An aircraft deicing / electromagnetic low scattering compatible structure and its design method

Abstract

The present invention relates to the field of electronic materials, specifically to an aircraft de-icing / electromagnetic low-scattering compatible structure and a design method thereof. The present invention adopts a periodic high-wave-transmittance electric heating impedance network, conforms the impedance network to the wave-absorbing structure, and realizes high-wave-transmittance performance through patterned design of the impedance network; and by optimizing the key ice formation areas on the surface of the wave-absorbing structure to perform resistance control in partitioned areas, the purpose of zoned heating and concentrated heat adjustment is achieved, thereby achieving more targeted de-icing. The present invention can simultaneously realize the functions of electric heating de-icing and electromagnetic low-scattering, which can not only meet the needs of aircraft de-icing at high altitude or in cold environments, thereby protecting the flight safety of the aircraft, but also ensure that the aircraft's electromagnetic low-scattering function is not affected. It is expected to be applied to the safety protection and electromagnetic compatibility of future low-detectable aircraft, and become its important technical accumulation.

Description

Technical Field

[0001] The present invention relates to the field of electronic materials, and in particular to an aircraft deicing / electromagnetic low scattering compatible structure and a design method thereof. Background Art

[0002] When an aircraft encounters clouds containing supercooled water droplets during flight, exposed surfaces are susceptible to rapid accumulation and freezing of these droplets, forming ice. This phenomenon is known as aircraft icing. Aircraft icing can seriously endanger flight safety, disrupting the aerodynamic shape of the airfoil, reducing the lift coefficient and lift-to-drag ratio, and potentially causing premature stall. Aircraft anti-icing and de-icing technology is a key tool in reducing icing incidents and improving aircraft safety.

[0003] Current aircraft de-icing systems are primarily categorized by their method of operation: liquid, mechanical, thermal, and other types of systems. The basic principle of electrothermal de-icing systems is to convert electrical energy into thermal energy through heating elements on the aircraft's interior, which is then transferred to the aircraft's exterior to prevent and remove ice. Due to its high efficiency, reliability, and ease of control, it is currently one of the most widely used and promising aircraft de-icing methods.

[0004] For low-observable aircraft targets, the complexity of their mission environments requires consideration not only of potential icing issues but also of electromagnetic compatibility issues arising from the complex electromagnetic environment. To prevent the de-icing components of the aircraft from interfering with electromagnetic scattering, a de-icing / electromagnetic low-scattering compatible structure is required. While electric heating de-icing systems are suitable for aircraft de-icing due to their simple structure and high de-icing efficiency, the introduction of a metal network in electric heating systems inevitably introduces electromagnetic scattering issues. Therefore, achieving both high-efficiency de-icing and electromagnetic low-scattering compatibility becomes a key and challenging design challenge. Currently, there is limited research on the appropriate use of electric heating de-icing systems on such aircraft, and research on de-icing / electromagnetic low-scattering compatible structures is lacking. Summary of the Invention

[0005] In response to the above-mentioned problems or shortcomings, and to solve the problem of existing low-observable aircraft achieving both electromagnetic compatibility and efficient de-icing functions, the present invention provides an aircraft de-icing / electromagnetic low-scattering compatible structure and a design method thereof. A periodic high-transmittance electric heating impedance network is adopted, and the impedance network is conformally attached to the absorbing structure. The patterned design of the impedance network achieves high transmittance performance, and its heating power is adjusted by optimizing the size and resistance of the pattern.

[0006] The technical solutions of the present invention are as follows:

[0007] A design method for an aircraft deicing / electromagnetic low scattering compatible structure comprises the following steps:

[0008] Step 1: Design a high-transmittance periodically patterned impedance network. The basic unit of the impedance network is a square, and it is a double-layer structure consisting of a bottom substrate and a patterned upper resistive film. The same basic units are periodically arranged in a matrix to form the impedance network. The -1dB transmission frequency band is 2-16GHz, and the relative bandwidth is 155.56%. The entire frequency range from 2-18GHz is a low-reflection band below -10dB.

[0009] The patterned impedance network design can obtain a high-transmittance structure. The use of a periodic impedance network pattern can distribute the heat generated by power as evenly as possible over the entire covering surface of the wing while achieving high transmittance.

