Active Cooling Structure for the Sharp Leading Edge of a High-Speed Aircraft with Designable Microchannels

Abstract

The present invention discloses an active cooling structure for the sharp leading edge of a high-speed aircraft with a designable microchannel, which includes a microchannel material layer with a sharp leading edge, and a coolant cavity is arranged therein; a first microchannel with a slit structure is arranged on the microchannel material layer, the first end of the first microchannel is communicated with the coolant cavity, and the second end is located at the position of the leading edge stagnation point; a plurality of cylindrical second microchannels are arranged on the microchannel material layer, the first end of the second microchannel is communicated with the coolant cavity, and the second ends are distributed at intervals on the outer wall of the microchannel material layer; along the chord direction of the microchannel material layer, the flow channel cross-sectional areas of the second microchannels decrease in sequence. The present invention is applied to the thermal protection field of aerospace vehicles. By analyzing the pressure and heat flux distribution of the external flow environment, precisely designed cooling microchannels are constructed inside the leading edge near the stagnation point, so that the coolant distributes the flow rate according to the set path, and a more effective and uniform cooling effect is achieved on the premise of better meeting the structural stability.

Description

Technical Field

[0001] The invention relates to the technical field of thermal protection of aerospace vehicles, and in particular to a sharp leading edge active cooling structure of a high-speed vehicle containing a designable microchannel. Background Art

[0002] During high-speed flight, the leading edge of an aerospace vehicle will be subjected to extreme heat flow shocks, causing the structural temperature to rise sharply, even up to thousands of degrees Celsius. This high temperature environment will not only cause the degradation of material properties, but may also lead to the destruction of structural integrity, seriously affecting the safety and sustainable flight capability of the aircraft. Therefore, how to effectively reduce the heat load on the leading edge and improve the cooling efficiency of the thermal protection system is a key issue that needs to be solved in the design of hypersonic aircraft.

[0003] From the perspective of aerodynamic design, reducing the leading edge radius helps to reduce wave resistance, but it will significantly increase the peak heat flux, which in turn affects the thermal environment. Existing studies have shown that the heat flux density of a sharp leading edge is proportional to the inverse of the square of the radius. Therefore, when the leading edge radius is in the range of 1 mm or even micrometers, the heat flux density can reach extremely high levels, which makes it difficult for traditional passive thermal protection methods (such as high-temperature heat-resistant materials and ablative materials) to withstand high temperature shocks for a long time. The heat flux density and pressure at the stagnation point of the sharp leading edge are much higher than those in other areas, and vary dramatically along the way, which leads to uneven cooling effects of conventional sweat cooling, and even further exacerbates the problem of local overheating.

[0004] At present, porous material sweat cooling is a common active thermal protection solution, which fully exchanges heat between the internal porous medium and the coolant, and at the same time penetrates the surface to form a layer of air film to reduce the surface temperature. However, in the application of sharp leading edges, this technology still has certain limitations, mainly including:

[0005] Insufficient structural strength: Porous materials with high porosity have difficulty maintaining sufficient mechanical strength at the leading edge with a small radius; Excessive coolant consumption and uneven cooling: There is insufficient coolant supply in the stagnation area, while the trailing edge may be over-cooled, resulting in poor overall thermal protection effect. Summary of the invention

[0006] In view of the deficiencies in the above-mentioned prior art, the present invention provides an active cooling structure for a sharp leading edge of a high-speed aircraft containing a designable microchannel. Different from traditional porous materials, the active cooling structure constructs a precisely designed cooling microchannel inside the leading edge near the stagnation point according to the distribution of heat flow and surface pressure, so that the coolant flows along a set path, and the coolant flow rate is controlled by the microchannel cross-section, thereby achieving a more effective and uniform cooling effect while better satisfying the structural stability.

[0007] To achieve the above object, the present invention provides an active cooling structure for the sharp leading edge of a high-speed aircraft with a designable microchannel, which includes a microchannel material layer with a sharp leading edge, and a coolant cavity is provided in the microchannel material layer;

[0008] A first microchannel with a slit structure is provided on the microchannel material layer. The first end of the first microchannel communicates with the coolant cavity, and the second end of the first microchannel is located at a position corresponding to the leading edge stagnation point on the microchannel material layer;

[0009] A plurality of second microchannels with cylindrical structures are further provided on the microchannel material layer. The first end of each second microchannel communicates with the coolant cavity; in the spanwise and chordwise directions of the microchannel material layer, the second ends of each second microchannel are spaced apart and distributed on the outer wall of the microchannel material layer;

[0010] The flow channel cross-sectional areas of the second microchannels located in the same spanwise direction of the microchannel material layer are the same, and along the chordwise direction of the microchannel material layer, the flow channel cross-sectional areas of each second microchannel gradually decrease.

