A prediction method, device, equipment and medium for three-dimensional distribution of aerodynamic heat environment of an aircraft based on a composite model

Abstract

The present application discloses a prediction method, device, equipment and medium for three-dimensional distribution of the aerodynamic heat environment of an aircraft based on a composite model, which relates to the field of computer technology. The method includes: a processing device obtains the geometric features of the aircraft, then determines N components, constructs a three-dimensional composite model of the aircraft using the N components, divides the surface mesh of the model, then obtains the oncoming flow conditions, obtains the geometric relationship between the Mth component and the (M-1)th component, calculates the geometric parameters and flow parameters of the mth grid point of the Mth component, then calculates the aerodynamic heat environment parameters of the mth grid point of the Mth component, and then determines whether the mth grid point is the last grid point of the Mth component. If so, it determines whether the Mth component is the last component among the N components. If so, it outputs the aerodynamic heat environment parameters of each grid point of the N components. This method can obtain the distribution of the aerodynamic heat environment on each surface of the aircraft.

Description

Technical Field

[0001] The present application relates to the field of computer technology, and particularly to a method, device, equipment and medium for predicting the three-dimensional distribution of the aerodynamic heat environment of an aircraft based on a composite model. Background Art

[0002] Aircraft include missiles, spaceships, space shuttles, and aerospace planes. When an aircraft is flying at high speed, the air around the aircraft will be strongly compressed and generate intense friction. Most of the kinetic energy is converted into heat energy, causing the air temperature to rise sharply and transfer heat to the surface of the aircraft. This heat transfer method is called aerodynamic heat. Aerodynamic heat protection is one of the key problems in the success or failure of the entire aircraft. Therefore, research on aerodynamic heat calculation is very necessary.

[0003] The traditional calculation of aerodynamic heat is based on the position and geometric characteristics of the target point on the surface of the aircraft. The aircraft is simplified to a single-point heat flux calculation on a certain simple component. The single-point heat flux calculation obtains single-point data, which is difficult to reflect the distribution of the aerodynamic heat environment on each surface of the aircraft. Summary of the Invention

[0004] The present application provides a method, device, equipment and medium for predicting the three-dimensional distribution of the aerodynamic heat environment of an aircraft based on a composite model, which can obtain the distribution of the aerodynamic heat environment on each surface of the aircraft.

[0005] To achieve the above object, the present application adopts the following technical solutions:

[0006] In a first aspect, the present application provides a method for predicting the three-dimensional distribution of the aerodynamic heat environment of an aircraft based on a composite model, the method comprising:

[0007] Obtain the geometric characteristics of the aircraft;

[0008] According to the geometric characteristics of the aircraft, determine N components, and use the N components to construct a three-dimensional composite model of the aircraft, where N is an integer greater than 1;

[0009] Divide the surface mesh of the three-dimensional composite model;

[0010] Obtain the oncoming flow conditions, obtain the geometric relationship between the Mth component and the (M - 1)th component, and calculate the geometric parameters of the mth grid point of the Mth component. The geometric relationship includes the surface inclination angle, and the geometric parameters include the equivalent angle and the streamline length, where 1 < M ≤ N and M is an integer;

[0011] According to the oncoming flow conditions, the geometric relationship between the Mth component and the (M - 1)th component, and the geometric parameters of the mth grid point of the Mth component, calculate the flow parameters of the mth grid point of the Mth component;

[0012] Calculate the aerothermal environment parameters of the m-th grid point of the M-th component according to the flow parameters of the m-th grid point of the M-th component;

[0013] Determine whether the m-th grid point is the last grid point of the M-th component;

[0014] If the m-th grid point is the last grid point of the M-th component, determine whether the M-th component is the last component among the N components;

[0015] If the M-th component is the last component among the N components, output the aerothermal environment parameters of each grid point of the N components.

[0016] In some possible implementation manners, the method further includes:

[0017] If the m-th grid point is not the last grid point of the M-th component, calculate the aerothermal environment parameters of the (m + 1)-th grid point of the M-th component.

[0018] In some possible implementation manners, the method further includes:

[0019] If the M-th component is not the last component among the N components, calculate the aerothermal environment parameters of the first grid point of the (M + 1)-th component.

[0020] In some possible implementation manners, the flow parameters include: shock wave angle, flow parameters behind the wave, surface pressure coefficient, and boundary outer edge parameters.

[0021] In some possible implementation manners, the calculating the aerothermal environment parameters of the m-th grid point of the M-th component according to the flow parameters of the m-th grid point of the M-th component includes:

[0022] Obtain the component type of the M-th component;

[0023] According to the calculation strategy corresponding to the component type of the M-th component, use the flow parameters of the m-th grid point of the M-th component to calculate the aerothermal environment parameters of the m-th grid point of the M-th component.

[0024] In some possible implementation manners, the determining the N components according to the geometric characteristics of the aircraft includes:

[0025] Determine N sub-characteristics in the geometric characteristics of the aircraft;

[0026] According to the sub-characteristics and the mapping relationship between the preset reference characteristics and reference components, determine the components corresponding to each of the N sub-characteristics.

[0027] In some possible implementations, the aircraft includes a symmetric aircraft and a body of revolution aircraft.

