WO2016127932A1 - 一种超高速飞行器的热防护与减阻方法和系统 - Google Patents

一种超高速飞行器的热防护与减阻方法和系统 Download PDF

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Publication number
WO2016127932A1
WO2016127932A1 PCT/CN2016/073710 CN2016073710W WO2016127932A1 WO 2016127932 A1 WO2016127932 A1 WO 2016127932A1 CN 2016073710 W CN2016073710 W CN 2016073710W WO 2016127932 A1 WO2016127932 A1 WO 2016127932A1
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Prior art keywords
cold source
speed aircraft
high speed
cavity
thermal protection
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PCT/CN2016/073710
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English (en)
French (fr)
Inventor
张文武
向树红
郭春海
童靖宇
张天润
杨旸
宋涛
Original Assignee
中国科学院宁波材料技术与工程研究所
北京卫星环境工程研究所
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Application filed by 中国科学院宁波材料技术与工程研究所, 北京卫星环境工程研究所 filed Critical 中国科学院宁波材料技术与工程研究所
Priority to EP16748733.9A priority Critical patent/EP3257756B1/en
Priority to US15/550,777 priority patent/US20180057191A1/en
Priority to RU2017131642A priority patent/RU2671064C1/ru
Priority to JP2017560861A priority patent/JP2018505099A/ja
Publication of WO2016127932A1 publication Critical patent/WO2016127932A1/zh

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    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
    • B64C1/38—Constructions adapted to reduce effects of aerodynamic or other external heating
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D13/00—Arrangements or adaptations of air-treatment apparatus for aircraft crew or passengers, or freight space, or structural parts of the aircraft
    • B64D13/006—Arrangements or adaptations of air-treatment apparatus for aircraft crew or passengers, or freight space, or structural parts of the aircraft the air being used to cool structural parts of the aircraft
    • 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
    • B64G1/52—Protection, safety or emergency devices; Survival aids
    • B64G1/58—Thermal protection, e.g. heat shields
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F42—AMMUNITION; BLASTING
    • F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
    • F42B10/00—Means for influencing, e.g. improving, the aerodynamic properties of projectiles or missiles; Arrangements on projectiles or missiles for stabilising, steering, range-reducing, range-increasing or fall-retarding
    • F42B10/32—Range-reducing or range-increasing arrangements; Fall-retarding means
    • F42B10/38—Range-increasing arrangements
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F42—AMMUNITION; BLASTING
    • F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
    • F42B10/00—Means for influencing, e.g. improving, the aerodynamic properties of projectiles or missiles; Arrangements on projectiles or missiles for stabilising, steering, range-reducing, range-increasing or fall-retarding
    • F42B10/32—Range-reducing or range-increasing arrangements; Fall-retarding means
    • F42B10/38—Range-increasing arrangements
    • F42B10/40—Range-increasing arrangements with combustion of a slow-burning charge, e.g. fumers, base-bleed projectiles
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F42—AMMUNITION; BLASTING
    • F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
    • F42B10/00—Means for influencing, e.g. improving, the aerodynamic properties of projectiles or missiles; Arrangements on projectiles or missiles for stabilising, steering, range-reducing, range-increasing or fall-retarding
    • F42B10/32—Range-reducing or range-increasing arrangements; Fall-retarding means
    • F42B10/38—Range-increasing arrangements
    • F42B10/42—Streamlined projectiles
    • F42B10/46—Streamlined nose cones; Windshields; Radomes
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F42—AMMUNITION; BLASTING
    • F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
    • F42B15/00—Self-propelled projectiles or missiles, e.g. rockets; Guided missiles
    • F42B15/34—Protection against overheating or radiation, e.g. heat shields; Additional cooling arrangements
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C2230/00—Boundary layer controls
    • B64C2230/16—Boundary layer controls by blowing other fluids over the surface than air, e.g. He, H, O2 or exhaust gases
    • 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
    • B64G1/46—Arrangements or adaptations of devices for control of environment or living conditions
    • B64G1/50—Arrangements or adaptations of devices for control of environment or living conditions for temperature control
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F15—FLUID-PRESSURE ACTUATORS; HYDRAULICS OR PNEUMATICS IN GENERAL
    • F15D—FLUID DYNAMICS, i.e. METHODS OR MEANS FOR INFLUENCING THE FLOW OF GASES OR LIQUIDS
    • F15D1/00—Influencing flow of fluids
    • F15D1/002—Influencing flow of fluids by influencing the boundary layer
    • F15D1/0065—Influencing flow of fluids by influencing the boundary layer using active means, e.g. supplying external energy or injecting fluid
    • F15D1/008—Influencing flow of fluids by influencing the boundary layer using active means, e.g. supplying external energy or injecting fluid comprising fluid injection or suction means
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/10—Drag reduction
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/50—On board measures aiming to increase energy efficiency

Definitions

  • the invention belongs to the technical field of ultra-high speed aircraft, and particularly relates to a method and system for thermal protection and drag reduction of a super high speed aircraft.
