Aircraft icing prediction method considering temperature influence of front parts

Abstract

The invention discloses an aircraft icing prediction method considering the temperature influence of front components, and belongs to the field of aircraft flight safety. The method mainly comprises the following steps: neglecting the influence of temperature change of front parts by adopting a traditional method, and performing preliminary icing calculation on the whole parts of the aircraft; obtaining the surface temperature distribution of the frozen front part; the temperature distribution serves as the temperature boundary of the front part, and the convective heat transfer coefficient of the rear part is calculated; calculating the icing process again to obtain the icing ice shape of the rear part; and circularly calculating until the temperature of the whole part and the icing ice shape are updated. According to the method, the influence of temperature change caused by icing of the front part of the airplane on icing of the rear part is considered, the icing calculation precision of the rear part is improved, and the method better conforms to the real icing process.

Description

Technical Field

[0001] The present invention relates to the technical field of aircraft flight safety design, and particularly relates to an aircraft icing prediction method considering the influence of front components. Background Art

[0002] When an aircraft flies in clouds containing supercooled water droplets, the supercooled water droplets will impact the outer surface of the windward components of the aircraft and icing will occur, thus endangering flight safety. The existing methods for studying aircraft icing mainly include two methods: experiment and numerical simulation. In recent years, with the continuous development of computer technology, the numerical simulation method with low calculation cost and high speed has gradually become an important means for studying aircraft icing.

[0003] The aircraft shape is very complex, and the supercooled water droplets in the cloud will impact the windward surfaces of various different components and freeze. Traditional icing simulation methods usually conduct separate analyses for the target components of aircraft icing, such as the windshield, wing or tail, etc., or consider the overall aircraft shape for synchronous icing calculation. When there is an interaction between the flow processes on the surfaces of the front and rear components of the aircraft, the air flow and the water droplet movement impact process are determined by their interaction. At the same time, the release of the freezing latent heat of the rear component icing is also affected by the surface temperature distribution of the front component, thereby determining the icing rate of the rear component. When the front component is also impacted by supercooled water droplets, its surface temperature changes during the freezing process of the supercooled water droplets, affecting the convective heat transfer process of the air, and thus changing the icing characteristics of the rear component. When considering the icing on the surface of the rear component, the existing methods do not pay attention to the temperature change caused by the icing of the front component. Usually, all the components under study are set to be equal to or higher than the stagnation temperature, or the front component is directly set as an adiabatic surface to obtain the convective heat transfer coefficient of the rear component surface, and this is used to solve the icing thermodynamics model to obtain the icing characteristics. This is somewhat different from the actual icing process and may lead to inaccurate prediction of the icing rate and ice shape of the rear component. Summary of the Invention

[0004] Aiming at the above problems existing in the prior art, in order to overcome the deficiencies of the prior art, the present invention proposes an aircraft icing prediction method considering the influence of the front component temperature, which can more realistically describe the aircraft icing process and accurately predict the ice shape on the surface of the rear component.

[0005] An aircraft icing prediction method considering the influence of the front component temperature includes the following steps:

[0006] (1) Geometrically model the overall components of the aircraft used in the simulation, mesh the computational domain, and import the overall component mesh into the computational fluid dynamics software;

[0007] (2) Combine the flight environment parameters, select a computational fluid dynamics model and set the air flow field boundary conditions of the computational domain to solve the air flow field under the icing meteorological environment of the overall aircraft components, and obtain the air velocity distribution and the convective heat transfer coefficient distribution on the surface of the overall components; based on the air velocity distribution, determine the movement trajectory of the water droplets and the situation of impacting the component surface, and obtain the local water collection coefficient distribution; based on the air flow field, the surface convective heat transfer coefficient distribution, and the local water collection coefficient distribution, calculate the icing process on the surface of the overall components to obtain the preliminary ice shape prediction result and the temperature distribution after icing;

[0008] (3) Divide the overall components into a front component and several rear components, extract the temperature distribution after icing of the front component as the temperature boundary of the front surface, and perform the air flow field calculation again to obtain the convective heat transfer coefficients of several rear components affected by the front icing;

[0009] (4) Based on the convective heat transfer coefficients of several rear components affected by the front icing, calculate the icing process of several rear components again to obtain the ice shape prediction results of several rear components considering the temperature influence of the front component;

[0010] (5) Integrate the ice shape of the front component in the preliminary ice shape prediction result of step (2) and the ice shape prediction results of several rear components considering the temperature influence of the front component in step (4), and output the final ice shape of the overall aircraft components.

