Model design method for icing wind tunnel test of large aircraft component

Abstract

The invention provides a model design method for a large aircraft component icing wind tunnel test. The model design method comprises the steps that the large aircraft component for the icing wind tunnel test and the size of an icing wind tunnel test section are determined; according to the size of an icing wind tunnel test section, segmenting a part model used for a test in the full-aircraft model; rectifying the segmented part model; key parameters of the rectified component model are calculated; key parameters of the whole-aircraft model and the component model are compared, and whether the component model can effectively represent the icing characteristics of the whole aircraft in the area or not is evaluated; and after repeated iterative optimization is carried out until the component model meets the test requirement, it is considered that the test model is successfully designed, and the current component model is determined as the model shape for the icing test. According to the invention, the complex flow field of the complex part of the large aircraft can be accurately simulated, so that the icing condition of the complex part of the large aircraft can be accurately simulated.

Description

Technical Field

[0001] The invention relates to the technical field of model design for icing wind tunnel tests, and in particular to a model design method for icing wind tunnel tests of large aircraft components. Background Art

[0002] When an aircraft is in flight, water droplets in the clouds at a certain altitude will condense. When the temperature is zero degrees Celsius or below, ice will form on the aircraft surface and may further form ice accumulation. Aircraft icing is one of the main hidden dangers threatening aircraft flight safety and has received widespread attention in recent years. The research methods for aircraft icing include flight tests, icing wind tunnel tests and numerical simulations, among which icing wind tunnel tests are the most important means to study aircraft icing phenomena.

[0003] Currently, the world's largest icing wind tunnel test section is less than 8 meters in length and width, while large aircraft are over 30 meters in length and width. Therefore, it is not possible to conduct a 1:1 scale icing wind tunnel test of the entire aircraft under the current test equipment conditions.

[0004] At the same time, since the theory of wind tunnel tests on icing of scaled models is still imperfect, when studying the icing phenomenon of large aircraft at the current technical level, only 1:1 scale icing tests of aircraft components can be conducted. There is currently no unified design specification for the design of wind tunnel test models for icing of large aircraft components.

[0005] For aircraft parts with regular shapes such as wings or horizontal tails, there are usually two methods for designing icing wind tunnel test models:

[0006] 1. Directly cut a section of the outer wing of the wing or horizontal tail as the test model. This scheme is called a 3D test model scheme;

[0007] 2. Extract a typical airfoil from the wing or horizontal tail, and then directly stretch it in the vertical direction of the airfoil (called 2D test model scheme) or stretch it with a sweep angle (called 2.5D test model scheme) to form a test model.

[0008] However, for irregularly shaped aircraft components such as the nose and landing gear, due to the complex flow field around the components, it is difficult to accurately simulate the icing conditions of these complex components if the model is designed according to the above two methods. Summary of the invention

[0009] In order to solve the design problem of icing wind tunnel test models for irregularly shaped aircraft components, the purpose of the present invention is to provide a model design method for icing wind tunnel tests of large aircraft components. The model designed using this method can accurately simulate the complex flow field around the complex components, and thus can accurately simulate the icing conditions of the complex components.

[0010] The present invention achieves the above-mentioned purpose through the following technical solutions:

[0011] A model design method for icing wind tunnel testing of large aircraft components, the method comprising the following steps:

[0012] Calculate the key parameters of the full machine model, including pressure distribution, droplet collection efficiency and ice shape;

[0013] Determine the size of large aircraft components and icing wind tunnel test sections for icing wind tunnel tests;

[0014] According to the size of the icing wind tunnel test section, the component models used for testing are divided from the full aircraft model;

[0015] Rectify the segmented component models;

[0016] Calculate the key parameters of the rectified component model;

[0017] Compare the key parameters of the whole aircraft model and the component model to evaluate whether the component model can effectively represent the icing characteristics of the whole aircraft in this area;

[0018] After repeated iterative optimization, when the component model meets the test requirements, the test model design is considered successful, and the current component model is determined to be the model shape for the icing experiment.