  • Ultra-high-speed aircraft refers to aircraft with a flight rate of Mach 5 or higher, including rockets, missiles, spacecraft, space shuttles, and space planes.
  • the super-high-speed aircraft faces the air viscosity problem when entering and leaving the atmosphere, and needs a lot of energy to overcome the aerodynamic drag;
  • the ultra-high-speed aircraft will face severe during the flight.
  • the aerodynamic shock wave friction heat generation phenomenon, the occurrence of thermal barriers, in severe cases, can generate thousands of degrees of high temperature plasma, resulting in communication interruption, this stage is a high-risk period of the aircraft.
  • thermal protection of ultra-high-speed aircraft
  • the research at home and abroad is mainly divided into six types of thermal protection: heat sinking heat protection; radiation heat protection; ablation heat protection; sweat cooling and heat prevention; surface heat insulation; heat pipe cooling .
  • ablative heat protection and sweat cooling and heat protection have better heat prevention effects, and are suitable for aircraft with serious thermal phenomena (for example, friction generating heat generating plasma).
  • both of these methods are difficult to perform thermal protection for a long time, resulting in expensive aircraft requiring frequent overhaul or facing the situation of being eliminated several times.
  • it is difficult to control the internal temperature of the aircraft by these two methods but the continuous increase of the internal temperature of the aircraft will seriously endanger the safety of the portable system.
  • the structure of the relevant protection system is complex and prone to unexpected failures.
  • the drag reduction technology capable of effectively reducing the air viscosity resistance and the heat protection technology capable of effectively mitigating and overcoming the thermal barrier and avoiding excessive thermal erosion are urgently needed for research on ultra-high speed aircraft.
  • the present invention provides a method for thermal protection and drag reduction of a high-speed aircraft, in particular, a method for thermal protection and drag reduction of a super-high-speed aircraft, which can avoid excessive thermal erosion of a super-high-speed aircraft and can reduce super Air resistance of high speed aircraft.
  • the technical scheme adopted by the invention is: a method for thermal protection and drag reduction of a super high speed aircraft, a cold source is arranged inside the cavity of the ultra high speed aircraft, and a plurality of micro holes are arranged in the wall surface of the ultra high speed aircraft, under the driving force
  • the cold source is emitted from the micropores in a high-pressure gas form, and a gas film is formed on the outer surface of the cavity.
  • the arrangement position of the micropores is not limited, and is preferably disposed at a portion of a nose cone (or a head) and/or a tail of a super high speed aircraft cavity.
  • the distribution of the micropores in the wall surface of the ultrahigh-speed aircraft cavity is not limited.
  • the micropores are regularly distributed on the cavity wall surface of the ultrahigh speed aircraft, and further preferably, the micro holes are in accordance with a pneumatic characteristic rule. Distributed on the cavity wall of a super high speed aircraft.
  • the pore shape is not limited, and may be a straight hole or a shaped hole, and the microporous cross section may be a regular shape (for example, a circle or the like) or an irregular shape (for example, a butterfly shape, a dome shape, or the like).
  • Numerical simulations show that when the micropores are shaped holes
  • the cold source which is favorable for spraying forms a gas film on the surface of the cavity, and can achieve superior cooling effect with less micropores, thereby improving the cooling effect of the gas film while better ensuring the structural strength.
  • the diameter of the micropores is not limited.
  • the diameter of the micropores is designed to take into account the structural strength of the hypervelocity aircraft cavity and the coverage of the cavity wall by the cold source.
  • the micropores are circular straight holes having a diameter of 0.05 mm to 2.0 mm.
  • the source of the cold source is not limited, and may be a cooling source such as liquid nitrogen, dry ice, compressed air, or other cooling substances generated by a chemical reaction.
  • the driving force is not limited, and includes pressure, elasticity, electric power, and the like.
  • the ultra high speed aircraft has a flight rate of more than 5 Mach.
  • the ultra-high speed aircraft includes rockets, missiles, spaceships, space shuttles, space planes, and the like.
  • the cavity material of the ultra-high speed aircraft is not limited, and includes a carbon-carbon composite material resistant to high temperature corrosion, a composite material of carbon and silicon carbide, and the like.
  • the method of the present invention is applicable to high speed aircraft, especially ultra high speed aircraft.
