A design method for controllable ablation heat protection structure for aircraft

Abstract

The present invention discloses a controllable ablative heat protection structure design method for aircraft, including: obtaining the first total thickness of a single-layer heat protection structure based on aerodynamic heating environment parameters and the first initial thickness of a first ablative heat protection material; obtaining the second total thickness of a multi-layer heat protection structure based on aerodynamic heating environment parameters, the second initial thickness of the first layer of heat protection material, the third initial thickness of the second layer of heat protection material and the interface characteristic temperature constraint condition; judging whether the difference between the second total thickness of the multi-layer heat protection structure and the first total thickness of the single-layer heat protection structure meets a preset condition, and if so, obtaining the optimal thickness of the multi-layer heat protection structure. The multi-layer heat protection structure designed by the controllable ablative heat protection structure design method of the present invention has different material compositions of each heat protection layer, and different materials can have different temperature resistance and thermal conductivity, which meets the requirements of both high temperature resistance and low average thermal conductivity as a whole.

Description

Technical Field

[0001] The present invention belongs to the technical field of hypersonic aircraft thermal protection system design, and more specifically, relates to a controllable ablation heat protection structure design method, electronic equipment and medium for aircraft. Background Art

[0002] Thermal protection systems are the cornerstone for the development and guarantee of safe service of hypersonic vehicles and reusable vehicles in extreme environments. In order to improve the survivability of aircraft and break through the "new thermal barrier" problem of thermal protection, efficient thermal protection systems must improve their ability to resist extreme service environments and meet the temperature resistance, durability and reliability requirements of aircraft thermal protection systems. In recent years, with the rapid development of related aerospace and military needs, the research and application of thermal protection systems have made great progress and gradually become a relatively independent and growing research field. The emergence of ablative heat protection technology can be seen as a revolutionary new technical approach to fundamentally overcome the aerodynamic heating problem of high-speed aircraft. Ablative heat protection refers to the fact that in a heat flow environment, heat protection materials can undergo a variety of physical and chemical changes that absorb heat energy, such as decomposition, melting, and sublimation, and consume a large amount of heat through the mass loss of the material itself, so as to prevent heat flow from entering the structure.

[0003] At present, ablative thermal protection systems mainly adopt single-layer thermal protection structures. However, with the increasing requirements of modern high-speed aircraft for lightweight thermal protection systems, new requirements are put forward for the density and thermal protection performance of thermal protection systems, that is, the thermal insulation performance is getting higher and higher, and the density is getting lower and lower. Traditional thermal insulation materials with low thermal conductivity have relatively low temperature resistance, while materials that can withstand high and ultra-high temperatures have relatively high thermal conductivity. Therefore, it is very difficult to achieve ultra-high temperature resistance and ultra-low thermal conductivity by single-layer thermal protection materials.

[0004] Therefore, it is expected to invent a controllable ablation heat protection structure design method for aircraft. The controllable ablation heat protection structure designed by the present invention can solve the problem that the heat protection materials used in the single-layer heat protection structure in the prior art are difficult to achieve ultra-high temperature resistance and ultra-low thermal conductivity at the same time. Summary of the invention

[0005] The purpose of the present invention is to propose a design method for a controllable ablation heat protection structure for an aircraft. The controllable ablation heat protection structure designed by the present invention solves the problem that the heat protection materials used in the single-layer heat protection structure in the prior art are difficult to achieve ultra-high temperature resistance and ultra-low thermal conductivity at the same time.

[0006] In order to achieve the above object, the present invention provides a controllable ablation heat protection structure design method for an aircraft, comprising:

[0007] Step 1: Obtain the aerodynamic heating environment parameters of the aircraft and set the constraint conditions for the thermal protection structure design;

[0008] Step 2: Selecting a first ablative heat protection material for designing a single-layer heat protection structure from a heat protection material library and setting a first initial thickness of the first ablative heat protection material;

[0009] Step 3: obtaining a first total thickness of a single-layer heat protection structure based on the aerodynamic heating environment parameters and the first initial thickness of the first ablative heat protection material;

[0010] Step 4: using the first ablative heat-protective material as a first-layer heat-protective material for designing a multi-layer heat-protective structure, and selecting a second ablative heat-protective material from the heat-protective material library as a second-layer heat-protective material;

