A windward cavity structure based on microstructure seepage and its cooling method

Abstract

The present invention belongs to the field of cooling the nose of an aircraft, and specifically relates to a windward concave cavity structure based on microstructure seepage and a cooling method thereof, comprising an air supply system and a concave cavity disposed at the nose of the aircraft; the sidewalls of the concave cavity are provided with periodically arranged microstructures, the microstructures having recessed portions, and the surfaces of the microstructures are also provided with micropores connected to the air supply system, the air supply system being used to output gas to the micropores, the gas flowing from the surface of the microstructure into the recessed portions through the micropores, forming microscale vortices in the recessed portions and entering the concave cavity. The present invention achieves cooling of the nose of the aircraft by providing a microstructure within the concave cavity and delivering gas to the microstructure on the basis of existing windward concave cavity thermal protection. The overall structure is simple and has no impact on the original windward concave cavity thermal protection structure; in addition, the gas working medium requirement is small, and on this basis, the cooling effect on the nose of the aircraft can be significantly improved.

Description

Technical Field

[0001] The present invention belongs to the field of aircraft head cooling, and in particular relates to a windward concave cavity structure based on microstructure seepage and a cooling method thereof. Background Art

[0002] When an aircraft flies at high speed in the air, the friction between the body and the air generates a large amount of heat energy, resulting in severe aerodynamic heating. Thermal protection technology for high-speed aircraft is a key technology that needs to be broken through for its engineering application. Re-entry high-speed aircraft such as space shuttles and return capsules are subject to severe aerodynamic heating, and most of them use ablation protection methods, which is relatively costly. For long-term high-temperature environments, passive cooling methods using heat-resistant materials or material ablation are difficult to meet the requirements, and active cooling methods are required. A needle-shaped rod is installed axially at the head of the aircraft to push the shock wave away from the object surface and form a low-pressure recirculation zone at the head. This is a shock wave needle thermal protection method that can effectively reduce drag and prevent heat. It has been actually applied in flight devices, but the needle rod needs to be replaced frequently due to local high heat burning. Designing a concave cavity structure facing the incoming flow at the front end of the aircraft and using the shock wave oscillation characteristics to reduce the aerodynamic heating of the stagnation area is called upwind cavity thermal protection. This structure is simple in design and has excellent heat protection effect. The reverse jet technology ejects cooling gas from a reverse high-speed nozzle arranged facing the horizontal flow, which squeezes the free flow and pushes the bow shock wave away from its original position. The cooling airflow isolates the leading edge of the aircraft from the free flow to achieve cooling.

[0003] With the continuous deepening of research on the cooling mechanism of thermal protection systems, several combined active thermal protection methods have emerged. Geng Yunfei of the Beijing University of Aeronautics and Astronautics proposed a thermal protection method that combines a shock needle with a reverse jet. A nozzle is installed at the head of a high-speed aircraft, through which a cooling medium is injected in reverse. The nozzle position does not generate strong aerodynamic heating in the stagnation area, thus avoiding the ablation problem of a single shock needle. However, this method requires a large amount of cooling medium, and thus a large air source, which is inconsistent with the need for lightweight aircraft.

[0004] Lu Haibo of the National University of Defense Technology proposed a design method for a heat shield structure combining a windward cavity and a counter-flow jet, which can effectively cool the nose of a high-speed aircraft. However, this method requires a large amount of cooling fluid, and thus a large air source, which conflicts with the need for lightweight aircraft.

[0005] Due to the curved surface effect of the leading edge structure, the temperature and pressure in the stagnation point area near the curved surface are too high, and the coolant mass is relatively low in the head area with the highest heat flux density, which affects the normal operation of the aircraft. Luo Shibin of Central South University combined the aircraft's leading edge air film and divergent cooling technologies to achieve complementary advantages through structural design. Luan Yun of the University of Science and Technology of China proposed a combined cavity and divergent cooling method, which improved the cooling effect near the stagnation point and reduced the overall structural temperature. However, the cooling effect of these two methods still needs to be improved. Summary of the Invention

[0006] The technical problem to be solved by the present invention is to provide a windward cavity structure based on microstructure seepage and having a simple structure, low gas consumption and good cooling effect, and a cooling method thereof.

[0007] The present invention provides a windward concave cavity structure based on microstructure seepage, comprising an air source supply system and a concave cavity arranged at the head of an aircraft;

[0008] The side wall of the cavity is provided with a periodically arranged microstructure, the microstructure has a recessed portion, and the surface of the microstructure is also provided with micropores connected to a gas supply system. The gas supply system is used to output gas to the micropores. The gas flows from the surface of the microstructure into the recessed portion through the micropores, forms a microscale vortex in the recessed portion, and enters the cavity.