[0019] According to a model design method for icing wind tunnel testing of large aircraft components provided by the present invention, the key parameters of the whole aircraft model are calculated, including:

[0020] The whole aircraft model was numerically simulated using aircraft icing simulation software to calculate the pressure distribution, water droplet collection rate and ice shape on the model surface;

[0021] According to a model design method for large aircraft component icing wind tunnel testing provided by the present invention, the component model for testing is segmented from the full aircraft model according to the size of the icing wind tunnel test section, comprising:

[0022] According to the size of the icing wind tunnel test section and the original size of the full aircraft model, the components of concern in the icing test are identified in the full aircraft model, and the selected components are segmented from the full aircraft model using segmentation tools in the 3D model.

[0023] According to a model design method for large aircraft component icing wind tunnel test provided by the present invention, the segmented component model is rectified, comprising:

[0024] Movable control points 1-5 are set, and control curves are determined by each control point and the corresponding fixed point on the component model. The control curve forms a surface through a multi-section surface or a filling function, and the surfaces are connected to obtain a modified shape. If the modified shape needs to be changed, some or all of the control points 1-5 need to be moved, thereby changing the control curve and then changing the surface.

[0025] According to a model design method for icing wind tunnel tests of large aircraft components provided by the present invention, when comparing key parameters of the full aircraft model and the component model, the pressure distribution, water droplet collection efficiency and ice shape are compared in sequence, that is, the pressure distribution is compared first, then the water droplet collection efficiency is compared, and finally the ice shape is compared.

[0026] According to a model design method for icing wind tunnel tests of large aircraft components provided by the present invention, if, during comparison, it is found that there are large differences in key parameters between the full-aircraft model and the component model, it is necessary to readjust the rectification design of the component model and perform a new round of calculations and comparisons; through continuous iterative optimization, until there is no obvious difference in key parameters between the two, it is ensured that the test model can accurately reflect the icing conditions under the full-aircraft environment.

[0027] It can be seen that the present invention has the following beneficial effects:

[0028] 1. The present invention calculates the key parameters of the whole aircraft model (such as pressure distribution, water droplet collection efficiency and ice shape) and determines the specific test components and sizes accordingly, thereby ensuring that the test is directly carried out on the aircraft's icing-sensitive or key areas, improving the pertinence and effectiveness of the test, and can effectively solve the design difficulties of icing wind tunnel test models for complex components of large aircraft, with high practical value.

[0029] 2. In the design process, the present invention can restore the airflow state of the component in actual flight to the greatest extent by rectifying the component model and recalculating the key parameters, thereby ensuring the accuracy and reliability of the test results. In addition, by comparing the key parameters of the whole aircraft model and the component model, the rationality of the model design is further verified and the accuracy of the model is improved.

[0030] 3. Compared with the full aircraft model, the component model significantly reduces the test size and complexity, and reduces the cost and time of wind tunnel testing. At the same time, iterative optimization ensures that the model meets the test requirements, avoids unnecessary repeated tests, and improves test efficiency.

[0031] 4. This method allows for flexible design for different aircraft components and icing conditions, and has strong adaptability and scalability. For example, the test model and test plan can be quickly adjusted according to different aircraft models, different flight phases or different climate conditions.

[0032] 5. Standardized design process and key parameter calculation method ensure the repeatability of the test. No matter which team conducts the test or at what time, as long as the same design process and standards are followed, relatively consistent results can be obtained.

[0033] 7. The model design method proposed in the present invention can be completed using any commercial 3D design software and 3D ice shape calculation software, and does not require special usage skills, so it has the advantages of lower cost and simpler operation.

[0034] 8. The model design method proposed in the present invention is implemented in a digital way, which can accurately control the geometric parameters of the test model, so the designed test model is more accurate.

[0035] The present invention is further described in detail below in conjunction with the accompanying drawings and specific embodiments. BRIEF DESCRIPTION OF THE DRAWINGS

[0036] Figure 1 The present invention is a flow chart of an embodiment of a model design method for icing wind tunnel testing of large aircraft components.

[0037] Figure 2 The schematic diagram of the structure of a front fuselage model of a large aircraft is provided in an embodiment of a model design method for an icing wind tunnel test of a large aircraft component of the present invention.

[0038] Figure 3 The schematic diagram of the structure of the wave suppression trough component model after rectification is provided in an embodiment of the model design method for large aircraft component icing wind tunnel test of the present invention.

[0039] Figure 4 It is a pressure distribution comparison diagram of a full aircraft model (Example 1) and a wave suppression trough component model (Example 2) provided in an embodiment of a model design method for an icing wind tunnel test of a large aircraft component of the present invention.

