CN114810430A - A low-ablation rocket engine nozzle structure and cooling method for active cooling throat lining - Google Patents
A low-ablation rocket engine nozzle structure and cooling method for active cooling throat lining Download PDFInfo
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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
- F02K9/97—Rocket nozzles
- F02K9/972—Fluid cooling arrangements for nozzles
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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
- F02K9/97—Rocket nozzles
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Abstract
Description
技术领域technical field
本发明涉及火箭发动机喷管设计技术领域,主要涉及一种主动冷却喉衬的低烧蚀火箭发动机喷管结构及冷却方法。The invention relates to the technical field of rocket engine nozzle design, in particular to a low ablation rocket engine nozzle structure and a cooling method for actively cooling throat linings.
背景技术Background technique
火箭发动机的喷管是将热能转换成动能的装置,其中喉部的尺寸决定着火箭发动机的工作点,可控制燃烧室的压力和燃气流率,对发动机的性能和工作安全性都有重要影响。喷管喉部在工作时流通的高温高压热流和粒子冲蚀会引起内型面烧蚀,导致型面退移、尺寸变化,喉径的扩大最终降低发动机性能,降低发动机比冲。现有技术中的喷管隔热结构主要采用增加绝热层厚度的方式提高发动机的可靠性,增加了消极质量,不利于提高发动机的质量比。另外专利2020116067168是通过设置腹板结构来引导冷却气进入喷管后部位置,同时通过型面对冷却气进行压缩,但这种技术无法实时控制冷却气的压力,无法提升系统鲁棒性。The nozzle of a rocket engine is a device that converts thermal energy into kinetic energy. The size of the throat determines the working point of the rocket engine, which can control the pressure and gas flow rate of the combustion chamber, which has an important impact on the performance and safety of the engine. . The high temperature and high pressure heat flow and particle erosion in the nozzle throat during operation will cause the internal profile ablation, resulting in profile retreat and size change. The expansion of the throat diameter will eventually reduce the engine performance and reduce the engine specific impulse. The thermal insulation structure of the nozzle in the prior art mainly adopts the method of increasing the thickness of the thermal insulation layer to improve the reliability of the engine, which increases the negative mass, which is not conducive to improving the mass ratio of the engine. In addition, the patent 2020116067168 uses a web structure to guide the cooling gas into the rear of the nozzle, and at the same time compress the cooling gas through the profile, but this technology cannot control the pressure of the cooling gas in real time, and cannot improve the robustness of the system.
发明内容SUMMARY OF THE INVENTION
发明目的:针对上述背景技术中存在的问题,本发明提供了一种可以主动冷却喉衬的低烧蚀火箭发动机喷管结构及冷却方法,采用主动冷却系统对喷管喉部进行冷却,并在喉径内部形成一定厚度的气膜,降低了喉衬部位的温度,从而减缓了喉径的增大。此外,在喉部出现严重烧蚀导致喉径变大时,通过增大冷却气的压力,减小冷却气与热气流形成界面的直径,即保证喉径有效直径,提高的喷管的鲁棒性。同时,冷却气膜对喷管内型面的冷却作用,可以减小喷管绝热层的设计厚度,减小了消极质量。Purpose of the invention: In view of the problems existing in the above-mentioned background technology, the present invention provides a low-ablation rocket engine nozzle structure and cooling method that can actively cool the throat lining. The active cooling system is used to cool the throat of the nozzle. A certain thickness of air film is formed inside the diameter of the throat, which reduces the temperature of the throat lining, thereby slowing down the increase of the throat diameter. In addition, when the throat diameter becomes larger due to severe ablation of the throat, the diameter of the interface formed by the cooling gas and the hot air flow is reduced by increasing the pressure of the cooling gas, that is, the effective diameter of the throat diameter is guaranteed, and the robustness of the nozzle is improved. sex. At the same time, the cooling effect of the cooling gas film on the inner profile of the nozzle can reduce the design thickness of the thermal insulation layer of the nozzle and reduce the negative mass.
技术方案:为实现上述目的,本发明采用的技术方案为:Technical scheme: In order to realize the above-mentioned purpose, the technical scheme adopted in the present invention is:
一种主动冷却喉衬的低烧蚀火箭发动机喷管结构,包括喷管本体、主动冷却系统和控制系统;所述喷管本体包括沿轴线方向一体成型的裙部和喉衬;所述喉衬内部绕轴线方向设有环形共轨空腔;喉衬内壁面设有若干与共轨空腔相连的进气孔;所述主动冷却系统包括电动空气泵、低压进气管和高压出气管;所述电动空气泵通过低压进气管抽取外部低压冷却空气并进行压缩,通过与共轨空腔相连的高压出气管,将高压气体输入共轨空腔;所述控制系统用于控制主动冷却系统工作。A low-ablation rocket engine nozzle structure with active cooling throat lining, comprising a nozzle body, an active cooling system and a control system; the nozzle body includes a skirt and a throat liner integrally formed along an axis direction; the throat liner is inside An annular common rail cavity is arranged around the axis; the inner wall of the throat lining is provided with a number of air intake holes connected to the common rail cavity; the active cooling system includes an electric air pump, a low-pressure air inlet pipe and a high-pressure air outlet pipe; the electric air The pump extracts and compresses the external low-pressure cooling air through the low-pressure air inlet pipe, and inputs the high-pressure gas into the common rail cavity through the high-pressure air outlet pipe connected to the common rail cavity; the control system is used to control the active cooling system to work.
