Aircraft design material selection method based on part design requirements

Abstract

The invention belongs to the field of aircraft design and manufacturing, and relates to an aviation key casting airworthiness verification method. And according to the weight coefficient of each performance index, the comprehensive matching score of the calculation material and the cost-effectiveness ratio of the calculation material, sorting the cost-effectiveness ratios of all candidate materials, and selecting the material with the highest cost-effectiveness ratio as a recommended material selection scheme of the part. The method is more accurate in demand matching. Aiming at functional positioning of different parts (for example, structural parts need high strength, and avionic parts need insulativity), weight distribution is dynamically adjusted to ensure that the material performance is highly matched with a design target, and performance redundancy or insufficiency is avoided.

Description

Technical Field

[0001] This invention belongs to the field of aircraft design and manufacturing, and relates to a method for verifying the airworthiness of key aviation castings. Background Technology

[0002] In aircraft design, component design requirements are complex and diverse. Different components have varying material performance requirements due to differences in function, service environment, and load-bearing capacity. For example, high-temperature components in aircraft engines require materials with excellent high-temperature resistance, high strength, and good oxidation resistance; while structural components such as aircraft wings require materials with high specific strength, high specific stiffness, and good fatigue performance. Traditional aircraft design material selection methods have many problems. On the one hand, material selection often relies on the designer's experience and limited material handbook data, lacking comprehensive and efficient material analysis methods. This results in the inability to fully consider all design requirements of components during the material selection process, leading to the selection of suboptimal materials that may affect the overall performance, safety, and service life of the aircraft. On the other hand, existing material selection methods struggle to quickly identify materials that meet the specific design requirements of components when faced with a large amount of material types and performance data. This makes the material selection process time-consuming and labor-intensive, severely impacting the progress and efficiency of aircraft design. Furthermore, with the continuous development of aviation technology, new materials are constantly emerging, further increasing the difficulty of traditional material selection methods. Therefore, there is a need for a method that can accurately and efficiently select materials for aircraft design based on component design requirements, and achieve the highest cost-effectiveness ratio through reasonable optimization and matching. Summary of the Invention

[0003] Purpose of the invention: The purpose of this invention is to address the shortcomings of existing material selection methods by proposing an aircraft design material selection method based on component design requirements, thereby improving material selection efficiency, the matching degree between material properties and component design requirements, and achieving a high cost-effectiveness ratio for material selection in case of failure.

[0004] The technical solution of the present invention is as follows: An aircraft design material selection method based on component design requirements includes the following steps: Step 1: Collect performance data and cost information for various materials.

[0005] Step 2 involves cleaning and verifying the collected data to ensure its accuracy and consistency.

[0006] Step 3: Regularly update the material performance database to promptly incorporate performance data and cost information of new materials, as well as changes in the performance and cost of existing materials.

[0007] Step 4 involves defining the functions and design requirements of each component of the aircraft in detail.

[0008] Step 5 uses auxiliary engineering software to perform mechanical analysis, thermal analysis, and fluid analysis on the parts to determine the stress, temperature distribution, and the influence of other environmental factors on the parts under different working conditions.

[0009] Step 6: Based on the analysis results, and in conjunction with the overall design goals and safety standards of the aircraft, formulate specific material performance requirements for each part and generate a standardized requirements report.

[0010] Step 7: Determine the weighting coefficients of each material performance index based on the functional importance of the part, the severity of the working environment, and the impact of different performance indicators on the reliability and safety of the part.

[0011] Step 8: Adjust and optimize the weighting coefficients.

[0012] Step 9: Extract performance and cost data of all potentially applicable materials for the part from the material properties database.

[0013] Step 10 compares the actual performance data of each material with the performance requirements of the part design to calculate the matching degree. For performance exceeding the requirements, an upper limit can be appropriately set to avoid excessive pursuit of high performance leading to increased costs.

