Heat protection structure integrating heat protection and heat dredging

Abstract

The invention discloses a thermal protection structure integrating heat protection and heat dredging, which is used for an aircraft nose and comprises a porous carbon matrix, a carbon fiber bundle and an anti-oxidation coating. The porous carbon substrate is adaptive to the appearance of the nose, is arranged on the outer surface of the nose of the aircraft, and is of a three-dimensional network structure with pores; the plurality of carbon fiber bundles are distributed in the circumferential direction of the machine head and are radially embedded into the porous carbon matrix, and each carbon fiber bundle extends from the machine head to the machine tail to form a heat conduction channel; the anti-oxidation coating is arranged on the outer surface of the porous carbon substrate. Through the coupling effect of three mechanisms of heat radiation shielding of the anti-oxidation coating, directional heat conduction of the carbon fiber bundles and heat capacity buffering of the porous carbon matrix, the heat protection structure realizes heat redistribution in space and a heat buffering effect in time, so that concentrated heat in a stationary point area is quickly dispersed to a larger surface area, and the heat capacity of the surface area is greatly improved. Therefore, the local heat flow density is greatly reduced, and the instantaneous thermal shock is converted into a slow temperature rise process.

Description

Technical Field

[0001] This application relates to the field of aircraft thermal protection technology, and in particular to a thermal protection structure that integrates heat protection and heat conduction. Background Technology

[0002] Hypersonic vehicles face extremely severe thermal environment challenges during flight, especially when speeds exceed Mach 5, where aerodynamic heating significantly intensifies, causing surface temperatures to rapidly rise to 1000-2000°C. Unlike traditional spacecraft, hypersonic vehicles need to cruise within the atmosphere for extended periods, experiencing sustained and uneven heat loads, which places more stringent demands on thermal protection systems. Existing thermal protection technologies for hypersonic vehicles mainly fall into two categories: active cooling systems and passive thermal insulation materials.

[0003] Active cooling systems, such as recycle fuel cooling technology, are highly efficient, but they are complex, heavy, and have limited reliability. Passive insulation materials, such as ceramic matrix composites, have good heat resistance, but their thermal conductivity is insufficient, and they are prone to heat accumulation problems in long-term high-temperature environments.

[0004] A common problem with current hypersonic vehicle thermal protection systems is the separation of "heat protection" and "heat dissipation," lacking an integrated design approach. This results in complex structures, increased weight, and low thermal management efficiency. Especially during hypersonic cruise, different parts of the aircraft face different thermal environments, and traditional single-function thermal protection materials are insufficient to meet the overall thermal management requirements of the aircraft.

[0005] Therefore, developing a thermal protection structure that can simultaneously achieve efficient heat protection and directional heat conduction functions and adapt to the complex thermal environment of hypersonic flight is of great significance for improving the performance and reliability of hypersonic aircraft. Summary of the Invention

[0006] This application provides a thermal protection structure that integrates heat protection and heat conduction, which can solve the technical problems of complex thermal protection structures, large weight, and low thermal management efficiency that are common in existing hypersonic vehicle thermal protection systems. The technical solution is as follows:

[0007] A thermal protection structure integrating heat protection and heat conduction for use in the nose of an aircraft includes: a porous carbon matrix, multiple carbon fiber bundles, and an anti-oxidation coating.

[0008] The porous carbon matrix is ​​adapted to the shape of the nose and is disposed on the outer surface of the nose of the aircraft. The porous carbon matrix is ​​a three-dimensional mesh structure with pores. Multiple carbon fiber bundles are distributed along the circumference of the nose and are radially embedded in the porous carbon matrix. Each carbon fiber bundle extends from the nose to the tail to form a heat conduction channel. An anti-oxidation coating is disposed on the outer surface of the porous carbon matrix.

[0009] Optionally, the porosity of the porous carbon matrix is ​​40%-60%.

[0010] Optionally, the head is a quasi-conical shape that gradually expands from the head to the tail and has a blunt-tipped body configuration, and the porous carbon matrix is ​​also a quasi-conical shape that gradually expands from the head to the tail and has a blunt-tipped body configuration; the number of carbon fiber bundles is 8-12, and the multiple carbon fiber bundles are evenly distributed circumferentially along the central axis of the head, and two carbon fiber bundles in the same axial section form a parabolic shape.