进一步地,所述电动空气泵通过支架安装在喷管本体上;喷管本体下部还安装有动力电池;所述动力电池通过导线与电动空气泵相连。Further, the electric air pump is installed on the nozzle body through a bracket; a power battery is also installed at the lower part of the nozzle body; the power battery is connected with the electric air pump through a wire.
进一步地,所述控制系统在控制火箭点火后,进一步控制动力电池连通电动空气泵,使电动空气泵工作。Further, after controlling the ignition of the rocket, the control system further controls the power battery to communicate with the electric air pump to make the electric air pump work.
进一步地,从所述进气孔均匀喷出冷气流,工作时与高温高压高速气流在喉衬位置的交界面直径即为等效喉径;根据实际喷管烧蚀情况,通过电动空气泵增加或减小压力,可以保证等效喉径处于有效直径范围。Further, the cold airflow is uniformly ejected from the air inlet, and the diameter of the interface between the high-temperature, high-pressure and high-speed airflow at the throat lining position during operation is the equivalent throat diameter; according to the actual nozzle ablation, the electric air pump increases Or reduce the pressure to ensure that the equivalent throat diameter is within the effective diameter range.
一种采用上述主动冷却喉衬的低烧蚀火箭发动机喷管结构的主动冷却方法,包括以下步骤:An active cooling method for a low-ablation rocket engine nozzle structure using the above-mentioned active cooling throat lining, comprising the following steps:
步骤S1、控制系统控制火箭点火后,高温高压的高速气体流过喉道;此时控制系统控制动力电池和电动空气泵连通,电动空气泵开始工作;Step S1, after the control system controls the ignition of the rocket, the high-temperature and high-speed gas flows through the throat; at this time, the control system controls the power battery to communicate with the electric air pump, and the electric air pump starts to work;
步骤S2、电动空气泵通过低压进气管吸入空气,进行压缩后,将压缩气体通过高压出气管送入共轨空腔,环形共轨空腔内的高压空气对喉衬进行冷却;高压空气通过与共轨空腔相连的进气孔流入喉衬内壁面,并沿高速气体流向扩散至裙部;沿整个喷管本体内型面形成一层冷却气膜,对喷管本体进一步进行冷却,有效防止高温高压气体对喷管本体内型面的烧蚀。Step S2, the electric air pump inhales air through the low-pressure air intake pipe, and after compression, sends the compressed gas into the common rail cavity through the high-pressure air outlet pipe, and the high-pressure air in the annular common rail cavity cools the throat lining; The air intake holes connected to the rail cavity flow into the inner wall of the throat lining, and spread to the skirt along the high-speed gas flow direction; a cooling gas film is formed along the inner surface of the entire nozzle body, which further cools the nozzle body and effectively prevents high temperature. The ablation of the inner surface of the nozzle body by the high pressure gas.
有益效果:Beneficial effects:
(1)本发明提供的有主动冷却系统的火箭发动机喷管结构,通过设置环状共轨空腔,主动冷却系统能够对喉衬进行可控压力的空气冷却,降低喉衬位置温度。(1) The rocket engine nozzle structure with the active cooling system provided by the present invention, by setting the annular common rail cavity, the active cooling system can perform air cooling on the throat liner with a controllable pressure and reduce the temperature of the throat liner position.
(2)本发明提供的有主动冷却系统的火箭发动机喷管结构,喉部的冷却气为空气且气流方向与高温高压高速气流垂直,冷却气流有效降低了粒子速度,减缓了粒子对喉部的冲击,同时空气进一步氧化粒子,更小的粒子对喉部冲蚀作用更小。(2) The rocket engine nozzle structure with active cooling system provided by the present invention, the cooling gas of the throat is air and the airflow direction is perpendicular to the high-temperature, high-pressure and high-speed airflow, and the cooling airflow effectively reduces the particle velocity and slows down the particle to the throat. impact, while the air further oxidizes the particles, smaller particles have less effect on the throat erosion.