[0028] In a second aspect, the present application provides a prediction device for the three-dimensional distribution of the aerodynamic heat environment of an aircraft based on a composite model. The device includes:

[0029] An acquisition module, configured to acquire the geometric features of the aircraft; determine N components according to the geometric features of the aircraft, and use the N components to construct a three-dimensional composite model of the aircraft, where N is an integer greater than 1; divide the surface mesh of the three-dimensional composite model; acquire the oncoming flow conditions, acquire the geometric relationship between the Mth component and the (M - 1)th component, and calculate the geometric parameters of the mth grid point of the Mth component, where the geometric relationship includes the surface inclination angle, and the geometric parameters include the equivalent angle and the streamline length, 1 < M ≤ N, and M is an integer;

[0030] A calculation module, configured to calculate the flow parameters of the mth grid point of the Mth component according to the oncoming flow conditions, the geometric relationship between the Mth component and the (M - 1)th component, and the geometric parameters of the mth grid point of the Mth component; calculate the aerodynamic heat environment parameters of the mth grid point of the Mth component according to the flow parameters of the mth grid point of the Mth component;

[0031] A judgment module, configured to judge whether the mth grid point is the last grid point of the Mth component; if the mth grid point is the last grid point of the Mth component, judge whether the Mth component is the last component among the N components; if the Mth component is the last component among the N components, output the aerodynamic heat environment parameters of each grid point of the N components.

[0032] In a third aspect, the present application provides a computing device, including a memory and a processor;

[0033] Wherein, one or more computer programs are stored in the memory, and the one or more computer programs include instructions; when the instructions are executed by the processor, the computing device is enabled to execute the method according to any one of the first aspect.

[0034] In a fourth aspect, the present application provides a computer-readable storage medium, which is used to store a computer program, and the computer program is used to execute the method according to any one of the first aspect.

[0035] Fifth aspect, the present application provides a computer program product, the computer program product includes one or more computer instructions, when the computer instructions are executed by a computer, the computer executes the method according to any one of the first aspect.

[0036] As can be seen from the above technical solutions, the present application has at least the following beneficial effects:

[0037] In the present application, the processing device obtains the geometric features of the aircraft, then determines N components according to the geometric features of the aircraft, constructs a three-dimensional combined model of the aircraft using the N components, then divides the surface mesh of the three-dimensional combined model, and then the processing device obtains the oncoming flow conditions, obtains the geometric relationship between the Mth component and the (M - 1)th component, calculates the geometric parameters of the mth grid point of the Mth component, and calculates the flow parameters of the mth grid point of the Mth component according to the oncoming flow conditions, the geometric relationship between the Mth component and the (M - 1)th component, and the geometric parameters of the mth grid point of the Mth component. Then, according to the flow parameters of the mth grid point of the Mth component, the aerothermal environment parameters of the mth grid point of the Mth component are calculated. After calculating the aerothermal environment parameters of the mth grid point of the Mth component, it is judged whether the mth grid point is the last grid point of the Mth component. If the mth grid point is the last grid point of the Mth component, it is judged whether the Mth component is the last component among the N components. If the Mth component is the last component among the N components, the aerothermal environment parameters of each grid point of the N components are output. In the existing solution, the traditional engineering algorithm simplifies the aircraft to the single-point heat flux calculation on a certain simple component according to the position and geometric features of the target point on the aircraft surface, but ignores the mutual influence between components, such as the influence of the previous component on the subsequent component, resulting in difficulty in dealing with complex-shaped aircraft with multi-stage compression and expansion. Moreover, the existing technology only calculates a single point and cannot obtain the three-dimensional distribution of the thermal environment parameters on the aircraft surface. If data of multiple points are to be obtained, manual point selection must be repeated, and then it is simplified to the single-point calculation on a single component, and this calculation process is cumbersome, time-consuming and laborious. It can be seen that the present application considers the mutual influence between components, obtains the flow expansion or compression relationship between components, can handle complex-shaped aircraft with multi-stage compression and expansion, and can obtain the three-dimensional distribution of the pressure and thermal environment parameters on the aircraft surface.

[0038] It should be understood that the description of technical features, technical solutions, beneficial effects or similar language in this application does not imply that all features and advantages can be achieved in any single embodiment. On the contrary, it can be understood that the description of features or beneficial effects means that specific technical features, technical solutions or beneficial effects are included in at least one embodiment. Therefore, the description of technical features, technical solutions or beneficial effects in this specification does not necessarily refer to the same embodiment. Furthermore, the technical features, technical solutions and beneficial effects described in this embodiment can be combined in any appropriate manner. Those skilled in the art will understand that an embodiment can be implemented without one or more specific technical features, technical solutions or beneficial effects of a specific embodiment. In other embodiments, additional technical features and beneficial effects can also be identified in specific embodiments that do not embody all embodiments. Brief Description of the Drawings

[0039] Figure 1 It is a flowchart of a prediction method for three-dimensional distribution of aerodynamic heat environment of an aircraft based on a combined body model provided by an embodiment of this application;

[0040] Figure 2 It is a schematic diagram of the position of reference features of a spin vehicle such as a projectile provided by an embodiment of this application;

[0041] Figure 3 It is a schematic diagram of node numbering of a three-dimensional combined body model provided by an embodiment of this application;

[0042] Figure 4 It is a schematic diagram of surface mesh division of a three-dimensional combined body model provided by an embodiment of this application;

[0043] Figure 5 It is a schematic diagram of judging the compression relationship between adjacent components provided by an embodiment of this application;

[0044] Figure 6 It is a schematic diagram of judging the expansion relationship between adjacent components provided by an embodiment of this application;

[0045] Figure 7 It is a schematic diagram of the geometric dimensions of a blunt double-cone test model provided by an embodiment of this application;

[0046] Figure 8 It is a schematic diagram of the three-dimensional distribution result of surface heat flux of a blunt double-cone experimental model provided by an embodiment of this application;

[0047] Figure 9 It is a schematic diagram of a comparison curve of a blunt double-cone test model on different meridians provided by an embodiment of this application;

[0048] Figure 10Schematic diagram of a prediction device for three-dimensional distribution of aerodynamic heat environment of an aircraft based on a composite model provided by an embodiment of the present application;

[0049] Figure 11 Schematic diagram of a computing device provided by an embodiment of the present application. Detailed implementation manners

[0050] Terms such as "first", "second", and "third" in the description and drawings of the present application are used to distinguish different objects, rather than to limit a specific order.