[0011] Optionally, the distance from the starting point of the carbon fiber bundle to the tip of the aircraft nose is greater than the distance from the tip of the porous carbon matrix.

[0012] Optionally, the axial thermal conductivity of the carbon fiber bundle is 300-1500 W / (m·K), and the thermal conductivity of the porous carbon matrix is ​​80-120 W / (m·K).

[0013] Optionally, the antioxidant coating is a silicon carbide and hafnium dioxide composite material, and the thickness of the silicon carbide and hafnium dioxide composite material is 0.1-0.3 mm.

[0014] Optionally, a pyrolytic carbon interface layer is provided between the carbon fiber bundle and the porous carbon matrix, the thickness of which is 100-500 nm.

[0015] Optionally, the pyrolytic carbon interface layer is formed by chemical vapor deposition.

[0016] Optionally, forming a pyrolytic carbon interface layer via chemical vapor deposition includes the following steps:

[0017] Step 1: Place the prefabricated thermal protection structure, which includes carbon fiber bundles and porous carbon matrix, in a reactor and heat it to 900-1100°C under an inert atmosphere.

[0018] Step 2: A hydrocarbon precursor is introduced, which undergoes a pyrolysis reaction at high temperature to generate carbon atoms, which are then deposited on the surface of the carbon fiber bundle to form a pyrolytic carbon interface layer.

[0019] Furthermore, in step 2, the hydrocarbon precursor is methane or propane.

[0020] The beneficial effects of the technical solutions provided in this application include at least the following:

[0021] A thermal protection structure integrating heat protection and heat conduction, used in the nose of an aircraft, comprises a porous carbon matrix and multiple carbon fiber bundles disposed within the porous carbon matrix. First, when high-temperature airflow impacts the surface of the aircraft nose, the anti-oxidation coating acts as the first line of thermal protection, effectively reflecting incident radiant heat flow and significantly reducing the net heat absorption of the thermal protection structure. Second, heat penetrating the anti-oxidation coating and entering the interior of the thermal protection structure encounters a radially arranged network of carbon fiber bundles. These carbon fiber bundles possess high axial thermal conductivity and low radial thermal conductivity, with a high ratio of axial to radial thermal conductivity, resulting in highly anisotropic thermal conduction characteristics. Driven by a high temperature gradient, according to Fourier's law, the heat flux density generated within the carbon fiber bundles far exceeds that reaching the porous carbon matrix, thus constructing a "rapid heat channel" from the high-temperature stagnation region to the low-temperature rear region. Based on the thermal diffusion equation, the high thermal diffusivity of the carbon fiber bundles shortens the thermal response time, achieving rapid heat redistribution. Third, on the one hand, the porous carbon matrix plays a dual role as a thermal capacity buffer and thermal barrier layer throughout the entire thermal protection process. The high specific heat capacity and high density of the porous carbon matrix enable it to absorb and temporarily store a large amount of transient heat, effectively mitigating the thermal shock effect. On the other hand, the thermal conductivity of the porous carbon matrix is ​​much lower than that of the carbon fiber bundle. According to thermal resistance theory, a high thermal resistance region is formed perpendicular to the fiber direction, significantly slowing down the rate of heat transfer to the deeper parts of the structure. Through the coupled effect of the above three mechanisms—thermal radiation shielding, directional heat conduction, and thermal capacity buffering—this thermal protection structure achieves spatial heat redistribution and temporal thermal buffering effects, rapidly dispersing the concentrated heat in the stagnation area to a larger surface area, thereby significantly reducing the local heat flux density. This transforms instantaneous thermal shock into a slow temperature rise process, preventing thermal shock damage to the material.