[0010] Step 2: Use the Navier-Stokes equations to solve the target aircraft wing model in the wing air flow field analysis; the three equations including conservation of mass, conservation of momentum, and conservation of energy are integrated into the integral form of the Navier-Stokes equations:

[0011]

[0012] Among them F C represents the convective flux, F v represents the viscous flux, W is the conserved variable, Ω is the control volume, Q is the heat flux, and S is the surface of the airfoil (control volume).

[0013] Then, introducing the perfect gas thermodynamic relationship and the Stokes ideal gas hypothesis to close the above Navier-Stokes equation, we get:

[0014]

[0015] ρH=ρE+P

[0016] γ is the adiabatic coefficient, which is 1.4 for air, (μ, v, ω) represent the three components of velocity, ρ represents density, P represents pressure, H is total enthalpy, and E is energy.

[0017] The pressure (flow parameters) of the air flow field of the wing is obtained under different working conditions, providing data for the calculation of water droplet impact characteristics and wing icing conditions.

[0018] Then, the water drop impact characteristics are calculated and expressed by the water drop collection coefficient, which refers to the actual impact amount W of water droplets impacting per unit span length per unit time. m With its maximum possible impact W max The ratio of E mIndicates that E m =W m / W, where W m The definition of is:

[0019] W m =V0(Y ou -Y ol )LWC=V0(y ou -y ol )LWC·L

[0020] V0 is the aircraft speed, Y ou 、Y ol is the starting coordinate of the chord length direction of the water droplet trajectory tangent to the upper and lower surfaces of the wing in the far field, y ou 、y ol Y ou 、Y ol The ratio of LWC to the chord length, LWC is the liquid water content, and L is the characteristic size of the object (chord length).

[0021] W max The definition of is:

[0022] W max =V0·M·LWC=V0·h·LWC·L

[0023] M is the projected height of the wing in the chord direction, and h is the ratio of M to the chord length.

[0024] Substituting the liquid water content, average droplet diameter, and calculated air flow pressure into the droplet impact calculation yields the droplet impact coefficient at the wing's leading edge (the droplet impact coefficient is highest at the wing's leading edge), correspondingly aligning with the icing's most severe condition at the leading edge. By calculating various operating conditions, the chord-length length, L1, of the target wing's iced area is determined.

[0025] Step 3: Conformally distribute the impedance network designed in step 1 on the surface of the target wing's absorbing structure. Then, based on L1 solved in step 2, draw a perpendicular line in the chord direction to obtain the corresponding wing leading edge area S+, and redesign the basic unit of the impedance network. Increase the resistance value R of the impedance film in the basic unit of the S+ area.

[0026] Based on the fact that ice accumulation is most severe on the leading edge of the wing, the present invention proposes a design method for concentrating heat in the S+ region of the wing's leading edge. This is because the length L2 of the wing's absorbing structure in the chord direction is greater than L1, and the electric heating system inevitably introduces electromagnetic scattering problems due to the introduction of a metal network. Therefore, it cannot be directly connected to the absorbing structure's key icing areas, but must be connected to the end of the absorbing structure. Therefore, the present invention achieves this goal by designing voltage division in different areas of the impedance network.

[0027] According to the formula of resistance R:

[0028]

[0029] Where S1 is the cross-sectional area of ​​the impedance network through which the current flows, ρ′ is the resistivity, and L3 is the length of the resistor film. If the cross-sectional area S1 is reduced, the resistance of this part will increase, thereby increasing the DC power supply voltage divided by this part. According to the average power density formula:

[0030]

[0031] P d Where is the average power per unit area, and U is the voltage. At this point, the heat output in this area increases. When the heat output in the S+ area increases relative to the non-S+ area, heat concentration is achieved. This area of ​​increased heat output precisely covers the target wing's most severely iced area (the S+ area), enabling more targeted de-icing.

[0032] Step 4: Perform electromagnetic simulation on the impedance network model redesigned in Step 3 to ensure that the target wing still meets the performance requirements of electromagnetic low scattering. If not, return to Step 3 to redesign the pattern shape of the impedance film in the basic unit of the wing leading edge area S+ until the target wing meets the performance requirements of electromagnetic low scattering.

