Aircraft leading edge thermal protection structure

Abstract

The thermal protection structure comprises a first porous medium layer arranged on the arc-shaped wall face of the front edge of the aircraft and a plurality of second porous medium layers arranged on the upper wall face and the lower wall face of the front edge of the aircraft at intervals in the flow direction. Solid walls are arranged between the first porous medium layers and the second porous medium layers and between every two adjacent second porous medium layers; the porosity of each second porous medium layer is firstly increased and then reduced along the flow direction, and the porosity of the first porous medium layer is smaller than that of the second porous medium layer at the most upstream. The method is applied to the field of thermal protection of aircrafts, the problem of local overheating existing in a traditional sweating cooling technology is effectively solved through the synergistic effect of the intermittent sweating layout and the gradient porosity material, and the situation that the cooling efficiency is reduced due to overheating of a cooling agent is avoided.

Description

Technical Field

[0001] The invention relates to the technical field of aircraft thermal protection, in particular to an aircraft leading edge thermal protection structure. Background Art

[0002] With the rapid development of aerospace technology, advanced equipment such as hypersonic vehicles and reusable spacecraft have put forward more stringent requirements on thermal protection systems. The leading edge of the aircraft (such as the leading edge of the wing, the nose cone, etc.) is the area with the most intense aerodynamic heating and is subjected to extreme thermal-mechanical coupling loads during high-speed flight. When the aircraft flies in the atmosphere at a speed of more than Mach 5, the surface temperature of the leading edge can reach more than 1500°C, accompanied by severe mechanical shock and oxidative 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 below 1200°C) can no longer meet the needs of the new generation of aircraft; although ceramic-based composite materials (such as C / SiC, ZrB2-SiC) have excellent high temperature resistance, they have problems such as insufficient impact resistance and interface oxidation failure. Microcracks are easily generated 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 needs of reusable aircraft. Active cooling technologies (such as sweat cooling and convection cooling) achieve thermal protection through internal cooling medium circulation, but there are problems such as complex systems, large mass, and limited coolant carrying capacity. Summary of the invention

[0004] In view of the above-mentioned deficiencies in the prior art, the present invention provides a thermal protection structure for the leading edge of an aircraft, which effectively solves the problem of local overheating existing in 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.

[0005] To achieve the above-mentioned object, the present invention provides a thermal protection structure for the leading edge of an aircraft, comprising a first porous medium layer arranged at the curved 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.

[0006] In one embodiment, 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.

[0007] In one embodiment, the second porous medium layer and the solid wall adjacent to the second porous medium layer downstream are defined as a discontinuity; The length of the discontinuity in the flow direction is 0.2 mm to 2 mm.

[0008] In one of the embodiments, in the discontinuity, the length of the second porous medium layer in the flow direction accounts for 50% to 75%.

[0009] In one embodiment, 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.

[0010] In one embodiment, shape memory alloy meshes are provided on the inner walls of the first porous medium layer and the second porous medium layer.

[0011] Compared with the prior art, the present invention has the following beneficial technical effects: 1. The present invention realizes discontinuous discharge of coolant on the leading edge surface by spatially discontinuously arranging the porous media on the leading edge of the aircraft, which not only effectively reduces the entrainment effect of the mainstream on the coolant, enhances the coverage of the coolant on the hot end wall surface, but also optimizes the distribution of the coolant, so that the coolant is more effectively utilized in the high heat flux area, and significantly improves the overall cooling efficiency; 2. According to the spatial distribution characteristics of heat flux density, the present invention adjusts the flow and distribution of the coolant by designing gradient porosity materials. In the stagnation point area, a smaller porosity material is used to enhance the convective heat exchange between the coolant and the solid skeleton in the porous medium. In the transition area from the stagnation point to the low heat flux, a method of rapidly increasing the porosity is used to form a thicker air film. The temperature gradient is reduced by adjusting the intermittent rhythm. In the low heat flux area, a gradually decreasing porosity material 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; 3. The present invention effectively 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 reduction of cooling efficiency caused by overheating of the coolant. BRIEF DESCRIPTION OF THE DRAWINGS

[0012] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on the structures shown in these drawings without paying creative work.

[0013] Figure 1 Schematic diagram of the structure of the leading edge thermal protection structure of an aircraft in an embodiment of the present invention; Figure 2 It is a schematic diagram of another implementation of the aircraft leading edge thermal protection structure in an embodiment of the present invention.

[0014] Reference numerals: first porous medium layer 1 , second porous medium layer 2 , solid wall 3 , shape memory alloy wire mesh 4 .

[0015] The realization of the purpose, functional features and advantages of the present invention will be further explained in conjunction with embodiments and with reference to the accompanying drawings. DETAILED DESCRIPTION

[0016] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0017] 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 components under a certain specific posture (as shown in the accompanying drawings). If the specific posture changes, the directional indication will also change accordingly.