[0014] Step 11: Calculate the overall matching score of the material based on the weight coefficients of each performance index. The calculation formula is: Overall matching score = matching degree of performance index 1 × weight coefficient of performance index 1 + matching degree of performance index 2 × weight coefficient of performance index 2 + matching degree of performance index 3 × weight coefficient of performance index 3 + ... + matching degree of performance index n × weight coefficient of performance index n.

[0015] Step 12 calculates the cost-effectiveness ratio of the materials. The formula is: Cost-effectiveness ratio = Overall matching score ÷ Material cost.

[0016] Furthermore, step 13 includes sorting all candidate materials by their cost-effectiveness and selecting the material with the highest cost-effectiveness as the recommended material for the part.

[0017] Furthermore, step 14 involves a comprehensive evaluation of the selected materials, taking into account factors such as material availability, processing technology maturity, and ease of maintenance.

[0018] Furthermore, step 15 uses computer simulation technology to simulate and verify the performance of the selected material under actual working conditions of the part, ensuring that the selected material can meet the design requirements and safety requirements of the part.

[0019] Furthermore, step 16 may involve material performance testing as needed to further verify and optimize the simulation results, ultimately determining the optimal material selection scheme.

[0020] Furthermore, in step 1, the performance data and cost information are obtained through material manuals, experimental testing, and cooperation with material suppliers.

[0021] Furthermore, in step 5, the auxiliary engineering software is specifically CAE.

[0022] Furthermore, in step 10, the matching degree is calculated specifically in percentage form.

[0023] Furthermore, in step 15, the computer simulation technology specifically refers to finite element analysis.

[0024] Technical effects: This method establishes a quantitative evaluation system, transforming key indicators such as material performance, cost, processability, and reliability into calculable weighting coefficients, thereby achieving a scientific ratio of multi-dimensional parameters. Its advantages are as follows: 1. More precise demand matching. By dynamically adjusting the weight allocation based on the functional positioning of different parts (such as high strength for structural components and insulation for avionics components), we ensure that the material performance is highly consistent with the design goals and avoid performance redundancy or insufficiency.

[0025] 2. More efficient cost control. While meeting minimum performance requirements, the cost of material procurement, processing, and lifecycle maintenance are balanced using weighted coefficients, prioritizing the most cost-effective solution.

[0026] 3. Enhanced systematicness and traceability in material selection. By clarifying the decision-making logic of each indicator's weight, reliance on subjective experience is reduced, making the material selection process more transparent and reproducible. Simultaneously, it facilitates rapid parameter adjustments during design iterations to adapt to different operating conditions or cost constraints, ultimately achieving the optimal balance between performance, safety, and economy in the overall aircraft design. Attached Figure Description

[0027] Figure 1 This is a flowchart illustrating the implementation of the present invention. Detailed Implementation

[0028] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below in conjunction with the embodiments of the present invention. Obviously, the described embodiments are only some embodiments of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0029] The features and illustrative embodiments of various aspects of the present invention will now be described in detail. In the following detailed description, numerous specific details are set forth in order to provide a thorough understanding of the invention. However, it will be apparent to those skilled in the art that the invention may be practiced without requiring some of these specific details. The following description of embodiments is merely intended to provide a better understanding of the invention by illustrating examples of the invention. The invention is by no means limited to any specific setup and method set forth below, but covers any improvements, substitutions, and modifications to the structures, methods, and devices without departing from the spirit of the invention. In the following description, well-known structures and techniques are not shown to avoid unnecessarily obscuring the invention.

[0030] In the description of this invention, it should be noted that the directions or positional relationships indicated by terms such as "center," "upper," "lower," "left," "right," "vertical," "horizontal," "inner," and "outer" are based on the stated directions or positional relationships and are only for the convenience of describing and simplifying the invention, and should not be construed as limiting the invention. Furthermore, the use of ordinal numbers (e.g., "first and second," etc.) is for distinguishing objects and is not limited to this order, and should not be construed as indicating or implying relative importance.