[0009] Furthermore, the side wall of the cavity is a conical cylindrical structure, and the diameter of the side close to the opening of the cavity is large, and the diameter of the side close to the bottom wall of the cavity is small.

[0010] Furthermore, the outer surface of the side wall of the aircraft head connected to the cavity is provided with periodically arranged microstructures;

[0011] The gas supply system is used to output gas to the microstructure on the outer surface of the aircraft head.

[0012] Furthermore, the bottom wall of the cavity is provided with periodically arranged microstructures;

[0013] The gas source supply system is used to output gas to the microstructure on the bottom wall of the cavity.

[0014] Furthermore, a lip is formed at the connection between the side wall of the cavity and the outer surface of the aircraft head, and the lip is rounded.

[0015] Furthermore, the height of the microstructure is no higher than the wall surface.

[0016] Furthermore, the gas source supply system includes a gas supply cavity and a gas supply device connected to the gas supply cavity;

[0017] The air supply cavity is arranged inside the aircraft head, and the wall surface of the air supply cavity is communicated with the micropores.

[0018] Furthermore, the gas supply device includes a connecting pipeline, a mass flow controller and a high-pressure gas cylinder arranged in sequence, and the other end of the connecting pipeline is connected to the gas supply cavity.

[0019] Furthermore, the sidewall of the cavity is made of porous medium material;

[0020] The micropores are pores in the porous medium material.

[0021] The present invention also provides a windward concave cavity cooling method based on microstructure seepage, using the above-mentioned windward concave cavity structure based on microstructure seepage, comprising the following steps:

[0022] S1, the gas supply system continuously delivers cooling gas to the micropores;

[0023] S2, cooling gas is ejected from the surface of the microstructure to the recessed portion, and the cooling gas forms a microscale vortex in the recessed portion due to the interaction between the incoming flow and the backflow vortex in the cavity;

[0024] First, all the micro-scale vortices form a micro-scale vortex layer, and the micro-scale vortex layer and the convex part of the microstructure constitute the equivalent wall surface of the side wall of the cavity;

[0025] Then, all the micro-scale vortices enter the cavity, enhancing the backflow vortex in the cavity;

[0026] In S3, affected by the unsteady characteristics of the cavity, the recirculating vortex that has undergone enhancement and heat exchange overflows through the lip to form overflow, and flows downstream along the outer surface of the aircraft head.

[0027] The beneficial effects of the present invention are that, based on the existing windward cavity thermal protection, the present invention adds a microstructure on the side wall of the cavity and transmits gas to the microstructure, thereby achieving the following effects:

[0028] First, the arrangement of the microstructure and micropores gives the sidewall of the cavity a larger specific surface area. The transported gas can conduct more sufficient convection heat exchange with the sidewall of the cavity through the micropores and microstructure, thereby improving the cooling effect of the sidewall of the cavity.

[0029] Secondly, the gas will form micro-scale vortices in the concave parts of the microstructure. The micro-scale vortices formed in a single concave part form a micro-scale vortex layer at the scale of the periodically arranged microstructure. At this time, the incoming flow entering the cavity and the return vortex formed in the cavity will intermittently contact the convex parts of the microstructure and the micro-scale vortex layer. That is, the micro-scale vortex layer constitutes the equivalent wall surface of the side wall of the cavity. At this time, the effective contact area between the incoming flow and return flow vortex and the side wall of the cavity can be greatly reduced, changing the original solid wall boundary condition.

[0030] Third, the seepage of the micro-scale vortex layer and the convex part will form a seepage air film on the side wall of the cavity. The seepage air film effectively isolates the side wall of the cavity from the hot gas of the incoming and return vortices, which can effectively protect the side wall of the cavity from ablation.

[0031] Fourthly, the gas will flow into the inner space of the cavity after forming micro-scale vortices, and interact with the return vortex to increase the strength of the return vortex, driving sufficient heat exchange between the gas and the hot gas inside the cavity, further improving the cooling effect of the cavity;

[0032] Fifth, affected by the unsteady characteristics of the cavity, under the action of the pressure gradient, the gas in the cavity will overflow from the lip of the cavity to form an overflow, and the overflow will flow downstream along the outer surface of the aircraft head. At this time, the temperature of the overflowing gas is lower than the temperature of the outer surface of the aircraft head, which means that it can exchange heat with the outer surface of the aircraft head and serve as an insulating layer to reduce the friction of the incoming flow on the outer surface of the aircraft head.