[0011] Further, in step (2), the flight and environment parameters include environmental pressure, far-field temperature, far-field incoming flow velocity, and flight angle of attack.

[0012] Further, in step (2), based on the air velocity distribution result, set the supercooled water droplet parameters in the cloud, and use the Lagrangian method to solve the force and motion equation of the water droplets, so as to obtain the movement trajectory of the water droplets and the impact points, and obtain the water droplet impact characteristics of the overall components.

[0013] Further, in step (2), use the Messinger theory to analyze the mass and energy conservation of the control volume on the icing surface of the aircraft, solve the thermodynamic model of aircraft component icing, and obtain the preliminary ice shape prediction result and the temperature distribution after icing.

[0014] Further, the criterion for dividing the front component and the rear components in step (3) is: divide the front component and the rear components based on the peak interval of the local water collection coefficient distribution curve in step (2), and each peak interval corresponds to a front component or a rear component, where the interval with the peak in the front is defined as the front component, and the interval with the peak in the rear is defined as the rear component.

[0015] Further, step (5) is specifically:

[0016] When there is one rear component, the ice shape of the front component in the preliminary ice shape prediction result of integration step (2) is integrated with the ice shape prediction result of the rear component in step (4) after considering the temperature influence of the front component, and the final ice shape of the overall aircraft components is output;

[0017] When there are two or more rear components, return to step (3) for cycling. In each cycle, take the rear component calculated in the previous time as the new front component, execute step (3) to obtain the temperature distribution after the new front component freezes; execute step (4) to obtain the ice shape prediction result of the rear component after considering the freezing temperature of the new front component. Cycle like this until the ice shape prediction of all rear components is completed; integrate the ice shape of the front component in the preliminary ice shape prediction result of step (2) with the ice shape prediction results of several rear components in step (4) after considering the temperature influence of the front component, and output the final ice shape of the overall aircraft components.

[0018] Advantages of the present invention over the prior art:

[0019] Aiming at the deficiencies of traditional icing technologies, when calculating the icing of the rear components of an aircraft, the present invention considers the influence of the temperature change caused by the icing of the front components on the icing rate of the rear components. By calculating and setting the surface temperature during the icing process of the front components, the convective heat transfer coefficient of the rear components affected by the icing of the front components is re-solved, and then the ice shape of the rear components of the aircraft is obtained, making it more in line with the actual aircraft icing process. Description of the Drawings

[0020] Figure 1 is a flowchart of the icing prediction method of the present invention.

[0021] Figure 2 is a schematic diagram of a nose example of an aircraft.

[0022] Figure 3 is a schematic diagram of the icing thermodynamic model.

[0023] Figure 4 is a curve graph of the local water collection coefficient.

[0024] Figure 5 is a comparison graph of the windshield ice shapes calculated by the present invention and traditional technologies. Detailed Embodiments

[0025] In order to more clearly understand the above-mentioned objects, features and advantages of the present invention, the following further elaborates on parts such as the setting of thermal boundary conditions, aircraft icing calculation, icing thermodynamic model, and comparison with the prior art in conjunction with the drawings. It should be noted that, without conflict, the embodiments of the present invention and the features in the embodiments can be combined with each other.

[0026] In the following description, many specific details are set forth in order to provide a thorough understanding of the present invention. However, the present invention may be practiced in other ways than those specifically described herein, and thus, the scope of protection of the present invention is not limited by the specific embodiments disclosed below.

[0027] An aircraft icing prediction method considering the influence of the temperature of the front components is as Figure 1 shown in the flowchart, and includes the following steps:

[0028] (1) Geometric modeling is performed on the overall components of the aircraft to be simulated, the calculation region is selected and meshed, and the overall component mesh is input into the computational fluid dynamics (CFD) software. The overall components of the aircraft simulated in this example are the nose of a certain aircraft, including sub-components such as the radome and windshield, as Figure 2 shown. The CFD software selected is FLUENT software. (2) Combining the flight environment parameters, select the CFD model and set the air flow field boundary conditions of the calculation region to solve the air flow field and the convective heat transfer coefficient on the surface of the overall components of the aircraft in the icing meteorological environment, determine the movement trajectory of the water droplets and the situation of impacting the component surface, calculate the icing process on the surface of the overall components, and obtain the preliminary ice shape prediction result and the temperature distribution after icing. Specifically as follows:

[0029] (2.1) In the FLUENT software, select the turbulent CFD model containing the energy equation, and set the far-field boundary conditions of the air flow field in the calculation region, including the environmental pressure, far-field temperature, far-field incoming flow velocity, and flight angle of attack; the parameters of a certain aircraft nose example are shown in Table 1. The surface temperature boundary of the overall components is set to be higher than the stagnation temperature of the incoming air; here, the surface temperature boundaries of the radome and windshield of the aircraft nose are both set to 300K. On this basis, the spatial grid is discretized by the finite volume method, and the Reynolds-averaged continuity equation, momentum conservation equation, and energy conservation equation of the air flow are solved using the FLUENT software. The Spalart-Allmaras model is selected for the turbulence model to obtain the air velocity distribution, temperature distribution, and surface heat flux density distribution. Through the calculated surface heat flux density, the set surface temperature, and the reference temperature, the surface convective heat transfer coefficient can be calculated:

[0030]

[0031] In the formula, h c is the convective heat transfer coefficient, q is the heat flux density, T w is the wall temperature, T ref is the reference temperature, and here the reference temperature is selected as the far-field temperature T ∞ .

[0032] Table 1 Flight icing condition table

[0033]

[0034] (2.2) Based on the air velocity distribution results, set the supercooled water droplet parameters in the cloud layer, that is, the droplet diameter in Table 1. Use the Lagrangian method to solve the force and motion equation of the water droplet, and then obtain the water droplet motion trajectory and impact point, and acquire the water droplet impact characteristics of the overall component, that is, the impact area and the local water collection coefficient.

[0035] The following simplifications are made for the water droplet motion:

[0036] 1) Do not consider the influence of supercooled water droplets in the air on the air flow field, that is, air-water droplet one-way coupling;

[0037] 2) The water droplets are spherical rigid bodies of uniform size without deformation, that is, do not consider the large water droplet effect;

[0038] 3) The water droplets will not break or coagulate during motion and will not splash after hitting the aircraft surface;

[0039] 4) Parameters such as temperature, viscosity, and density remain unchanged during the water droplet motion;

[0040] 5) Only consider the air resistance for the force acting on the water droplet.

[0041] Conduct a force analysis on a single water droplet to obtain the force and motion equation of the water droplet:

[0042]

[0043] In the formula, t is time; u is the velocity vector of the water droplet; u a is the velocity vector of the air; ρ is the density of the water droplet; μ a is the dynamic viscosity of the air; C D is the drag coefficient of the water droplet; d p is the diameter of the water droplet; Re is the relative Reynolds number, and its calculation formula is as follows:

[0044]

[0045] In the formula, ρ a is the density of the air.

[0046] According to the air velocity distribution solved by CFD, the force and motion equation of the water droplet can be solved to obtain the water droplet motion trajectory and the impact point situation on the aircraft component surface.

[0047] The local water collection coefficient β refers to the ratio of the actual water collection rate W β on the micro-element control volume on the aircraft component surface to the maximum possible water collection rate W βmax on this micro-element, as shown in the following formula, and it is an important parameter used to characterize the water droplet impact characteristics and water collection ability of the micro-element control volume on the surface.

[0048]

[0049] The Lagrangian method is used to obtain the local water collection coefficient: A large number of supercooled water droplets are released at a position far ahead of the aircraft component, and the distance between adjacent water droplets is equal, all being dy; the number of supercooled water droplets hitting the micro - control volume on the aircraft surface is counted and denoted as N. Then, the local water collection coefficient β is calculated by the following formula:

[0050]

[0051] In the formula, Δs is the area of the micro - control volume on the surface, which is the length in two - dimensional cases. According to the obtained distribution of the local water collection coefficient, the mass q of the water hitting the control volume on the aircraft component surface during the icing time can be calculated. m,im :

[0052] q m,im = |u d,∞ |·LWC·β·Δs·Δt

[0053] In the formula, Δt is the icing time, |u d,∞ | represents the magnitude of the water droplet velocity in the far - field, which is usually equal to the far - field air velocity, and LWC is the liquid water content in the air, that is, the liquid water content in Table 1.

[0054] (2.3) Based on the air flow field and the convective heat transfer coefficient distribution obtained in step (2.1) and the mass distribution of the water hitting the surface calculated from the water droplet force and motion in step (2.2), solve the icing thermodynamics model of the aircraft component surface, that is, the mass conservation equation and energy conservation equation of the water film, to obtain the icing area on the component surface, the mass of the water frozen into ice in the surface control volume, and the surface temperature distribution after icing.