[0051] In the embodiments of the present application, words such as "exemplary" or "for example" are used to represent examples, illustrations, or explanations. Any embodiment or design solution described as "exemplary" or "for example" in the embodiments of the present application should not be construed as being more preferred or having more advantages than other embodiments or design solutions. Rather, the use of words such as "exemplary" or "for example" is intended to present related concepts in a specific manner.

[0052] Currently, traditional engineering algorithms simplify an aircraft to a single-point heat flux calculation on a certain simple component according to the position and geometric features of the target points on the aircraft surface, but ignore the mutual influence between components, such as the influence of the previous component on the subsequent component, resulting in difficulty in dealing with complex-shaped aircraft with multi-stage compression and expansion.

[0053] Moreover, the existing technology only calculates a single point and cannot obtain the three-dimensional distribution of the heat environment parameters on the aircraft surface. If data for multiple points are to be obtained, manual point selection must be repeated and then simplified to a single-point calculation on a single component, which is a cumbersome, time-consuming, and laborious process.

[0054] In view of this, an embodiment of the present application provides a prediction method for the three-dimensional distribution of the aerodynamic heat environment of an aircraft based on a combined body model. In this method, a processing device obtains the geometric features of the aircraft, and then determines N components according to the geometric features of the aircraft. The three-dimensional combined body model of the aircraft is constructed by using the N components. Then, the surface mesh of the three-dimensional combined body model is divided. The processing device then obtains the oncoming flow conditions, obtains the geometric relationship between the Mth component and the (M - 1)th component, calculates the geometric parameters of the mth grid point of the Mth component, and calculates the flow parameters of the mth grid point of the Mth component according to the oncoming flow conditions, the geometric relationship between the Mth component and the (M - 1)th component, and the geometric parameters of the mth grid point of the Mth component. Then, according to the flow parameters of the mth grid point of the Mth component, the aerodynamic heat environment parameters of the mth grid point of the Mth component are calculated. After calculating the aerodynamic heat environment parameters of the mth grid point of the Mth component, it is determined whether the mth grid point is the last grid point of the Mth component. If the mth grid point is the last grid point of the Mth component, it is determined whether the Mth component is the last component among the N components. If the Mth component is the last component among the N components, the aerodynamic heat environment parameters of each grid point of the N components are output. It can be seen that the present application considers the mutual influence between components, obtains the flow expansion or compression relationship between components, can process aircraft with complex shapes with multiple stages of compression and expansion, and can obtain the three-dimensional distribution of the pressure and heat environment parameters on the aircraft surface.

[0055] To make the technical solution of the present application clearer and easier to understand, the following will introduce a prediction method for the three-dimensional distribution of the aerodynamic heat environment of an aircraft based on a combined body model provided by an embodiment of the present application with reference to the accompanying drawings. As Figure 1 shown, this figure is a flowchart of a prediction method for the three-dimensional distribution of the aerodynamic heat environment of an aircraft based on a combined body model provided by an embodiment of the present application.

[0056] This method is applied to a processing device, and the prediction method for the three-dimensional distribution of the aerodynamic heat environment of the aircraft with the combined body model includes:

[0057] S101. The processing device obtains the geometric features of the aircraft.

[0058] The aircraft includes a face-symmetric aircraft and a body-of-revolution aircraft.

[0059] For cruise-type face-symmetric aircraft, the face-symmetric aircraft is simplified into a two-dimensional combination of a cylinder, a flat plate, and a wedge; for missile and rocket body-of-revolution aircraft, the body-of-revolution aircraft is simplified into a combination of a stagnation point, a spherical surface, a cone, and a cylinder.

[0060] The first component of the axisymmetric aircraft is a two-dimensional cylinder (semi-infinite cylinder), and the subsequent components are flat plates or wedges; the first component of the axisymmetric body aircraft is a spherical surface, and the subsequent components are conical surfaces or cylindrical surfaces.

[0061] S102. The processing device determines N components according to the geometric characteristics of the aircraft, and constructs a three-dimensional combined body model of the aircraft by using the N components.

[0062] According to the geometric characteristics of the aircraft, N sub-characteristics in the geometric characteristics of the aircraft are determined; then, according to the sub-characteristics and the mapping relationship between the reference characteristics and the reference components preset in advance, the component corresponding to each sub-characteristic among the N sub-characteristics is determined, where N is an integer greater than 1.

[0063] In some embodiments, taking the axisymmetric body aircraft such as projectiles as an example, a typical combination of a spherical head, a cone, and a cylinder is shown in Table 1.

[0064] Table 1:

[0065]

[0066] Among them, the positional relationship of the reference characteristics in the axisymmetric body aircraft such as projectiles is as Figure 2 shown. This figure is a schematic diagram of the position of the reference characteristics of an axisymmetric body aircraft provided by an embodiment of the present application.