[0011] Step 5: setting the interface characteristic temperature constraint condition, and setting the second initial thickness of the first layer of heat protection material and the third initial thickness of the second layer of heat protection material;

[0012] Step 6: Based on the aerodynamic heating environment parameters, the second initial thickness of the first layer of heat protection material, the third initial thickness of the second layer of heat protection material and the interface characteristic temperature constraint condition, the first target thickness of the first layer of heat protection material and the second target thickness of the second layer of heat protection material are obtained, thereby obtaining a second total thickness of the multi-layer heat protection structure;

[0013] Step 7: Determine whether the difference between the second total thickness of the multi-layer heat protection structure and the first total thickness of the single-layer heat protection structure meets a preset condition, and if so, obtain the optimal thickness of the multi-layer heat protection structure.

[0014] Optionally, the step 3 includes:

[0015] Step 31: Based on the pneumatic heating environment parameters and the first initial thickness of the first ablative heat-protective material, a pneumatic heating ablation thermal response calculation program is used to calculate a first ablation retreat and a first temperature response of the first ablation heat-protective material;

[0016] Step 32: Based on the calculation results of the first ablation retreat and the first temperature response and the constraint conditions, determine whether the single-layer heat protection structure meets the preset heat protection requirements. If so, use the first initial thickness as the first total thickness. If not, change the first initial thickness and repeat step 31 until the preset heat protection requirements are met.

[0017] Optionally, step 6 includes:

[0018] Step 61: Based on the aerodynamic heating environment parameters and the second initial thickness of the first layer of heat protection material, a second ablation retreat and a second temperature response of the first layer of heat protection material are calculated using an aerodynamic heating ablation thermal response calculation program;

[0019] Step 62: Based on the calculation results of the second ablation retreat and the second temperature response and the interface characteristic temperature constraint condition, determine whether the first layer of heat protection material meets the preset heat protection requirement; if so, use the second initial thickness as the first target thickness; if not, change the second initial thickness and repeat step 61 until the preset heat protection requirement is met;

[0020] Step 63: Based on the aerodynamic heating environment parameters, the first target thickness of the first layer of heat protection material and the third initial thickness of the second layer of heat protection material, the aerodynamic heating ablation thermal response calculation program is applied to calculate the third ablation setback and the third temperature response of the multi-layer heat protection structure;

[0021] Step 64: Based on the calculation results of the third ablation setback and the third temperature response and the constraints of the heat protection structure design, determine whether the multi-layer heat protection structure meets the preset heat protection requirement; if so, use the third initial thickness as the second target thickness; if not, change the third initial thickness and repeat step 63 until the preset heat protection requirement is met;

[0022] Step 65: Add the first target thickness and the second target thickness to obtain the second total thickness of the multi-layer heat protection structure.

[0023] Optionally, after step 7, the following steps are further included:

[0024] If the difference between the second total thickness of the multi-layer heat protection structure and the first total thickness of the single-layer heat protection structure does not meet the preset condition, the interface characteristic temperature constraint condition is changed, and steps 5 to 6 are repeated until the difference meets the preset condition.

[0025] Optionally, the aerodynamic heating environment parameters include the heat flux density q along the time history t or 、High temperature gas recovery enthalpy h r , wall pressure p and surface shear force τ.

[0026] Optionally, the heat protection material library includes carbon / phenolic composite materials, lightweight carbon / phenolic composite materials, quartz / phenolic composite materials, lightweight quartz / phenolic composite materials and honeycomb reinforced heat protection materials.

[0027] Optionally, the constraint condition includes the surface density ρ of the aircraft thermal protection structuresur , total thickness h and maximum temperature t on the back of the vehicle max at least one of them.

[0028] Optionally, in step 4, the density of the first layer of heat-proof material is greater than the density of the second layer of heat-proof material;

[0029] The thermal conductivity of the first layer of heat-proof material is greater than the thermal conductivity of the second layer of heat-proof material.

[0030] An electronic device, comprising:

[0031] A memory storing executable instructions;

[0032] A processor runs the executable instructions in the memory to implement the controllable ablation heat protection structure design method for an aircraft.