[0022] The beneficial effects of the thermal protection structure that integrates heat protection and heat conduction in this application are as follows:

[0023] 1. The integrated "heat protection-heat dissipation" design solves the problem of functional fragmentation in traditional thermal protection systems, reducing system weight and simplifying structural complexity. 2. Utilizing the high thermal conductivity and directional arrangement of carbon fiber bundles, effective heat dissipation is achieved, significantly reducing the temperature peak in the stagnation area, minimizing the temperature gradient on the nose surface, and improving the thermal stability of the thermal protection structure. 3. The porous carbon matrix not only fixes the carbon fiber bundles but also absorbs some heat through its thermal capacity, acting as a buffer while reducing the overall structural density. 4. The arrangement density and direction of the carbon fiber bundles can be adjusted according to the thermal environment requirements of different parts of the aircraft, achieving customized thermal management.

[0024] Other features and advantages of this disclosure will be described in detail in the following detailed description section. Attached Figure Description

[0025] To more clearly illustrate the technical solutions in the embodiments of this application, the accompanying drawings used in the description of the embodiments will be briefly introduced below. Obviously, the accompanying drawings described below are only some embodiments of this application. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0026] Figure 1 This is a three-dimensional schematic diagram of a heat protection structure that integrates heat protection and heat conduction, as provided in the embodiments of this application.

[0027] Figure 2 This is a three-dimensional perspective view of the thermal protection structure integrating heat protection and heat conduction provided in the embodiments of this application;

[0028] Figure 3 This is a cross-sectional view of a thermal protection structure integrating heat protection and heat conduction provided in the embodiments of this application;

[0029] Figure 4 This is an exploded view of the thermal protection structure that integrates heat protection and heat conduction, as provided in the embodiments of this application.

[0030] Figure 5 Thermal simulation diagram of a simple porous carbon matrix thermal protection structure;

[0031] Figure 6 A thermal simulation diagram of a porous carbon matrix with a carbon fiber bundle thermal protection structure provided in the embodiments of this application.

[0032] Explanation of reference numerals in the attached figures

[0033] 1- Head; 2- Porous carbon matrix; 3- Carbon fiber bundle; 4- Antioxidant coating. Detailed Implementation

[0034] The specific embodiments of this disclosure will be described in detail below with reference to the accompanying drawings. It should be understood that the specific embodiments described herein are for illustration and explanation only and are not intended to limit this disclosure.

[0035] In this disclosure, unless otherwise stated, directional terms such as "upper" and "lower" generally refer to the relative positions of the corresponding components in the direction of gravity when they are in use, and "inner" and "outer" refer to the relative positions of the corresponding components with respect to their own contours. In the following description, when referring to the accompanying drawings, unless otherwise explained, the same reference numerals in different drawings denote the same or similar elements.

[0036] A thermal protection structure integrating heat protection and heat conduction is used in the nose of an aircraft, comprising: a porous carbon matrix 2 and multiple carbon fiber bundles 3 disposed within the porous carbon matrix 2.

[0037] The porous carbon matrix 2 is adapted to the shape of the nose 1 and is disposed on the outer surface of the nose 1 of the aircraft. The porous carbon matrix 2 is a three-dimensional mesh structure with pores. Multiple carbon fiber bundles 3 are distributed circumferentially along the nose 1. The multiple carbon fiber bundles 3 are radially embedded in the porous carbon matrix 2. Each carbon fiber bundle 3 extends from the nose 1 to the tail to form a heat conduction channel. An anti-oxidation coating 4 is disposed on the outer surface of the porous carbon matrix 2.