[0011] In one embodiment, along the direction from the first end to the second end, the flow channel cross-sectional area of the first microchannel gradually decreases.

[0012] In one embodiment, a throttling hole is provided at the first end of the second microchannel;

[0013] Along the direction from the first end to the second end, the flow channel cross-sectional area of the second microchannel gradually increases.

[0014] In one embodiment, the second microchannel is a cylindrical channel.

[0015] In one embodiment, in the chordwise direction of the microchannel material layer, the first end of the first second microchannel directly communicates with the coolant cavity, and the first end of the i-th second microchannel is opened on the channel wall of the (i - 1)-th second microchannel;

[0016] Wherein, i = 2 to N, and N is the number of second microchannels along the chordwise direction on the upper wall or lower wall of the microchannel material layer.

[0017] In one embodiment, according to the heat flux and pressure distribution, the outer surface of the leading edge region of the microchannel material layer is divided into a high heat flux and high pressure region, a medium heat flux and medium pressure region, and a low heat flux and low pressure region;

[0018] The high heat flux and high pressure region is located in the region corresponding to the leading edge stagnation point. The medium heat flux and medium pressure region is located on the upper and lower walls of the microchannel material layer and is adjacent to the high heat flux and high pressure region. The low heat flux and low pressure region is located on the upper and lower walls of the microchannel material layer and is adjacent to the medium heat flux and medium pressure region.

[0019] Define the boundary line between the high heat flux and high pressure region and the medium heat flux and medium pressure region as the first boundary, define the boundary line between the medium heat flux and medium pressure region and the low heat flux and low pressure region as the second boundary, and define the leading edge boundary on the microchannel material layer as the third boundary.

[0020] On the upper or lower wall of the microchannel material layer, the number of the second microchannels along the chordal direction is three. The second end of the first second microchannel is located at the first boundary, the second end of the second second microchannel is located at the second boundary, and the second end of the third second microchannel is located at the third boundary.

[0021] In one embodiment, the included angle between the outflow direction of the second end of each second microchannel and the oncoming flow direction is less than .

[0022] In one embodiment, the microchannel material layer is made of a nickel-based superalloy material.

[0023] In one embodiment, the coolant in the coolant cavity is liquid water.

[0024] Compared with the prior art, the present invention has the following beneficial technical effects:

[0025] 1. The present invention arranges the first microchannel with a slit structure in the stagnation point region of the sharp leading edge, and at the same time arranges the second microchannels with cylindrical structures and gradually decreasing flow channel cross-sectional areas in other regions of the leading edge according to the heat flux density. Through the stepped distribution of the microchannel design, the coolant flow rate ratio can be accurately distributed according to the heat flux and pressure characteristics of different regions, ensuring the effective cooling of the stagnation point of the sharp leading edge of the aircraft and the efficient uniform temperature cooling of the adjacent area.

[0026] 2. In the preferred embodiment of the present invention, the first microchannel with a slit structure is designed as a tapered flow channel, so that the coolant can be reasonably distributed according to the heat flux through the diversion design in the high heat flux region of the stagnation point, greatly reducing the amount of coolant used and improving the heat transfer efficiency of the coolant.

[0027] 3. In the preferred embodiment of the present invention, the included angle between the outflow direction of the second end of the second microchannel and the oncoming flow direction is designed to be less than 90°, so that the coolant forms a gas film on the surface of the aircraft, raising the boundary layer, and further effectively isolating the heat flux transfer, further improving the cooling efficiency and ensuring the stability of the surface temperature of the aircraft.

[0028] 4. In a preferred embodiment of the present invention, a nickel-based superalloy material is used to make the microchannel material layer, so that the active cooling result has excellent high-temperature resistance and corrosion resistance, and can operate stably for a long time in a high-temperature environment. BRIEF DESCRIPTION OF THE DRAWINGS

[0029] In order to more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the drawings required for the description of the embodiments or the prior art. Obviously, the drawings in the following description are only some embodiments of the present invention. For those of ordinary skill in the art, without creative efforts, other drawings can be obtained based on the structures shown in these drawings.