[0033] Furthermore, in step 3, the resistance value of the impedance film is increased by reducing the width of the impedance film.

[0034] Furthermore, the resistance value of the impedance film is increased in step 3 by redesigning the pattern of the resistance film.

[0035] Furthermore, the resistance value of the impedance film in step 3 is increased by reducing the resistance value of the basic unit in the non-wing leading edge area S+.

[0036] Furthermore, the resistance value of the basic unit in the non-wing leading edge area S+ is reduced by increasing the width of the impedance film.

[0037] Furthermore, the resistance value of the basic unit in the non-wing leading edge area S+ is reduced by redesigning the pattern of the resistance film.

[0038] An aircraft de-icing / electromagnetic low-scattering compatible structure is designed using the above method. When in use, the wing leading edge area S+ is adaptively and conformally covered with the basic unit redesigned in step 3, and the non-wing leading edge area S+ of the absorbing structure is also adaptively and conformally covered, and then electrically heated for de-icing.

[0039] In summary, this invention has designed a multifunctional electromagnetic structure that simultaneously implements electric heating deicing and electromagnetic low-scattering functions. This not only meets the de-icing requirements of aircraft at high altitudes or in cold environments, thereby protecting flight safety, but also ensures that the aircraft's electromagnetic low-scattering function is not compromised. This structure is expected to be applied to the safety protection and electromagnetic compatibility of future low-observable aircraft, becoming a key technological advancement. BRIEF DESCRIPTION OF THE DRAWINGS

[0040] Figure 1 Schematic diagram of the structure of the periodic unit used in the embodiment;

[0041] Figure 2 The nephogram of the water droplet impact coefficient distribution of the NACA0012 airfoil in the embodiment;

[0042] Figure 3 : is a water droplet impact coefficient distribution curve of the NACA0012 airfoil in the embodiment;

[0043] Figure 4 The ice shape diagram (left) and icing condition cloud diagram (right) of the NACA0012 airfoil in the embodiment are shown;

[0044] Figure 5 : is an icing intensity distribution curve of the NACA0012 airfoil in the embodiment;

[0045] Figure 6 A three-dimensional structural diagram of the final non-uniform electric heating impedance network designed in the embodiment;

[0046] Figure 7 is a temperature distribution diagram of the non-uniform impedance network in the embodiment;

[0047] Figure 8 Graph showing the electromagnetic performance simulation results of the non-uniform impedance network in the embodiment;

[0048] Figure 9 Graph showing the RCS simulation results of a wing loaded with an impedance network in an embodiment;

[0049] Figure 10 A sample physical diagram of a non-uniform impedance network according to an embodiment;

[0050] Figure 11 A sample test environment for the embodiment;

[0051] Figure 12 Comparison between the sample test curve and the simulation curve of the embodiment;

[0052] Figure 13 This is an infrared image of the electrical heating of the non-uniform impedance network under room temperature conditions in the embodiment;

[0053] Figure 14FIG. 1 is a melting curve diagram of ice on a non-uniform impedance network in a cold environment in an embodiment. DETAILED DESCRIPTION

[0054] The technical solution of the present invention is described in detail below with reference to the accompanying drawings and embodiments.

[0055] This example uses the NACA0012 aircraft wing airfoil as an example to simulate water droplet impact and icing. The specific distribution of ice accumulation on the wing surface is studied. The ice exhibits a trumpet-shaped characteristic of being "thick at both ends and thinner in the middle." A partitioned impedance loading design is adopted for the impedance network. By changing the size of the structure in the middle region, the resistance value in this region is increased to achieve partitioned heating. The impedance network structure designed by this method has a -1dB transmission frequency band of 2-16.7GHz and a relative bandwidth of 157.22%. The entire frequency range from 2-18GHz is a low-reflection band below -10dB. The simulation and experimental structures are consistent, proving that this design can effectively de-ice while maintaining high wave transmittance. This design provides an effective solution for the integrated de-icing / electromagnetic low-scattering design of aircraft, and it is highly operational and easy to implement.