[0067] According to the components corresponding to each sub-characteristic among the N sub-characteristics determined in the above steps, the combined body of the aircraft is obtained by using the components corresponding to each sub-characteristic among the N sub-characteristics, and the combined body is combined to obtain a three-dimensional combined body model of the aircraft.

[0068] S103. The processing device divides the surface mesh of the three-dimensional combined body model.

[0069] After obtaining the three-dimensional combined body model of the aircraft, the surface mesh of the three-dimensional combined body model of the aircraft is divided, and the node numbers (i, j) are used to define the grid points. Among them, i increases along the axial direction, and j increases clockwise along the circumferential direction (front view); i = 0 is located at the front end face of each component, and j = 0 is located on the back side in the longitudinal symmetry plane. Different i values correspond to different axial stations, and different j values correspond to different circumferential meridians, as Figure 3 shown. This figure is a schematic diagram of the node numbers of a three-dimensional combined body model provided by an embodiment of the present application. A typical combined body model of a spherical head, a cone, and a cylinder and its surface mesh are as Figure 4 shown. This figure is a schematic diagram of the surface mesh division of a three-dimensional combined body model provided by an embodiment of the present application.

[0070] S104. The processing device obtains the oncoming flow conditions, obtains the geometric relationship between the Mth component and the (M - 1)th component, and calculates the geometric parameters of the mth grid point of the Mth component.

[0071] The oncoming flow conditions are divided into ballistic conditions and wind tunnel conditions. The ballistic oncoming flow parameters include altitude, velocity, etc., which are calculated based on the 1976 US Standard Atmosphere Model. The wind tunnel oncoming flow parameters include total temperature, total pressure, and total density, etc., which are converted based on the flow stagnation parameter relationship. The user inputs the parameters in the state where the aircraft is located into the 1976 US Standard Atmosphere Model and the flow stagnation parameter relationship for calculation to obtain the oncoming flow conditions of the aircraft.

[0072] During the flight of the aircraft, to calculate the thermal environment of the aircraft at all times along the entire ballistic trajectory, in this application, based on the oncoming flow conditions at time t i subsequent calculations are carried out.

[0073] Obtain the geometric relationship between the Mth component and the (M - 1)th component. The geometric relationship is the surface inclination angle. Therefore, compare the surface inclination angle of the Mth component with the surface inclination angle of the (M - 1)th component to obtain the flow state of the Mth component.

[0074] If the surface inclination angle of the Mth component is greater than the surface inclination angle of the (M - 1)th component , then the flow of the Mth component relative to the (M - 1)th component is in a compressed state, and the relative inclination angle between the components is the compression angle , as Figure 5 shown. This figure is a schematic diagram for judging the compression relationship between adjacent components provided by an embodiment of this application. Among them, in the figure represents the oncoming flow velocity, that is, the velocity at infinity.

[0075] If the surface inclination angle of the Mth component is less than the surface inclination angle of the (M - 1)th component , then the flow of the Mth component relative to the (M - 1)th component is in an expansion state, and the relative inclination angle between the components is the expansion angle , as Figure 6 shown. This figure is a schematic diagram for judging the expansion relationship between adjacent components provided by an embodiment of this application.

[0076] Then calculate the geometric parameters of the mth grid point of the Mth component. Among them, the geometric parameters include the equivalent angle and the streamline length, where 1 < M ≤ N and M is an integer. The calculation methods for the equivalent angle and the streamline length are introduced in detail below.

[0077] Obtain the original half-cone angle, oncoming flow angle of attack, and meridian angle of the m-th grid point of the M-th component. The oncoming flow angle of attack is input by the user in the oncoming flow conditions. The original half-cone angle is automatically converted by the program after the user inputs the geometric parameters. The meridian angle is an intermediate variable automatically calculated by the program. Based on the original half-cone angle, oncoming flow angle of attack, and meridian angle of the m-th grid point of the M-th component, calculate the equivalent angle of the m-th grid point of the M-th component, that is, the equivalent half-cone angle, as shown in formula (1):

[0078] (1)

[0079] Wherein, is the equivalent half-cone angle, is the original half-cone angle, is the oncoming flow angle of attack, is the meridian angle.

[0080] For the windward side and the leeward side, the above formula is simplified to formula (2):

[0081] (2)

[0082] That is, the equivalent half-cone angle of the windward side is the original half-cone angle plus the angle of attack, and the equivalent half-cone angle of the leeward side is the original half-cone angle minus the angle of attack.

[0083] The streamline length of the m-th grid point of the M-th component is the sum of the distances from the first grid point of the first component to the m-th grid point of the M-th component.

[0084] In some embodiments, as Figure 2 shown, the streamline length of the m-th grid point of the component is .

[0085] S105. The processing device calculates the flow parameters of the m-th grid point of the M-th component according to the oncoming flow conditions, the geometric relationship between the M-th component and the (M - 1)-th component, and the geometric parameters of the m-th grid point of the M-th component.

[0086] The flow parameters include: shock wave angle, flow parameters behind the wave, surface pressure coefficient, and boundary outer edge parameters. The calculation methods for the flow parameters are introduced in detail.

