Thermal protection and drag reduction method and system for ultra high-speed aircraft

Abstract

Disclosed are the thermal protection and drag reduction method and system for an ultra high-speed aircraft. A cold source is and a cold source driving device are arranged inside a cavity of the ultra high-speed aircraft. A plurality of micropores are arranged on a wall surface of the cavity. The cold source driving device comprises an air pump, a cold source reservoir and a buffer. The air pump supplies compressed air to a cold source reservoir during operation. The cold source enters the buffer and is vaporized under the action of air pressure. High-pressure gas is ejected from the micropores to form a gas film on the outer surface of the cavity. The gas film not only can perform thermal protection on the ultra high-speed aircraft, but also can effectively reduce viscous drag between the aircraft and the external gas, by virtue of which the thermal barrier phenomenon is alleviated or eliminated. Therefore, security of the ultra high-speed aircraft is improved and service life is prolonged.

Description

TECHNICAL FIELD

The present invention relates to the technical field of ultra high-speed aircraft, and more particularly, to a thermal protection and drag reduction method and system for ultra high-speed aircraft.

BACKGROUND

Ultra high-speed aircrafts refers to aircrafts with flight speed of 5 Mach or more, including rockets, missiles, spacecrafts, space shuttles, aerospace planes and the like. Ultra high-speed aircrafts have two main problems during the flight in that: (1) ultra high-speed aircrafts may face a problem of air viscous drag while entering or exiting the atmospheric layer, and a lot of energy is required to overcome the aerodynamic drag; (2) Ultra high-speed aircrafts may face violent heat generation phenomenon by friction due to aerodynamic shock wave during the flight, thermal barrier occurs, and in serious cases, plasma with high temperature of thousands of degrees may be generated, leading to communication interruption, and thus this stage is the high risk period of the aircraft.

As for air viscous drag, the current ultra high-speed aircrafts can generally reduce air drag through a design of a streamlined profile.

As for the thermal protection of ultra high-speed aircrafts, the current domestic and foreign researches are divided into six types of thermal protection, i.e., heat sink thermal protection, radiation thermal protection, ablation thermal protection, transpiration cooling thermal protection, surface thermal insulation thermal protection, and heat pipe heat dissipation. Among them, the ablation thermal protection and transpiration cooling thermal protection have relatively better effect and are suitable for aircrafts suffering from serious thermal phenomenon (such as plasma generated by the generation of heat by friction). However, both of these methods are difficult to carry out long-term thermal protection, resulting in that expensive aircrafts are required to be overhauled frequently or obsoleted after several times of use. Secondly, it is difficult to control the internal temperature of the aircrafts by using such two methods, but the continuous increase in the internal temperature of the aircrafts will seriously endanger the safety of the carried system. In addition, the relevant protective system has a complicated structure, and accidental failure is easy to occur.

Therefore, drag reduction technology for effectively reducing air viscous drag and thermal protection technology for effectively retarding and overcoming thermal barrier as well as avoiding excessive heat erosion are the issues to be studied in urgent need for ultra high-speed aircrafts.

SUMMARY

In view of the above technical state, the invention provides a thermal protection and drag reduction method for high-speed aircraft, especially a thermal protection and drag reduction method for ultra high-speed aircraft, the application of this method can avoid excessive heat erosion of the ultra high-speed aircraft, while reducing air viscous drag of the ultra high-speed aircraft.

The technical solution adopted by the invention is: a thermal protection and drag reduction method for ultra high-speed aircraft, wherein a cold source is provided inside a cavity of the ultra high-speed aircraft, a plurality of micropores are arranged on a wall surface of the cavity of the ultra high-speed aircraft, and the cold source is ejected from the micropores in the form of high pressure gas under the action of driving force, so as to form a gas film on the outer surface of the cavity.

The position of the micropores is not limited, and preferably, the micropores are located at the nose cone (or head) and/or empennage portion and the like of the cavity of the ultra high-speed aircraft.

The distribution of micropores on the wall surface of the cavity of the ultra high-speed aircraft are not limited, and preferably, the micropores are regularly distributed on the wall surface of the cavity of the ultra high-speed aircraft. It is further preferred that the micropores are regularly distributed on the wall surface of the cavity of the ultra high-speed aircraft in accordance with aerodynamic characteristics.

