A leading-edge thermal protection structure for an aircraft

Abstract

The present invention discloses a thermal protection structure for the leading edge of an aircraft, which includes a first porous medium layer provided at the arc-shaped wall surface of the leading edge of the aircraft, and a plurality of second porous medium layers spaced along the flow direction at the upper and lower wall surfaces of the leading edge of the aircraft; a solid wall is provided between the first porous medium layer and the second porous medium layers, and between adjacent two of the second porous medium layers; the porosity of each of the second porous medium layers first increases and then decreases along the flow direction, and the porosity of the first porous medium layer is less than the porosity of the second porous medium layer at the most upstream. The present invention is applied to the field of aircraft thermal protection. Through the synergistic effect of the intermittent sweating layout and the gradient porosity material, the problem of local overheating existing in the traditional sweating cooling technology is effectively solved, and the decrease in cooling efficiency caused by overheating of the coolant is avoided.

Description

Technical Field

[0001] The present invention relates to the technical field of aircraft thermal protection, and specifically to a thermal protection structure for the leading edge of an aircraft. Background Art

[0002] With the rapid development of aerospace technology, advanced equipment such as hypersonic aircraft and reusable spacecraft have put forward more stringent requirements for thermal protection systems. The leading edge parts of the aircraft (such as the leading edge of the wing, nose cone, etc.), as the areas with the most intense aerodynamic heating, bear extreme thermal-mechanical coupling loads during high-speed flight. When the aircraft flies in the atmosphere at a speed above Mach 5, the surface temperature of the leading edge part can reach above 1500°C, accompanied by severe mechanical shock and oxidation corrosion environment, which poses a major challenge to the high-temperature resistance, thermal shock resistance and structural reliability of the thermal protection structure.

[0003] Traditional thermal protection technologies are mainly divided into two categories: passive protection and active cooling. In terms of passive protection, although the early metal thermal protection structures (such as nickel-based alloy honeycomb panels) have good structural strength, their temperature resistance limit (usually lower than 1200°C) can no longer meet the requirements of new-generation aircraft; although ceramic matrix composites (such as C / SiC, ZrB2-SiC) have excellent high-temperature resistance, they have problems such as insufficient impact resistance and interface oxidation failure, and are prone to generating microcracks under thermal cycle loads, resulting in protection failure. Ablative materials (such as carbon / phenolic composites) achieve thermal protection by absorbing heat through material decomposition, but they are disposable consumable materials and cannot meet the requirements of reusable aircraft. Active cooling technologies (such as transpiration cooling, convective cooling) achieve thermal protection through the circulation of internal cooling media, but they have problems such as complex systems, large mass, and limited coolant carrying capacity. Summary of the Invention

[0004] Aiming at the deficiencies in the above-mentioned prior art, the present invention provides a thermal protection structure for the leading edge of an aircraft. Through the synergistic effect of intermittent transpiration layout and gradient porosity materials, the problem of local overheating existing in traditional transpiration cooling technology is effectively solved, and the decrease in cooling efficiency caused by overheating of the coolant is avoided.

[0005] To achieve the above object, the present invention provides a thermal protection structure for the leading edge of an aircraft, including a first porous medium layer provided at the arc-shaped wall surface of the leading edge of the aircraft, and a plurality of second porous medium layers provided at intervals along the flow direction on the upper and lower wall surfaces of the leading edge of the aircraft;

[0006] A solid wall is provided between the first porous medium layer and the second porous medium layers, and between adjacent two of the second porous medium layers;

[0007] The porosity of each of the second porous medium layers first increases and then decreases along the flow direction, and the porosity of the first porous medium layer is less than that of the most upstream second porous medium layer.

[0008] In one embodiment, the positions on the upper and lower wall surfaces of the leading edge of the aircraft where the heat flux density is 45% - 55% of the maximum heat flux density are defined as the demarcation points;

[0009] Upstream of the demarcation point, the porosity of each of the second porous medium layers increases in sequence along the flow direction;

[0010] Downstream of the demarcation point, the porosity of each of the second porous medium layers decreases in sequence along the flow direction.