[0056] The design of this embodiment adopts a non-uniform pattern arrangement, and prints the following on a 0.1 mm thick polyimide film: Figure 1 The cross hollow pattern shown (yellow is the substrate) has a thermal conductivity of 2W·m -1 ·K -1 , using carbon paste as printing material, the square resistance of the printed pattern is The pattern width in the center is 0.8mm, while the width on the sides is 1.5mm, showing a non-uniform distribution. After connecting electrodes on both sides, they are applied to the leading edge of the wing (i.e., the surface of the absorbing structure), ultimately achieving the integrated functions of electric heating and electromagnetic scattering.

[0057] An integrated electric heating / electromagnetic scattering design for aircraft deicing includes the following steps:

[0058] Step 1: The basic unit structure of the impedance network pattern designed in this embodiment is as follows: Figure 1 As shown, a cross-shaped hollow impedance network is used, and the square resistance is Printed on polyimide film, it achieves performance with a transmittance of more than 80% and a reflectivity of less than 10% in the range of 2-18GHz.

[0059] Step 2: For the study of aircraft wing icing, this embodiment uses the NACA0012 wing model as the research object and uses the Navier-Stokes equations to solve the wing air flow field analysis.

[0060] Select the working conditions, determine the ambient temperature, water content, flight speed, and average diameter of water droplets. The ambient temperature is -5°C and the water content is 1g / m 3 , flight speed of 80 m / s, and average water droplet diameter of 20 microns, the pressure of the air flow field on the wing is obtained. These parameters and operating conditions are used to calculate the droplet impact characteristics and wing icing conditions. In this embodiment, the wing chord length is 1 m.

[0061] The calculated water droplet impact area of ​​the NACA0012 wing is concentrated on the leading edge, as shown in Figure 2 As shown, it shows a trend of being the largest in the middle and gradually decreasing to both sides. The curve changes as shown in Figure 3 As shown in (Y direction is the chord length direction). The icing situation of the wing is also concentrated in the leading edge, as shown in Figure 4 As shown in the figure, it presents a trumpet shape with thick ends and thinner middle, and no ice is accumulated when it extends to the rear of the wing. At this time, the ice area is within the area perpendicular to the chord length (-0.04m, 0.04m), as shown in the figure. Figure 5 shown.

[0062] Step 3: After obtaining the situation of the icing area, based on the premise of step 2, the non-uniform design of the impedance network pattern is carried out. The width of the impedance network pattern in the middle area is reduced. The basic unit width of the original model is 1.35mm. The width of the unit in the middle area of ​​the impedance network is reduced to 0.55mm, and the resistance value of this area is increased. The width of the unit in the area on both sides of the impedance network is increased to 1.55mm, and the resistance value of this area is reduced. Thus, a non-uniformly distributed impedance network structure is formed, such as Figure 6 As shown. In the comsol multi-physics field simulation calculation, the results can be obtained. The temperature rise is basically concentrated in the middle area. The temperature distribution is as follows Figure 7 shown.

[0063] Furthermore, the width of the impedance network pattern in other areas (non-S+ regions) can be simultaneously increased, increasing their cross-sectional area. This reduces the resistance in these areas, further reducing the DC power divider within them and thus reducing the heat generated. When the heat generation power in the central S+ region increases while that in the non-S+ regions on either side decreases, better heat concentration is achieved. This area with increased heat generation precisely covers the target wing's most severely iced area (S+ region), enabling more targeted de-icing.

[0064] Step 4: Use CST electromagnetic simulation software to perform electromagnetic simulation on the designed non-uniform structure, such as Figure 8As shown in the simulation of a planar structure, the -1dB transmission band of this structure is 2-16.7GHz, with a relative bandwidth of 157.22%. The entire frequency range from 2-18GHz is within the low-reflection band, below -10dB. This structure possesses high transmission performance and low reflectivity, making it an impedance film with excellent electromagnetic wave transmission performance.