[0038] In the above embodiments, firstly, when the high-temperature airflow impacts the surface of the aircraft nose 1, the anti-oxidation coating 4 acts as the first thermal protection barrier, effectively reflecting the incident radiative heat flow and significantly reducing the net heat absorption of the thermal protection structure. Secondly, the heat that penetrates the anti-oxidation coating 4 and enters the interior of the thermal protection structure encounters the radially arranged carbon fiber bundles 3. These carbon fiber bundles 3 have a high axial thermal conductivity and a low radial thermal conductivity, with a high ratio of axial to radial thermal conductivity, forming highly anisotropic thermal conduction characteristics. Driven by a high temperature gradient, according to Fourier's law, the heat flux density generated within the carbon fiber bundles 3 far exceeds that reaching the porous carbon matrix 2, thus constructing a "rapid channel" for heat from the high-temperature stagnation region to the low-temperature rear region. Based on the thermal diffusion equation, the high thermal diffusion coefficient of the carbon fiber bundles 3 shortens the thermal response time, achieving rapid heat redistribution. Third, on the one hand, the porous carbon matrix 2 plays a dual role as a thermal capacity buffer and a thermal barrier layer throughout the entire thermal protection process. The high specific heat capacity and high density of the porous carbon matrix 2 can absorb and temporarily store a large amount of transient heat, effectively mitigating the thermal shock effect. On the other hand, the thermal conductivity of the porous carbon matrix 2 is much lower than that of the carbon fiber bundle 3. According to the thermal resistance theory, a high thermal resistance region is formed in the direction perpendicular to the fiber, which significantly slows down the rate of heat transfer to the depth of the structure. Through the coupling effect of the above three mechanisms of thermal radiation shielding, directional heat conduction, and thermal capacity buffering, the thermal protection structure achieves spatial heat redistribution and temporal thermal buffering effects, so that the concentrated heat in the stagnation area is quickly dispersed to a larger surface area, thereby greatly reducing the local heat flux density and transforming the instantaneous thermal shock into a slow temperature rise process, avoiding thermal shock damage to the material.

[0039] The beneficial effects of the thermal protection structure that integrates heat protection and heat conduction in this application are as follows:

[0040] 1. The integrated "heat protection-heat dissipation" design solves the problem of functional fragmentation in traditional thermal protection systems, reducing system weight and simplifying structural complexity. 2. Utilizing the high thermal conductivity and directional arrangement of carbon fiber bundles, effective heat dissipation is achieved, significantly reducing the temperature peak in the stagnation area, decreasing the temperature gradient on the nose surface, and improving the thermal stability of the thermal protection structure. 3. The porous carbon matrix not only fixes the carbon fiber bundles but also absorbs some heat through its thermal capacity, acting as a buffer while reducing the overall structural density. 4. The arrangement density and direction of the carbon fiber bundles can be adjusted according to the thermal environment requirements of different parts of the aircraft, achieving customized thermal management.

[0041] According to an embodiment of this application, the porosity of the porous carbon matrix 2 is 40%-60%. In this case, the porous carbon matrix 2 can both ensure the lightweight of the material and provide a stable support environment for the carbon fiber bundles 3.

[0042] According to an embodiment of this application, the nose 1 can be a quasi-conical shape that gradually expands from the nose 1 to the tail, and the nose 1 of the aircraft has a blunt-nosed configuration. The porous carbon matrix 2 is also a quasi-conical shape that gradually expands from the nose 1 to the tail. The number of carbon fiber bundles 3 is 8-12, and the multiple carbon fiber bundles 3 are evenly distributed circumferentially along the central axis of the nose 1. Two carbon fiber bundles 3 within the same axial section form a parabolic shape. In the above embodiment, the nose 1 can be a quasi-conical shape with a blunt-nosed configuration, and its shape is similar to a bullet. Correspondingly, the porous carbon matrix 2 is also a quasi-conical shape with a blunt-nosed configuration. This structure has a better aerodynamic shape, which can optimize the shock wave structure and reduce the peak heat flux by 15-25%. The parabolic carbon fiber bundles 3 can improve the axial heat conduction efficiency of the carbon fiber bundles 3.

[0043] According to an embodiment of this application, the distance from the starting point of the carbon fiber bundle 3 to the tip of the aircraft nose 1 is greater than the distance from the tip of the porous carbon matrix 2. Since the highest temperature of the nose 1 is at the tip of the nose 1, in the above embodiment, setting the starting point of the carbon fiber bundle 3 closer to the tip of the porous carbon matrix 2 can improve heat conduction efficiency.

[0044] According to embodiments of this application, the axial thermal conductivity of the carbon fiber bundle 3 is 300-1500 W / (m·K), and the thermal conductivity of the porous carbon matrix 2 is 80-120 W / (m·K). The high ratio of axial to radial thermal conductivity of the carbon fiber bundle 3 results in highly anisotropic thermal conductivity. Driven by a high temperature gradient, according to Fourier's law, the heat flux density generated within the carbon fiber bundle 3 far exceeds that reaching the porous carbon matrix 2, creating a "rapid channel" for heat from the high-temperature stagnation region to the low-temperature rear region. Based on the thermal diffusion equation, the high thermal diffusion coefficient of the carbon fiber bundle 3 shortens the thermal response time, achieving rapid heat redistribution.