[0040] Figure 5 It is a comparison diagram of the droplet collection efficiency distribution of the full aircraft model (Example 1) and the wave suppression trough component model (Example 2) provided in an embodiment of a model design method for a large aircraft component icing wind tunnel test of the present invention.

[0041] Figure 6 It is a comparison diagram of calculated ice shapes of a full aircraft model (calculation example 1) and a wave suppression trough component model (calculation example 2) provided in an embodiment of a model design method for a large aircraft component icing wind tunnel test of the present invention.

[0042] Figure 7 This is a physical picture of an icing wind tunnel test of a wave suppression trough component model provided in an embodiment of a model design method for an icing wind tunnel test of a large aircraft component of the present invention. DETAILED DESCRIPTION

[0043] In order to make the purpose, technical solution and advantages of the present invention clearer, the technical solution of the present invention will be clearly and completely described below in conjunction with the drawings of the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0044] Reference to "embodiments" herein means that a particular feature, structure, or characteristic described in conjunction with the embodiments may be included in at least one embodiment of the present application. The appearance of the phrase in various locations in the specification does not necessarily refer to the same embodiment, nor is it an independent or alternative embodiment that is mutually exclusive with other embodiments. It is explicitly and implicitly understood by those skilled in the art that the embodiments described herein may be combined with other embodiments.

[0045] See also Figures 1 to 7 The present invention provides a model design method for icing wind tunnel test of large aircraft components, the method comprising the following steps:

[0046] Step S1, calculating key parameters of the whole machine model, the key parameters including pressure distribution, water droplet collection efficiency and ice shape;

[0047] Step S2, determining large aircraft components for icing wind tunnel testing;

[0048] Step S3, determining the size of the icing wind tunnel test section;

[0049] Step S4, dividing the component models used for the test from the full aircraft model according to the size of the icing wind tunnel test section;

[0050] Step S5, rectifying the segmented component model;

[0051] Step S6, calculating key parameters of the rectified component model;

[0052] Step S7, comparing key parameters of the whole aircraft model and the component model to evaluate whether the component model can effectively represent the icing characteristics of the whole aircraft in the area;

[0053] Step S8, after repeated iterative optimization, until the component model meets the test requirements, the test model design is considered successful, and the current component model is determined to be the model shape for the icing experiment.

[0054] In the above step S1, the key parameters of the whole machine model are calculated, including:

[0055] Use aircraft icing simulation software (such as FENSAP-ICE software) to perform numerical simulation on the full aircraft model and calculate the pressure distribution, water droplet collection efficiency and ice shape on the model surface;

[0056] In the above step S4, the component models for the test are segmented from the full aircraft model according to the size of the icing wind tunnel test section, including:

[0057] According to the size of the icing wind tunnel test section and the original size of the full aircraft model, the components of concern in the icing test are identified in the full aircraft model, and the selected components are segmented from the full aircraft model using segmentation tools in the 3D model.

[0058] In the above step S5, the segmented component model is rectified, including:

[0059] Movable control points 1-5 are set, and control curves are determined by each control point and the corresponding fixed point on the component model. The control curve forms a surface through a multi-section surface or a filling function, and the surfaces are connected to obtain a modified shape. If the modified shape needs to be changed, some or all of the control points 1-5 need to be moved, thereby changing the control curve and then changing the surface.

[0060] In the above step S7, when comparing the key parameters of the whole machine model and the component model, the pressure distribution, water droplet collection efficiency and ice shape are compared in sequence, that is, the pressure distribution is compared first, then the water droplet collection efficiency is compared, and finally the ice shape is compared.

[0061] Specifically, when comparing, if it is found that there are large differences in key parameters between the whole-machine model and the component model, it is necessary to readjust the segmentation method or rectification design of the component model and conduct a new round of calculation and comparison; through continuous iterative optimization, until there is no obvious difference in key parameters between the two, it is ensured that the test model can accurately reflect the icing conditions under the whole-machine environment.

[0062] In this embodiment, pressure distribution is the basis of fluid dynamics analysis, which directly affects the aerodynamic performance and stability of the aircraft. Under icing conditions, the pressure distribution of key components such as wings and tail fins may change, which in turn affects lift, drag and control performance. By comparing the pressure distribution of the full aircraft model and the component model under the same or similar conditions, it can be preliminarily determined whether there are significant differences between the models and whether these differences may lead to changes in aerodynamic performance.