  • the invention forms a low temperature air film on the surface of the ultra high speed aircraft cavity, and has the following beneficial effects:
  • the low-temperature gas film is located on the surface of the ultra-high-speed aircraft cavity, and the external gas interacts with the gas film, thereby effectively avoiding direct friction between the external gas and the ultra-high-speed aircraft to generate a large amount of heat; meanwhile, the external gas firstly contacts the gas film Layer friction effectively reduces the gas viscosity between the ultra-high-speed aircraft and the outside air, and the reduction of the gas viscosity is beneficial to reduce the surface temperature of the ultra-high-speed aircraft;
  • the low-temperature gas film is formed by cold injection from the inside of the ultra-high-speed aircraft cavity, in which the cold source takes away a large amount of heat inside the high-speed aircraft cavity, so the method can effectively control the ultra-high speed
  • the internal temperature of the aircraft effectively avoids the hazards caused by the continuous increase of the internal temperature of the aircraft;
  • the method of the invention can not only protect the ultra-high speed aircraft from thermal protection, but also effectively reduce the sticking resistance of the high-speed aircraft and the outside air, thereby improving the energy efficiency and the limit speed of the ultra-high speed aircraft, and can reduce or avoid the thermal barrier phenomenon. It reduces the ablation of the material of the thermal protection layer, improves the safety of the ultra-high speed aircraft, prolongs the service life, and has a good application prospect.
  • the invention also provides a drag reduction and thermal protection system for a high speed aircraft, in particular a drag reduction and thermal protection system for a super high speed aircraft, comprising a cold source disposed inside the sealed cavity of the ultra high speed aircraft, and a cold source driving device that converts a cold source into a high pressure gas and emits it;
  • At least part of the wall surface of the cavity wall of the ultra high speed aircraft is a sandwich structure, the sandwich structure comprises a transition layer through which a cold source gas can pass, and an outer surface layer on the surface of the transition layer, the outer surface layer is provided with a plurality of micropores For connecting the transition layer to the outside of the cavity;
  • the cold source driving device comprises a cold source reservoir, an air pump and a buffer; the air pump is connected to the cold source reservoir; the buffer comprises a buffer inlet and a buffer outlet, the buffer inlet is connected to the cold source reservoir phase, and the buffer outlet Connecting a transition layer of the wall surface of the cavity, and providing a sealing valve at a communication portion between the buffer outlet and the transition layer;
  • the air pump supplies compressed air to the cold source reservoir, and the cold source enters the buffer under the action of air pressure.
  • the gas is vaporized, the sealing valve is opened, the gas is sprayed from the buffer outlet into the transition layer of the wall surface of the cavity, and then the cavity is ejected from the micropores of the outer surface layer to form a gas film.
  • the transition layer is used to pass the cold source gas to the outer surface layer, and the transition layer may be a hollow layer, or may be another dielectric layer through which a cold source gas can pass.
  • the number of outlets of the buffer is greater than or equal to two, each outlet is connected to the transition layer of the wall surface of the cavity, and the communication portion is provided with a sealing valve.
  • the cold source driving device further includes a flow splitter including at least one inlet and two or more outlets, the splitter inlet and The buffer outlets are connected, each of the splitter outlets communicates with the transition layer of the wall surface of the cavity, and a sealing valve is arranged at a communication portion between each of the splitter outlets and the transition layer; the cold source is vaporized and enters the splitter through the inlet of the splitter, and the flow is divided.
  • the road gas is sprayed into the transition layer of the cavity wall from the outlet of each shunt, and finally the cavity is ejected from the micropores of the outer surface layer to form a gas film.
  • an electric valve and a one-way valve are disposed between the air pump and the cold source reservoir.
  • the electric valve and the one-way valve are opened, the compressed air enters the cold source reservoir, and the electric valve is adjusted to control the air flow. .
  • a check valve is arranged between the cold source reservoir and the buffer, and in the working state, the electric valve is opened, and the cold source enters the buffer.
  • said cold source drive further comprises a temperature sensor for monitoring the temperature of the cold source in the buffer.
  • a pressure sensor for detecting the pressure in the cold source reservoir and a safety valve for regulating the gas pressure in the cold source reservoir are provided on the cold source reservoir.
  • the wall surface of the cavity having the sandwich structure is a wall surface of a nose cone portion and/or a tail portion of the cavity.
  • the micropores are regularly distributed over the cavity wall of the ultra high speed aircraft.
  • the micropores are non-circular pores; further preferably, the micropores have a diameter of from 0.05 mm to 2.0 mm.
  • the source of the cold source is not limited, and may be a cooling source such as liquid nitrogen, dry ice, compressed air, or other cooling substances generated by a chemical reaction.
  • the ultra high speed aircraft has a flight rate of more than 5 Mach.