[0033] A computer-readable storage medium stores a computer program, which, when executed by a processor, implements the controllable ablation heat protection structure design method for an aircraft.

[0034] The beneficial effects of the present invention are:

[0035] Through the controllable ablation heat protection structure design method of the present invention, firstly, a first ablation heat protection material for single-layer heat protection structure design is selected and a first total thickness of the single-layer heat protection structure is obtained, and then a first layer of heat protection material for multi-layer heat protection structure design is selected, wherein the first ablation heat protection material is used as the first layer of heat protection material, and then a second ablation heat protection material is selected from the heat protection material library as the second layer of heat protection material, and the first target thickness of the first layer of heat protection material and the second target thickness of the second layer of heat protection material are obtained by setting the interface characteristic temperature constraint condition and the constraint condition, thereby obtaining the second total thickness of the multi-layer heat protection structure, and finally, the second total thickness of the multi-layer heat protection structure is subtracted from the first total thickness of the single-layer heat protection structure, and if the difference obtained meets the preset conditions, the optimal thickness of the multi-layer heat protection structure is obtained. The multi-layer heat protection structure designed by the controllable ablation heat protection structure design method of the present invention has different material compositions of each heat protection layer, and different materials can have different temperature resistance and thermal conductivity, thereby meeting the requirements of both high temperature resistance and low average thermal conductivity as a whole.

[0036] Other features and advantages of the present invention will be described in detail in the following detailed description. BRIEF DESCRIPTION OF THE DRAWINGS

[0037] The above and other objects, features and advantages of the present invention will become more apparent through a more detailed description of exemplary embodiments of the present invention in conjunction with the accompanying drawings, wherein like reference numerals generally represent like components throughout the exemplary embodiments of the present invention.

[0038] Figure 1 A flow chart of a method for designing a controllable ablation heat protection structure for an aircraft according to an embodiment of the present invention is shown. DETAILED DESCRIPTION

[0039] The preferred embodiments of the present invention will be described in more detail below. Although the preferred embodiments of the present invention are described below, it should be understood that the present invention can be implemented in various forms and should not be limited by the embodiments set forth herein. On the contrary, these embodiments are provided to make the present invention more thorough and complete, and to fully convey the scope of the present invention to those skilled in the art.

[0040] A controllable ablation heat protection structure design method for an aircraft according to the present invention comprises:

[0041] Step 1: Obtain the aerodynamic heating environment parameters of the aircraft and set the constraints for the thermal protection structure design;

[0042] Step 2: Selecting a first ablative heat protection material for designing a single-layer heat protection structure from a heat protection material library and setting a first initial thickness of the first ablative heat protection material;

[0043] Step 3: Based on the aerodynamic heating environment parameters and the first initial thickness of the first ablative heat protection material, a first total thickness of the single-layer heat protection structure is obtained;

[0044] Step 4: Use the first ablative heat protection material as the first layer of heat protection material for the multi-layer heat protection structure design, and select the second ablative heat protection material from the heat protection material library as the second layer of heat protection material;

[0045] Step 5: Set the interface characteristic temperature constraint condition, and set the second initial thickness of the first layer of heat protection material and the third initial thickness of the second layer of heat protection material;

[0046] Step 6: Based on the aerodynamic heating environment parameters, the second initial thickness of the first layer of heat protection material, the third initial thickness of the second layer of heat protection material and the interface characteristic temperature constraint condition, the first target thickness of the first layer of heat protection material and the second target thickness of the second layer of heat protection material are obtained, thereby obtaining the second total thickness of the multi-layer heat protection structure;

[0047] Step 7: Determine whether the difference between the second total thickness of the multi-layer heat protection structure and the first total thickness of the single-layer heat protection structure meets a preset condition. If so, obtain the optimal thickness of the multi-layer heat protection structure.