[0055] The icing thermodynamics model is established based on Messinger's mass and energy conservation model. The following simplifications are made for the icing process:

[0056] 1) When ice and water co - exist in the surface control volume, it is considered that the temperatures of both are 0 °C and there is no heat transfer between them;

[0057] 2) Ice is regarded as a heat - insulating material, and the heat conduction heat transfer between the ice and the aircraft surface is ignored;

[0058] 3) The temperature of the icing surface of the aircraft component is relatively low, and the radiation heat transfer on the surface is ignored.

[0059] As Figure 3 shown, for any control volume on the icing surface, the mass of the water entering the control volume is equal to the mass of the water leaving the control volume, and the mass conservation equation of the water film is established:

[0060] q m,im + qm,in = q m,va + q m,out + q m,so

[0061] where q m,im is the mass of water in the impacting water droplet, q m,in is the mass of water overflowing from the previous control volume, q m,va is the mass carried away by water evaporation, q m,out is the mass of water overflowing into the next control volume, q m,so is the mass of water that freezes.

[0062] For calculation purposes, the freezing coefficient f is defined as follows:

[0063]

[0064] As Figure 3 shown, the energy conservation in the control volume is:

[0065] E so + H im + H in = Q c + H out + H va

[0066] where E so is the heat released by the freezing of the water droplet, which is divided into three cases according to the different freezing coefficient f:

[0067] If f = 1, E so = q m,so [c ice (T0 - T s ) + L f ;

[0068] If 0 < f < 1, E so = q m,so L f ;

[0069] If f = 0, E so = 0.

[0070] where T0 is the reference temperature of the enthalpy, taken as 273.15 K. c ice is the specific heat of ice. T s is the equilibrium temperature reached during the energy balance of the control volume, that is, the surface temperature after freezing. L f is the latent heat of fusion of ice.

[0071] H im is the enthalpy brought by the impacting water droplet. Since the velocity of the water droplet becomes 0 after impact, the stagnation enthalpy (total enthalpy) is taken:

[0072] In the formula, c w is the specific heat of water, T ∞ and V ∞ are the temperature and velocity of the far-field oncoming flow, respectively.

[0073] H in is the enthalpy brought by the overflow water from the previous control volume (with an equilibrium temperature of T in ), and is calculated by the following formula:

[0074] H in = q m,in c w (T in - T0)

[0075] Q c is the heat taken away by convective heat transfer, and is calculated by the following formula:

[0076] Q c = h c Δs(T s - T ∞ )

[0077] In the formula, h c is the convective heat transfer coefficient, and its calculation has been given in step (2.1).

[0078] H out is the enthalpy taken away by the overflow water flowing out of this control volume, H va is the heat taken away by evaporation (convective mass transfer). The calculation of convective mass transfer is based on the Chilton-Colburn analogy, which relates convective heat transfer and convective mass transfer:

[0079]

[0080] In the formula, k c is the convective mass transfer coefficient, Sc is the Schmidt number, and its physical meaning is the ratio of momentum diffusion to mass diffusion:

[0081]

[0082] Finally, the mass of convective mass transfer is calculated by the following formula:

[0083]

[0084] In the formula, Le is the Lewis number, Le ≡ Sc / Pr, MW is the molecular weight, p v,s,sat is the saturation vapor pressure at the wall equilibrium temperature, p T and T T are the total temperature and total pressure, respectively, p e is the air pressure at the outer boundary of the boundary layer, rh is the relative humidity.

[0085] while H va and H out According to the different freezing coefficients f, it is also divided into three cases:

[0086] If f = 1, that is, all water droplets freeze:

[0087] H out = 0

[0088] H va = q m,va [L s c ice (T0 - T s ) - L f

[0089] If 0 < f < 1, some water droplets freeze:

[0090]

[0091] H va = q m,va L evap

[0092] If f = 0, all water droplets overflow into the subsequent control volume

[0093] H out = q m,out c w (T s - T0)

[0094] H va = q m,va [L evap + c w (T s - T0)]

[0095] In the formula, L s is the latent heat of sublimation, and L evap is the latent heat of evaporation.

[0096] The total number of equations provided by the mass conservation and energy conservation equations of the water film in the surface control volume is 2, while the number of unknowns is 4, which are the freezing mass q m,so , the mass of water flowing into the current control volume q m,in , the mass of water flowing out of the current control volume q m,out and the surface equilibrium temperature T s .