  • the ultra-high speed aircraft includes rockets, missiles, spaceships, space shuttles, space planes, and the like.
  • the cavity material of the ultra-high speed aircraft is not limited, and includes a carbon-carbon composite material resistant to high temperature corrosion, a composite material of carbon and silicon carbide, and the like.
  • the system of the invention can form a low temperature air film on the surface of the ultra high speed aircraft cavity, not only can protect the ultra high speed aircraft, but also can effectively reduce the sticking resistance of the high speed aircraft and the outside air, thereby improving the energy efficiency of the ultra high speed aircraft.
  • the limit speed can reduce or avoid the thermal barrier phenomenon, reduce the ablation of the thermal protection layer material, improve the safety of the ultra-high speed aircraft, prolong the service life, and have a good application prospect.
  • Embodiment 1 is a schematic structural view of a thermal protection and drag reduction system of a super high speed aircraft in Embodiment 1 of the present invention
  • Figure 2 is a perspective view showing the wall surface of the cavity head of Figure 1;
  • Figure 3 is a top plan view of Figure 2;
  • Figure 4 is a schematic view of the structure of Figure 3 along the A-A section;
  • Fig. 5 is an enlarged view of a portion B in Fig. 4.
  • the reference numerals in FIGS. 1-3 are: cold source driving device 100, cold source 200, microhole 300, cold source reservoir 210, air pump 110, electric valve 120, check valve 130, check valve 140, buffer The device 150, the temperature sensor 160, the liquid separator 170, the safety valve 220, the pressure sensor 230, the cavity head wall surface 310, the transition layer 320, and the outer surface layer 330.
  • the ultra-high-speed aircraft includes a sealed cavity
  • the drag reduction and thermal protection system of the ultra-high-speed aircraft includes a cold source 200 disposed inside the sealed cavity of the ultra-high-speed aircraft, and is used for cooling
  • the source 200 is converted to a high pressure gas and injected into the cold source drive unit 100.
  • the wall surface 310 of the head of the ultra high speed aircraft sealing cavity is a sandwich structure.
  • 2 is a schematic perspective view of the wall surface of the cavity head portion
  • FIG. 3 is a schematic plan view of FIG. 2
  • FIG. 4 is a schematic structural view of FIG. 3 along the A-A section
  • FIG. 5 is an enlarged view of a portion B of FIG. As can be seen from FIG. 2 to FIG.
  • the interlayer structure includes a transition layer 320 and an outer surface layer 330 on the surface of the transition layer 320 from the inside of the cavity to the outside of the cavity.
  • the surface layer 330 is provided with a plurality of micro holes 300 for The transition layer 320 is connected to the outside of the cavity.
  • the micro-holes 300 are distributed in a divergent manner on the head wall 310 of the ultra-high-speed aircraft sealing cavity, each micro-hole being of a meandering shape, and an angle between a normal of each micro-hole and a normal of the wall surface 310 of the cavity head. In the range of 0-90 degrees.
  • the cold source drive device 100 includes a cold source reservoir 210, an air pump 110, a buffer 150, and a flow splitter 170.
  • the air pump 110 is in communication with the cold source reservoir 210.
  • Buffer 150 includes a buffer inlet and a buffer outlet.
  • the splitter 170 includes at least one inlet and two or more outlets.
  • the buffer inlet communicates with the cold source reservoir 210, the buffer outlet communicates with the splitter inlet, and each splitter outlet communicates with the transition layer 320 of the cavity wall (as shown, three shunt outlets are shown in Fig. 1 to communicate with the cavity wall)
  • the transition layer 320) is provided with a sealing valve (not shown in FIG. 1) at the communication portion of each of the splitter outlets and the transition layer 320 of the cavity wall.
  • An electric valve 120 and a check valve 130 for air entering the cold source reservoir 210 are disposed between the air pump 110 and the cold source reservoir 210.
  • a one-way valve 140 for the cold source 200 to enter the damper 150 is disposed between the cold source reservoir 210 and the damper 150.
  • a pressure sensor 230 and a safety valve 220 are disposed on the cold source reservoir 210.
  • the cold source 200 is liquid nitrogen.
  • the electric valve 120 and the check valve 130 are opened, the air pump 110 is activated, the compressed air enters the cold source reservoir 210, the electric valve 120 is adjusted to control the air flow rate; the check valve 140 is opened, and the liquid nitrogen is applied under the air pressure.
  • the buffer 150 After entering the buffer 150, after the buffer 150 is vaporized into nitrogen gas, it enters the flow divider through the inlet of the flow divider 170 under pressure, is divided into multiple nitrogen gas, opens the sealing valve, and the nitrogen gas is injected into the cavity from each outlet of the flow divider 170.