[0011] In one embodiment, the second porous medium layer and the adjacent downstream solid wall are defined as a discontinuity;

[0012] The length of the discontinuity in the flow direction is 0.2 mm - 2 mm.

[0013] In one embodiment, in the discontinuity, the proportion of the length of the second porous medium layer in the flow direction is 50% - 75%.

[0014] In one embodiment, the first porous medium layer and the second porous medium layer are made of a superalloy or composite ceramic material with a certain porosity;

[0015] The solid wall is made of a non-porous superalloy.

[0016] In one embodiment, shape memory alloy wire meshes are provided on the inner walls of the first porous medium layer and the second porous medium layer.

[0017] Compared with the prior art, the present invention has the following beneficial technical effects:

[0018] 1. By spatially discontinuously arranging the porous medium on the leading edge of the aircraft, the present invention realizes the discontinuous discharge of the coolant on the leading edge surface, not only effectively reducing the entrainment effect of the mainstream on the coolant, enhancing the coverage ability of the coolant on the hot end wall surface, but also optimizing the distribution of the coolant, enabling the coolant to be more effectively utilized in the high heat flux region, and significantly improving the overall cooling efficiency;

[0019] 2. According to the spatial distribution characteristics of the heat flux density, the present invention adjusts the flow rate and distribution of the coolant by designing a gradient porosity material. A material with a smaller porosity is used in the stagnation region to enhance the convective heat transfer between the coolant and the solid skeleton in the porous medium. In the region where the stagnation point transitions to a low heat flux, a relatively thick gas film is formed by rapidly increasing the porosity, and the temperature gradient is reduced by adjusting the intermittent rhythm. In the low heat flux region, a material with a gradually decreasing porosity is used to maintain stable cooling protection and avoid overcooling, making the coolant distribution more reasonable and effectively solving the problem of uneven cooling of traditional thermal protection structures in complex thermal environments.

[0020] 3. Through the synergistic effect of the intermittent sweating layout and the gradient porosity material, the present invention effectively solves the problem of local overheating existing in the traditional sweating cooling technology and avoids the decrease in cooling efficiency caused by overheating of the coolant. BRIEF DESCRIPTION OF THE DRAWINGS

[0021] 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.

[0022] Figure 1 It is a schematic structural diagram of the thermal protection structure of the leading edge of an aircraft in an embodiment of the present invention;

[0023] Figure 2 It is a schematic diagram of another implementation manner of the thermal protection structure of the leading edge of an aircraft in an embodiment of the present invention.

[0024] Reference numerals in the drawings: the first porous medium layer 1, the second porous medium layer 2, the solid wall 3, the shape memory alloy wire mesh 4.

[0025] The realization of the objectives, functional features and advantages of the present invention will be further described in conjunction with the embodiments and with reference to the drawings. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0026] 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.

[0027] 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 positional relationship, movement conditions, etc. between components in a specific posture (as shown in the attached drawings). If the specific posture changes, the directional indications will also change accordingly.

[0028] In addition, in the present invention, descriptions such as "first" and "second" are only for descriptive purposes 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 of such features. In the description of the present invention, "a plurality of" means at least two, such as two, three, etc., unless otherwise specifically defined.

[0029] In the present invention, unless otherwise clearly defined and limited, terms such as "connection" and "fixation" should be understood in a broad sense. For example, "fixation" can be a fixed connection, a detachable connection, or integrated; it can be a mechanical connection, an electrical connection, a physical connection or a wireless communication connection; it can be directly connected or indirectly connected through an intermediate medium, and can be the communication inside two elements or the interaction relationship between two elements, unless otherwise clearly limited. 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.

[0030] 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 protection scope required by the present invention.

[0031] As Figure 1 Shown is a thermal protection structure for the leading edge of an aircraft disclosed in this embodiment, which mainly includes a first porous medium layer 1 provided at the arc-shaped wall surface of the leading edge of the aircraft, and a plurality of second porous medium layers 2 provided at intervals along the flow direction on the upper and lower wall surfaces of the leading edge of the aircraft. A solid wall 3 is provided between the first porous medium layer 1 and the second porous medium layers 2, and between adjacent second porous medium layers 2. The porosity of each second porous medium layer 2 first increases and then decreases along the flow direction, and the porosity of the first porous medium layer 1 is less than the porosity of the second porous medium layer 2 at the most upstream.