The shape of the micropores are not limited, and the micropores may be straight holes or shaped holes, and the cross section thereof may be regular shapes (e.g., circular or the like) or irregular shapes (e.g., butterfly-shaped, dustpan-shaped or the like). The numerical simulation shows that when the micropores are shaped holes, it is advantageous to eject the cold source to cover the surface of the cavity so as to form the gas film, and excellent cooling effect may be achieved by less micropores, thereby improving the cooling effect of the gas film while better ensuring structural strength.

The diameters of the micropores is not limited, and preferably, the design of diameters of the micropores takes into account the structural strength of the cavity of the ultra high-speed aircraft and the coverage extent of the cold source to the wall surface of the cavity. As one embodiment, the micropores are circular straight holes with diameters of 0.05 mm to 2.0 mm.

The source of the cold source is not limited, and the cold source may be a cooling source such as liquid nitrogen, dry ice, compressed air, or other cooling material obtained by a chemical reaction.

The driving force is not limited, including pressure, elastic force, electric power, and the like.

The flight speed of the ultra high-speed aircraft is 5 Mach or more. The ultra high-speed aircraft comprises a rocket, a missile, a spacecraft, a space shuttle, an aerospace plane and the like.

The material of the cavity of the ultra high-speed aircraft is not limited, including high temperature corrosion-resistant C—C composites, C—SiC composites, and the like.

In summary, the method of the present invention is applicable to a high speed aircraft, particularly to an ultra high-speed aircraft. By application of the present invention, a low-temperature gas film may be formed on the surface of the cavity of the ultra high-speed aircraft, and the present invention has the following advantageous effects.

• (1) The low-temperature gas film is located on the surface of the cavity of the ultra high-speed aircraft, and the external gas interacts with the gas film, thereby effectively avoiding the generation of a lot of heat due to the direct friction between the external gas and the ultra high-speed aircraft. Meanwhile, the external gas is firstly subjected to friction with the gas film layer, which effectively reduces the gas viscous drag between the ultra high-speed aircraft and the external gas, and the reduction of the gas viscous drag is conducive to reduce the surface temperature of ultra high-speed aircraft.
• (2) The low-temperature gas film is formed by the ejection of the cold source from the inside of the cavity of the ultra high-speed aircraft, and in this process, the cold source takes away a lot of heat inside the cavity of the ultra high-speed aircraft, and thus, such method can effectively control the internal temperature of the ultra high-speed aircraft, and can effectively avoid the damage due to continuously rising internal temperature of the aircraft.

Therefore, the application of the method according to the present invention can not only perform thermal protection on the ultra high-speed aircraft, but also effectively reduce viscous drag between the high-speed aircraft and the external gas, thereby improving the energy efficiency and ultimate speed of the ultra high-speed aircraft. The method can retard or avoid the thermal barrier phenomenon, reduce ablation of the thermal protective layer material, improve the safety of the ultra high-speed aircraft and prolong the service life, and thus, it has a good application prospects.

The invention also provides a drag reduction and thermal protection system for high-speed aircraft, especially a drag reduction and thermal protection system for ultra high-speed aircraft comprising a cold source disposed inside a sealed cavity of the ultra high-speed aircraft, and a cold source driving device for converting the cold source into high pressure gas and ejecting the cold source.

At least part of a wall surface of a cavity wall of the ultra high-speed aircraft has a sandwich structure, wherein the sandwich structure comprises a transition layer through which cold source gas passes and an outer surface layer located at a surface of the transition layer, and the outer surface layer is provided with a plurality of micropores for communicating the transition layer with 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 in communication with the cold source reservoir; the buffer comprises a buffer inlet and a buffer outlet, the buffer inlet is in communication with the cold source reservoir, the buffer outlet is in communication with the transition layer of the wall surface of the cavity, and a sealing valve is provided at a portion where the buffer outlet is in communication with the transition layer.

During the operation, the air pump supplies a compressed air to the cold source reservoir, the cold source enters the buffer and is vaporized under air pressure, and the gas is ejected into the transition layer of the wall surface of the cavity from the buffer outlet when the sealing valve is open, and then ejected out of the cavity from the micropores of the outer surface layer so as to form a gas film.

The transition layer serves to direct the cold source gas to the outer surface layer, and may be a hollow layer, or other dielectric layer through which the cold source gas may pass.

In order to improve the ejection effect of the cold source, as a preferred embodiment, the number of the buffer outlet is two or more, and each outlet is in communication with the transition layer of the wall surface of the cavity, and sealing valves are provided at the communicating portion.