[0063] In this embodiment, the water droplet collection efficiency (WDCE) is an important indicator to measure the icing potential of the aircraft surface. It reflects the ability of the aircraft surface to capture and retain supercooled water droplets. After confirming that there is no significant difference in the pressure distribution, the water droplet collection efficiency is further analyzed. High water droplet collection efficiency means a higher risk of icing, so this step is crucial for evaluating the icing sensitivity of the model. By comparison, it is possible to identify which areas or components are more prone to icing, providing a basis for subsequent anti-icing and de-icing design.

[0064] In this embodiment, ice shape directly affects the aerodynamic characteristics and flight safety of the aircraft. Different types of ice (such as frost ice, clear ice, mixed ice) and different ice shapes (such as smooth ice layer, ice ridge, ice nodule, etc.) have different effects on the aircraft. Based on the analysis of pressure distribution and water droplet collection efficiency, the ice shapes of the whole aircraft model and the component model are finally compared.

[0065] In summary, comparative analysis in the order of pressure distribution, droplet collection efficiency, and ice shape can help to comprehensively evaluate the performance differences and safety of the full-machine model and component models under icing conditions.

[0066] Furthermore, in this embodiment, the pressure distribution, water droplet collection efficiency and ice shape of the whole machine model are used as a comparison benchmark and only need to be calculated once, while the pressure distribution, water droplet collection efficiency and ice shape of the component model need to be iteratively calculated multiple times according to the rectification effect.

[0067] Furthermore, in this embodiment, the three-dimensional design software used to rectify the aircraft components may be any commercial three-dimensional design software.

[0068] Furthermore, in this embodiment, the software for calculating the model pressure distribution, water droplet collection efficiency and ice shape can be any commercial software with a three-dimensional ice shape calculation function.

[0069] Furthermore, in this embodiment, when judging the difference in pressure distribution between the whole aircraft model and the component model, the aircraft icing simulation software is used to perform numerical simulation on the whole aircraft model and the component model to obtain pressure distribution diagrams of the two under different working conditions. By comparing the pressure distribution diagrams, it is observed whether there are significant differences in key parameters such as the pressure distribution curve, pressure gradient, and pressure extreme points.

[0070] Furthermore, in this embodiment, when determining the difference in water droplet collection efficiency between the whole machine model and the component model, the water droplet impact characteristics and the collection coefficient are calculated, and the water droplet collection efficiencies of the two under different working conditions are compared.

[0071] Furthermore, in this embodiment, when determining the difference in ice shape between the whole machine model and the component model, software is used to perform icing prediction on the whole machine model and the component model to obtain the ice shapes of the two under different icing conditions. The ice shape diagrams are compared to observe whether there are significant differences in ice shape, thickness, distribution and other characteristics.

[0072] In practical applications, for step S1, this embodiment uses FENSAP-ICE software to calculate the pressure distribution, water droplet collection efficiency and ice shape of a large amphibious aircraft full-machine model. The calculation model is a 1:1 scale full-machine model, and the model is placed in the free stream. The calculation state is: flight altitude 4500m, flight speed 105.56m / s, angle of attack 4.7°, total temperature -2.5℃, static temperature -8.05℃, water droplet diameter 20μm, liquid water content 0.47g / m3, and ice formation time 45 minutes. The calculation results are shown in Figure 4 , 5 , in “Example 1” in 6.

[0073] For step S2, a large amphibious aircraft is about 40 meters long and 38 meters wide. It is obviously impossible to conduct a 1:1 scale icing wind tunnel test on the whole aircraft under the current test equipment conditions. Therefore, a 1:1 scale wave suppression trough component is selected for the icing wind tunnel test.

[0074] For step S3, an icing test is performed in the FL-16 wind tunnel sub-test section of the China Aerodynamics Research and Development Center, where the size of the sub-test section is 4.8 m (width) × 3.2 m (height) × 9 m (length).

[0075] For step S4, Figure 2 As shown, Figure 2 The darker area is the wave suppression trough part cut from the front fuselage. The total length of the wave suppression trough is 10.5 meters, and the aircraft shape is symmetrical. In order to meet the test model's limit on the blockage of the icing wind tunnel, the right half of the 3.5-meter front end of the wave suppression trough was cut for testing. The icing intensity in this area is the most serious.