[0097] However, according to the definition of the freezing coefficient, 0 ≤ f ≤ 1, and:

[0098] ​When 0 < f < 1, the control volume is an ice - water mixture, so T s = T0;

[0099] When f = 0, no ice forms in the control volume, so q m,so = 0;

[0100] When f = 1, all the water in the control volume freezes, so q m,out = 0.

[0101] One unknown can be reduced by the above - mentioned constraints. Additionally, at the stagnation point, it is considered that q m,in = 0. The water droplets flowing out of this control volume are evenly divided into two parts and overflow from the upper and lower surfaces respectively. So for each downstream control volume, q m,in can be obtained from the q m,out of the previous control volume. Therefore, the equation is solvable.

[0102] The specific solution method is to start calculating from the control volume at the stagnation point. First, assume that the wall equilibrium temperature T s = T0, solve the mass - conservation and energy - conservation equations to obtain q m,so and q m,out , and then calculate the freezing coefficient f.

[0103] If 0 < f < 1, it means that the assumption of T s = T0 is correct. The q m,so and q m,out obtained here are the final ice - forming mass and overflow mass of this control volume.

[0104] If f < 0, it means that no ice is formed. Let q m,so = 0, re - solve the equations for calculation, and obtain the surface equilibrium temperature T s and the overflow water mass q m,out .

[0105] If f > 1, it means that all the water freezes and no water overflows to the next control volume. So let q m,out = 0, re - solve the equations, and obtain the surface equilibrium temperature T s and the ice - forming mass q m,so .

[0106] After solving the control volume where the stagnation point is located, divide the overflow water mass q m,out into two equal parts. One - half is used as the q m,in of the control volume on the lower surface of the airfoil adjacent to it, and one - half is used as the q m,in of the adjacent control volume on the upper surface, and then calculate downstream respectively using a similar method.

[0107] After obtaining the ice - forming mass q m,soAfter that, the ice thickness ΔH is calculated by the following formula:

[0108]

[0109] In the formula, ρ ice is the density of ice.

[0110] Based on the ice thickness distribution on the surface, the ice shape is calculated, and the surface equilibrium temperature obtained by solving the icing thermodynamic model is the temperature after icing.

[0111] Using this icing model to calculate the ice shape of the overall aircraft components, the results are as shown in Figure 5 the ice shape calculated by the traditional method. For the surface control volume in the icing area, the freezing coefficient f is 0 < f < 1, and the temperature after icing is 273.15K.

[0112] (3) Divide the overall component into a front component and several rear components. Take the surface temperature distribution after icing of the front component obtained in step (2) as the temperature boundary of the front surface. Assume that the front component has been iced at this time, and calculate the air flow field again to obtain the convective heat transfer coefficients of several rear components affected by the front icing;

[0113] (3.1) When dividing the front and rear of the component, the core basis is the distribution characteristics of the water droplet collection coefficient in the direction perpendicular to the oncoming flow. The specific division criteria are as follows:

[0114] When the local water collection coefficient distribution on the surface of a single component shows a typical single-peak curve in the direction perpendicular to the oncoming flow, if there are two components, the overall distribution curve will show a double-peak characteristic, that is, each component corresponds to an independent single-peak interval, and so on for multiple components. At this time, by analyzing the local water collection coefficient distribution curve, identify the ranges corresponding to the two peaks. The valley area between the peaks is the boundary position of the components: the interval with the front peak is defined as the front component, and the interval with the rear peak is defined as the rear component. For multiple components, according to the position relationship, select two adjacent components for the front and rear component division. In the example Figure 4 the local water collection coefficient distribution shows a double-peak characteristic. By analyzing the curve, two peaks are identified. The left peak is in the front, and the corresponding radome part is defined as the front component; the right peak is in the rear, and the corresponding windshield part is defined as the rear component, as shown in Figure 2 shown. The valley area between the two peaks is determined as the boundary position of the components.

[0115] (3.2) Calculate the air flow field again. In the FLUENT software, select the turbulent CFD model with the energy equation, set the far-field boundary conditions of the air flow field in the calculation region, and the example parameters are the same as those in Table 1. The calculated surface temperature of the radome remains at 273.15K after icing, so the temperature boundary of the radome at the aircraft nose is updated to 273.15K, and the windshield surface is set to 300K. The remaining settings are the same as in step (2.1), and the updated surface convective heat transfer coefficient considering the influence of the icing temperature of the front components is calculated.