  • a transition layer 320 of the head wall through which the gas is ejected from the micropores 300 in the outer surface layer 330 to form a cavity Air film.
  • the pressure sensor 230 detects the gas pressure in the cold source reservoir 210, and adjusts the safety valve 220 in real time by observing the pressure sensor 230 to adjust the gas pressure in the cold source reservoir 210 to realize the discharge of liquid nitrogen from the cold source reservoir 210 to the buffer 150. Rate regulation.
  • the buffer 150 is coupled to the temperature sensor 160, and the temperature of the nitrogen in the buffer 150 is monitored by the temperature sensor 160.

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Abstract

一种超高速飞行器的热防护与减阻系统和方法。在超高速飞行器的腔体内部设置冷源（200）和冷源驱动装置（100），腔体壁面（310）设置若干微孔（300），冷源驱动装置（100）包括空气泵（110）、冷源储存器（210）和缓冲器（150），工作时，空气泵（110）向冷源储存器（210）提供压缩空气，在空气压力作用下冷源（200）进入缓冲器（150）并气化，高压气体自微孔（300）射出，在腔体外表面形成气膜。该气膜不仅能够对超高速飞行器进行热防护，而且能够有效减少飞行器与外界气体的粘阻，有助于减缓或消除热障现象，从而提高超高速飞行器的安全性，延长使用寿命。

Description

一种超高速飞行器的热防护与减阻方法和系统 技术领域
本发明属于超高速飞行器技术领域，特别涉及一种超高速飞行器的热防护与减阻方法和系统。
背景技术
超高速飞行器是指飞行速率在5马赫以上的飞行器，包括火箭、导弹、飞船、航天飞机、空天飞机等。超高速飞行器在飞行过程中存在两个主要问题：(1)超高速飞行器进出大气层时面临空气粘阻问题，需要耗费大量能量克服气动阻力；(2)超高速飞行器在飞行过程中还会面临剧烈的气动激波摩擦生热现象，出现热障，严重情况下，可以产生数千度高温的等离子体，导致通讯中断，这个阶段是飞行器的高危期。
关于空气粘阻，目前的超高速飞行器一般通过流线型的外形设计降低空气阻力。
关于超高速飞行器的热防护，目前国内外研究主要分为六类热防护方式：热沉式防热；辐射防热；烧蚀式防热；发汗冷却防热；表面隔热防热；热管散热。其中，烧蚀式防热与发汗冷却防热的防热效果较好，适用于热现象严重(例如摩擦生热产生等离子体)的飞行器。但是，这两种方式都难以长时间持续地进行热防护，导致昂贵的飞行器需要频繁地进行大修或者面临使用数次就被淘汰的现状。其次，利用这两种方式很难控制飞行器的内部温度，但是飞行器内部温度持续升高将严重危及携带系统的安全性。另外，相关防护系统的结构复杂，容易出现意外故障。
因此，能够有效减小空气粘阻的减阻技术以及能够有效减缓、克服热障，避免过度热蚀的热防护技术是目前超高速飞行器急需研究的课题。
发明内容
针对上述技术现状，本发明提供了一种高速飞行器的热防护与减阻方法，尤其是超高速飞行器的热防护与减阻方法，利用该方法能够避免超高速飞行器过度热蚀，同时能够减少超高速飞行器的空气粘阻。
本发明采用的技术方案为：一种超高速飞行器的热防护与减阻方法，在超高速飞行器的腔体内部设置冷源，超高速飞行器的腔体壁面设置若干微孔，在驱动力作用下，该冷源呈高压气体状自微孔射出，在腔体外表面形成气膜。
所述的微孔的设置位置不限，优选设置在超高速飞行器腔体的鼻锥(或称头部)和/或尾翼等部位。