[0030] Figure 1 Is an axonometric view of the active cooling structure in the embodiment of the present invention;

[0031] Figure 2 Is a sectional view of the active cooling structure in the embodiment of the present invention;

[0032] Figure 3 Is a schematic diagram of the heat flux density and surface pressure distribution near the stagnation point in the embodiment of the present invention.

[0033] Reference numerals in the drawings: microchannel material layer 1, coolant cavity 2, first microchannel 3, second microchannel 4, high heat flux and high pressure region 5, medium heat flux and medium pressure region 6, low heat flux and low pressure region 7, first boundary 8, second boundary 9, third boundary 10, throttle hole 11.

[0034] The realization of the object, functional features and advantages of the present invention will be further described with reference to the embodiments and the drawings. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0035] The following will clearly and completely describe the technical solutions in the embodiments of the present invention with reference to the drawings in the embodiments of the present invention. Obviously, the described embodiments are only some embodiments of the present invention, rather than all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative efforts belong to the scope of protection of the present invention.

[0036] It should be noted that all directional indications (such as up, down, left, right, front, back...) in the embodiments of the present invention are only used to explain the relative position relationship and movement conditions between components in a specific posture (as shown in the drawings). If the specific posture changes, the directional indications will also change accordingly.

[0037] In addition, in the present invention, descriptions such as "first" and "second" are for descriptive purposes only, and should not be construed as indicating or implying their relative importance or implicitly specifying the quantity of the indicated technical features. Thus, features defined with "first" and "second" may explicitly or implicitly include at least one such feature. In the description of the present invention, "a plurality of" means at least two, such as two, three, etc., unless otherwise specifically and clearly defined.

[0038] In the present invention, unless otherwise clearly specified and defined, terms such as "connected" and "fixed" shall be understood in a broad sense. For example, "fixed" may be a fixed connection, a detachable connection, or integrated; it may be a mechanical connection, an electrical connection, a physical connection or a wireless communication connection; it may be directly connected, or indirectly connected through an intermediate medium, and may be the internal communication of two components or the interaction relationship between two components, unless otherwise clearly defined. For those of ordinary skill in the art, the specific meanings of the above terms in the present invention can be understood according to specific circumstances.

[0039] In addition, the technical solutions between various embodiments of the present invention can be combined with each other, but it must be based on the ability of those of ordinary skill in the art to implement. When the combination of technical solutions results in contradictions or cannot be implemented, it should be considered that such a combination of technical solutions does not exist and is not within the scope of protection required by the present invention.

[0040] This embodiment discloses a high-speed aircraft sharp leading-edge active cooling structure with a designable microchannel (hereinafter referred to as "active cooling structure"). By optimizing the geometric distribution of the cooling channels on the leading edge, the coolant can be directionally transported in the leading-edge area according to different heat flux density requirements, thereby effectively reducing the high-temperature peak value, balancing the cooling effect, improving the heat resistance and structural safety of the aircraft, effectively coping with the extreme heat load in the stagnation area of the sharp leading edge of the high-speed aircraft, and significantly improving the survivability and mission execution ability of the aircraft in a high-temperature environment.

[0041] Reference Figure 1, in this embodiment, the active cooling structure mainly includes a microchannel material layer 1 with a sharp leading edge and a coolant cavity 2 arranged in the microchannel material layer 1. The coolant in the coolant cavity 2 is liquid water. The coolant cavity 2 is connected to a pressurized delivery system, and through an accurate coolant flow control system, the coolant flow can be determined according to different flight conditions, ensuring maximum cooling effect and reducing resource waste. Among them, a first microchannel 3 with a slit structure is provided on the microchannel material layer 1. The first end of the first microchannel 3 communicates with the coolant cavity 2, and the second end of the first microchannel 3 is located at a position corresponding to the leading edge stagnation point on the microchannel material layer 1. A plurality of second microchannels 4 with cylindrical structures are also provided on the microchannel material layer 1. The first end of each second microchannel 4 communicates with the coolant cavity 2; the second ends of each second microchannel 4 are distributed at intervals in the spanwise and chordwise directions on the outer wall of the microchannel material layer 1. At the same time, each second microchannel 4 is designed correspondingly according to the heat flux density, that is, the flow channel cross-sectional areas of the second microchannels 4 located in the same span of the microchannel material layer 1 are the same, and along the chordwise direction of the microchannel material layer 1, the flow channel cross-sectional areas of each second microchannel 4 decrease in sequence. Among them, taking the active cooling structure on the wing as an example, the spanwise direction refers to the direction from the wing root to the wing tip, and the chordwise direction refers to the direction from the leading edge to the trailing edge.