[0033] In summary, by setting up microstructures in the cavity and transporting gas to the microstructures on the basis of the existing windward cavity thermal protection, the aircraft head can be cooled. The overall structure is simple and has no impact on the original windward cavity thermal protection structure. In addition, the gas working fluid is transported through the micropores, and the required gas working fluid is small. On this basis, the cooling effect on the aircraft head can be significantly improved, so that the temperature of the cavity and the aircraft head is lower, and ablation can be avoided in a long-term environment. BRIEF DESCRIPTION OF THE DRAWINGS

[0034] Attachment Figure 1 It is a structural schematic diagram of the present invention;

[0035] Attachment Figure 2 It is a schematic diagram of the principle of the present invention;

[0036] Attachment Figure 3 It is a partial structural diagram of the gas source supply system in the present invention;

[0037] Attachment Figure 4 Schematic diagram of the microstructure of the present invention;

[0038] Attachment Figure 5 Schematic diagram of a microscale vortex in the present invention;

[0039] Attachment Figure 6 It is a partial enlarged schematic diagram of the microstructure in the present invention;

[0040] Attachment Figure 7 This is a schematic diagram of the layout of the temperature measurement points in the experiment;

[0041] Attachment Figure 8 The temperature variation curve of each train at measuring point 1 in the experiment over time;

[0042] Attachment Figure 9 The temperature variation curve of each train at measuring point 2 in the experiment over time;

[0043] Attachment Figure 10The temperature variation curve of each train at measuring point 3 in the experiment over time;

[0044] Attachment Figure 11 The temperature variation curve of each train at measuring point 4 in the experiment over time;

[0045] Attachment Figure 12 This is the temperature variation curve of each train at measuring point 5 in the experiment over time.

[0046] In the figure, 1-aircraft head; 101-outer surface; 2-cavity; 21-microstructure; 211-recessed portion; 212-convex portion; 22-micropore; 201-side wall; 202-bottom wall; 3-air supply cavity; 4-connecting pipeline; 5-mass flow controller; 6-high-pressure gas cylinder; 7-microscale vortex; 8-recirculation vortex; 9-bow shock wave; 10-overflow. DETAILED DESCRIPTION

[0047] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. All other embodiments obtained by ordinary technicians in this field based on the embodiments of the present invention without making any creative efforts shall fall within the scope of protection of the present invention.

[0048] It should be noted that all directional indications (such as up, down, left, right, front, back, etc.) in the embodiments of the present invention are only used to explain the relative position relationship, movement status, etc. between the various components under a certain specific posture (as shown in the accompanying drawings). If the specific posture changes, the directional indication will also change accordingly.

[0049] In the present invention, unless otherwise specified or limited, the terms "connection" and "fixation" should be understood in a broad sense. For example, "fixation" can mean fixed connection, detachable connection, or integration; it can mean mechanical connection, electrical connection, physical connection, or wireless communication connection; it can mean direct connection or indirect connection through an intermediate medium; it can mean internal communication between two elements or interaction between two elements, unless otherwise specified. For those skilled in the art, the specific meanings of the above terms in the present invention can be understood according to specific circumstances.

[0050] In addition, the technical solutions between the various embodiments of the present invention can be combined with each other, but it must be based on the fact that ordinary technicians in this field can implement it. When the combination of technical solutions is mutually contradictory or cannot be implemented, it should be deemed that such a combination of technical solutions does not exist and is not within the scope of protection required by the present invention.

[0051] As attached Figure 1 -Attached Figure 12 As shown, the present invention provides a windward concave cavity structure based on microstructure seepage, comprising an air source supply system and a concave cavity 2 provided at the head 1 of an aircraft, wherein the head 1 of the aircraft can be the nose portion of the aircraft, or the head portion of a needle-shaped shock wave rod axially mounted on the nose of the aircraft, that is, the head 1 of the aircraft is the first portion of the aircraft to contact the incoming flow, and the concave cavity 2 is provided on the head 1 of the aircraft, and the concave cavity 2 is preferably arranged facing the incoming flow, and the shock wave oscillation characteristics can be used to reduce aerodynamic heating in the stagnation area, thereby forming thermal protection for the windward concave cavity;

[0052] For conventional windward cavity thermal protection, the shock wave oscillations of cavity 2 in the supersonic flow field are the source of its thermal protection effect. The violent longitudinal pressure fluctuations within cavity 2 cause longitudinal oscillations in the shock wave position, which oscillates between the cavity lip and a certain position upstream. When the shock wave moves toward the windward cavity lip, the flow within cavity 2 is primarily aligned with the incoming flow. The space between the bow shock wave 9 and the cavity 2 wall narrows, the shock wave compression becomes severe, and the flow parameters on both sides of the shock wave change more dramatically. Under unchanged free flow conditions, the Mach number of the afterflow decreases as the shock wave approaches the wall, while the temperature increases as the shock wave approaches the wall. The real gas constant-pressure specific heat increases, and under unchanged total enthalpy, the stagnation temperature of the afterflow decreases. The reduced energy carried by the afterflow weakens the aerodynamic heating of the aircraft nose 1. As the shock wave moves upstream, the airflow primarily moves in the opposite direction of the incoming flow. The space between the bow shock wave 9 and the wall of the cavity 2 widens, weakening the compression of the shock wave. The stagnation temperature behind the wave rises. However, as the shock wave moves upstream, the excess space between the shock wave and the cavity wall is filled by the downstream airflow. The downstream gas is drawn between the shock wave and the wall, separating the flow to form a recirculation vortex 8. The total temperature of the downstream air is lower than that of the downstream air behind the wave. The introduction of downstream air and the formation of the recirculation zone protect the outer surface 101 of the aircraft nose 1 from the effects of extremely hot air, reducing aerodynamic heating in the stagnation area of ​​the aircraft nose 1.