[0076] For step S5, Figure 3 As shown in the figure, the dark area is the prototype of the wave suppression trough component (the long plate-shaped component is the wave suppression plate), and the transparent area is the volume increased by the rectification process. The intercepted wave suppression trough model is rectified and reshaped using the three-dimensional design software CATIA.

[0077] For step S6, the pressure distribution, droplet collection efficiency and ice shape of the rectified trough model are calculated using FENSAP-ICE software. The calculation model is a 1:1 rectified and shaped trough model. The model is placed in the FL-16 wind tunnel. The calculation conditions are: flight altitude 4500m, flight speed 105.56m / s, angle of attack 4.7°, total temperature -2.5℃, static temperature -8.05℃, droplet diameter 20μm, liquid water content 0.47g / m3, and ice formation time 45 minutes. The calculation results are shown in Figure 4 , 5 , “Example 2” in 6.

[0078] For step S7, the pressure distribution, droplet collection efficiency and ice shape of the full machine model and the rectified wave suppression tank model are compared, such as Figure 4 As shown, Figure 4 A comparison of pressure distribution is shown, without the wave suppression plate: in the front area of ​​the two models, the pressure at the inlet of the wave suppression trough is the smallest, the pressure at the lower part of the inlet of the wave suppression trough is the largest, and the pressure in the rear area of ​​the two models is relatively small. Therefore, it can be judged that the pressure distribution of the two models is similar.

[0079] like Figure 5 As shown, Figure 5 The figure shows the comparison of water droplet collection efficiency. The wave suppression plate is not shown in the figure. The water droplet collection efficiency is relatively large at the lower part of the wave suppression trough inlet and the middle and rear part of the wave suppression trough in the front area of ​​the two models. The water droplet collection efficiency is relatively small in other areas of the two models. Therefore, it can be judged that the water droplet collection efficiency distribution of the two models is similar.

[0080] like Figure 6 As shown, Figure 6 A comparison of ice shapes is shown. The wave suppression plate is not shown in the figure. The amount of ice is large at the lower part of the wave suppression trough entrance and the middle and rear part of the wave suppression trough in the front area of ​​the two models, and the amount of ice is small in other areas of the two models. Therefore, it can be judged that the ice shape distribution of the two models is similar.

[0081] See also Figure 7The purpose of the icing wind tunnel test in this embodiment is to study whether the icing phenomenon affects the function of the wave suppression trough. Based on conservative considerations, the test ice shape is required to be more severe than the actual ice shape, but not too severe to lead to erroneous conclusions. Therefore, the judgment criteria for whether the test scheme is feasible include, in addition to the similarity of pressure, droplet collection efficiency and ice shape distribution, the requirement that the ice mass of the wave suppression plate pillars of the rectified wave suppression trough model is greater than the ice mass of the wave suppression plate pillars of the full-machine model, and not more than 20%. In this example, the ice mass of the wave suppression plate pillars of the rectified wave suppression trough model is 2.01 kg, and the ice mass of the wave suppression plate pillars of the full-machine model is 1.73 kg. The ice mass of the wave suppression plate pillars of the wave suppression trough model is 16.2% greater than the ice mass of the wave suppression plate pillars of the full-machine model, which can meet the test requirements.

[0082] If the comparison results are quite different, continue to use CATIA software to rectify and reshape the intercepted trough model, and then use FENSAP-ICE software to calculate the pressure distribution, droplet collection efficiency and ice shape of the trough model for comparison.

[0083] For step S8, when the pressure distribution, droplet collection efficiency and ice shape differences of the full machine model and the rectified trough model are reduced to an acceptable level to meet the test requirements, the current trough model can be determined as the model shape for the icing wind tunnel test.

[0084] like Figure 7 As shown, Figure 7 The actual picture of the icing wind tunnel test of the wave suppression tank model is shown. The test model is designed completely according to the method provided by the present invention, and the test results are consistent with expectations.