In order to improve the ejection effect of the cold source, as another preferred embodiment, the cold source driving device further comprises a splitter comprising at least one inlet and two or more outlets, the inlet of the splitter is in communication with the buffer outlet, each outlet of the splitter is in communication with the transition layer of the wall surface of the cavity, and a sealing valve is provided at the portion where each outlet of the splitter is in communication with the transition layer; the cold source enters the splitter through the inlet of the splitter after vaporized, and is ejected into the transition layer of the wall surface of the cavity from each outlet of the splitter after being split into gases in multi-channels, and finally, ejected out of the cavity from the micropores of the outer surface layer so as to form the gas film.

Preferably, an electric valve and a check valve are provided between the air pump and the cold source reservoir. During operation, the compressed air enters the cold source reservoir when the electric valve and the check valve are open, and the air flow can be controlled by adjusting the electric valve.

Preferably, a check valve is provided between the cold source reservoir and the buffer, and during operation, the cold source enters the buffer when the check valve is open.

Preferably, the cold source driving device further comprises a temperature sensor for monitoring the temperature of the cold source in the buffer.

In order to adjust the rate of cold source entered from the cold source reservoir into the buffer, a pressure sensor for detecting the gas pressure in the cold source reservoir and a safety valve for adjusting the gas pressure in the cold source reservoir are provided on the cold source reservoir.

Preferably, the wall surface of the cavity having the sandwich structure locates the nose cone portion and/or the empennage portion and the like of the cavity.

Preferably, the micropores are regularly distributed on the wall surface of the cavity of the ultra high-speed aircraft.

Preferably, the micropores are non-circular pores; further preferably, the diameters of the micropores range 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 material obtained by a chemical reaction.

The flight speed of the ultra high-speed aircraft is 5 Mach or more. The ultra high-speed aircraft comprises a rocket, a missile, a spacecraft, a space shuttle, an aerospace plane and the like.

The material of the cavity of the ultra high-speed aircraft is not limited, including high temperature corrosion-resistant C—C composites, C—SiC composites and the like.

The method according to the present invention can form a low-temperature gas film on the surface of the cavity of the ultra high-speed aircraft, which can not only perform thermal protection on the ultra high-speed aircraft, but also effectively reduce the viscous drag between the high-speed aircraft and the external gas, thereby improving the energy efficiency and ultimate speed of the ultra high-speed aircraft. The method can retard or avoid the thermal barrier phenomenon, reduce ablation of the thermal protective layer material, improve the safety of the ultra high-speed aircraft and prolong the service life, and thus, it has a good application prospects.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic structural view of a thermal protection and drag reduction system for ultra high-speed aircraft according to embodiment 1 of the present invention;

FIG. 2 is a schematic view of the three-dimensional structure of the wall surface at the head portion of the cavity in FIG. 1;

FIG. 3 is a schematic top-view of the structure in FIG. 2;

FIG. 4 is a schematic structural view of the section taken along A-A in FIG. 3; and

FIG. 5 is an enlarged view of the portion B in FIG. 4.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

The present invention is described in connection with the accompanying drawings and embodiments, it should be noted that the following embodiment is intended to be convenient for understanding the present invention, but does not limit the present invention.

Reference numerals in FIGS. 1-3: cold source driving device 100, cold source 200, micropores 300, cold source reservoir 210, air pump 110, electric valve 120, check valve 130, check valve 140, buffer 150, temperature sensor 160, dispenser 170, safety valve 220, pressure sensor 230, wall surface 310 of the head of the cavity, transition layer 320, outer surface layer 330.

Embodiment 1

In order to make the technical solution of the present invention clearer, the thermal protection and drag reduction system for ultra high-speed aircraft of the present invention will be described in more detail with reference to the accompanying drawings. It will be understood that the specific embodiments described are only used for explaining the present invention, but not for limiting the present invention.