[0053] The concave cavity 2 includes an opening, a side wall 201 and a bottom wall 202. The opening is arranged toward the incoming flow, the bottom wall 202 is arranged opposite to the opening, and the side wall 201 connects the bottom wall 202 and the opening. The side wall 201 of the concave cavity 2 is provided with a periodically arranged microstructure 21. Preferably, the microstructure 21 is a strip structure. The length direction of the microstructure 21 of the strip structure is perpendicular to the incoming flow direction of the aircraft, that is, the microstructure 21 of the strip structure is an arc or a circular strip. When the microstructure 21 arranged in the side wall 201 of the concave cavity 2 is ring-shaped, multiple microstructures 21 is arranged in an array along the incoming flow direction, so that the microstructure 21 is arranged in a periodic arrangement; when the microstructure 21 provided in the side wall 201 of the cavity 2 is in an arc shape, multiple microstructures 21 are arranged in a circular array with the axis of the cavity 2, and are arranged in an array along the incoming flow direction, so that the microstructure 21 is arranged in a periodic arrangement. It should be noted that the structure of the microstructure 21 can be two-dimensional or three-dimensional, such as grooves, dune-shaped protrusions, bionic shark skin microgrooves, etc., and only the concave portion 211 that can form the microscale vortex 7 is required. For a specific The structural design of the microstructure 21 is not an improvement of the present invention and will not be described in detail here. The microstructure 21 has a recessed portion 211, and correspondingly, the microstructure 21 has a raised portion 212. The height of the raised portion 212 is higher than the recessed portion 211. Specifically, the recessed portion 211 can be formed by enclosing the raised portions 212 on one microstructure 21 or by enclosing the raised portions 212 on adjacent microstructures 21. The surface of the microstructure 21 is also provided with micropores 22 connected to the gas supply system. The micropores 22 are preferably arranged The surface of the recessed portion 211 can also be provided on the surface of the raised portion 212. The gas source supply system is used to output gas to the micropores 22. The gas is not limited. The temperature of the gas can be lower than the temperature of the outer surface 101 of the aircraft head 1 to be cooled. Compared with conventional heavy gas injection control, the present invention does not rely on high-temperature vibration excitation of heavy gas, has more selectivity and lower cost. The gas flows from the surface of the microstructure 21 through the micropores 22 into the recessed portion 211, and forms a microscale vortex 7 in the recessed portion 211 and enters the cavity 2.

[0054] On the basis of existing windward cavity thermal protection, the present invention adds a microstructure 21 on the sidewall 201 of the cavity 2 and delivers gas to the microstructure 21, thereby achieving the following effects:

[0055] First, the arrangement of the microstructure 21 and the micropores 22 allows the sidewall 201 of the cavity 2 to have a larger specific surface area. The transported gas can pass through the micropores 22 and the microstructure 21 and conduct more sufficient convection heat exchange with the sidewall 201 of the cavity 2, thereby improving the cooling effect of the sidewall 201 of the cavity 2.

[0056] Secondly, the gas will form a microscale vortex 7 in the recessed portion 211 of the microstructure 21. The microscale vortex 7 formed in a single recessed portion 211 forms a microscale vortex layer at the scale of the periodically arranged microstructure 21. At this time, the incoming flow entering the cavity 2 and the return vortex 8 formed in the cavity 2 will intermittently contact the raised portion 212 of the microstructure 21 and the microscale vortex layer, that is, the microscale vortex layer constitutes the equivalent wall surface of the side wall 201 of the cavity 2. At this time, the effective contact area between the incoming flow and the return vortex 8 and the side wall 201 of the cavity 2 can be greatly reduced, changing the original solid wall boundary condition.

[0057] Third, the seepage of the micro-scale vortex layer and the protrusion 212 will form a seepage air film on the side wall 201 of the cavity 2. The seepage air film effectively isolates the side wall 201 of the cavity 2 from the hot gas of the incoming and return vortices 8, and can effectively protect the side wall 201 of the cavity 2 from ablation.