[0087] Calculate the shock wave angle. For the stagnation point or the spherical surface, the shock wave type in the flow field is a normal shock wave, and the shock wave angle is 90 degrees; for the swept cylinder, the shock wave type in the flow field is a parallel shock wave, and the shock wave angle is the complementary angle of the sweep angle, that is, 90 degrees minus the sweep angle; for the wedge, solve the shock wave angle through the oblique shock wave relation; for the cone, calculate the shock wave angle by iteratively solving the conical flow equation.

[0088] Calculate the flow parameters behind the wave. Using the freestream parameters, shock angle, equivalent angles of each component, and compression / expansion parameters as input conditions, calculate the flow parameters behind the wave based on the relations of normal shock, oblique shock, and Prandtl-Meyer expansion flow. For a normal shock, when using a calorically perfect gas for calculation, solve the normal shock equations to obtain a series of simplified relations that are functions of the Mach number ahead of the wave; for an oblique shock, when using a calorically perfect gas for calculation, take the velocity component of the freestream perpendicular to the shock ahead of the wave as the basic quantity to obtain simplified relations that are functions of the normal Mach number ahead of the wave; for Prandtl-Meyer expansion flow, when using a calorically perfect gas for calculation, iteratively solve the P-M relations. In some embodiments using a high-temperature gas for calculation, for normal shocks and oblique shocks, simple formulas cannot be obtained, and the system of equations must be numerically solved.

[0089] Calculate the surface pressure coefficient, and calculate the surface pressure coefficient of the aircraft based on the modified Newton theory, tangent wedge method, and tangent cone method.

[0090] The modified Newton theory uses the Lees formula, as shown in Equation (3):

[0091] (3)

[0092] where is the surface pressure coefficient, is the angle between the freestream vector and the surface normal vector, is the pressure coefficient at the stagnation point behind the normal shock, and the calculation formula is as shown in Equation (4):

[0093] (4)

[0094] where is the stagnation pressure behind the shock, is the freestream pressure, is the freestream density, is the freestream velocity, is the freestream Mach number, is the specific heat ratio of the gas.

[0095] Tangent wedge method and tangent cone method: At any point i on the surface of the object, define its local deflection angle as the angle between the local surface tangent and the freestream vector. Extend the tangent at this point to obtain an equivalent sharp wedge or equivalent sharp cone with a half-wedge angle or half-cone angle of . The tangent wedge method or tangent cone method assumes that the pressure at point i is equal to the pressure on the corresponding equivalent wedge or equivalent cone.

[0096] Calculate the outer edge parameters of the boundary layer. The first-order boundary layer approximation allows the pressure at the outer edge of the boundary layer to be equal to the corresponding wall pressure. In the hypersonic case, the boundary layer is very thin, and it can be assumed that the entropy at the outer edge of the boundary layer is equal to the entropy after the normal shock wave. Other parameters can be calculated from the pressure and the flow parameters after the shock wave.

[0097] S106. The processing device calculates the aerothermal environment parameters of the m-th grid point of the M-th component according to the flow parameters of the m-th grid point of the M-th component.

[0098] Calculate the aerothermal environment parameters of the m-th grid point of the M-th component according to the flow parameters of the m-th grid point of the M-th component, including:

[0099] Obtain the component type of the M-th component. The component types include stagnation point, spherical surface, flat plate, wedge, and cone.

[0100] For the stagnation point, the aerothermal environment parameters of the spherical head stagnation point can be calculated by the Fay Riddell method. The calculation formula is shown in formula (5):

[0101] (5)

[0102] Where, is the stagnation point heat flux, is the Nusselt number, is the Reynolds number, is the wall density, is the viscosity coefficient, is the enthalpy, is the Prandtl number, is the velocity gradient at the outer edge of the boundary layer, is the enthalpy at the outer edge of the boundary layer.

[0103] In formula (5), can be calculated by formula (6), can be calculated by formula (7). Formulas (6) and (7) are shown as follows:

[0104] (6)

[0105] (7)

[0106] Where, is the Nusselt number, is the Reynolds number, is the Prandtl number, is the viscosity coefficient, is the density at the outer edge of the boundary layer, is the pressure, is the viscosity coefficient, is the Lewis number, is the dissociation enthalpy as the radius.

[0107] For a spherical surface, which has laminar and turbulent layers, for the laminar layer, the Lees method is used to calculate the aerothermal environment parameters at the m-th grid point of the M-th component. For the turbulent layer, the Detra method is used to calculate the aerothermal environment parameters at the m-th grid point of the M-th component.

[0108] The calculation formula of the Lees method is shown in formulas (8) and (9):

[0109] (8)

[0110] (9)

[0111] Wherein, is the laminar heat transfer coefficient, is the stagnation heat transfer coefficient, is the central angle of the sphere, is the intermediate variable expression.

[0112] The calculation formula of the Detra method is shown in formula (10):

[0113] (10)

[0114] Wherein, is the turbulent heat transfer coefficient, is the streamline length at the calculation point position, is the local tangent angle of the spherical surface, is the Prandtl number, is the dissociation enthalpy at the outer edge of the boundary layer, is the total enthalpy at the outer edge of the boundary layer, is the viscosity coefficient at the outer edge of the boundary layer, is the density at the outer edge of the boundary layer, is the velocity at the outer edge of the boundary layer.

[0115] According to the calculation strategy corresponding to the component type of the M-th component, using the flow parameters at the m-th grid point of the M-th component, calculate the aerothermal environment parameters at the m-th grid point of the M-th component.

[0116] For flat plates, wedges, and cones, the Eckert reference enthalpy method can be used to calculate the aerothermal environment parameters at the m-th grid point of the M-th component.