所述的微孔在超高速飞行器腔体壁面的分布不限，作为优选，所述的微孔规则分布在超高速飞行器的腔体壁面，进一步优选，所述的微孔按照气动力学的特征规则分布在超高速飞行器的腔体壁面。
所述的微孔形状不限，可以是直孔或者异型孔，微孔横截面可以是规则形状(例如圆形等)或者不规则形状(例如蝶形、簸箕形等)。数值仿真表明，当微孔为异型孔时 有利于喷射出的冷源覆盖在腔体表面形成气膜，可以用较少的微孔实现优越的冷却效果，从而提高气膜的冷却效果同时更好地保障结构强度。
所述的微孔直径不限，作为优选，微孔的直径设计兼顾超高速飞行器腔体的结构强度以及冷源对腔体壁面的覆盖程度。作为一种实现方式，所述的微孔是直径为0.05毫米-2.0毫米的圆形直孔。
所述的冷源的来源不限，可以是液氮、干冰、压缩空气等冷却源，还可以是其他通过化学反应产生的冷却物质等。
所述的驱动力不限，包括压力、弹力、电力等。
所述的超高速飞行器的飞行速率在5马赫以上。所述的超高速飞行器包括火箭、导弹、飞船、航天飞机、空天飞机等。
所述的超高速飞行器的腔体材料不限，包括耐高温腐蚀的碳碳复合材料，碳与碳化硅的复合材料等。
综上所述，本发明的方法适用于高速飞行器，尤其是超高速飞行器。本发明在超高速飞行器腔体表面形成低温气膜，具有如下有益效果：
(1)该低温气膜位于超高速飞行器腔体表面，外界气体与该气膜相互作用，因此有效避免了外界气体与超高速飞行器直接摩擦而产生大量热量；同时，外界气体首先与该气膜层摩擦，有效降低了超高速飞行器与外界气体间的气体粘阻，气体粘阻的降低又有利于降低超高速飞行器的表面温度；
(2)该低温气膜是由冷源自超高速飞行器腔体内部喷射出而形成的，在此过程中冷源带走了高速飞行器腔体内部的大量热量，因此该方法能够有效控制超高速飞行器内部温度，有效避免飞行器内部温度持续升高而导致的危害；
因此，利用本发明的方法不仅能够对超高速飞行器进行热防护，而且能够有效减少高速飞行器与外界气体的粘阻，从而提高了超高速飞行器的能量效率和极限速度，可以减缓或避免热障现象，降低对热防护层材料的烧蚀，提高超高速飞行器的安全性，延长使用寿命，具有良好的应用前景。
本发明还提供了一种高速飞行器的减阻与热动防护系统，尤其是超高速飞行器的减阻与热动防护系统，包括设置在超高速飞行器的密封腔体内部的冷源，以及用于将冷源转化为高压气体并射出的冷源驱动装置；
所述超高速飞行器的腔体壁中至少部分壁面为夹层结构，所述夹层结构包括冷源气体可通过的过渡层以及位于该过渡层表面的外表面层，所述外表面层设置若干微孔，用于连通过渡层与腔体外部；
所述冷源驱动装置包括冷源储存器、空气泵以及缓冲器；空气泵连通冷源储存器；缓冲器包括缓冲器入口与缓冲器出口，缓冲器入口连通冷源储存器相，缓冲器出口连通腔体壁面的过渡层，并且缓冲器出口与过渡层的连通部位设置密封阀；
工作状态时，空气泵向冷源储存器提供压缩空气，在空气压力作用下冷源进入缓冲 器并气化，打开密封阀，气体自缓冲器出口喷入腔体壁面的过渡层，然后自外表面层的微孔喷射出腔体形成气膜。
所述过渡层用于将冷源气体通向外表面层，该过渡层可以是中空层，也可以是冷源气体可通过的其他介质层等。
为了提高冷源的喷射效果，作为一种优选方式，缓冲器的出口数目大于或等于两个，每个出口连通腔体壁面的过渡层，连通部位设置密封阀门。
为了提高冷源的喷射效果，作为另一种优选方式，所述冷源驱动装置还包括分流器，所述分流器包括至少一个入口与两个或者两个以上个出口，所述分流器入口与缓冲器出口连通，每个分流器出口连通腔体壁面的过渡层，并且每个分流器出口与过渡层的连通部位设置密封阀；冷源气化后经分流器入口进入分流器，分流为多路气体后自各分流器出口喷入腔体壁面的过渡层，最后自外表面层的微孔喷射出腔体形成气膜。
作为优选，所述的空气泵与冷源储存器之间设置电动阀与单向阀，工作状态时，打开电动阀与单向阀，压缩空气进入冷源储存器，调节电动阀可控制空气流量。
作为优选，所述的冷源储存器与缓冲器之间设置单向阀，工作状态时，打开该电动阀，冷源进入缓冲器。
作为优选，所述的冷源驱动装置还包括温度传感器，用于监测缓冲器中冷源的温度。
为了调节冷源自冷源储存器进入缓冲器的速率，冷源储存器上设置用于检测冷源储存器内压力的压力传感器以及用于调节冷源储存器内气体压力的安全阀。
作为优选，具有夹层结构的腔体壁面为腔体的鼻锥部位和/或尾翼等部位的壁面。