In the present embodiment, as shown in FIG. 1, the ultra high-speed aircraft comprises a sealed cavity, the thermal protection and drag reduction system for ultra high-speed aircraft comprises a cold source 200 disposed inside the sealed cavity of the ultra high-speed aircraft, and a cold source driving device 100 for converting the cold source 200 into a high pressure gas and emitting the high pressure gas. The wall surface 310 of the head portion of the sealed cavity of the ultra high-speed aircraft has a sandwich structure. FIG. 2 is a schematic view of the three-dimensional structure of the wall surface at the head portion of the cavity; FIG. 3 is a schematic top-view of the structure in FIG. 2; FIG. 4 is a schematic structural view of the section taken along A-A in FIG. 3; and FIG. 5 is an enlarged view of the portion B in FIG. 4. As can be seen from FIG. 2 to FIG. 5, the sandwich structure comprises a transition layer 320 and an outer surface layer 330 on the surface of the transition layer 320 when observed in the direction from the inside of the cavity to the outside of the cavity, and the surface layer 330 is provided with a plurality of micropores 300 for communicating the transition layer 320 with the outside of the cavity. The micropores 300 are distributed on the wall surface 310 of the head portion of the sealed cavity of the ultra high-speed aircraft in a divergent form, each of the micropores is dustpan-shaped, and the angle between the normal of each of the micropores and the normal of the wall surface 310 of the head portion of the cavity is in the range of 0-90 degree.

The cold source driving device 100 comprises a cold source reservoir 210, an air pump 110, a buffer 150, and a splitter 170. The air pump 110 is in communication with the cold source reservoir 210. The buffer 150 comprises a buffer inlet and a buffer outlet. The splitter 170 comprises at least one inlet and two or more outlets. The buffer inlet is in communication with the cold source reservoir 210, the buffer outlet is in communication with the inlet of the splitter, and each outlet of the splitter is in communication with the transition layer 320 of the wall surface of the cavity (as indicated, FIG. 1 shows that the transition layer 320 of the wall surface of the cavity is communicated with the three outlets of the splitter), and a sealing valve (not shown in FIG. 1) is provided at the portion where each outlet of the splitter is in communication with the transition layer 320 of the wall surface of the cavity).

An electric valve 120 and a check valve 130 are provided between the air pump 110 and the cold source reservoir 210, and the check valve 130 is used for air to enter the cold source reservoir 210.

A check valve 140 is provided between the cold source reservoir 210 and the buffer 150, and the check valve 140 is used for the cold source 200 to enter the buffer 150.

The cold source reservoir 210 is provided with a pressure sensor 230 and a safety valve 220.

In the present embodiment, the cold source 200 is liquid nitrogen.

During operation, the compressed air enters the cold source reservoir 210 when the electric valve 120 and the check valve 130 are open and the air pump 110 is actuated, and the air flow can be controlled by adjusting the electric valve 120. The liquid nitrogen enters the buffer 150 under the air pressure when the check valve 140 is opened, and enters the splitter through the inlet of the splitter 170 under the pressure after vaporized into nitrogen gas at the buffer 150, and then the nitrogen gas is split into gases in multi-channels. The nitrogen gas is ejected into the transition layer 320 of the wall surface of the head of the cavity from each outlet of the splitter 170 when the sealing valves are open, and ejected out of the cavity from the micropores 300 in the outer surface layer 330 after passing through the transition layer 320 so as to form the gas film.

The pressure sensor 230 detects the gas pressure in the cold source reservoir 210, and the safety valve 220 may be adjusted in real time by observing the pressure sensor 230 so as to adjust the gas pressure in the cold source reservoir 210, so that the rate control of the liquid nitrogen discharged from the cold source reservoir 210 to the buffer 150 can be realized.

The buffer 150 is connected to the temperature sensor 160, and the temperature of the nitrogen gas in the buffer 150 is monitored by the temperature sensor 160.

The technical solutions of the present invention are specifically explained through the above embodiments, and it will be understood that the above mentioned are only specific embodiments of the present embodiment, but not for limiting the present invention, and any modifications, supplements and the like within the principle of the present invention should be incorporated into the scope of protection of the present invention.

Claims

1. A thermal protection and drag reduction method for ultra high-speed aircraft, comprising providing a cold source inside a cavity of the ultra high-speed aircraft,arranging a plurality of micropores on a wall surface of the cavity of the ultra high-speed aircraft, wherein the cold source is ejected from the micropores in the form of high pressure gas under the action of driving force, so as to form a gas film on an outer surface of the cavity.

2. The thermal protection and drag reduction method for ultra high-speed aircraft according to claim 1, wherein the micropores are provided at a nose cone portion and/or an empennage portion of the cavity of the ultra high-speed aircraft.

3. The thermal protection and drag reduction method for ultra high-speed aircraft according to claim 1, wherein the micropores are regularly distributed on the wall surface of the cavity of the ultra high-speed aircraft.

4. The thermal protection and drag reduction method for ultra high-speed aircraft according to claim 1, wherein the micropores are non-circular holes.