[0031] In the description of this invention, it should be noted that, unless otherwise explicitly specified and limited, the terms "installation," "connection," and "linking" should be interpreted broadly, encompassing both direct connection and indirect connection via an intermediate medium. Those skilled in the art can understand the specific meaning of these terms in this invention based on the specific circumstances.

[0032] It should be noted that, unless otherwise specified, the embodiments of the present invention and the features thereof can be combined with each other, and the various embodiments can be referenced and cited in each other. The present invention will now be described in detail with reference to and in conjunction with the embodiments.

[0033] Example 1: Taking the selection of materials for aircraft fuselage skin as an example: 1. Part Design Requirements Analysis: The aircraft fuselage skin primarily serves to maintain the aerodynamic shape of the fuselage, bear some aerodynamic loads, and protect the internal structure. CAE analysis was used to determine the aerodynamic pressure, temperature range, and vibration experienced by the fuselage skin during flight. Based on the analysis results, the performance requirements for the fuselage skin materials were determined to be: sufficient strength and stiffness (to withstand aerodynamic loads), good corrosion resistance (to adapt to different atmospheric environments), relatively light weight (to reduce the overall weight of the aircraft and improve fuel efficiency), and low cost (to reduce aircraft manufacturing costs). Specific performance requirements are as follows: tensile strength ≥ 400 MPa, elastic modulus ≥ 70 GPa, density ≤ 2.8 g / cm³, salt spray corrosion resistance ≥ 500 hours, and cost ≤ 100 RMB / kg.

[0034] 2. Determination of weighting coefficients: The weighting coefficients for each performance index are determined, with tensile strength having a weighting coefficient of 0.25, elastic modulus having a weighting coefficient of 0.2, density having a weighting coefficient of 0.25, salt spray corrosion resistance having a weighting coefficient of 0.2, and cost having a weighting coefficient of 0.1.

[0035] 3. Material optimization and cost-effectiveness calculation: Several candidate materials were selected from the material performance database, such as 2024-T3 aluminum alloy sheet, 7075-T6 aluminum alloy sheet, and magnesium alloy AZ31B. The matching degree of each material in various performance indicators was calculated. 1) 2024-T3 aluminum alloy sheet: Tensile strength 427MPa (100% matching), elastic modulus 72GPa (100% matching), density 2.78g / cm³ (100% matching), salt spray corrosion resistance 300 hours (60% matching), cost 78 yuan / kg (100% matching). Overall matching score = 100%×0.25 + 100%×0.2 + 100%×0.25 + 60%×0.2 + 100%×0.1 = 25 + 20 + 25 + 12 + 10 = 92 points. Cost-effectiveness ratio = 92 ÷ 78 ≈ 1.18.

[0036] 2) Aluminum Alloy 7075-T6: Tensile strength 537MPa (100% matching), elastic modulus 71GPa (100% matching), density 2.796g / cm³ (100% matching), salt spray corrosion resistance 550 hours (100% matching), cost 95 yuan / kg (100% matching). Overall matching score = 100%×0.25 + 100%×0.2 + 100%×0.25 + 100%×0.2 + 100%×0.1 = 25 + 20 + 25 + 20 + 10 = 100 points. Cost-effectiveness ratio = 100 ÷ 95 ≈ 1.05.

[0037] 3) Magnesium Alloy AZ31B: Tensile strength 260MPa (matching degree 65%), elastic modulus 45GPa (matching degree 64.3%), density 1.78g / cm³ (matching degree 100%), salt spray corrosion resistance 100 hours (matching degree 20%), cost 78 yuan / kg (matching degree 100%). Overall matching score = 65%×0.25 + 64.3%×0.2 + 100%×0.25 + 20%×0.2 + 100%×0.1 = 16.25 + 12.86 + 25 + 4 + 10 = 68.11 points. Cost-effectiveness ratio = 68.11 ÷ 78 ≈ 0.873.

[0038] 4. Material selection plan determined: By comparing the cost-effectiveness of various materials, aluminum alloy 2024-T3 has the highest cost-effectiveness (1.18), so aluminum alloy 2024-T3 was selected as the material for the aircraft fuselage skin.