[0032] The gaseous working medium as the coolant flows from Figure 1The arrow direction enters the interior of the leading edge. The aircraft leading edge thermal protection structure solves the local overheating problem existing in the traditional sweating cooling technology through the synergistic effect of the intermittent sweating layout and the gradient porosity material, and avoids the decrease in cooling efficiency caused by overheating of the coolant. Specifically, by setting the solid wall 3 to realize the spatial discontinuous layout of the porous medium, the discontinuous discharge of the coolant on the leading edge surface is realized, which can not only effectively reduce the entrainment effect of the mainstream on the coolant, enhance the coverage of the coolant on the hot end wall, but also optimize the distribution of the coolant, so that the coolant is more effectively utilized in the high heat flow area, and the overall cooling efficiency is significantly improved.

[0033] In the specific implementation process, the position where the heat flux density on the upper and lower walls of the leading edge of the aircraft is 45% to 55% of the maximum heat flux density is defined as the boundary, for example, the position where the heat flux density on the upper and lower walls of the leading edge of the aircraft is 50% of the maximum heat flux density is defined as the boundary. In the area upstream of the boundary on the upper and lower walls of the leading edge of the aircraft, the porosity of each second porous medium layer 2 increases in sequence along the flow direction; in the area downstream of the boundary on the upper and lower walls of the leading edge of the aircraft, the porosity of each second porous medium layer 2 decreases in sequence along the flow direction downstream of the boundary.

[0034] On the basis of intermittent sweating, this embodiment also uses porous media with gradient porosity to further improve the cooling efficiency, that is, a first porous medium layer 1 with lower porosity (for example, a porosity of 10%) is selected at the leading edge stagnation point. Because the mechanism of single-phase sweating cooling consists of two aspects, on the one hand, the convective heat exchange between the coolant and the solid skeleton in the porous medium, and on the other hand, the thermal blocking effect formed on the surface of the medium after flowing out of the porous medium, and the position of the bow shock wave is closer to the nose cone stagnation point, at this time, the pressure difference between the inside and outside of the porous medium is the smallest, so that the thermal blocking effect is almost lost. In this case, sweating cooling relies more on convective heat transfer. The increase in porosity will lead to a decrease in the specific surface area inside the porous medium, insufficient heat exchange between the coolant and the solid skeleton, a decrease in the effective thermal conductivity of the porous medium, a decrease in the amount of heat that the coolant can take away, and a weakened cooling effect. Therefore, in this embodiment, the porosity of the first porous medium layer 1 is less than the porosity of the second porous medium layer 2 at the most upstream.

[0035] For the leading-edge region near the stagnation point, the porosity distribution of the second porous medium layer 2 shows a rapidly increasing trend along the flow direction. Specifically, the porosity of the first second porous medium layer 2 along the flow direction can be directly set to 2-4 times that of the first porous medium layer 1, and the porosity gradient of the subsequent second porous medium layers 2 is increased until it reaches the boundary and then decreased gradually. For example, assume that there are three second porous medium layers 2 on the upper wall surface of the leading edge of the aircraft upstream of the boundary and three second porous medium layers 2 downstream. The porosity of the first porous medium layer 1 can be 10%. For the six second porous medium layers 2 on the upper wall surface of the leading edge of the aircraft, the porosities along the flow direction can be set to 40%, 50%, 60%, 45%, 30%, and 15% respectively. Since the permeability increases with the increase of porosity, the coolant is more likely to enter the porous medium with a larger porosity, rapidly increasing the thickness of the gas film layer in this region, protecting the wall surface by the coolant film, and reducing the amount of heat that the mainstream can transfer to the wall surface. Finally, for the middle and rear parts of the leading edge, a gradually decreasing porosity arrangement is adopted. Because the internal and external pressure difference of the porous medium gradually increases along the flow direction and the flow resistance of the coolant in the porous medium gradually decreases, the trend of increasing coolant flow is offset by gradually reducing the porosity, so that more coolant flows to the leading-edge stagnation region, improving the temperature uniformity on the surface of the porous medium, reducing the temperature gradient, and achieving an overall improvement in the thermal protection effect of the leading-edge wall surface. This method can basically increase the temperature reduction of the leading-edge peak temperature by about 5% and the temperature reduction of the average temperature by 10%-15% on the basis of intermittent transpiration cooling.