[0042] In this embodiment, the active cooling structure arranges a first microchannel 3 with a slit structure in the stagnation point area of the leading edge, and at the same time arranges a second microchannel 4 with a cylindrical structure and a gradually decreasing flow channel cross-sectional area in other areas of the leading edge according to the heat flux density. Through the stepped distribution of the microchannel design, the coolant flow ratio can be accurately allocated according to the heat flux and pressure characteristics of different regions, enabling the coolant to increase the flow supply in the high heat flux region and maintain an appropriate coolant flow in the low heat flux region, so as to improve the cooling efficiency and reduce unnecessary coolant loss, ensuring efficient and uniform temperature cooling of each region at the leading edge of the aircraft.

[0043] Reference Figure 2 , in this embodiment, the outer surface of the leading edge area of the microchannel material layer 1 is divided into a high heat flux and high pressure area 5, a medium heat flux and medium pressure area 6, and a low heat flux and low pressure area 7 according to the heat flux distribution. Among them, the high heat flux and high pressure area 5 is located in the area corresponding to the leading edge stagnation point. This area is near the stagnation point, where the heat flux and pressure are the largest, and the coolant flow demand is also the most vigorous; the medium heat flux and medium pressure area 6 is located on the upper and lower walls of the microchannel material layer 1 and adjacent to the high heat flux and high pressure area 5. The heat flux and pressure in this area are moderate, and the cooling demand is slightly lower than that of the high heat flux and high pressure area 5; the low heat flux and low pressure area 7 is located on the upper and lower walls of the microchannel material layer 1 and adjacent to the medium heat flux and medium pressure area 6. The heat flux and pressure in this area are small, and the cooling demand is the lowest. In the specific application process, the high heat flux and high pressure area 5, the medium heat flux and medium pressure area 6, and the low heat flux and low pressure area 7 can be divided according to the proportion of the heat flux density. For example:

[0044] Starting from the stagnation point, the region that extends symmetrically from the upper wall and the lower wall along the chord direction and whose heat flux density accounts for 60% of the total heat flux density within the 180° range of the leading edge is defined as the high heat flux and high pressure region 5;

[0045] Starting from the boundary line of the high heat flux and high pressure region 5, the region that extends symmetrically from the upper wall and the lower wall along the chord direction and whose heat flux density accounts for 30% of the total heat flux density within the 180° range of the leading edge is defined as the medium heat flux and medium pressure region 6;

[0046] The region within the 180° range of the leading edge of the microchannel material layer 1, excluding the high heat flux and high pressure region 5 and the medium heat flux and medium pressure region 6, is defined as the low heat flux and low pressure region 7.

[0047] The boundary line between the high heat flux and high pressure region 5 and the medium heat flux and medium pressure region 6 is defined as the first boundary 8, the boundary line between the medium heat flux and medium pressure region 6 and the low heat flux and low pressure region 7 is defined as the second boundary 9, and the leading edge boundary on the microchannel material layer 1 is defined as the third boundary 10. In this embodiment, the number of the first microchannels 3 is one, which is directly opened at the stagnation point in the high heat flux and high pressure region 5. The number of the second microchannels 4 along the chord direction on the upper wall or the lower wall of the microchannel material layer 1 is three. The second end of the first second microchannel 4 is located at the first boundary 8, the second end of the second second microchannel 4 is located at the second boundary 9, and the second end of the third second microchannel 4 is located at the third boundary 10.