[0045] According to an embodiment of this application, the anti-oxidation coating 4 is a silicon carbide and hafnium dioxide composite material with a thickness of 0.1-0.3 mm. In the above embodiment, the silicon carbide and hafnium dioxide composite material has a high emissivity of 0.7-0.8, which can effectively reflect 30-40% of the incident radiative heat flux, significantly reducing the net heat absorption of the thermal protection structure. In a high-temperature oxidation environment above 1500°C, the silicon carbide component in the anti-oxidation coating 4 can oxidize to form a dense silica glass phase, while hafnium dioxide provides excellent high-temperature stability due to its ultra-high melting point of 2810°C. The two components synergistically construct a self-healing oxidation barrier layer, reducing the oxygen diffusion coefficient by 3-4 orders of magnitude, thereby effectively preventing oxygen from eroding the substrate material.

[0046] In the thermal protection structure of this application, due to the difference in thermal expansion coefficients between the carbon fiber bundle 3 and the porous carbon matrix 2, interfacial debonding and crack propagation are prone to occur during repeated thermal cycling, severely affecting the thermal conductivity and mechanical integrity of the structure. The interface between the carbon fiber bundle 3 and the porous carbon matrix 2 is a critical weak point in the entire system. To improve the bonding strength between the carbon fiber bundle 3 and the porous carbon matrix 2, a pyrolytic carbon interface layer is provided between the carbon fiber bundle 3 and the porous carbon matrix 2, the thickness of which is 100-500 nm.

[0047] According to one embodiment of this application, the pyrolytic carbon interface layer can be formed by chemical vapor deposition (CVD). The formation mechanism of the pyrolytic carbon interface layer is based on the high-temperature cracking reaction of hydrocarbons. The CVD process requires the introduction of a hydrocarbon precursor, such as methane or propane, to form the pyrolytic carbon interface layer. At high temperatures, carbon atoms generated from the pyrolysis of methane molecules nucleate and epitaxially grow on the surface of carbon fibers, forming a pyrolytic carbon layer with a disordered graphite structure. This pyrolytic carbon layer has good lattice matching with the carbon fibers and can partially penetrate into the surface pores of the porous carbon matrix 2, forming a mechanically interlocked structure. The interfacial bonding strength between the porous carbon matrix 2 and the carbon fiber bundle 3 is enhanced by CVD technology, ensuring the stability of the overall structure.

[0048] According to one embodiment of this application, forming a pyrolytic carbon interface layer by chemical vapor deposition includes the following steps:

[0049] Step 1: Place the prefabricated thermal protection structure, which includes carbon fiber bundles and porous carbon matrix, in a reactor and heat it to 900-1100°C under an inert atmosphere.

[0050] Step 2: A hydrocarbon precursor is introduced, which undergoes a pyrolysis reaction at high temperature to generate carbon atoms, which are then deposited on the surface of the carbon fiber bundle to form a pyrolytic carbon interface layer.

[0051] Furthermore, in step 2, the hydrocarbon precursor is methane or propane.

[0052] To verify the heat dissipation capability of the thermal protection structure of this application, numerical simulation calculations were performed. The temperature at the stagnation point (the tip of the thermal protection structure) was set to a fixed value, and the heat dissipation capability of the thermal protection structure was determined by comparing the temperature at the tail end.

[0053] Figure 5 This is a thermal simulation diagram of a simple porous carbon matrix thermal protection structure. Figure 6 A thermal simulation diagram of a porous carbon matrix with a carbon fiber bundle thermal protection structure provided in the embodiments of this application.

[0054] In the numerical simulation, the stagnation temperature of both structures was set to 400℃. The numerical simulation results show that the thermal protection structure with carbon fiber bundles in the porous carbon matrix of this application can better and more evenly conduct heat from the porous carbon matrix cone than the simple carbon matrix thermal protection structure, and the stagnation temperature is lower. Therefore, the thermal protection structure of this application, which integrates heat protection and heat conduction, can improve the heat protection capability of the stagnation point.