[0048] Specifically, the controllable ablation heat protection structure design method of the present invention first selects the first ablation heat protection material for the single-layer heat protection structure design and obtains the first total thickness of the single-layer heat protection structure, and then selects the first layer of heat protection material for the multi-layer heat protection structure design, wherein the first ablation heat protection material is used as the first layer of heat protection material, and then the second ablation heat protection material is selected from the heat protection material library as the second layer of heat protection material, and the first target thickness of the first layer of heat protection material and the second target thickness of the second layer of heat protection material are obtained by setting the interface characteristic temperature constraint condition and the constraint condition, thereby obtaining the second total thickness of the multi-layer heat protection structure, and finally subtracting the second total thickness of the multi-layer heat protection structure from the first total thickness of the single-layer heat protection structure, and if the difference obtained meets the preset conditions, the optimal thickness of the multi-layer heat protection structure is obtained. The multi-layer heat protection structure designed by the controllable ablation heat protection structure design method of the present invention has different material compositions of each heat protection layer, and different materials can have different temperature resistance and thermal conductivity, which meets the requirements of both high temperature resistance and low average thermal conductivity as a whole.

[0049] Furthermore, in the design process of the multi-layer heat protection structure, the thickness of each heat protection layer can be set according to the needs of the specific application environment, which is difficult to achieve with a single-layer heat protection structure.

[0050] In one example, step 3 includes:

[0051] Step 31: Based on the aerodynamic heating environment parameters and the first initial thickness of the first ablative heat protection material, a first ablation retreat and a first temperature response of the first ablative heat protection material are calculated using an aerodynamic heating ablation thermal response calculation program;

[0052] Step 32: Based on the calculation results and constraints of the first ablation setback and the first temperature response, determine whether the single-layer heat protection structure meets the preset heat protection requirements. If so, use the first initial thickness as the first total thickness. If not, change the first initial thickness and repeat step 31 until the preset heat protection requirements are met.

[0053] Specifically, based on the calculation results and constraints of the first ablation setback and the first temperature response, judging whether the single-layer heat protection structure meets the preset heat protection requirements can be understood as: based on the calculation results of the first ablation setback and the first temperature response, judging whether the calculation results of the first ablation setback and the first temperature response meet the constraints, wherein, in actual applications, the calculation results of the first ablation setback and the first temperature response are the total ablation setback amount of the first ablation heat protection material and the maximum temperature on the back of the first ablation heat protection material.

[0054] In one example, step 6 includes:

[0055] Step 61: Based on the aerodynamic heating environment parameters and the second initial thickness of the first layer of heat protection material, a second ablation retreat and a second temperature response of the first layer of heat protection material are calculated using an aerodynamic heating ablation thermal response calculation program;

[0056] Step 62: Based on the calculation results of the second ablation setback and the second temperature response and the interface characteristic temperature constraint condition, determine whether the first layer of heat protection material meets the preset heat protection requirement. If so, use the second initial thickness as the first target thickness. If not, change the second initial thickness and repeat step 61 until the preset heat protection requirement is met.

[0057] Step 63: Based on the aerodynamic heating environment parameters, the first target thickness of the first layer of heat protection material and the third initial thickness of the second layer of heat protection material, the aerodynamic heating ablation thermal response calculation program is used to calculate the third ablation setback and the third temperature response of the multi-layer heat protection structure;

[0058] Step 64: based on the calculation results after the third ablation and the third temperature response and the constraints of the heat protection structure design, determine whether the multi-layer heat protection structure meets the preset heat protection requirement; if so, use the third initial thickness as the second target thickness; if not, change the third initial thickness and repeat step 63 until the preset heat protection requirement is met;

[0059] Step 65: Add the first target thickness and the second target thickness to obtain a second total thickness of the multi-layer heat protection structure.

[0060] Specifically, based on the second ablation setback and the second temperature response calculation results and the interface characteristic temperature constraint conditions, judging whether the first layer of heat protection material meets the preset heat protection requirements can be understood as: based on the second ablation setback and the second temperature response calculation results, judging whether the second ablation setback and the second temperature response calculation results meet the interface characteristic temperature constraint conditions, wherein, in actual applications, the second ablation setback and the second temperature response calculation results are the total ablation setback amount of the first layer of heat protection material and the maximum temperature on the back side of the heat protection material; based on the third ablation and the third temperature response calculation results and the constraints of the heat protection structure design, judging whether the multi-layer heat protection structure meets the preset heat protection requirements can be understood as: based on the third ablation and the third temperature response calculation results, judging whether the third ablation and the third temperature response calculation results meet the constraints, wherein, in actual applications, the third ablation and the third temperature response calculation results are the total ablation setback amount of the second layer of heat protection material and the maximum temperature on the back side of the second layer of heat protection material.