5. The thermal protection and drag reduction method for ultra high-speed aircraft according to claim 1, wherein the cold source is liquid nitrogen, dry ice, compressed air, or other cooling material obtained by chemical reactions.

6. The thermal protection and drag reduction method for ultra high-speed aircraft according to claim 1, wherein the flight speed of the ultra high-speed aircraft is 5 Mach or more.

7. A thermal protection and drag reduction system for ultra high-speed aircraft, comprising a cold source disposed inside a sealed cavity of the ultra high-speed aircraft, and a cold source driving device for converting the cold source into high pressure gas and emitting the high pressure gas; wherein, at least part of a wall surface of a cavity wall of the ultra high-speed aircraft has a sandwich structure comprising a transition layer through which cold source gas passes and an outer surface layer located at a surface of the transition layer, the outer surface layer is provided with a plurality of micropores for communicating the transition layer with 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 in communication with the cold source reservoir; the buffer comprises a buffer inlet and a buffer outlet, the buffer inlet is in communication with the cold source reservoir, the buffer outlet is in communication with the transition layer of the wall surface of the cavity, and a sealing valve is provided at a portion where the buffer outlet is in communication with the transition layer; andduring operation, the air pump supplies compressed air to the cold source reservoir, the cold source enters the buffer and is vaporized under air pressure, and the gas is ejected into the transition layer from the buffer outlet when the sealing valve is open, and then ejected out of the cavity from the micropores of the outer surface layer so as to form a gas film.

8. The thermal protection and drag reduction system for ultra high-speed aircraft according to claim 7, wherein a number of the buffer outlets is two or more.

9. The thermal protection and drag reduction system for ultra high-speed aircraft according to claim 7, wherein the cold source driving device further comprises a splitter comprising at least one inlet and two or more outlets, the inlet of the splitter is in communication with the buffer outlet, each outlet of the splitter is in communication with the transition layer of the wall surface of the cavity, and a sealing valve is provided at the portion where each outlet of the splitter is in communication with the transition layer; and the cold source enters the splitter through the inlet of the splitter after vaporized, and is ejected into the transition layer of the wall surface of the cavity from each outlet of the splitter after being split into gases in multi-channels, and then ejected out of the cavity from the micropores so as to form the gas film.

10. The thermal protection and drag reduction system for ultra high-speed aircraft according to claim 7, wherein an electric valve and a check valve are provided between the air pump and the cold source reservoir, and during operation, the compressed air enters the cold source reservoir when the electric valve and the check valve are open, and the air flow is controlled by adjusting the electric valve.

11. The thermal protection and drag reduction system for ultra high-speed aircraft according to claim 7, wherein the cold source driving device further comprises a temperature sensor for monitoring the temperature of the cold source in the buffer.

12. The thermal protection and drag reduction system for ultra high-speed aircraft according to claim 7, wherein flight speed of the ultra high-speed aircraft is 5 Mach or more.

13. The thermal protection and drag reduction system for ultra high-speed aircraft according to claim 7, wherein the wall surface of the cavity having the sandwich structure locates a nose cone portion and/or an empennage portion of the cavity.

14. The thermal protection and drag reduction system for ultra high-speed aircraft according to claim 7, wherein the micropores are regularly distributed on the wall surface of the cavity of the ultra high-speed aircraft.

15. The thermal protection and drag reduction system for ultra high-speed aircraft according to claim 7, wherein the cold source is liquid nitrogen, dry ice, compressed air, or other cooling material produced by chemical reactions.

16. The thermal protection and drag reduction method for ultra high-speed aircraft according to claim 1, wherein the ultra high-speed aircraft is a rocket, a missile, a spacecraft, a space shuttle, or an aerospace plane.

17. The thermal protection and drag reduction system for ultra high-speed aircraft according to claim 7, wherein a check valve is provided between the cold source reservoir and the buffer, and during operation, the cold source enters the buffer when the check valve is open.

18. The thermal protection and drag reduction system for ultra high-speed aircraft according to claim 7, wherein a pressure sensor for detecting gas pressure in the cold source reservoir and a safety valve for adjusting the gas pressure in the cold source reservoir are provided on the cold source reservoir.

19. The thermal protection and drag reduction system for ultra high-speed aircraft according to claim 7, wherein the ultra high-speed aircraft is a rocket, a missile, a spacecraft, a space shuttle, or an aerospace plane.

20. The thermal protection and drag reduction system for ultra high-speed aircraft according to claim 7, wherein the micropores are non-circular pores.