[0048] Taking the case where the leading edge radius is 1 mm, the oncoming flow condition is MA = 8, and H = 36 km as an example, the wall temperature is set to 1000 K, and the heat flux density distribution and the surface pressure distribution along the outer wall are calculated by a commercial finite element software, as Figure 3 shown. In order to adapt to Figure 3 the extremely uneven heat flux density and surface pressure along the path shown, the active cooling structure in this embodiment adopts a stepped distribution of microchannels design, and optimizes the supply and flow of the coolant by reasonably adjusting the flow channel structure of each region. Among them, the liquid water reaches the saturation temperature and undergoes a phase change within about 1 mm in the coolant cavity 2 at the designed flow rate, and is in a gaseous state in both the first microchannel 3 and the second microchannel 4.

[0049] For the high heat flux and high pressure region 5 near the stagnation point, this embodiment selects to adopt the first microchannel 3 with a slit shape, that is, a slit channel with a larger cross-sectional area is used to distribute a large amount of coolant. Preferably, in order to reduce the flow velocity loss, the cross-sectional area of the first microchannel 3 is in a tapered shape, that is, the flow channel cross-sectional area of the first microchannel 3 gradually decreases along the direction from the first end to the second end, thereby improving the heat transfer ability of the coolant and optimizing the flow velocity distribution.

[0050] For the medium heat flux medium pressure region 6 and low heat flux low pressure region 7 far from the stagnation point, in this embodiment, the second microchannels 4 with a micro straight cylindrical structure and arranged in parallel are selected, a throttling hole 11 is provided at the first end of the second microchannels 4, and the cross-sectional area of each second microchannel 4 is gradually reduced along the chord direction to adapt to the change of heat flux density. Among them, the selection of micro straight channels as the second microchannels 4 can effectively increase the convective heat transfer area and improve the heat transfer efficiency between the solid and the coolant. In addition, the second microchannels 4 are preferably cylindrical gradually expanding channels, that is, the flow channel cross-sectional area of the second microchannels 4 gradually expands along the direction from the first end to the second end. Through the balanced design of increasing flow resistance and flow rate, the coolant flow rate is effectively controlled, and the flow inside the channel is ensured to be more stable.

[0051] In the specific application process, the calculation process of the coolant flow rates of the first microchannels 3 and each second microchannel 4 is as follows:

[0052] First, calculate the average heat flux densities of the high heat flux high pressure region 5, medium heat flux medium pressure region 6, low heat flux low pressure region 7, and the non-leading edge region on the microchannel material layer 1 respectively;

[0053] Then, based on the average heat flux density, according to the heat balance formula, calculate the coolant heat sinks required for the four regions, and then determine the flow rate distribution ratios of each region;

[0054] In addition, to ensure that the coolant in each channel flows out smoothly according to the flow rate distribution ratio, adjust the flow resistance of each outlet to compensate for different pressure differences, so that the flow rates of each outlet meet the target ratio. Therefore, a throttling hole 11 is designed at the first end of each second microchannel 4 to increase the flow resistance, and the diameter of the throttling hole 11 is adjusted so that the resistances of different outlets meet the flow rate ratio.

[0055] As a preferred implementation manner, in the chord direction of the microchannel material layer 1, the first end of the first second microchannel 4 is directly connected to the coolant cavity 2, and the first end of the i-th second microchannel 4 is opened on the channel wall of the (i - 1)-th second microchannel 4, thereby forming a second microchannel 4 with a stepped distribution. Among them, i = 2 to N, and N is the number of second microchannels 4 along the chord direction on the upper wall or lower wall of the microchannel material layer 1. In this embodiment, by arranging the second microchannels 4 with a stepped distribution, while ensuring the structural strength, compared with the traditional parallel inlet which is prone to excessive flow rate (local subcooling) in the near inlet region and insufficient flow rate at the far end, in this embodiment, the flow rate is balanced by step-by-step shunting, making the chordal temperature distribution more uniform. At the same time, the inlet pressure drop of each second microchannel 4 is shared by its previous stage channel, avoiding a concentrated high pressure drop at a single inlet.

[0056] As a preferred implementation manner, the included angle between the outflow direction of the second end of each second microchannel 4 and the oncoming flow direction is less than , which helps the coolant form a gas film on the aircraft surface and spread on the solid surface, raise the boundary layer, thus forming a heat insulation barrier and further improving the cooling efficiency.

[0057] In this embodiment, the microchannel material layer 1 is made of a nickel-based superalloy material, so that the active cooling result has excellent high-temperature resistance and corrosion resistance, and can operate stably for a long time in a high-temperature environment.

