倾斜核心喷嘴23连接燃烧室25，其中所述倾斜核心喷嘴23的朝向与环形二次喷嘴22的朝向呈夹角布置，所述夹角为0～90°。其中倾斜核心喷嘴23、涡流喷嘴24分别沿燃烧室25的周向布置有多个，优选采用倾斜核心喷嘴23、涡流喷嘴24分别均匀布置在燃烧室25的周向，如图2所示，为第一子流道以及第二子流道在可燃剂流道12上的布置示意图，相应的的匹配有倾斜核心喷嘴23、涡流喷嘴24。
具体地，如图1所示，在一个优选例中，所述倾斜核心喷嘴23包括多个第二喷孔，第二喷孔的长径比为2～4，每相邻的两个第二喷孔的撞击角A为60～90°，如图2所示，所述电极15延伸端的端面与撞击点沿电极15的轴线方向的高度为3～5mm，撞击点高度用Lc表示，撞击点高度的布置使得所述的电极15的端头离开核心高温燃烧区，可以有效降低高温燃气向所述的电极15的热传输。
具体地，所述涡流喷嘴24包括多个第三喷孔，优选4～8个，多个所述第三喷孔沿燃烧室25沿周向的均匀布置且所述第三喷孔朝向燃烧室25周向的切线方向，所述第三喷孔长径比为2～4，涡流喷嘴24的设置使发动机燃料总流量的80％以上沿所述的推力室2内壁切向喷入形成贴壁的均匀膜冷却，保障所述的推力室2的热结构安全，如图5所示，多个涡流喷嘴24朝向燃烧室25沿顺时针或逆时针方向旋转形成涡流式的气流布置，使燃烧剂和助燃剂能够充分混合。
具体地，所述倾斜核心喷嘴23位于燃烧室25中助燃剂流动的上游，所述涡流喷嘴24相比于倾斜核心喷嘴23位于燃烧室25中助燃剂流动的下游。
具体地，如图1所示，所述自击式一次喷嘴21朝向容纳空间的方向上设置有汽化隔板13，所述汽化隔板13连接所述喷注器本体1。所述的喷注器本体1在所述的自击式一次喷嘴21的出口设置汽化隔板13，液氧通过所述的自击式一次喷嘴21后进入所述的雾化汽化室16，在所述的汽化隔板13的导流作用下，液氧雾化后的汽化蒸发面积显著增大，发动机点火瞬态过程中，所述雾化汽化室16的初始低压环境使雾化的液氧闪蒸汽化，汽化的气氧通过所述的环形二次喷嘴22时因比液氧更容易被击穿电离，有效地提高了发动机的点火可靠性；发动机稳态工作过程中，在所述的雾化汽化室16内利用所述的电极15的热端回热使雾化的液氧汽化，汽化的气氧通过所述的环形二次喷嘴22形成均匀的氧气膜，使下游燃烧更加均匀稳定，提高了发动机的燃烧性能。
进一步地，所述环形二次喷嘴22两侧通过所述的电极15和所述的内凸阴极14构成几千伏的高压击穿回路，并通过所述的内凸阴极14控制环形击穿气隙Le大小，在发动机点火过程中通过所述的电极15施加高频击穿电压(典型频率值100～200Hz，典型电压值2000V)，持续击穿流经所述的环形二次氧喷嘴22的气氧形成纯氧等离子体流进入核心可燃剂与助燃剂的混合区即燃烧室25，以实现发动机可靠点火，同时这种环形击穿气隙结构有利于形成随机多点击穿，几乎同时产生多条纯氧等离子体流，进一步提高点火可靠性；发动机稳态工作过程中，气氧加速通过所述的环形二次喷嘴22时既可以冷却所述的电极15的侧面，还可以在所述的电极15的端面形成纯氧涡流区进行冷却，因此发动机稳态工作期间所述的电极15的端头始终被纯氧包围，结合所述倾斜核心喷嘴23的布置方式，可以有效地防止所述电极15的端头过热和烧蚀。
本发明解决现有技术中存在的小推力低温发动机点火可靠性和热结构安全性不佳等问题，所述发动机能够满足空间环境几万次脉冲点火工作以及长寿命使用等特殊需求。本发明适用于小推力的液氧与甲烷姿态控制发动机，同样也适用于液氧/液氢或者液氧/煤油等低温双组元非自燃推进剂组合的小推力火箭发动机。