[0061] In one example, step 7 further includes:

[0062] If the difference between the second total thickness of the multi-layer heat protection structure and the first total thickness of the single-layer heat protection structure does not meet the preset conditions, the interface characteristic temperature constraint is changed and steps 5 to 6 are repeated until the difference meets the preset conditions.

[0063] In one example, the aerodynamic heating environment parameters include the heat flux density q along the time history t or 、High temperature gas recovery enthalpy h r , wall pressure p and surface shear force τ.

[0064] In one example, the thermal protection material library includes carbon / phenolic composites, lightweight carbon / phenolic composites, quartz / phenolic composites, lightweight quartz / phenolic composites, and honeycomb reinforced thermal protection materials.

[0065] In one example, the constraints include the surface density ρ of the aircraft thermal protection structure sur , total thickness h and maximum temperature t on the back of the vehicle max at least one of them.

[0066] In one example, in step 4, the density of the first layer of heat protection material is greater than the density of the second layer of heat protection material;

[0067] The thermal conductivity of the first layer of heat protection material is greater than the thermal conductivity of the second layer of heat protection material.

[0068] An electronic device, the electronic device comprising:

[0069] A memory storing executable instructions;

[0070] A processor runs executable instructions in the memory to implement the controllable ablation heat protection structure design method for aircraft.

[0071] A computer-readable storage medium stores a computer program, which, when executed by a processor, implements the controllable ablation heat protection structure design method for an aircraft.

[0072] Embodiment 1

[0073] like Figure 1 As shown,

[0074] A controllable ablation heat protection structure design method for an aircraft, comprising:

[0075] Step 1: Obtain the aerodynamic heating environment parameters of the aircraft and set the constraints for the thermal protection structure design;

[0076] Step 2: Selecting a first ablative heat protection material for designing a single-layer heat protection structure from a heat protection material library and setting a first initial thickness of the first ablative heat protection material;

[0077] Step 3: Based on the aerodynamic heating environment parameters and the first initial thickness of the first ablative heat protection material, a first total thickness of the single-layer heat protection structure is obtained;

[0078] Step 4: Use the first ablative heat protection material as the first layer of heat protection material for the multi-layer heat protection structure design, and select the second ablative heat protection material from the heat protection material library as the second layer of heat protection material;

[0079] Step 5: Set the interface characteristic temperature constraint condition, and set the second initial thickness of the first layer of heat protection material and the third initial thickness of the second layer of heat protection material;

[0080] Step 6: Based on the aerodynamic heating environment parameters, the second initial thickness of the first layer of heat protection material, the third initial thickness of the second layer of heat protection material and the interface characteristic temperature constraint condition, the first target thickness of the first layer of heat protection material and the second target thickness of the second layer of heat protection material are obtained, thereby obtaining the second total thickness of the multi-layer heat protection structure;

[0081] Step 7: Determine whether the difference between the second total thickness of the multi-layer heat protection structure and the first total thickness of the single-layer heat protection structure meets a preset condition. If so, obtain the optimal thickness of the multi-layer heat protection structure.

[0082] Specifically, step 11: obtaining the aerodynamic heating environment parameters of the aircraft, wherein the aerodynamic heating environment parameters include the heat flux density q along the time history t=200s or =1.0MW / m 2 、High temperature gas recovery enthalpy h r =20.0MW / m 2 , wall pressure p = 10.0 kPa and surface shear force τ = 100 Pa;

[0083] Step 12: Set the constraints for heat protection structure design. The constraints are the maximum temperature t on the back of the heat protection material at t = 200s. max =200±2℃;

[0084] Step 13: Design a single-layer heat protection structure, select a carbon / phenolic composite material from the heat protection material library as the first ablative heat protection material and set the first initial thickness of the carbon / phenolic composite material to The density of carbon / phenolic composite material is 1420kg / m 3 , specific heat capacity is 1.2kJ / (kg·K), thermal conductivity is 0.75W / (m·K);

[0085] Step 14: Based on the aerodynamic heating environment parameters, the initial temperature of the carbon / phenolic composite material is set to 20°C, and the boundary condition on the back of the heat protection material is set to an adiabatic boundary;