[0058] The above are only the preferred embodiments of the present invention, and thus do not limit the protection scope of the present invention. Any equivalent structural transformation made under the inventive concept of the present invention by using the content of the specification and drawings of the present invention, or any direct / indirect application in other related technical fields is included in the protection scope of the present invention.

Claims

1. An active cooling structure for the sharp leading edge of a high-speed aircraft with a designable microchannel, characterized in that It includes a microchannel material layer with a sharp leading edge, and a coolant cavity is provided inside the microchannel material layer; A first microchannel with a slit structure is provided on the microchannel material layer. The first end of the first microchannel communicates with the coolant cavity, and the second end of the first microchannel is located at a position corresponding to the leading edge stagnation point on the microchannel material layer; A plurality of second microchannels with cylindrical structures are also provided on the microchannel material layer. The first end of each second microchannel communicates with the coolant cavity; in the spanwise and chordwise directions of the microchannel material layer, the second ends of each second microchannel are spaced apart and distributed on the outer wall of the microchannel material layer; The flow channel cross-sectional areas of the second microchannels located in the same spanwise direction of the microchannel material layer are the same, and along the chordwise direction of the microchannel material layer, the flow channel cross-sectional areas of each second microchannel gradually decrease.

2. The active cooling structure for the sharp leading edge of a high-speed aircraft with a designable microchannel according to claim 1, characterized in that In the direction from the first end to the second end, the flow channel cross-sectional area of the first microchannel gradually shrinks.

3. The active cooling structure for the sharp leading edge of a high-speed aircraft with a designable microchannel according to claim 1, characterized in that, A throttle hole is provided at the first end of the second microchannel; In the direction from the first end to the second end, the flow channel cross-sectional area of the second microchannel gradually expands.

4. The active cooling structure for the sharp leading edge of a high-speed aircraft with a designable microchannel according to claim 1 or 2, characterized in that, The second microchannel is a cylindrical channel.

5. The active cooling structure for the sharp leading edge of a high-speed aircraft with a designable microchannel according to claim 1 or 2 or 3, characterized in that, In the chordwise direction of the microchannel material layer, the first end of the first second microchannel directly communicates with the coolant cavity, and the first end of the i-th second microchannel is opened on the channel wall of the (i - 1)-th second microchannel; wherein, i = 2 to N, and N is the number of second microchannels along the chordwise direction on the upper wall or lower wall of the microchannel material layer.

6. The active cooling structure for the sharp leading edge of a high-speed aircraft with a designable microchannel according to claim 1 or 2 or 3, characterized in that, According to the heat flux and pressure distribution, the outer surface of the leading edge region of the microchannel material layer is divided into a high heat flux and high pressure region, a medium heat flux and medium pressure region, and a low heat flux and low pressure region; The high heat flux and high pressure region is located in the region corresponding to the leading edge stagnation point, the medium heat flux and medium pressure region is located on the upper wall surface and the lower wall surface of the microchannel material layer and is adjacent to the high heat flux and high pressure region, and the low heat flux and low pressure region is located on the upper wall surface and the lower wall surface of the microchannel material layer and is adjacent to the medium heat flux and medium pressure region; Define the boundary line between the high heat flux and high pressure region and the medium heat flux and medium pressure region as the first boundary, define the boundary line between the medium heat flux and medium pressure region and the low heat flux and low pressure region as the second boundary, and define the leading edge boundary on the microchannel material layer as the third boundary; The number of the second microchannels along the chordwise direction on the upper wall surface or the lower wall surface of the microchannel material layer is three. Among them, the second end of the first second microchannel is located at the first boundary, the second end of the second second microchannel is located at the second boundary, and the second end of the third second microchannel is located at the third boundary.

7. The active cooling structure for the sharp leading edge of a high-speed aircraft with a designable microchannel according to claim 1 or 2 or 3, characterized in that, The included angle between the outflow direction of the second end of each second microchannel and the oncoming flow direction is less than 90°.

8. The active cooling structure for the sharp leading edge of a high-speed aircraft with a designable microchannel according to claim 1 or 2 or 3, characterized in that, The microchannel material layer is made of a nickel-based superalloy material.

9. The active cooling structure for the sharp leading edge of a high-speed aircraft with a designable microchannel according to claim 1 or 2 or 3, characterized in that, The coolant in the coolant cavity is liquid water.

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