在本申请的描述中，需要理解的是，术语“上”、“下”、“前”、“后”、“左”、“右”、“竖直”、“水平”、“顶”、“底”、“内”、“外”等指示的方位或位置关系为基于附图所示的方位或位置关系，仅是为了便于描述本申请和简化描述，而不是指示或暗示所指的装置或元件必须具有特定的方位、以特定的方位构造和操作，因此不能理解为对本申请的限制。
以上对本发明的具体实施例进行了描述。需要理解的是，本发明并不局限于上述特定实施方式，本领域技术人员可以在权利要求的范围内做出各种变化或修改，这并不影响本发明的实质内容。在不冲突的情况下，本申请的实施例和实施例中的特征可以任意相互组合。

Claims (10)

  1. 一种适用于空间装置的低温发动机，其特征在于，包括喷注器本体(1)、推力室(2)、火花塞(3)；
    所述喷注器本体(1)中设置有容纳空间，所述火花塞(3)安装在喷注器本体(1)的一侧且火花塞(3)上所具有的电极(15)延伸到容纳空间中；
    所述推力室(2)安装在喷注器本体(1)的另一侧且与容纳空间连通；
    所述喷注器本体(1)上设置有助燃剂流道(11)以及可燃剂流道(12)，所述助燃剂流道(11)、可燃剂流道(12)都与容纳空间连接。
  2. 根据权利要求1所述的适用于空间装置的低温发动机，其特征在于，还包括第一阀门(4)以及第二阀门(5)；
    所述第一阀门(4)安装在助燃剂流道(11)上，所述第二阀门(5)安装在可燃剂流道(12)上。
  3. 根据权利要求1所述的适用于空间装置的低温发动机，其特征在于，所述助燃剂流道(11)通过自击式一次喷嘴(21)与容纳空间连接。
  4. 根据权利要求3所述的适用于空间装置的低温发动机，其特征在于，所述自击式一次喷嘴(21)包括多个第一喷孔，每两个第一喷孔组成一个自击对，所述自击对的数量为2～8个，所述第一喷孔的长径比为2～4，在自击对中的自击对撞击角为60～90°。
  5. 根据权利要求1所述的适用于空间装置的低温发动机，其特征在于，所述喷注器本体(1)的内壁上设置有内凸阴极(14)，所述内凸阴极(14)将所述容纳空间分割为雾化汽化室(16)以及燃烧室(25)，所述电极(15)沿延伸方向的一端间隙安装在内凸阴极(14)中，其中所述间隙形成环形二次喷嘴(22)，所述环形二次喷嘴(22)连通雾化汽化室(16)和燃烧室(25)。
  6. 根据权利要求5所述的适用于空间装置的低温发动机，其特征在于，所述可燃剂流道(12)上延伸出第一子流道以及第二子流道，所述第一子流道通过倾斜核心喷嘴(23)与容纳空间连接，所述第二子流道通过涡流喷嘴(24)与容纳空间连接。
  7. 根据权利要求5所述的适用于空间装置的低温发动机，其特征在于，所述自击式一次喷嘴(21)朝向容纳空间的方向上设置有汽化隔板(13)，所述汽化隔板 (13)连接所述喷注器本体(1)。
  8. 根据权利要求6所述的适用于空间装置的低温发动机，其特征在于，所述倾斜核心喷嘴(23)连接燃烧室(25)，其中所述倾斜核心喷嘴(23)的朝向与环形二次喷嘴(22)的朝向呈夹角布置，所述夹角为0～90°。
  9. 根据权利要求6所述的适用于空间装置的低温发动机，其特征在于，所述倾斜核心喷嘴(23)包括多个第二喷孔，第二喷孔的长径比为2～4，每相邻的两个第二喷孔的撞击角为60～90°，所述电极(15)延伸端的端面与撞击点沿电极(15)的轴线方向的高度为3～5mm。
  10. 根据权利要求6所述的适用于空间装置的低温发动机，其特征在于，所述涡流喷嘴(24)包括多个第三喷孔，多个所述第三喷孔沿燃烧室(25)沿周向的均匀布置且所述第三喷孔朝向燃烧室(25)周向的切线方向，所述第三喷孔长径比为2～4。
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