[0058] Fourthly, after forming the micro-scale vortex 7, the gas will flow into the inner space of the cavity 2 and interact with the recirculation vortex 8 to increase the strength of the recirculation vortex 8, driving the gas and the hot gas inside the cavity 2 to fully exchange heat, further improving the cooling effect of the cavity 2. Specifically, conventional linear jet technology has strict restrictions on the pressure ratio before and after the nozzle, and generally due to the fast ejection speed, it cannot be fully mixed with the incoming hot gas in the cavity, and the mass of the jet working fluid is large. This seepage forms a micro-scale vortex 7, which can be fully mixed with the incoming hot gas in the recirculation zone of the cavity 2 in a small area. Due to the characteristics of rotation, it can be more fully mixed, and the required seepage flow rate is small. At the same time, it will combine with the recirculation vortex 8 during the heat exchange process, increasing the strength of the recirculation vortex 8 and further improving the heat exchange effect.

[0059] Fifth, influenced by the unsteady characteristics of cavity 2 and the pressure gradient, the gas within cavity 2 overflows from its lip (arc c) to form overflow 10. Overflow 10 flows downstream along the outer surface 101 of the vehicle head 1. At this time, if the outer surface 101 of the vehicle head 1 lacks microstructures 21, the temperature of the gas in overflow 10 is lower than that of the outer surface 101 of the vehicle head 1 and the bow shock wave 9. This allows heat exchange with the outer surface 101 of the vehicle head 1 and acts as a thermal insulation layer to reduce friction between the incoming flow (including the bow shock wave 9) and the outer surface 101 of the vehicle head 1. When the outer surface 101 of the vehicle head 1 has microstructures 21, the temperature of the gas in overflow 10 is higher than the seepage film on the outer surface 101, but lower than the temperature of the bow shock wave 9, thereby preventing the bow shock wave 9 from transferring heat to the seepage film.

[0060] In summary, by arranging a microstructure 21 in the cavity 2 and delivering gas to the microstructure 21 on the basis of the existing windward cavity thermal protection, the aircraft head 1 can be cooled. The overall structure is simple and has no effect on the original windward cavity thermal protection structure. In addition, the gas working fluid is delivered through the micropores 22, and the required gas working fluid is small. On this basis, the cooling effect on the aircraft head 1 can be significantly improved, so that the temperature of the cavity 2 and the aircraft head 1 is lower, and ablation can be avoided in a long-term environment.

[0061] It should be noted that the reference Figure 1 In the present application, various dimensions of the cavity 2, such as the depth L of the cavity 2, the bottom diameter D of the cavity 2, the radius C of the lip of the cavity 2, and the expansion angle θ of the side wall 201 of the cavity 2 can be adjusted according to different flow conditions and can be selected according to actual needs.

[0062] In one embodiment, the side wall 201 of the cavity 2 is a conical cylindrical structure, and the diameter is larger on the side close to the opening of the cavity 2, and smaller on the side close to the bottom wall 202 of the cavity 2, that is, the side wall 201 of the cavity 2 has an expansion angle θ relative to the axial direction of the cavity 2, which is more conducive to the outflow of the overflow 10, that is, after the gas seeps out through the microstructure 21, it can be discharged from the cavity 2 area in a timely manner in the form of overflow 10 after heat exchange with the hot gas, which can further improve the cooling effect.

[0063] In one embodiment, the outer surface 101 of the aircraft head 1, which connects to the sidewall 201 of the cavity 2, is provided with periodically arranged microstructures 21. Similarly, the microstructures 21 have recessed portions 211. The gas supply system is used to deliver gas to the microstructures 21 on the outer surface 101 of the aircraft head 1. In this embodiment, the microstructures 21 are also provided on the outer wall of the aircraft head 1, which has the following effects:

[0064] First, the outer surface 101 of the aircraft head 1 has a larger specific surface area, and the transported gas can more fully conduct convective heat exchange with the outer surface 101 of the aircraft head 1 through the micropores 22 and the microstructures 21, thereby improving the cooling effect of the outer surface 101 of the aircraft head 1;

[0065] Secondly, the gas will form micro-scale vortices 7 in the recessed portions 211 of the microstructure 21. The micro-scale vortices 7 formed in a single recessed portion 211 form a micro-scale vortex layer at the scale of the periodically arranged microstructure 21. At this time, the incoming flow contacting the outer surface 101 of the aircraft head 1 will intermittently contact the raised portions 212 of the microstructure 21 and the micro-scale vortex layer, that is, the micro-scale vortex layer constitutes the equivalent wall surface of the outer surface 101 of the aircraft head 1. At this time, the effective contact area between the incoming flow and the outer surface 101 of the aircraft head 1 can be greatly reduced, changing the original solid wall boundary condition.