[0055] This application's integrated thermal protection structure, combining heat protection and heat dissipation, not only solves the problem of the separation between "heat protection" and "heat dissipation" in traditional thermal protection systems, but also achieves efficient thermal management through the synergistic effect of carbon fiber bundles and porous carbon matrix, significantly improving the thermal protection performance of hypersonic aircraft.

[0056] The preferred embodiments of this disclosure have been described in detail above with reference to the accompanying drawings. However, this disclosure is not limited to the specific details of the above embodiments. Within the scope of the technical concept of this disclosure, various simple modifications can be made to the technical solutions of this disclosure, and these simple modifications all fall within the protection scope of this disclosure.

[0057] It should also be noted that the various specific technical features described in the above embodiments can be combined in any suitable manner without contradiction. To avoid unnecessary repetition, this disclosure will not describe the various possible combinations separately.

[0058] Furthermore, various different embodiments of this disclosure can be combined in any way, as long as they do not violate the spirit of this disclosure, they should also be regarded as the content disclosed in this disclosure.

Claims

1. A thermal protection structure integrating heat protection and heat conduction, used in the nose of an aircraft (1), characterized in that, include: A porous carbon matrix (2) adapted to the shape of the head (1) is disposed on the outer surface of the head (1). The porous carbon matrix (2) is a three-dimensional network structure with pores. Multiple carbon fiber bundles (3) are distributed circumferentially along the head (1). The multiple carbon fiber bundles (3) are radially embedded in the porous carbon matrix (2). Each carbon fiber bundle (3) extends from the head (1) to the tail to form a heat conduction channel. And an antioxidant coating (4) disposed on the outer surface of the porous carbon matrix (2).

2. The thermal protection structure according to claim 1, characterized in that, The porous carbon matrix (2) has a porosity of 40%-60%.

3. The thermal protection structure according to claim 1, characterized in that, The head (1) is a quasi-conical shape that gradually expands from the head (1) to the tail and has a blunt-tipped body configuration. The porous carbon matrix (2) is also a quasi-conical shape that gradually expands from the head (1) to the tail and has a blunt-tipped body configuration. The number of carbon fiber bundles (3) is 8-12. Multiple carbon fiber bundles (3) are evenly distributed circumferentially along the central axis of the machine head (1). Two carbon fiber bundles (3) in the same axial section form a parabolic shape.

4. The thermal protection structure according to claim 3, characterized in that, The distance from the starting point of the carbon fiber bundle (3) to the tip of the aircraft nose (1) is greater than the distance from the tip of the porous carbon matrix (2).

5. The thermal protection structure according to claim 1, characterized in that, The axial thermal conductivity of the carbon fiber bundle (3) is 300-1500 W / (m·K), and the thermal conductivity of the porous carbon matrix (2) is 80-120 W / (m·K).

6. The thermal protection structure according to claim 1, characterized in that, The antioxidant coating (4) is a silicon carbide and hafnium dioxide composite material with a thickness of 0.1-0.3 mm.

7. The thermal protection structure according to claim 1, characterized in that, A pyrolytic carbon interface layer is provided between the carbon fiber bundle (3) and the porous carbon matrix (2), and the thickness of the pyrolytic carbon interface layer is 100-500 nm.

8. The thermal protection structure according to claim 7, characterized in that, The pyrolytic carbon interface layer is formed by chemical vapor deposition.

9. The thermal protection structure according to claim 8, characterized in that, The formation of a pyrolytic carbon interface layer via chemical vapor deposition includes the following steps: Step 1: Place the prefabricated thermal protection structure, which includes carbon fiber bundles and porous carbon matrix, in a reactor and heat it to 900-1100°C under an inert atmosphere. Step 2: A hydrocarbon precursor is introduced, which undergoes a pyrolysis reaction at high temperature to generate carbon atoms, which are then deposited on the surface of the carbon fiber bundle to form a pyrolytic carbon interface layer.

10. The thermal protection structure according to claim 9, characterized in that, In step 2, the hydrocarbon precursor is methane or propane.

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