[0018] In addition, in the present invention, descriptions such as "first", "second", etc. are only used for descriptive purposes and cannot be understood as indicating or implying their relative importance or implicitly indicating the number of the indicated technical features. Therefore, the features defined as "first" or "second" may explicitly or implicitly include at least one of the features. In the description of the present invention, the meaning of "plurality" is at least two, such as two, three, etc., unless otherwise clearly and specifically defined.

[0019] In the present invention, unless otherwise clearly specified and limited, the terms "connection", "fixation", etc. should be understood in a broad sense. For example, "fixation" can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection, an electrical connection, a physical connection, or a wireless communication connection; it can be a direct connection, or an indirect connection through an intermediate medium, or it can be the internal connection of two elements or the interaction relationship between two elements, unless otherwise clearly defined. For ordinary technicians in this field, the specific meanings of the above terms in the present invention can be understood according to specific circumstances.

[0020] 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 the field can implement it. When the combination of technical solutions is contradictory or cannot be implemented, it should be deemed that such combination of technical solutions does not exist and is not within the scope of protection required by the present invention.

[0021] like Figure 1 As 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 arranged at the curved wall surface of the leading edge of the aircraft, and a plurality of second porous medium layers 2 arranged at intervals along the flow direction at the upper and lower walls 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 layer 2 and between two adjacent second porous medium layers 2. The porosity of each second porous medium layer 2 first increases and then decreases along the flow direction. The porosity of the first porous medium layer 1 is less than the porosity of the most upstream second porous medium layer 2.

[0022] The gaseous working fluid used as coolant is Figure 1 The 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.

[0023] 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.

[0024] 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.

[0025] For the leading edge region near the stagnation point, the porosity distribution of the second porous medium layer 2 shows a trend of rapid increase 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 second porous medium layer 2 in the future can be increased until it reaches the boundary and then the gradient is reduced. For example, assuming that there are three second porous medium layers 2 upstream of the boundary and three second porous medium layers 2 downstream on the upper wall of the leading edge of the aircraft, the porosity of the first porous medium layer 1 can be 10%, and the porosity of the six second porous medium layers 2 on the upper wall of the leading edge of the aircraft can be set to 40%, 50%, 60%, 45%, 30%, and 15% respectively along the flow direction. Since the permeability increases with the increase of porosity, the coolant is easier to enter the porous medium with a larger porosity. The thickness of the air film layer is rapidly increased in this area, so that the wall is protected by the coolant film, and the heat that the mainstream can transfer to the wall is less. Finally, a gradually decreasing porosity arrangement is used for the middle and rear parts of the leading edge. Because the internal and external pressure difference of the porous medium gradually increases along the flow direction, the flow resistance of the coolant in the porous medium gradually decreases. Therefore, the trend of increasing coolant flow is offset by gradually reducing the porosity, so that more coolant flows to the leading edge stagnation area, the temperature uniformity of the porous medium surface is improved, the temperature gradient is reduced, and the overall thermal protection effect of the leading edge wall is improved. This method can basically increase the reduction of the leading edge peak temperature by about 5% and the average temperature by 10%-15% on the basis of intermittent sweat cooling.

[0026] The second porous medium layer 2 and its downstream adjacent solid wall 3 are defined as a discontinuity. The length of the discontinuity in the present embodiment is 0.2 mm to 2 mm in the flow direction, and the length of the second porous medium layer 2 in the discontinuity in the flow direction accounts for 50% to 75%. When the high-speed airflow passes through the second porous medium layer 2, the coolant forms a continuous air film (thickness of about 0.1 to 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 air film coolant near the wall by 20% to 35%, enhancing the convective heat transfer efficiency. The length of the second porous medium layer 2 is set to 50% to 75%, which can ensure the effective coverage of the air film coolant at the discontinuity (coverage rate ≥ 85%), and prevent the solid wall 3 from being directly exposed to the high-temperature airflow.

[0027] In this embodiment, the first porous medium layer 1 and the second porous medium layer 2 are made of a high temperature alloy or a composite ceramic material with a certain porosity, and the solid wall 3 is made of a non-porous high temperature alloy. Preferably, a layer of shape memory alloy wire mesh 4 can be arranged on the inner wall of the first porous medium layer 1 and the second porous medium layer 2, and the sum of the thickness of the first porous medium layer 1 or the second porous medium layer 2 and the corresponding shape memory alloy wire mesh 4 is kept equal to the thickness of the adjacent solid wall 3, that is, Figure 2 In the specific application process, the phase change temperature of the shape memory alloy wire mesh 4 can be set according to the needs. When the temperature of the local area of ​​the leading edge exceeds the phase change 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 enhancing the structural strength of the first porous medium layer 1 and the second porous medium layer 2 while realizing adaptive heat flow regulation.

[0028] The above description is only a preferred embodiment of the present invention, and does not limit the protection scope of the present invention. All equivalent structural changes made by using the contents of the present invention specification and drawings under the inventive concept of the present invention, or directly / indirectly applied in other related technical fields are 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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