[0066] Third, the micro-scale vortex layer will form a seepage air film on the outer surface 101 of the aircraft head 1. The seepage air film will isolate the outer surface 101 of the aircraft head 1 from the hot gas, effectively protecting the outer surface 101 of the aircraft head 1 from ablation.

[0067] Fourthly, the overflow 10 flows outside the seepage air film, further protecting the aircraft head 1 from ablation.

[0068] In one embodiment, the bottom wall 202 of the cavity 2 is provided with periodically arranged microstructures 21; the gas supply system is configured to deliver gas to the microstructures 21 of the micropores 22 of the bottom wall 202 of the cavity 2. In this embodiment, this provides the same effect as the periodically arranged microstructures 21 provided on the sidewalls 201 of the cavity 2, enhancing the cooling effect of the bottom wall 202 of the cavity 2 and further protecting the bottom wall 202 of the cavity 2 from ablation.

[0069] In one embodiment, a lip is formed at the connection between the side wall 201 of the cavity 2 and the outer surface 101 of the aircraft head 1. The lip is rounded to facilitate the overflow 10 from the cavity 2 to flow in a guided manner toward the outer surface 101 of the aircraft head 1, thereby reducing air resistance, reducing stress concentration, and improving structural strength.

[0070] In one embodiment, the height of the microstructure 21 is no higher than the wall surface. That is, the raised portion 212 of the microstructure 21 cannot be higher than the outer surface 101 of the aircraft head 1, the sidewalls 201 of the cavity 2, and the bottom wall 202 of the cavity 2. This prevents the raised portion 212 of the microstructure 21 from creating resistance and affecting the aerodynamic shape of the wall surface. When the microstructure 21 is fully distributed over the outer surface 101 of the aircraft head 1, the sidewalls 201 of the cavity 2, and the bottom wall 202 of the cavity 2, the raised portion 212 of the microstructure 21 will correspond to the outer surface 101 of the aircraft head 1, the sidewalls 201 of the cavity 2, and the bottom wall 202 of the cavity 2.

[0071] In one embodiment, refer to the attached Figure 1 , the gas source supply system includes a gas supply cavity 3 and a gas supply device connected to the gas supply cavity 3;

[0072] The air supply cavity 3 is disposed within the vehicle head 1, and the wall of the air supply cavity 3 is connected to the micropores 22. Preferably, when microstructures 21 are provided on both the outer surface 101 of the vehicle head 1 and the bottom wall 202 of the cavity 2, the air supply cavity 3 is simultaneously connected to all microstructures 21, thereby simultaneously achieving seepage cooling on all surfaces. Preferably, the vehicle head 1 is a wall structure, i.e., the outer surface 101 of the vehicle head 1, the bottom wall 202 of the cavity 2, and the sidewalls 201 of the cavity 2 have the same thickness, and the inner wall of the wall serves as the cavity wall of the air supply cavity 3.

[0073] In one embodiment, the gas supply device includes a connecting pipe 4, a mass flow controller 5 and a high-pressure gas cylinder 6 arranged in sequence. The other end of the connecting pipe 4 is connected to the gas supply chamber 3. In this embodiment, the gas flow delivered to the gas supply chamber 3 can be adjusted by the mass flow controller 5, thereby realizing real-time adjustment of the seepage flow according to actual flow needs.

[0074] In one embodiment, refer to the attached Figure 4 The sidewall 201 of the cavity 2 is made of a porous medium material; the micropores 22 are pores in the porous medium material. In this embodiment, there is no need to separately process the micropores 22, which simplifies the processing difficulty. This embodiment is suitable for the embodiment where the microstructure 21 is a groove, that is, the recessed portion 211 occupies the entire surface of the microstructure 21. In this embodiment, the micropores 22 are irregular in shape. Of course, the wall surface can also be made of an ordinary non-porous material, and the microstructure 21 and micropores 22 can be processed on the non-porous material. For example, the microstructure 21 and micropores 22 can be processed by laser drilling or 3D printing. The structure of the micropores 22 can also be a regular cylindrical shape or other irregular shapes.