[0117] The calculation formula for laminar flow is shown in formula (11):

[0118] (11)

[0119] The calculation formula for turbulent flow is shown in formula (12):

[0120] (12)

[0121] Wherein, is the heat transfer coefficient, takes 32.174, is the Prandtl number under reference conditions, is the reference density, is the Mangler transformation factor, is the Reynolds number calculated under reference conditions.

[0122] In some embodiments, the flat plate, wedge, and cone all adopt the Eckert reference enthalpy method, but different Mangler factors are used to equivalently process the streamline lengths. For the cone, the Mangler factor is 3.0 in the laminar state and 2.0 in the turbulent state; for the flat plate and wedge, the Mangler factor is 1.0 in both the laminar state and the turbulent state.

[0123] S107. The processing device determines whether the m-th grid point is the last grid point of the M-th component.

[0124] The aerothermal environment parameters of the m-th grid point of the M-th component are obtained through the above steps. Next, it is determined whether the m-th grid point is the last grid point of the M-th component.

[0125] If the m-th grid point is the last grid point of the M-th component, then execute S108;

[0126] If the m-th grid point is not the last grid point of the M-th component, then execute S110.

[0127] S108. The processing device determines whether the M-th component is the last component among the N components.

[0128] If the M-th component is the last component among the N components, then execute S109; if the M-th component is not the last component among the N components, then execute S111.

[0129] S109. The processing device outputs the aerothermal environment parameters of each grid point of the N components.

[0130] If the M-th component is the last component among the N components, then the aerothermal environment parameters at time t i are calculated.

[0131] Next, the aerothermal environment parameters of each grid point of each component at the next time are calculated, that is, calculate at time t i+1The aerothermal environment parameters of each grid point of each component at each moment are calculated as shown in S104 to S106, which will not be elaborated here. Among them, the data of each grid point includes the data of all ballistic or test points. Specify the node numbers (i,j), then only the data of this grid point will be output. Specify the meridian number (j), then the data of all grid points on this meridian will be output.

[0132] S110. The processing device processes the next grid point of the Mth component.

[0133] The processing device calculates the aerothermal environment parameters of the (m + 1)th grid point of the Mth component, returns to S104 with m = m + 1, and the calculation method is as shown in S104 to S106, which will not be elaborated here.

[0134] S111. The processing device processes the next component.

[0135] The processing device calculates the aerothermal environment parameters of the first grid point of the (M + 1)th component, returns to S104 with M = M + 1 and m = 1, and the calculation method is as shown in S104 to S106, which will not be elaborated here.

[0136] Based on the above, the processing device obtains the geometric characteristics of the aircraft, then determines N components according to the geometric characteristics of the aircraft, constructs a three-dimensional combined model of the aircraft using the N components, then divides the surface grid of the three-dimensional combined model. The processing device then obtains the oncoming flow conditions, obtains the geometric relationship between the Mth component and the (M - 1)th component, calculates the geometric parameters of the mth grid point of the Mth component, calculates the flow parameters of the mth grid point of the Mth component according to the oncoming flow conditions, the geometric relationship between the Mth component and the (M - 1)th component, and the geometric parameters of the mth grid point of the Mth component, and then calculates the aerothermal environment parameters of the mth grid point of the Mth component according to the flow parameters of the mth grid point of the Mth component. After calculating the aerothermal environment parameters of the mth grid point of the Mth component, it is judged whether the mth grid point is the last grid point of the Mth component. If the mth grid point is the last grid point of the Mth component, it is judged whether the Mth component is the last component among the N components. If the Mth component is the last component among the N components, the aerothermal environment parameters of each grid point of the N components are output. It can be seen that this application considers the mutual influence between components, obtains the flow expansion or compression relationship between components, can process complex-shaped aircraft with multi-stage compression and expansion, and can obtain the three-dimensional distribution of pressure and thermal environment parameters on the aircraft surface.

[0137] In some embodiments, taking the blunt double-cone test model as an example, the above method is verified.

[0138] The blunt double-cone test model is a sphere-cone-cone combination, and the specific dimension parameters are shown in Table 2; the geometric dimension schematic diagram of the blunt double-cone test model is as Figure 7 shown. This figure is a schematic diagram of the geometric dimensions of a blunt double-cone test model provided by an embodiment of the present application.

[0139] Table 2 - Dimension parameters of the blunt double-cone combination

[0140]

[0141] The oncoming flow parameters are the wind tunnel conditions: , , , , .

[0142] The calculation results of the method of the present invention are as Figure 8 , 9 shown. Figure 8 This is a schematic diagram of the three-dimensional distribution result of the surface heat flux of the blunt double-cone experimental model provided by an embodiment of the present application, Figure 9 This is a schematic diagram of the comparison curves of the blunt double-cone test model on different meridians provided by an embodiment of the present application. Among them, exp is the wind tunnel test result, CFD is the numerical simulation result, and ZQheat is the calculation result of the present invention.

[0143] According to the above calculation results, it shows that a prediction method for the three-dimensional distribution of the aerodynamic heat environment of an aircraft based on a combined body model proposed in the present application has a certain calculation accuracy, can handle aircraft with complex shapes with multiple-stage compression and expansion, and can obtain the three-dimensional distribution of the pressure and heat environment parameters on the surface of the aircraft.