[0039] 5. Material selection verification: The selected aluminum alloy 2024-T3 was simulated and verified. Finite element analysis software was used to simulate the stress distribution and deformation of the fuselage skin under aerodynamic loads. The results showed that its strength and stiffness met the design requirements. Simultaneously, salt spray corrosion resistance tests were conducted. Although its salt spray corrosion resistance was slightly lower than the ideal requirements, it can meet the needs of practical applications with appropriate surface treatment (such as applying an anti-corrosion coating). After comprehensive evaluation, the material selection scheme was determined to be feasible.

[0040] The above detailed embodiments are a description of the present invention. It should not be considered that the specific embodiments of the present invention are limited to these descriptions. For those skilled in the art, several simple deductions and substitutions can be made without departing from the concept of the present invention, and all of these should be considered to fall within the protection scope of the present invention.

Claims

1. A method for selecting materials in aircraft design based on component design requirements, characterized in that, Includes the following steps: Step 1: Collect performance data and cost information for various materials; Step 2 involves cleaning and verifying the collected data to ensure its accuracy and consistency. Step 3: Regularly update the material performance database to promptly incorporate performance data and cost information of new materials, as well as changes in the performance and cost of existing materials. Step 4 involves defining the functions and design requirements of each component of the aircraft in detail. Step 5: Use auxiliary engineering software to perform mechanical analysis, thermal analysis, and fluid analysis on the parts to determine the stress, temperature distribution, and the influence of other environmental factors on the parts under different working conditions. Step 6: Based on the analysis results, and in conjunction with the overall design goals and safety standards of the aircraft, formulate specific material performance requirements for each part and generate a standardized requirements report; Step 7: Determine the weighting coefficients of each material performance index based on the functional importance of the part, the severity of the working environment, and the factors affecting the reliability and safety of the part according to different performance indicators. Step 8: Adjust and optimize the weighting coefficients; Step 9: Extract performance and cost data of all potentially applicable materials for the part from the material properties database; Step 10: For each material performance index, compare the actual performance data of the material with the performance index requirements of the part design and calculate the matching degree. For performance exceeding the requirements, an appropriate upper limit can be set to avoid excessive pursuit of high performance leading to increased costs. Step 11: Calculate the comprehensive matching score of the material based on the weight coefficients of each performance index. The calculation formula is: Comprehensive matching score = matching degree of performance index 1 × weight coefficient of performance index 1 + matching degree of performance index 2 × weight coefficient of performance index 2 + matching degree of performance index 3 × weight coefficient of performance index 3 + ... + matching degree of performance index n × weight coefficient of performance index n. Step 12 calculates the cost-effectiveness ratio of the materials. The formula is: Cost-effectiveness ratio = Overall matching score ÷ Material cost.

2. The method as described in claim 1, characterized in that, It also includes step 13, which sorts all candidate materials by their cost-effectiveness and selects the material with the highest cost-effectiveness as the recommended material for the part.

3. The method as described in claim 2, characterized in that, Step 14 involves further evaluating the selected materials by comprehensively considering factors such as material availability, processing technology maturity, and ease of maintenance.

4. The method as described in claim 3, characterized in that, Step 15 uses computer simulation technology to simulate and verify the performance of the selected material under actual working conditions of the part, to ensure that the selected material can meet the design requirements and safety requirements of the part.

5. The method as described in claim 4, characterized in that, Step 16: Material performance testing may be conducted as needed to further verify and optimize the simulation results, and finally determine the optimal material selection scheme.

6. The method as described in claim 1, characterized in that, In step 1, performance data and cost information are obtained through material manuals, experimental testing, and cooperation with material suppliers.

7. The method as described in claim 1, characterized in that, In step 5, the auxiliary engineering software is specifically CAE.

8. The method as described in claim 1, characterized in that, In step 10, the matching degree is calculated using a percentage.

9. The method as described in claim 1, characterized in that, In step 15, the computer simulation technology specifically refers to finite element analysis.

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