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.

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; and during 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.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0038] 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;

[0039] 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;

[0040] FIG. 3 is a schematic top-view of the structure in FIG. 2;

[0041] FIG. 4 is a schematic structural view of the section taken along A-A in FIG. 3; and

[0042] FIG. 5 is an enlarged view of the portion B in FIG. 4.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0043] 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.

[0044] 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

[0045] 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.

[0046] 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.

[0047] 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).

[0048] 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.

[0049] 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.

[0050] The cold source reservoir 210 is provided with a pressure sensor 230 and a safety valve 220.

[0051] In the present embodiment, the cold source 200 is liquid nitrogen.

[0052] 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.

[0053] 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.

[0054] 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.

[0055] 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.