[0144] The above has combined Figures 1 to 9 to introduce in detail a prediction method for the three-dimensional distribution of the aerodynamic heat environment of an aircraft based on a combined body model provided by an embodiment of the present application. Next, the devices and equipment provided by an embodiment of the present application will be introduced with reference to the accompanying drawings.

[0145] An embodiment of the present application also provides a prediction device for the three-dimensional distribution of the aerodynamic heat environment of an aircraft based on a combined body model, as Figure 10 shown. This figure is a schematic diagram of a prediction device for the three-dimensional distribution of the aerodynamic heat environment of an aircraft based on a combined body model provided by an embodiment of the present application. The device includes: an acquisition module 1001, a calculation module 1002, and a judgment module 1003;

[0146] An acquisition module 1001 is configured to acquire the geometric features of an aircraft; determine N components according to the geometric features of the aircraft, and construct a three-dimensional combined model of the aircraft by using the N components, where N is an integer greater than 1; divide the surface mesh of the three-dimensional combined model; acquire the oncoming flow conditions, acquire the geometric relationship between the Mth component and the (M - 1)th component, and calculate the geometric parameters of the mth grid point of the Mth component, where the geometric relationship includes the surface inclination angle, and the geometric parameters include the equivalent angle and the streamline length, 1 < M ≤ N, and M is an integer;

[0147] A calculation module 1002 is configured to calculate the flow parameters of the mth grid point of the Mth component according to the oncoming flow conditions, the geometric relationship between the Mth component and the (M - 1)th component, and the geometric parameters of the mth grid point of the Mth component; calculate the aerothermal environment parameters of the mth grid point of the Mth component according to the flow parameters of the mth grid point of the Mth component;

[0148] A judgment module 1003 is configured to judge whether the mth grid point is the last grid point of the Mth component; if the mth grid point is the last grid point of the Mth component, judge whether the Mth component is the last component among the N components; if the Mth component is the last component among the N components, output the aerothermal environment parameters of each grid point of the N components.

[0149] In some possible implementation manners, the judgment module 1003 is further configured to calculate the aerothermal environment parameters of the (m + 1)th grid point of the Mth component if the mth grid point is not the last grid point of the Mth component.

[0150] In some possible implementation manners, the judgment module 1003 is further configured to calculate the aerothermal environment parameters of the first grid point of the (M + 1)th component if the Mth component is not the last component among the N components.

[0151] In some possible implementation manners, the flow parameters include: shock wave angle, post-shock flow parameters, surface pressure coefficient, and boundary outer edge parameters.

[0152] In some possible implementation manners, the calculation module 1002 is further configured to acquire the component type of the Mth component; calculate the aerothermal environment parameters of the mth grid point of the Mth component by using the flow parameters of the mth grid point of the Mth component according to the calculation strategy corresponding to the component type of the Mth component.

[0153] In some possible implementations, the obtaining module 1001 is further configured to determine N sub-features of the geometric features of the aircraft; and determine a component corresponding to each of the N sub-features according to the sub-features and a mapping relationship between a reference feature and a reference component preset in advance.

[0154] In some possible implementations, the aircraft includes a face-symmetric aircraft and a body-of-revolution aircraft.

[0155] The embodiments of the present application further provide a computing device. As Figure 11 shown, this figure is a schematic diagram of a computing device provided by an embodiment of the present application. The computing device 1100 includes a bus 1101, a processor 1102, a communication interface 1103, and a memory 1104. The processor 1102, the memory 1104, and the communication interface 1103 communicate with each other through the bus 1101.

[0156] The bus 1101 may be a peripheral component interconnect (PCI) bus, an extended industry standard architecture (EISA) bus, or the like. The bus may be divided into an address bus, a data bus, a control bus, etc. For the sake of convenience of representation, Figure 11 only a thick line is used to represent it in the figure, but it does not mean that there is only one bus or one type of bus.

[0157] The processor 1102 may be any one or more of a central processing unit (CPU), a graphics processing unit (GPU), a microprocessor (MP), or a digital signal processor (DSP).

[0158] The communication interface 1103 is used for external communication.

[0159] The memory 1104 may include a volatile memory, such as a random access memory (RAM). The memory 1104 may further include a non-volatile memory, such as a read-only memory (ROM), a flash memory, a hard disk drive (HDD), or a solid state drive (SSD).

[0160] The executable code is stored in the memory 1104, and the processor 1102 executes the executable code to execute the foregoing prediction method for the three-dimensional distribution of the aerodynamic heat environment of the aircraft based on the composite model.

[0161] The embodiment of the present application also provides a computer-readable storage medium. The computer-readable storage medium may be any available medium that can be stored by a computing device or a data storage device such as a data center including one or more available media. The available medium may be a magnetic medium (such as a floppy disk, a hard disk, a magnetic tape), an optical medium (such as a DVD), or a semiconductor medium (such as a solid-state drive), etc. The computer-readable storage medium includes instructions that instruct the computing device to execute the foregoing prediction method for the three-dimensional distribution of the aerodynamic heat environment of the aircraft based on the composite model.

[0162] The embodiment of the present application also provides a computer program product, which includes one or more computer instructions. When the computer instructions are loaded and executed on a computing device, the processes or functions according to the embodiments of the present application are fully or partially generated.

[0163] The computer instructions may be stored in a computer-readable storage medium or transmitted from one computer-readable storage medium to another computer-readable storage medium. For example, the computer instructions may be transmitted from one website, computer, or data center to another website, computer, or data center in a wired manner (such as coaxial cable, optical fiber) or a wireless manner (such as infrared, wireless, microwave, etc.).