[0036] Define the second porous medium layer 2 and its adjacent downstream solid wall 3 as a discontinuity. The length of the discontinuity in the flow direction in this embodiment is 0.2 mm to 2 mm, and the length ratio of the second porous medium layer 2 in the flow direction in the discontinuity is 50% to 75%. When the high-speed airflow passes through the second porous medium layer 2, the coolant forms a continuous gas film (with a thickness of about 0.1-0.3 mm) through the pores. The sudden contraction of the solid wall 3 at the discontinuity causes local flow separation, forming a reverse vortex, which prolongs the residence time of the gas film coolant near the wall surface by 20% to 35%, enhancing the convective heat transfer efficiency. Setting the length ratio of the second porous medium layer 2 to 50% to 75% can ensure the effective coverage of the gas film coolant at the discontinuity (coverage rate ≥ 85%) and prevent the solid wall 3 from being directly exposed to the high-temperature airflow.

[0037] In this embodiment, the first porous medium layer 1 and the second porous medium layer 2 are made of a superalloy or composite ceramic material with a certain porosity, and the solid wall 3 is made of a non-porous superalloy. Preferably, a shape memory alloy wire mesh 4 can be arranged on the inner walls of the first porous medium layer 1 and the second porous medium layer 2, and the sum of the thicknesses of the first porous medium layer 1 or the second porous medium layer 2 and the corresponding shape memory alloy wire mesh 4 is equal to the thickness of the adjacent solid wall 3, that isFigure 2 As shown. In the specific application process, the phase transition temperature of the shape memory alloy wire mesh 4 can be set according to requirements. When the temperature of the local area at the leading edge exceeds the phase transition temperature of the shape memory alloy wire mesh 4, the shape memory alloy wire mesh 4 can drive the porosity of the first porous medium layer 1 or the second porous medium layer 2 to be dynamically adjusted, thereby realizing the self-adaptive regulation of heat flow while enhancing the structural strength of the first porous medium layer 1 and the second porous medium layer 2.

[0038] The above are only the preferred embodiments of the present invention, and 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 aircraft leading edge thermal protection structure, characterized in that: It comprises a first porous medium layer arranged at the arc-shaped wall surface of the leading edge of the aircraft, and a plurality of second porous medium layers arranged at intervals along the flow direction at the upper and lower walls of the leading edge of the aircraft; A solid wall is provided between the first porous medium layer and the second porous medium layer, and between two adjacent second porous medium layers; The porosity of each of the second porous medium layers increases first and then decreases along the flow direction, and the porosity of the first porous medium layer is smaller than the porosity of the second porous medium layer at the most upstream side.

2. The aircraft leading edge thermal protection structure according to claim 1, characterized in that: The position where the heat flux density on the upper and lower walls of the leading edge of the aircraft is 45% to 55% of the maximum heat flux density is defined as the boundary; Upstream of the boundary, the porosity of each of the second porous medium layers increases sequentially along the flow direction; Downstream of the boundary, the porosity of each of the second porous medium layers decreases successively along the flow direction.

3. The aircraft leading edge thermal protection structure according to claim 1 or 2, characterized in that: defining the second porous medium layer and the solid wall adjacent to it downstream as a discontinuity; The length of the discontinuity in the flow direction is 0.2 mm to 2 mm.

4. The aircraft leading edge thermal protection structure according to claim 3, characterized in that: In the discontinuity, the length of the second porous medium layer in the flow direction accounts for 50% to 75%.

5. The aircraft leading edge thermal protection structure according to claim 1 or 2, characterized in that: The first porous medium layer and the second porous medium layer are made of high-temperature alloy or composite ceramic material with a certain porosity; The solid wall is made of a non-porous high-temperature alloy.

6. The aircraft leading edge thermal protection structure according to claim 1 or 2, characterized in that: Shape memory alloy meshes are provided on the inner walls of the first porous medium layer and the second porous medium layer.

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