(3)本发明在喉部出现严重烧蚀导致喉径变大时,通过控制电动空气泵增大冷却气流的压力,减小冷却气与热气流形成界面的直径,保证等效喉径处于有效直径范围。(3) When the throat diameter becomes larger due to severe ablation at the throat, the present invention increases the pressure of the cooling airflow by controlling the electric air pump, reduces the diameter of the interface formed by the cooling gas and the hot airflow, and ensures that the equivalent throat diameter is effectively diameter range.
综合可见,本发明提供的主动冷却喉衬的低烧蚀火箭发动机喷管结构可以减缓喉径的增大,同时通过控制电动空气泵实现对烧蚀喉径的弥补,提高喷管鲁棒性,可以更持久的保证发动机最佳性能,提高火箭发动机比冲。It can be seen comprehensively that the low ablation rocket engine nozzle structure of the active cooling throat lining provided by the present invention can slow down the increase of the throat diameter, and at the same time, by controlling the electric air pump, the ablation throat diameter can be compensated, and the robustness of the nozzle can be improved. It can ensure the best performance of the engine for a longer time and improve the specific impulse of the rocket engine.
(4)在喷管喉衬形成的气膜可以对喉衬和喷管裙部有一定的降温作用,提高了可靠性,同等工况应用下,可以采用更薄的喷管绝热层,减小了喷管的消极质量。(4) The gas film formed in the nozzle throat lining can have a certain cooling effect on the throat lining and nozzle skirt, which improves the reliability. Under the same working conditions, a thinner nozzle insulation layer can be used to reduce the negative quality of the nozzle.
附图说明Description of drawings
图1是本发明提供的主动冷却喉衬的低烧蚀火箭发动机喷管结构示意图;Fig. 1 is the low ablation rocket engine nozzle structure schematic diagram of the active cooling throat lining provided by the present invention;
图2是本发明提供的主动冷却喉衬的低烧蚀火箭发动机喷管结构剖视图;2 is a cross-sectional view of a low-ablation rocket engine nozzle structure of an active cooling throat lining provided by the present invention;
图3是本发明提供的主动冷却喉衬的低烧蚀火箭发动机喷管控制原理示意图;3 is a schematic diagram of the control principle of the low ablation rocket engine nozzle of the active cooling throat lining provided by the present invention;
图4是本发明提供的主动冷却喉衬的低烧蚀火箭发动机喷管结构工作示意图。FIG. 4 is a schematic working diagram of the low-ablation rocket engine nozzle structure of the active cooling throat liner provided by the present invention.
附图标记说明:Description of reference numbers:
1-喷管本体;1.1-裙部;1.2-喉衬;1.3-共轨空腔;1.4-进气孔;2-主动冷却系统;2.1-电动空气泵;2.2-低压进气管;2.3-高压出气管;2.4-支架;3-动力电池;3.1-导线;4-控制系统;5-火箭点火;6-冷却气膜;6.1-冷却气方向;7-高速高温高压气流;7.1-热气流粒子气流方向。1-nozzle body; 1.1-skirt; 1.2-throat lining; 1.3-common rail cavity; 1.4-intake hole; 2-active cooling system; 2.1-electric air pump; 2.2-low pressure intake pipe; 2.3-high pressure Outlet pipe; 2.4-bracket; 3-power battery; 3.1-wire; 4-control system; 5-rocket ignition; 6-cooling film; 6.1-cooling gas direction; 7-high-speed, high-temperature and high-pressure airflow; 7.1-hot airflow particle Airflow direction.
具体实施方式Detailed ways
下面结合附图提供一份具体实施例,对本发明作更进一步的说明。A specific embodiment is provided below in conjunction with the accompanying drawings to further illustrate the present invention.
本发明提供的可以主动冷却喉衬的低烧蚀火箭发动机喷管结构如图1所示,包括喷管本体1、主动冷却系统2、动力电池3和控制系统4。喷管本体1包括沿轴线方向一体成型的圆台状裙部1.1和圆柱状喉衬1.2。喉衬1.2内部绕轴线方向设有圆环形共轨空腔1.3。喉衬1.2内壁面设有若干与共轨空腔1.3相连的进气孔1.4。特别地,本实施例中裙部和喉衬部均采用耐烧蚀结构制成。The structure of the low ablation rocket engine nozzle provided by the present invention that can actively cool the throat lining is shown in FIG. The
主动冷却系统2包括电动空气泵2.1、低压进气管2.2、高压出气管2.3和支架2.4。电动空气泵2.1设置于喷管本体1上端,通过支架2.4安装固定。电动空气泵2.1外部伸出有低压进气管2.2,通过低压进气管2.2抽取外部低压冷却空气并进行压缩,通过与共轨空腔1.3相连的高压出气管2.3,将高压气体输入共轨空腔1.3,如图2所示。The
电动空气泵控制进气孔均匀喷出冷气流,工作时与高温高压高速气流在喉衬位置的交界面直径即为等效喉径;根据实际喷管烧蚀情况,通过电动空气泵增加或减小压力,可以保证等效喉径处于有效直径范围。The electric air pump controls the air inlet to spray out cold air evenly, and the diameter of the interface with the high temperature, high pressure and high speed air flow at the throat lining position during operation is the equivalent throat diameter; according to the actual nozzle ablation, the electric air pump can increase or decrease it. Small pressure can ensure that the equivalent throat diameter is within the effective diameter range.