[0164] When the computer program product is executed by a computer, the computer executes any one of the foregoing prediction methods for the three-dimensional distribution of the aerodynamic heat environment of the aircraft based on the composite model. The computer program product may be a software installation package. In the case where any one of the foregoing prediction methods for the three-dimensional distribution of the aerodynamic heat environment of the aircraft based on the composite model is required, the computer program product may be downloaded and executed on the computer.

[0165] The descriptions of the processes or structures corresponding to the foregoing various drawings each have their own focuses. For the parts not detailed in a certain process or structure, reference may be made to the relevant descriptions of other processes or structures.

[0166] The above is only the specific implementation manner of the present application, but the protection scope of the present application is not limited thereto. Any changes or substitutions within the technical scope disclosed in the present application should be covered by the protection scope of the present application.

Claims

1. A prediction method for the three-dimensional distribution of the aerodynamic heat environment of an aircraft based on a composite model, characterized in that The method includes: Obtaining the geometric features of the aircraft, where the geometric features of the aircraft include the spherical head radius, spherical crown half angle, conical section length, rear radius of the conical section, and cylindrical section length; Determining N components according to the geometric features of the aircraft, and constructing a three-dimensional combined body model of the aircraft using the N components, where N is an integer greater than 1; Dividing the surface mesh of the three-dimensional combined body model; Obtaining the oncoming flow conditions, obtaining the geometric relationship between the Mth component and the (M - 1)th component, and calculating the geometric parameters of the mth grid point of the Mth component, where the geometric relationship includes the surface inclination angle, and the geometric parameters include the equivalent angle and streamline length, 1 < M ≤ N, and M is an integer; Calculating the flow parameters of the mth grid point of the Mth component according to the oncoming flow conditions, the geometric relationship between the Mth component and the (M - 1)th component, and the geometric parameters of the mth grid point of the Mth component; Calculating the aerothermal environment parameters of the mth grid point of the Mth component according to the flow parameters of the mth grid point of the Mth component; Determining whether the mth grid point is the last grid point of the Mth component; If the mth grid point is the last grid point of the Mth component, then determining whether the Mth component is the last component among the N components; If the Mth component is the last component among the N components, outputting the aerothermal environment parameters of each grid point of the N components.

2. The method according to claim 1, characterized in that The method further includes: If the mth grid point is not the last grid point of the Mth component, then calculating the aerothermal environment parameters of the (m + 1)th grid point of the Mth component.

3. The method according to claim 1, characterized in that The method further includes: If the Mth component is not the last component among the N components, calculating the aerothermal environment parameters of the first grid point of the (M + 1)th component.

4. The method according to claim 1, wherein The flow parameters include: shock wave angle, post-shock flow parameters, surface pressure coefficient, and boundary outer edge parameters.

5. The method according to claim 1, wherein The calculating the aerothermal environment parameters of the mth grid point of the Mth component according to the flow parameters of the mth grid point of the Mth component includes: Obtaining the component type of the Mth component; Calculating the aerothermal environment parameters of the mth grid point of the Mth component according to the calculation strategy corresponding to the component type of the Mth component, using the flow parameters of the mth grid point of the Mth component.

6. The method according to claim 1, wherein The determining N components according to the geometric features of the aircraft includes: Determining N sub-features in the geometric features of the aircraft; Determining the component corresponding to each sub-feature among the N sub-features according to the sub-features and the mapping relationship between the reference feature and the reference component set in advance.

7. The method according to claim 1, characterized in that The aircraft includes a face-symmetric aircraft and a body of revolution aircraft.

8. A prediction device for three-dimensional distribution of the aerodynamic heat environment of an aircraft based on a composite model, characterized in that, The device includes: An acquisition module, configured to acquire the geometric features of an aircraft, where the geometric features of the aircraft include the spherical head radius, the spherical crown half angle, the cone section length, the rear radius of the cone section, and the cylinder section length; determine N components according to the geometric features of the aircraft, and construct a three-dimensional combined body model of the aircraft by using the N components, where N is an integer greater than 1; divide the surface mesh of the three-dimensional combined body model; acquire the oncoming flow conditions, acquire the geometric relationship between the Mth component and the (M - 1)th component, and calculate the geometric parameters of the mth grid point of the Mth component, where the geometric relationship includes the surface inclination angle, and the geometric parameters include the equivalent angle and the streamline length, 1 < M ≤ N, and M is an integer; A calculation module, configured to calculate the flow parameters of the mth grid point of the Mth component according to the oncoming flow conditions, the geometric relationship between the Mth component and the (M - 1)th component, and the geometric parameters of the mth grid point of the Mth component; calculate the aerodynamic heat environment parameters of the mth grid point of the Mth component according to the flow parameters of the mth grid point of the Mth component; A judgment module, configured to judge whether the mth grid point is the last grid point of the Mth component; if the mth grid point is the last grid point of the Mth component, judge whether the Mth component is the last component among the N components; if the Mth component is the last component among the N components, output the aerodynamic heat environment parameters of each grid point of the N components.

9. A computing device, characterized in that, Comprising a memory and a processor; Wherein, one or more computer programs are stored in the memory, and the one or more computer programs include instructions; when the instructions are executed by the processor, the computing device is caused to execute the method according to any one of claims 1 to 7.

10. A computer-readable storage medium, characterized in that, The computer-readable storage medium is used to store a computer program, and the computer program is used to execute the method according to any one of claims 1 to 7.

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