[0116] (4) Combine the convective heat transfer coefficient of the rear components obtained in step (3) and other relevant parameters, and calculate the icing process of the rear components again to obtain the ice shape prediction result of the rear components considering the influence of the front component temperature;

[0117] In the specific implementation process, since the calculation results of water droplet impingement mainly depend on the oncoming flow conditions and the component shape structure. During the ice shape update process of the rear components, the changes in the oncoming flow conditions and the component shape are not considered, so the water droplet motion and impingement process are not recalculated. Use the original water droplet impingement calculation results in step (2), and change the convective heat transfer coefficient h c in the new icing calculation process to s update the ice shape and the surface equilibrium temperature T of the rear components.

[0118] (5) When the number of aircraft component divisions is two, that is, one front component and one rear component, integrate the ice shape of the front component obtained in step (2) and the ice shape of the rear component obtained in step (4) considering the influence of the front component temperature, and output the final ice shape of the overall aircraft components;

[0119] When the number of aircraft component divisions is three or more, use the current rear component as the new front component, and process the subsequent components in turn according to the methods in steps (3) and (4) above, and cycle through the icing conditions of each component until the icing prediction of all components is completed, and output the final ice shape of the overall aircraft components.

[0120] The ice shape result finally obtained using the method of this invention is as shown Figure 5 in the ice shape graph line calculated by this invention in this invention, and the ice shape calculated by this invention is more in line with the actual aircraft icing process.

Claims

1. An aircraft icing prediction method considering the influence of the temperature of the front component, characterized in that The following steps are involved: (1) Geometric modeling of the aircraft components used in the simulation, meshing of the calculation area, and importing the mesh of the components into the computational fluid dynamics software; (2) Combined with the flight environment parameters, the computational fluid dynamics model is selected and the air flow field boundary conditions of the calculation area are set to solve the air flow field of the aircraft's entire component in the icing meteorological environment, and the air velocity distribution and the distribution of the convective heat transfer coefficient on the surface of the entire component are obtained; based on the air velocity distribution, the movement trajectory of the water droplets and the impact on the component surface are determined to obtain the local water collection coefficient distribution; based on the air flow field and the surface convective heat transfer coefficient distribution and the local water collection coefficient distribution, the icing process on the surface of the entire component is calculated to obtain preliminary ice shape prediction results and the temperature distribution after icing; (3) The whole component is divided into a front component and several rear components, and the temperature distribution of the front component after icing is extracted as the temperature boundary of the front surface. The air flow field is calculated again to obtain the convective heat transfer coefficients of several rear components after being affected by the front icing; (4) Based on the convective heat transfer coefficients of the rear components affected by the front icing, the icing process of the rear components is calculated again to obtain the ice shape prediction results of the rear components after considering the temperature influence of the front components; (5) Integrate the ice shape of the front components in the preliminary ice shape prediction results of step (2) and the ice shape prediction results of several rear components after considering the temperature influence of the front components in step (4), and output the final ice shape of the entire aircraft components.

2. The aircraft icing prediction method considering the influence of the temperature of the front component according to claim 1, wherein: The criterion for dividing the front components and the rear components in step (3) is: dividing the front components and the rear components based on the peak interval of the local water collection coefficient distribution curve in step (2), each peak interval corresponds to a front component or a rear component, wherein the interval near the front peak is defined as the front component, and the interval near the back peak is defined as the rear component.

3. The aircraft icing prediction method considering the influence of the temperature of the front component according to claim 1, characterized in that: The step (5) is specifically as follows: When there is only one rear component, the ice shape of the front component in the preliminary ice shape prediction result of step (2) is integrated with the ice shape prediction result of the rear component after considering the influence of the temperature of the front component in step (4), and the final ice shape of the entire aircraft component is output; When there are two or more rear components, the process returns to step (3) and loops. In each loop, the rear component calculated in the previous loop is used as the new front component, and step (3) is executed to obtain the temperature distribution of the new front component after freezing. Step (4) is executed to obtain the ice shape prediction result of the rear component after considering the new front component freezing temperature. This loop is repeated until the ice shape prediction of all rear components is completed. The ice shape of the front component in the preliminary ice shape prediction result of step (2) is integrated with the ice shape prediction results of several rear components in step (4) after considering the influence of the temperature of the front component, and the final ice shape of the entire aircraft component is output.

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