动力电池3固定安装在喷管本体1下部,通过导线3.1将电动空气泵2.1连接在一起,为电动空气泵2.1供电。The
本发明提供的主动冷却喉衬的低烧蚀火箭发动机喷管结构控制方法如图3所示,控制系统4控制火箭点火5工作后,控制动力电池3与电动空气泵2.1电路连通,进而控制电动空气泵2.1正常工作。电动空气泵2.1将从低压进气管2.2吸入的空气进行加压,通过高压出气管2.3将高压空气送入共轨空腔1.3,高压共轨空腔1.3中的高压空气对喉衬1.2进行冷却,同时通过喉衬1.2上的进气口1.4将高压空气送入喷管本体1中,在喷管本体1内壁面高速高温高压气流7的作用下,在喷管本体1内型面形成一层冷却气膜6,对喉衬1.2以及喷管本体1内壁面进行冷却,减小喉径因高温烧蚀的扩大速度,具体如图4所示。The low-ablation rocket engine nozzle structure control method of the active cooling throat lining provided by the present invention is shown in FIG. 3 . After the control system 4 controls the rocket ignition 5 to work, it controls the
其中,喉衬1.2的冷却气为空气且冷却气方向6.1是垂直于热气流粒子方向7.1的,气流降低了粒子速度,减缓了粒子对喉衬1.2的冲击,同时空气进一步氧化粒子,更小的粒子对喉衬1.2冲蚀作用更小。Among them, the cooling air of the throat lining 1.2 is air and the cooling air direction 6.1 is perpendicular to the hot air particle direction 7.1. The airflow reduces the particle velocity and slows down the impact of the particles on the throat lining 1.2. At the same time, the air further oxidizes the particles, and the smaller the The particles have less erosion effect on the throat lining 1.2.
此外,当喉衬1.2出现严重烧蚀导致喉径变大时,控制系统4通过控制电动空气泵1.2提高冷却气的压力,减小冷却气与热气流形成界面的直径,保证等效喉径一直维持在有效喉径范围内,可以减缓喉径的增大,同时通过控制电动空气泵实现对烧蚀喉径的弥补,显著提高喷管鲁棒性,有效保证发动机更长久地处于最佳性能状态,提高了火箭发发动机的比冲。冷却气膜对喷管内型面的冷却作用,可以减小喷管绝热层的设计厚度,减小了消极质量。In addition, when the throat lining 1.2 is severely ablated and the throat diameter becomes larger, the control system 4 increases the pressure of the cooling gas by controlling the electric air pump 1.2, reduces the diameter of the interface formed by the cooling gas and the hot air flow, and ensures that the equivalent throat diameter is always Maintaining within the effective throat diameter range can slow down the increase in throat diameter. At the same time, the ablation throat diameter can be compensated by controlling the electric air pump, which can significantly improve the robustness of the nozzle and effectively ensure that the engine is in the best performance state for a longer time. , which increases the specific impulse of the rocket engine. The cooling effect of the cooling gas film on the inner profile of the nozzle can reduce the design thickness of the thermal insulation layer of the nozzle and reduce the negative mass.
以上所述仅是本发明的优选实施方式,应当指出:对于本技术领域的普通技术人员来说,在不脱离本发明原理的前提下,还可以做出若干改进和润饰,这些改进和润饰也应视为本发明的保护范围。The above is only the preferred embodiment of the present invention, it should be pointed out that: for those skilled in the art, without departing from the principle of the present invention, several improvements and modifications can also be made, and these improvements and modifications are also It should be regarded as the protection scope of the present invention.
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CN115434828A (en) * | 2022-10-17 | 2022-12-06 | 西安探火航天技术有限公司 | Variable expansion ratio rocket engine jet pipe |
CN115434829A (en) * | 2022-10-17 | 2022-12-06 | 西安探火航天技术有限公司 | Reusable rocket engine nozzle assembly with variable expansion ratio |
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CN115434828A (en) * | 2022-10-17 | 2022-12-06 | 西安探火航天技术有限公司 | Variable expansion ratio rocket engine jet pipe |
CN115434829A (en) * | 2022-10-17 | 2022-12-06 | 西安探火航天技术有限公司 | Reusable rocket engine nozzle assembly with variable expansion ratio |
CN115434828B (en) * | 2022-10-17 | 2023-08-29 | 西安探火航天技术有限公司 | Rocket engine spray pipe with variable expansion ratio |
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