[0065] The impedance network is covered to the leading edge of the wing loaded with the absorbing structure, such as Figure 9 As shown in the figure, the RCS reduction curve is obtained. It can be seen that at the azimuth angles of (-90°, +90°), the model loaded with the non-uniform impedance network not only does not cause an increase in RCS, but actually has a slight RCS reduction effect. This shows that the impedance network does not affect the electromagnetic scattering performance of the aircraft and is suitable for aircraft wings or similar curved structures.

[0066] Step 5: In order to verify the engineering feasibility of the model scheme, the model unit is designed to be 198mm*198mm in size and processed. Figure 10 The free space method is used to test the reflection and transmission performance of the sample. The test environment is as follows: Figure 11 As shown. Analyze the error between simulation results and test results, and get the simulation test comparison ( Figure 12 ) It can be seen that the difference is not big, and the expected effect is achieved, which verifies the effectiveness of the structure. Copper electrodes are attached to both ends of the impedance network, and then a DC source is loaded to test the electric heating characteristics of the sample. At room temperature, a voltage of 120V is applied. After the temperature stabilizes, the infrared image of the sample is measured as follows Figure 13 As shown, the calculated heat power of the middle part is 1450w / m 2 Then, the electric heating and de-icing of ice on the surface of the impedance network was tested at an ambient temperature of about -8°C. The temperature change curve is as follows: Figure 14 As shown in Figure 2, it fully demonstrates the effectiveness of the non-uniform impedance network design for temperature concentration.

[0067] As can be seen from the above examples, this embodiment simulates water droplet impact and icing on a NACA0012 aircraft wing. The specific distribution of ice accumulation on the wing surface is studied. The ice exhibits a characteristic "thick at the ends and thinner in the middle" trumpet shape. A zoned impedance loading design is employed for the impedance network. By varying the structural dimensions of the central region, the resistance in that region is increased, achieving zoned heating. The impedance network structure designed using this method exhibits a -1dB transmission frequency range of 2-16.7 GHz and a relative bandwidth of 157.22%. The entire frequency range from 2-18 GHz is within a low reflection band below -10 dB. Electromagnetic simulations of this zoned impedance loading model, performed on the leading edge of a NACA0012 airfoil, show that the model maintains high transmittance and low reflectivity, performing well in both TE and TM polarizations, with a slight decrease in RCS. This demonstrates that the impedance network does not affect the aircraft's electromagnetic scattering performance. The simulation and experimental structures are consistent, proving that this design can effectively de-ice while having high wave transmittance. This design provides an effective solution for the integrated de-icing / electromagnetic low-scattering design of aircraft, and it is highly operational and easy to implement.

[0068] In summary, the present invention utilizes a periodic, highly transparent, electrically heated impedance network, conforming it to an absorbing structure. The patterned design of the impedance network achieves high transmittance. Furthermore, by optimizing the resistance values ​​of key ice-forming areas on the absorbing structure's surface, zoned heating and concentrated heat are achieved, thereby enabling more targeted deicing. This invention simultaneously achieves both electric heating deicing and electromagnetic low-scattering functions, meeting the requirements for deicing aircraft at high altitudes or in cold environments, thereby protecting flight safety, while also ensuring the aircraft's electromagnetic low-scattering function is not compromised. It is expected to be applied to future low-observable aircraft safety protection and electromagnetic compatibility, becoming a key technological advancement.