[0085] In summary, the present invention calculates the key parameters of the whole aircraft model (such as pressure distribution, water droplet collection efficiency and ice shape) and determines the specific test components and sizes accordingly, thereby ensuring that the test is carried out directly on the aircraft's icing-sensitive or critical areas, improving the pertinence and effectiveness of the test, and can effectively solve the design difficulties of icing wind tunnel test models for complex components of large aircraft, and has high practical value.

[0086] During the design process, the present invention can restore the airflow state of the component in actual flight to the greatest extent by rectifying the component model and recalculating key parameters, thereby ensuring the accuracy and reliability of the test results. In addition, by comparing the key parameters of the full-aircraft model and the component model, the rationality of the model design is further verified and the accuracy of the model is improved. Compared with the full-aircraft model, the component model significantly reduces the test size and complexity, and reduces the cost and time of wind tunnel tests. At the same time, iterative optimization is used to ensure that the model meets the test requirements, avoid unnecessary repeated tests, and improve test efficiency.

[0087] This method allows for flexible design for different aircraft components and icing conditions, and has strong adaptability and scalability. Standardized design processes and key parameter calculation methods ensure the repeatability of the test. Regardless of which team conducts the test or at what time, as long as the same design process and standards are followed, relatively consistent results can be obtained.

[0088] The model design method proposed in the present invention can be completed using any commercial 3D design software and 3D ice shape calculation software, and does not require special usage skills, so it has the advantages of lower cost and simpler operation. The model design method proposed in the present invention is implemented in a digital way, and can accurately control the geometric parameters of the test model, so the designed test model is more accurate.

[0089] The technical features of the above embodiments may be combined arbitrarily. To make the description concise, not all possible combinations of the technical features in the above embodiments are described. However, as long as there is no contradiction in the combination of these technical features, they should be considered to be within the scope of this specification.

[0090] The above-mentioned embodiments are only preferred embodiments of the present invention and cannot be used to limit the scope of protection of the present invention. Any non-substantial changes and substitutions made by technicians in this field on the basis of the present invention shall fall within the scope of protection required by the present invention.

Claims

1. A model design method for icing wind tunnel testing of large aircraft components, characterized in that: The following steps are involved: Calculate the key parameters of the full machine model, including pressure distribution, droplet collection efficiency and ice shape; Determine the size of large aircraft components and icing wind tunnel test sections for icing wind tunnel tests; According to the size of the icing wind tunnel test section, the component models used for testing are divided from the full aircraft model; Rectify the segmented component models; Calculate the key parameters of the rectified component model; Compare the key parameters of the whole aircraft model and the component model to evaluate whether the component model can effectively represent the icing characteristics of the whole aircraft in this area; After repeated iterative optimization, when the component model meets the test requirements, the test model design is considered successful, and the current component model is determined to be the model shape for the icing experiment.

2. The method according to claim 1, characterized in that The key parameters of the calculation of the whole machine model include: The aircraft icing simulation software is used to perform numerical simulation on the entire aircraft model and calculate the pressure distribution on the model surface, water droplet collection efficiency and ice shape.

3. The method according to claim 1, characterized in that The component models for testing are segmented from the full aircraft model according to the size of the icing wind tunnel test section, including: According to the size of the icing wind tunnel test section and the original size of the full aircraft model, the components of concern in the icing test are identified in the full aircraft model, and the selected components are segmented from the full aircraft model using segmentation tools in the 3D model.

4. The method according to claim 1, characterized in that: The step of rectifying the segmented component models comprises: Movable control points 1-5 are set, and control curves are determined by each control point and the corresponding fixed point on the component model. The control curve forms a surface through a multi-section surface or a filling function, and the surfaces are connected to obtain a modified shape. If the modified shape needs to be changed, some or all of the control points 1-5 need to be moved, thereby changing the control curve and then changing the surface.

5. The method according to claim 1, characterized in that: When comparing key parameters of the full machine model and the component model, the pressure distribution, water droplet collection efficiency and ice shape are compared in sequence, that is, the pressure distribution is compared first, then the water droplet collection efficiency is compared, and finally the ice shape is compared.

6. The method according to claim 5, characterized in that: During the comparison, if it is found that there are large differences in key parameters between the whole-machine model and the component model, it is necessary to readjust the rectification design of the component model and conduct a new round of calculations and comparisons; through continuous iterative optimization, until there is no obvious difference in key parameters between the two, it is ensured that the test model can accurately reflect the icing conditions under the whole-machine environment.

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