[0086] Step 15: Based on the pneumatic heating environment parameters and the first initial thickness of the carbon / phenolic composite material The ablation thermal response calculation program of pneumatic heating ablation was used to calculate the ablation retreat and temperature response of carbon / phenolic composite materials, and the maximum temperature t on the back side at t = 200s was obtained. max =562℃;

[0087] Step 16: Maximum Temperature t on the Backside of Carbon / Phenolic Composite max =562℃ does not meet the constraints of the heat protection system t max =200±2℃, increase the first initial thickness of the carbon / phenolic composite material and repeat steps 14-15 until the constraints of the heat protection design are met, and obtain t max =The thickness of the carbon / phenolic composite material at 200°C is h1=27.26mm;

[0088] Step 17: Design a multi-layer heat protection structure. The heat protection material of the first layer is the same as that of the single-layer heat protection structure. Select a carbon / phenolic composite material. Select a lightweight carbon / phenolic composite material from the heat protection material library as the second layer of heat protection material. The selection principle of the second layer of heat protection material is: the density and thermal conductivity of the second layer of heat protection material are lower than those of the carbon / phenolic composite material. The density of the lightweight carbon / phenolic is 720kg / m 3 , specific heat capacity is 1.5kJ / (kg·K), thermal conductivity is 0.15W / (m·K);

[0089] Step 18: Set the interface characteristic temperature constraint condition, the interface characteristic temperature constraint condition is that the maximum temperature of the interface at t = 200s is 700 ± 2 °C;

[0090] Step 19: Set the thickness of the first layer of heat protection material, that is, the second initial thickness of the carbon / phenolic composite material Based on the aerodynamic heating environment, the back boundary condition is set as an adiabatic boundary;

[0091] Step 20: Based on the aerodynamic heating environment parameters and the first layer of heat protection material, the second initial thickness of the carbon / phenolic composite material The second ablation retreat and second temperature response of carbon / phenolic composites were calculated using the pneumatic heating ablation thermal response calculation program, and the maximum interface temperature at t = 200s was 1176℃;

[0092] Step 21: Backside temperature t of carbon / phenolic composite material at t = 200s max =1176℃ is greater than the interface characteristic temperature constraint at t=200s, that is, the maximum interface temperature at t=200s is 700±2℃, the thickness of the first layer of heat protection material is increased, and steps 19-20 are repeated until the interface characteristic temperature constraint is met, and t max= The first target thickness h of the carbon / phenolic composite material at 700°C 21 =18.83mm;

[0093] Step 21: Set the thickness of the second layer of heat protection material, that is, the third initial thickness of the lightweight carbon / phenolic composite material The surface of the first layer of heat protection material is set to aerodynamic heating environment parameters, and the back boundary condition of the second layer of heat protection material is set to adiabatic boundary;

[0094] Step 22: Based on the aerodynamic heating environment parameters and the first target thickness h of the first layer of heat protection material 21 =18.83mm and the third initial thickness of the second layer of heat protection material The third ablation retreat and the third temperature response of the multilayer heat protection structure are calculated by using the aerodynamic heating ablation thermal response calculation program, so as to obtain the maximum interface temperature t at t = 200s. max =319°C;

[0095] Step 23: The maximum interface temperature t of the second layer of heat protection material, i.e., the lightweight carbon / phenolic composite material at t = 200s max =319℃ does not meet the constraints of the heat protection system. Increase the thickness of the second layer of heat protection material and repeat steps 21-22 until the constraints of the heat protection design are met. Then, t max = The second target thickness h of the lightweight carbon / phenolic composite material at 201°C 22 =7.00mm;

[0096] Step 24: Obtain the second total thickness of the multi-layer heat protection structure:

[0097] h2=h 21 +h 22 =18.83mm+7.00mm=25.83mm;

[0098] Step 25: Determine whether the difference between the second total thickness of the multi-layer heat protection structure and the first total thickness of the single-layer heat protection structure meets the preset condition, that is, whether Δh = h1-h2 = 27.26mm-25.83mm = 1.43mm meets the preset condition. Based on the maximum interface temperature of 700±2℃ at t = 200s and the back surface temperature t max = 200 ± 2 ° C constraint, the second total thickness of the multilayer heat protection structure of "carbon / phenolic composite material + lightweight carbon / phenolic composite material" is 25.83, and the back surface temperature t at t = 200s max =200±2℃ constraint, the first total thickness of the single-layer heat protection structure of the "carbon / phenolic composite material" is 27.26mm, and the multi-layer heat protection design reduces the thickness of the heat protection layer by 1.43mm, and the surface density is reduced from 38.71kg / m2 to 31.78kg / m2 , thereby simultaneously achieving ultra-high temperature resistance and ultra-low thermal conductivity of the multi-layer heat protection structure.