作为优选，所述微孔规则分布在超高速飞行器的腔体壁面。
作为优选，所述微孔是非圆形孔；进一步优选，所述微孔直径为0.05毫米-2.0毫米。
所述的冷源的来源不限，可以是液氮、干冰、压缩空气等冷却源，还可以是其他通过化学反应产生的冷却物质等。
所述的超高速飞行器的飞行速率在5马赫以上。所述的超高速飞行器包括火箭、导弹、飞船、航天飞机、空天飞机等。
所述的超高速飞行器的腔体材料不限，包括耐高温腐蚀的碳碳复合材料，碳与碳化硅的复合材料等。
利用本发明的系统能够在超高速飞行器腔体表面形成低温气膜，不仅能够对超高速飞行器进行热防护，而且能够有效减少高速飞行器与外界气体的粘阻，从而可以提高超高速飞行器的能量效率和极限速度，可以减缓或避免热障现象，降低对热防护层材料的烧蚀，提高超高速飞行器的安全性，延长使用寿命，具有良好的应用前景。
附图说明
图1是本发明实施例1中超高速飞行器的热防护及减阻系统的结构示意图；
图2是图1中腔体头部的壁面的立体结构示意图；
图3是图2的俯视结构示意图；
图4是图3沿A-A切面的结构示意图；
图5是图4中的局部B的放大图。
具体实施方式
以下将结合附图及实施例对本发明做进一步说明，需要指出的是，以下所述实施例旨在便于对本发明的理解，而对其不起任何限定作用。
图1-3中的附图标记为：冷源驱动装置100、冷源200、微孔300、冷源储存器210、空气泵110、电动阀120、单向阀130、单向阀140、缓冲器150、温度传感器160、分液器170、安全阀220、压力传感器230、腔体头部壁面310、过渡层320、外表面层330。
实施例1：
为了使本发明的技术方案更加清楚，以下结合附图，对本发明超高速飞行器热防护及减阻系统作进一步详细的说明。应当理解，此处所描述的具体实施例仅用以解释本发明并不用于限定本发明。
本实施例中，如图1所示，超高速飞行器包括密封腔体，超高速飞行器的减阻与热动防护系统包括设置在超高速飞行器密封腔体内部的冷源200，以及用于将冷源200转化为高压气体并射出的冷源驱动装置100。超高速飞行器密封腔体的头部的壁面310为夹层结构。图2是该腔体头部的壁面的立体结构示意图，图3是图2的俯视结构示意图，图4是图3沿A-A切面的结构示意图，图5是图4中的局部B的放大图。从图2至图5可以看出，自腔体内部至腔体外部方向，该夹层结构包括过渡层320以及位于过渡层320表面的外表面层330，表面层330设置若干微孔300，用于连通过渡层320与腔体外部。该微孔300在超高速飞行器密封腔体头部壁面310呈发散状分布，每个微孔为簸箕型，各微孔的法线与该腔体头部壁面310的法线之间的夹角在0-90度范围。
冷源驱动装置100包括冷源储存器210、空气泵110、缓冲器150以及分流器170。空气泵110与冷源储存器210相连通。缓冲器150包括缓冲器入口与缓冲器出口。分流器170包括至少一个入口与两个或者两个以上个出口。缓冲器入口连通冷源储存器210，缓冲器出口连通分流器入口，每个分流器出口连通腔体壁面的过渡层320(作为示意，图1中示出三个分流器出口连通腔体壁面的过渡层320)，每个分流器出口与腔体壁面的过渡层320的连通部位设置密封阀门(图1中未示出)。
空气泵110与冷源储存器210之间设置电动阀120以及用于空气进入冷源储存器210的单向阀130。
冷源储存器210与缓冲器150之间设置用于冷源200进入缓冲器150的单向阀140。
冷源储存器210上设置压力传感器230与安全阀220。
本实施例中，冷源200为液氮。
工作状态时，打开电动阀120与单向阀130，启动空气泵110，压缩空气进入冷源储存器210，调节电动阀120可控制空气流量；打开单向阀140，在空气压力作用下液氮进入缓冲器150，在缓冲器150气化为氮气后在压力作用下经分流器170的入口进入分流器，分流为多路氮气，打开密封阀，氮气自分流器170的各出口喷入腔体头部壁面的过渡层320，通过该过渡层320后气体自外表面层330中的微孔300喷射出腔体形成 气膜。
压力传感器230检测冷源储存器210内气体压力，通过观察压力传感器230实时地调节安全阀220，以调节冷源储存器210内气体压力，实现液氮自冷源储存器210向缓冲器150排出的速率调控。
缓冲器150连接温度传感器160，通过温度传感器160监测缓冲器150中氮气的温度。