[0075] The present invention also provides a windward concave cavity cooling method based on microstructure seepage, using the above-mentioned windward concave cavity structure based on microstructure seepage, comprising the following steps:

[0076] S1, the gas supply system continuously delivers cooling gas to the micropores 22;

[0077] S2, cooling gas is ejected from the surface of the microstructure 21 to the recessed portion 211, and the cooling gas forms a microscale vortex 7 in the recessed portion 211 under the interaction of the incoming flow and the backflow vortex 8 in the cavity 2;

[0078] First, all the micro-scale vortices 7 form a micro-scale vortex layer, and the micro-scale vortex layer and the protrusion 212 of the microstructure 21 constitute an equivalent wall surface of the side wall 201 of the cavity 2;

[0079] Then, all the micro-scale vortices 7 enter the cavity 2, enhancing the return vortex 8 in the cavity 2;

[0080] In this step, the arrangement of the microstructure 21 and the micropores 22 makes the side wall 201 of the cavity 2 have a larger specific surface area at this time, and the transported gas will pass through the micropores 22 and the microstructure 21 to conduct more sufficient convection heat exchange with the side wall 201 of the cavity 2, thereby improving the cooling effect of the side wall 201 of the cavity 2; in addition, the gas will form microscale vortices 7 in the recessed portion 211 of the microstructure 21, and the microscale vortex 7 formed in a single recessed portion 211 forms a microscale vortex layer at the scale of the periodically arranged microstructure 21. At this time, the incoming flow entering the cavity 2 and the return vortex 8 formed in the cavity 2 will intermittently contact the raised portion 212 of the microstructure 21 and the microscale vortex layer, that is, the microscale vortex layer constitutes the equivalent wall surface of the side wall 201 of the cavity 2. At this time, the effective contact area between the incoming flow and the return vortex 8 and the side wall 201 of the cavity 2 can be greatly reduced, thereby changing the original solid wall boundary condition. In addition, the seepage of the micro-scale vortex layer and the protrusion 212 will form a seepage air film on the side wall 201 of the cavity 2. The seepage air film effectively separates the side wall 201 of the cavity 2 from the hot gas of the incoming flow and the return vortex 8, and can effectively protect the side wall 201 of the cavity 2 from ablation. In addition, after forming the micro-scale vortex 7, the gas will flow into the internal space of the cavity 2 and interact with the return vortex 8 to increase the strength of the return vortex 8, thereby driving the gas and the hot gas inside the cavity 2 to fully exchange heat, further improving the cooling effect of the cavity 2.

[0081] S3, affected by the unsteady characteristics of the cavity 2, the backflow vortex 8 after enhancement and heat exchange overflows through the lip to form an overflow 10, and the overflow 10 adheres to the outer surface 101 of the aircraft head 1 to form a flow downstream.

[0082] In this step, affected by the unsteady characteristics of the cavity 2, under the action of the pressure gradient, the gas in the cavity 2 will overflow from the lip of the cavity 2 to form an overflow 10, and the overflow 10 flows downstream along the outer surface 101 of the aircraft head 1. At this time, the gas temperature of the overflow 10 is lower than the temperature of the outer surface 101 of the aircraft head 1, that is, it can exchange heat with the outer surface 101 of the aircraft head 1, and can also serve as an insulating layer to reduce the friction of the incoming flow on the outer surface 101 of the aircraft head 1.

[0083] In the embodiment in which the outer surface 101 of the aircraft head 1 has a microstructure 21, the microstructure 21 of the outer surface 101 of the aircraft head 1 makes the outer surface 101 of the aircraft head 1 have a larger specific surface area, and the transported gas will conduct more sufficient convection heat exchange with the outer surface 101 of the aircraft head 1 through the micropores 22 and the microstructure 21, thereby improving the cooling effect of the outer surface 101 of the aircraft head 1; in addition, the gas will form microscale vortices 7 in the recessed portions 211 of the microstructure 21, and the microscale vortices 7 formed in a single recessed portion 211 will form a microscale vortex layer at the scale of the periodically arranged microstructure 21. At this time, the incoming flow that contacts the outer surface 101 of the aircraft head 1 will intermittently contact the raised portions 212 of the microstructure 21 and the microscale vortex layers. The layer, that is, the micro-scale vortex layer constitutes the equivalent wall surface of the outer surface 101 of the aircraft head 1. At this time, the effective contact area between the incoming flow and the outer surface 101 of the aircraft head 1 can be greatly reduced, changing the original solid wall boundary condition; in addition, the micro-scale vortex layer will form a seepage air film on the outer surface 101 of the aircraft head 1, and the seepage air film will separate the outer surface 101 of the aircraft head 1 from the hot gas, effectively protecting the outer surface 101 of the aircraft head 1 from ablation; in addition, the overflow 10 of the cavity 2 flows on the outside of the seepage air film. At this time, the gas temperature of the overflow 10 is higher than the seepage air film on the outer surface 101, but lower than the temperature of the bow shock wave 9, thereby preventing the bow shock wave 9 from transferring heat to the seepage air film, further improving the protection of the aircraft head 1 from ablation.

[0084] In the embodiment where the bottom wall 202 of the cavity 2 has the microstructure 21 , the microstructure 21 provided on the bottom wall 202 of the cavity 2 further improves the heat exchange capacity in the cavity 2 .