[0099] Embodiment 2

[0100] The present disclosure provides an electronic device, which includes: a memory storing executable instructions; a processor running the executable instructions in the memory to implement the above-mentioned controllable ablation heat protection structure design method for aircraft.

[0101] An electronic device according to an embodiment of the present disclosure includes a memory and a processor.

[0102] The memory is used to store non-temporary computer-readable instructions. Specifically, the memory may include one or more computer program products, which may include various forms of computer-readable storage media, such as volatile memory and / or non-volatile memory. The volatile memory may, for example, include random access memory (RAM) and / or cache memory (cache), etc. The non-volatile memory may, for example, include read-only memory (ROM), hard disk, flash memory, etc.

[0103] The processor may be a central processing unit (CPU) or other forms of processing units having data processing capabilities and / or instruction execution capabilities, and may control other components in the electronic device to perform desired functions. In one embodiment of the present disclosure, the processor is used to run the computer-readable instructions stored in the memory.

[0104] Those skilled in the art should be able to understand that in order to solve the technical problem of how to obtain a good user experience, the present embodiment may also include well-known structures such as a communication bus and an interface, and these well-known structures should also be included in the protection scope of the present disclosure.

[0105] For detailed description of this embodiment, reference may be made to the corresponding descriptions in the aforementioned embodiments, which will not be repeated here.

[0106] Embodiment 3

[0107] The present disclosure provides a computer-readable storage medium storing a computer program, which, when executed by a processor, implements the above-mentioned controllable ablation heat protection structure design method for an aircraft.

[0108] According to the computer-readable storage medium of the embodiment of the present disclosure, non-transitory computer-readable instructions are stored thereon. When the non-transitory computer-readable instructions are executed by a processor, all or part of the steps of the method of each embodiment of the present disclosure are executed.

[0109] The above-mentioned computer-readable storage media include, but are not limited to: optical storage media (e.g., CD-ROM and DVD), magneto-optical storage media (e.g., MO), magnetic storage media (e.g., magnetic tape or mobile hard disk), media with built-in rewritable non-volatile memory (e.g., memory card) and media with built-in ROM (e.g., ROM box).

[0110] The embodiments of the present invention have been described above, and the above description is exemplary, not exhaustive, and is not limited to the disclosed embodiments. Many modifications and changes will be apparent to those skilled in the art without departing from the scope and spirit of the described embodiments.

Claims

1. A method for designing a controllable ablation heat protection structure for an aircraft, characterized in that: include: Step 1: Obtain the aerodynamic heating environment parameters of the aircraft and set the constraint conditions for the thermal protection structure design; Step 2: Selecting a first ablative heat protection material for designing a single-layer heat protection structure from a heat protection material library and setting a first initial thickness of the first ablative heat protection material; Step 3: obtaining a first total thickness of a single-layer heat protection structure based on the aerodynamic heating environment parameters and the first initial thickness of the first ablative heat protection material; The step 3 specifically includes: step 31: based on the aerodynamic heating environment parameters and the first initial thickness of the first ablative heat protection material, applying the aerodynamic heating ablation thermal response calculation program to calculate the first ablation setback and the first temperature response of the first ablation heat protection material; step 32: based on the calculation results of the first ablation setback and the first temperature response and the constraint conditions, judging whether the single-layer heat protection structure meets the preset heat protection requirement, if so, taking the first initial thickness as the first total thickness, if not, changing the first initial thickness and repeating the step 31 until the preset heat protection requirement is met; Step 4: using the first ablative heat-protective material as a first-layer heat-protective material for designing a multi-layer heat-protective structure, and selecting a second ablative heat-protective material from the heat-protective material library as a second-layer heat-protective material; Step 5: setting the interface characteristic temperature constraint condition, and setting the second initial thickness of the first layer of heat protection material and the third initial thickness of the second layer of heat protection material; Step 6: Based on the aerodynamic heating environment parameters, the second initial thickness of the first layer of heat protection material, the third initial thickness of the second layer of heat protection material and the interface characteristic temperature constraint condition, the first target thickness of the first layer of heat protection material and the second target thickness of the second layer of heat protection material are obtained, thereby obtaining a second total thickness of the multi-layer heat protection structure; Step 7: Determine whether the difference between the second total thickness of the multi-layer heat protection structure and the first total thickness of the single-layer heat protection structure meets a preset condition, and if so, obtain the optimal thickness of the multi-layer heat protection structure.