Claims

1. A design method for an aircraft deicing / electromagnetic low scattering compatible structure, characterized in that: The following steps are involved: Step 1: Designing a periodically patterned impedance network with high transmittance; The basic unit of the impedance network is a square, a double-layer structure consisting of a bottom substrate and a patterned upper resistive film. The same basic units are periodically arranged in a matrix to form the impedance network. The -1dB transmission frequency band is 2-16GHz, with a relative bandwidth of 155.56%. The entire frequency range from 2-18GHz is a low-reflection band below -10dB. Step 2: Use the Navier-Stokes equations to solve the target aircraft wing model in the wing air flow field analysis; the three equations including conservation of mass, conservation of momentum, and conservation of energy are integrated into the integral form of the Navier-Stokes equations: Among them F c represents the convective flux, F v represents the viscous flux, W is the conserved variable, Ω is the airfoil, Q is the heat flux, and S is the surface of the airfoil; Then, introducing the perfect gas thermodynamic relationship and the Stokes ideal gas hypothesis to close the above Navier-Stokes equation, we get: pH=ρE+P γ is the adiabatic coefficient, which is 1.4 for air; μ, v, and ω represent the three components of velocity, ρ represents density, P represents pressure, H represents total enthalpy, and E represents energy; Obtain the pressure of the air flow field of the wing under different working conditions; Then, the water drop impact characteristics are calculated and expressed by the water drop collection coefficient, which refers to the actual impact amount W of water droplets impacting per unit span length per unit time. m With its maximum possible impact W max The ratio of E m Indicates that E m =W n / W, where W m The definition of is: W m =V0(Y ou -Y ol )LWC=V0(y ou -y ol )LUCK·L V0 is the aircraft speed, Y ou 、Y ol is the starting coordinate of the chord length direction of the water droplet trajectory tangent to the upper and lower surfaces of the wing in the far field, y ou 、y ol Y ou 、Y ol The ratio of LWC to the chord length is LWC, which is the liquid water content and L is the characteristic size of the object, i.e., the chord length. W max The definition of is: W max =V0·M·LWC=V0·h·LWC·L M is the projected height of the wing in the chord direction, and h is the ratio of M to the chord length; Substituting the liquid water content, average droplet diameter, and calculated air flow pressure into the droplet impact calculation, the droplet impact coefficient at the wing leading edge is obtained. By calculating different operating conditions, the length L1 of the target wing ice area in the chord direction is determined. Step 3: Conformally distribute the impedance network designed in step 1 on the surface of the target wing's absorbing structure. Then, based on the L1 solved in step 2, draw a perpendicular line in the chord direction to obtain the corresponding wing leading edge area S+, and redesign the basic unit of the impedance network. Increase the resistance value of the impedance film in the basic unit of the S+ area. According to the formula of the resistance value R: Where S1 is the cross-sectional area of ​​the current flowing through the impedance network, ρ′ is the resistivity, and L3 is the length of the resistor film; According to the average power density formula: P d is the average power per unit area, U is the voltage; Step 4: Perform electromagnetic simulation on the impedance network model redesigned in Step 3 to ensure that the target wing still meets the performance requirements of electromagnetic low scattering. If not, return to Step 3 to redesign the pattern shape of the impedance film in the basic unit of the wing leading edge area S+ until the target wing meets the performance requirements of electromagnetic low scattering.

2. The design method for an aircraft deicing / electromagnetic low scattering compatible structure according to claim 1, characterized in that: The resistance value of the impedance film in step 3 is increased by reducing the width of the impedance film.

3. The design method of an aircraft deicing / electromagnetic low scattering compatible structure according to claim 1, characterized in that: The design method for increasing the resistance value of the impedance film in step 3 is to redesign the pattern of the resistor film.

4. The design method of an aircraft deicing / electromagnetic low scattering compatible structure according to claim 1, characterized in that: The resistance value of the impedance film in step 3 is increased by reducing the resistance value of the basic unit in the non-wing leading edge area S+.

5. The design method of an aircraft deicing / electromagnetic low scattering compatible structure according to claim 4, characterized in that: The resistance value of the basic unit in the non-wing leading edge area S+ is reduced by increasing the width of the impedance film.

6. The design method of an aircraft deicing / electromagnetic low scattering compatible structure according to claim 4, characterized in that: The resistance value of the basic unit in the non-wing leading edge area S+ is reduced by redesigning the pattern of the resistance film.

7. The design method of an aircraft deicing / electromagnetic low scattering compatible structure according to claim 1, characterized in that: The substrate in the basic unit is a 0.1 mm thick polyimide film with a thermal conductivity of 2 W·m -1 ·K -1 ; Carbon paste is used as the printing material for the resistor film, and the square resistance of the printed pattern is 660Ω / □.

8. An aircraft deicing / electromagnetic low scattering compatible structure, characterized by: The design method according to any one of claims 1 to 7 is adopted. When in use, the wing leading edge area S+ is adaptively and conformally covered with the basic unit redesigned in step 3, and the non-wing leading edge area S+ of the absorbing structure is also adaptively and conformally covered, and then electrically heated for de-icing.

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