[0085] The present invention has been proven to be effective through experiments. Specifically, the experimental data are as follows:

[0086] In the hypersonic high temperature wind tunnel ( ), under the conditions of high total temperature and high total pressure, the temperature distribution of the bottom wall 202 of the seepage cavity 2 is obtained at different seepage flow rates.

[0087] See the attached diagram for temperature measurement points Figure 7 .

[0088] The experimental train numbers are shown in Table 1.

[0089] Table 1 Temperature distribution measurement experiment of seepage cavity 2 bottom wall 202

[0090]

[0091] The whole process of wind tunnel operation was recorded by Schlieren, and the flow field started normally. Figure 8 -Attached Figure 12, which are the temperature variation curves of each train at the five measuring points over time. It can be seen that under the action of seepage, the cooling effect is obvious, and the cooling effect is better as the seepage flow rate increases.

[0092] The above description is merely an embodiment and does not limit the present invention in any way. Any person skilled in the art can, without departing from the scope of the technical solution of the present invention, make many possible changes, modifications, or modifications to the technical solution of the present invention using the technical content disclosed above to make equivalent embodiments of equivalent changes. Therefore, any simple modifications, equivalent changes, and modifications made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solution of the present invention shall fall within the scope of protection of the technical solution of the present invention.

Claims

1. A windward cavity structure based on microstructure seepage, characterized by: It comprises an air source supply system and a concave cavity (2) arranged at the head of the aircraft (1); The sidewall (201) of the cavity (2) is provided with a periodically arranged microstructure (21), the microstructure (21) having a recessed portion (211), and the surface of the microstructure (21) is further provided with micropores (22) connected to a gas supply system, the gas supply system being used to output gas to the micropores (22), the gas flowing from the surface of the microstructure (21) into the recessed portion (211) through the micropores (22), and forming a microscale vortex (7) in the recessed portion (211) and entering the cavity (2); The side wall (201) of the cavity (2) is a conical cylindrical structure, and the side close to the opening of the cavity (2) has a larger diameter, and the side close to the bottom wall (202) of the cavity (2) has a smaller diameter; An outer surface (101) of a side wall (201) of the aircraft head (1) connected to the cavity (2) is provided with periodically arranged microstructures (21); The gas source supply system is used to output gas to the microstructure (21) on the outer surface (101) of the aircraft head (1); The bottom wall (202) of the cavity (2) is provided with periodically arranged microstructures (21); The gas source supply system is used to output gas to the microstructure (21) of the bottom wall (202) of the cavity (2); The gas source supply system comprises a gas supply cavity (3) and a gas supply device connected to the gas supply cavity (3); The air supply cavity (3) is arranged inside the aircraft head (1), and the wall surface of the air supply cavity (3) is in communication with the micropores (22).

2. The windward cavity structure based on microstructure seepage according to claim 1, characterized in that: A lip is formed at the connection between the side wall (201) of the cavity (2) and the outer surface (101) of the aircraft head (1), and the lip is rounded.

3. The windward concave cavity structure based on microstructure seepage according to claim 1 or 2, characterized in that: The height of the microstructure (21) is no higher than the wall surface.

4. The windward cavity structure based on microstructure seepage according to claim 1, characterized in that: The gas supply device comprises a connecting pipeline (4), a mass flow controller (5) and a high-pressure gas cylinder (6) which are arranged in sequence, and the other end of the connecting pipeline (4) is connected to the gas supply cavity (3).

5. The windward cavity structure based on microstructure seepage according to claim 1 or 2, characterized in that: The side wall (201) of the cavity (2) is made of porous medium material; The micropores (22) are pores in the porous medium material.

6. A windward cavity cooling method based on microstructure seepage, characterized in that: Using the windward concave cavity structure based on microstructure seepage as described in any one of claims 1 to 5 comprises the following steps: S1, the gas supply system continuously delivers cooling gas to the micropores (22); S2, cooling gas is ejected from the surface of the microstructure (21) to the recessed portion (211), and the cooling gas forms a microscale vortex (7) in the recessed portion (211) under the interaction of the incoming flow and the return vortex (8) in the cavity (2); First, all the micro-scale vortices (7) form a micro-scale vortex layer, and the micro-scale vortex layer and the protrusion (212) of the microstructure (21) constitute an equivalent wall surface of the side wall (201) of the cavity (2); Then, all the micro-scale vortices (7) enter the cavity (2), enhancing the return vortex (8) in the cavity (2); S3, affected by the unsteady characteristics of the cavity (2), the backflow vortex (8) after enhancement and heat exchange overflows through the lip to form overflow (10), and adheres to the outer surface (101) of the aircraft head (1) and flows downstream.

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