2. The method for designing a controllable ablation heat protection structure for an aircraft according to claim 1, characterized in that: The step 6 comprises: Step 61: Based on the aerodynamic heating environment parameters and the second initial thickness of the first layer of heat protection material, a second ablation retreat and a second temperature response of the first layer of heat protection material are calculated using an aerodynamic heating ablation thermal response calculation program; Step 62: Based on the calculation results of the second ablation retreat and the second temperature response and the interface characteristic temperature constraint condition, determine whether the first layer of heat protection material meets the preset heat protection requirement; if so, use the second initial thickness as the first target thickness; if not, change the second initial thickness and repeat step 61 until the preset heat protection requirement is met; Step 63: Based on the aerodynamic heating environment parameters, the first target thickness of the first layer of heat protection material and the third initial thickness of the second layer of heat protection material, the aerodynamic heating ablation thermal response calculation program is applied to calculate the third ablation setback and the third temperature response of the multi-layer heat protection structure; Step 64: Based on the calculation results of the third ablation setback and the third temperature response and the constraints of the heat protection structure design, determine whether the multi-layer heat protection structure meets the preset heat protection requirement; if so, use the third initial thickness as the second target thickness; if not, change the third initial thickness and repeat step 63 until the preset heat protection requirement is met; Step 65: Add the first target thickness and the second target thickness to obtain the second total thickness of the multi-layer heat protection structure.

3. The method for designing a controllable ablation heat protection structure for an aircraft according to claim 1, characterized in that: After step 7, the following steps are also included: If the difference between the second total thickness of the multi-layer heat protection structure and the first total thickness of the single-layer heat protection structure does not meet the preset condition, the interface characteristic temperature constraint condition is changed, and steps 5 to 6 are repeated until the difference meets the preset condition.

4. The method for designing a controllable ablation heat protection structure for an aircraft according to claim 1, characterized in that: The aerodynamic heating environment parameters include the heat flux density q along the time history t or 、High temperature gas recovery enthalpy h r , wall pressure p and surface shear force τ.

5. The method for designing a controllable ablation heat protection structure for an aircraft according to claim 1, characterized in that: The heat protection material library includes carbon / phenolic composite materials, lightweight carbon / phenolic composite materials, quartz / phenolic composite materials, lightweight quartz / phenolic composite materials and honeycomb reinforced heat protection materials.

6. The method for designing a controllable ablation heat protection structure for an aircraft according to claim 1, characterized in that: The constraints include the surface density ρ of the aircraft thermal protection structure sur , total thickness h and maximum temperature t on the back of the vehicle max at least one of them.

7. The method for designing a controllable ablation heat protection structure for an aircraft according to claim 1, characterized in that: In step 4, the density of the first layer of heat-proof material is greater than the density of the second layer of heat-proof material; The thermal conductivity of the first layer of heat-proof material is greater than the thermal conductivity of the second layer of heat-proof material.

8. An electronic device, characterized in that: The electronic device comprises: A memory storing executable instructions; A processor, wherein the processor runs the executable instructions in the memory to implement the controllable ablation heat protection structure design method for an aircraft according to any one of claims 1-7.

9. A computer-readable storage medium, characterized in that: The computer-readable storage medium stores a computer program, and when the computer program is executed by a processor, the method for designing a controllable ablation heat protection structure for an aircraft according to any one of claims 1 to 7 is implemented.

Images