以上所述的实施例对本发明的技术方案进行了详细说明，应理解的是以上所述仅为本发明的具体实施例，并不用于限制本发明，凡在本发明的原则范围内所做的任何修改、补充或类似方式替代等，均应包含在本发明的保护范围之内。

Claims (15)

  1. 一种超高速飞行器的热防护与减阻方法，其特征是：在超高速飞行器的腔体内部设置冷源，超高速飞行器的腔体壁面设置若干微孔，在驱动力作用下，该冷源呈高压气体状自微孔射出，在腔体外表面形成气膜。
  2. 如权利要求1所述的超高速飞行器的热防护与减阻方法，其特征是：所述微孔设置在超高速飞行器腔体的鼻锥部位和/或尾翼部位。
  3. 如权利要求1所述的超高速飞行器的热防护与减阻方法，其特征是：所述微孔规则分布在超高速飞行器的腔体壁面。
  4. 如权利要求1所述的超高速飞行器的热防护与减阻方法，其特征是：所述微孔是非圆形孔；所述微孔直径优选为0.05毫米-2.0毫米。
  5. 如权利要求1至4中任一权利要求所述的超高速飞行器的热防护与减阻方法，其特征是：所述冷源是液氮、干冰、压缩空气，或者是其他通过化学反应产生的冷却物质。
  6. 如权利要求1至4中任一权利要求所述的超高速飞行器的热防护与减阻方法，其特征是：所述的超高速飞行器的飞行速率在5马赫以上；作为优选，所述的超高速飞行器是火箭、导弹、飞船、航天飞机或者空天飞机。
  7. 一种超高速飞行器的热防护与减阻系统，其特征是：包括设置在超高速飞行器密封腔体内部的冷源，以及用于将冷源转化为高压气体并射出的冷源驱动装置；
    所述超高速飞行器的腔体壁中至少部分壁面为夹层结构，所述夹层结构包括冷源气体可通过的过渡层以及位于该过渡层表面的外表面层，所述外表面层设置若干微孔，用于连通过渡层与腔体外部；
    所述冷源驱动装置包括冷源储存器、空气泵以及缓冲器；空气泵连通冷源储存器；缓冲器包括缓冲器入口与缓冲器出口，缓冲器入口连通冷源储存器相，缓冲器出口连通腔体壁面的过渡层，并且缓冲器出口与过渡层的连通部位设置密封阀；
    工作状态时，空气泵向冷源储存器提供压缩空气，在空气压力作用下冷源进入缓冲器并气化，打开密封阀，气体自缓冲器出口喷入过渡层，然后自外表面层的微孔喷射出腔体形成气膜。
  8. 如权利要求7所述的超高速飞行器的热防护与减阻系统，其特征是：所述缓冲器出口的数目大于或等于两个。
  9. 如权利要求7所述的超高速飞行器的热防护与减阻系统，其特征是：所述冷源驱动装置还包括分流器，所述分流器包括至少一个入口与两个或者两个以上个出口，所述分流器入口与缓冲器出口连通，每个分流器出口连通腔体壁面的过渡层，并且每个分流器出口与过渡层的连通部位设置密封阀；冷源气化后经分流器入口进入分流器，分流为多路气体后自各分流器出口喷入腔体壁面的过渡层，然后自微孔喷射出腔体形成气膜。
  10. 如权利要求7、8或9所述的超高速飞行器的热防护与减阻系统，其特征是：所述空气泵与冷源储存器之间设置电动阀与单向阀，工作状态时，打开电动阀与单向阀，压缩空气进入冷源储存器，调节电动阀可控制空气流量；作为优选，所述冷源储存器与 缓冲器之间设置单向阀，工作状态时，打开该电动阀，冷源进入缓冲器。
  11. 如权利要求7、8或9所述的超高速飞行器的热防护与减阻系统，其特征是：所述冷源驱动装置还包括温度传感器，用于监测缓冲器中冷源的温度；作为优选，所述冷源储存器上设置用于检测冷源储存器内气体压力的压力传感器以及用于调节冷源储存器内气体压力的安全阀。
  12. 如权利要求7、8或9所述的超高速飞行器的热防护与减阻系统，其特征是：所述的超高速飞行器的飞行速率在5马赫以上；作为优选，所述的超高速飞行器是火箭、导弹、飞船、航天飞机或者空天飞机。
  13. 如权利要求7、8或9所述的超高速飞行器的热防护与减阻系统，其特征是：具有夹层结构的腔体壁面为腔体的鼻锥部位和/或尾翼部位的壁面。
  14. 如权利要求7、8或9所述的超高速飞行器的热防护与减阻系统，其特征是：所述微孔规则分布在超高速飞行器的腔体壁面；作为优选，所述微孔是非圆形孔；进一步优选，所述微孔直径为0.05毫米-2.0毫米。
  15. 如权利要求7、8或9所述的超高速飞行器的热防护与减阻系统，其特征是：所述冷源是液氮、干冰、压缩空气，或者是其他通过化学反应产生的冷却物质。
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