Method, device and equipment for predicting three-dimensional distribution of aerodynamic thermal environment of aircraft based on assembly model, and medium

Abstract

The invention discloses an aircraft aerodynamic thermal environment three-dimensional distribution prediction method and device based on an assembly model, equipment and a medium, and relates to the technical field of computers.The method comprises the steps that processing equipment obtains geometrical characteristics of an aircraft, N parts are determined, the N parts are used for constructing a three-dimensional assembly model of the aircraft, and the three-dimensional assembly model is used for predicting the aerodynamic thermal environment three-dimensional distribution of the aircraft; dividing surface grids of the model, then acquiring incoming flow conditions, acquiring a geometrical relationship between an Mth part and an (M-1) th part, calculating geometrical parameters and flow parameters of an mth grid point of the Mth part, and then calculating aerodynamic thermal environment parameters of the mth grid point of the Mth part, and then judging whether the mth grid point is the last grid point of the Mth part, if so, judging whether the Mth part is the last part of the N parts, and if so, outputting the aerodynamic thermal environment parameters of each grid point of the N parts. According to the method, the distribution of the aerodynamic thermal environment on each profile of the aircraft can be obtained.

Description

Technical Field

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

[0002] Aircraft include missiles, spacecraft, space shuttles, and aerospace planes. When an aircraft flies at high speed, the air around the aircraft will be strongly compressed and produce 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 thermal protection is one of the key problems in the success or failure of the entire aircraft, so research on aerodynamic heat calculation is very necessary.

[0003] The traditional calculation of aerodynamic heat is to simplify the aircraft into a single-point heat flow calculation on a simple component based on the position and geometric characteristics of the target point on the aircraft surface. The single-point heat flow calculation obtains single-point data, which is difficult to reflect the distribution of the aerodynamic thermal environment on various surfaces 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 thermal environment of an aircraft based on a combined model, which can obtain the distribution of the aerodynamic thermal environment on various surfaces of the aircraft.

[0005] In order to achieve the above objectives, this application adopts the following technical solutions: In a first aspect, the present application provides a method for predicting the three-dimensional distribution of the aerodynamic thermal environment of an aircraft based on a combined model, the method comprising: Obtaining 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 assembly model of the aircraft, where N is an integer greater than 1; Dividing the surface grid of the three-dimensional assembly model; Obtaining the incoming flow condition, obtaining the geometric relationship between the Mth component and the M-1th component, and calculating the geometric parameters of the mth grid point of the Mth component, wherein the geometric relationship includes the object plane inclination angle, and the geometric parameters include the equivalent angle and the streamline length, 1<M≤N, and M is an integer; Calculate the flow parameter of the mth grid point of the Mth component according to the incoming flow condition, the geometric relationship between the Mth component and the M-1th component, and the geometric parameter of the mth grid point of the Mth component; Calculating the aerodynamic thermal environment parameter of the mth grid point of the Mth component according to the flow parameter of the mth grid point of the Mth component; Determine whether the m-th grid point is the last grid point of the M-th component; If the m-th grid point is the last grid point of the M-th component, determining whether the M-th component is the last component among the N components; If the Mth component is the last component among the N components, the aerodynamic thermal environment parameters of each grid point of the N components are output.

[0006] In some possible implementations, the method further includes: If the mth grid point is not the last grid point of the Mth component, the aerodynamic thermal environment parameters of the m+1th grid point of the Mth component are calculated.

[0007] In some possible implementations, the method further includes: If the Mth component is not the last component among the N components, the aerodynamic thermal environment parameters of the first grid point of the M+1th component are calculated.

[0008] In some possible implementations, the flow parameters include: shock wave angle, post-wave flow parameters, surface pressure coefficient, and boundary outer edge parameters.

[0009] In some possible implementations, calculating the aerodynamic thermal environment parameter of the mth grid point of the Mth component according to the flow parameter of the mth grid point of the Mth component includes: Obtaining the component type of the Mth component; According to the calculation strategy corresponding to the component type of the Mth component, the aerodynamic thermal environment parameters of the mth grid point of the Mth component are calculated using the flow parameters of the mth grid point of the Mth component.

[0010] In some possible implementations, determining N components according to geometric features of the aircraft includes: Determining N sub-features of the geometric features of the aircraft; According to the sub-features and a mapping relationship between a preset reference feature and a reference component, a component corresponding to each sub-feature in the N sub-features is determined.

[0011] In some possible implementations, the aircraft includes a plane-symmetric aircraft and a rotational body aircraft.

[0012] In a second aspect, the present application provides a device for predicting the three-dimensional distribution of the aerodynamic thermal environment of an aircraft based on a combined model, the device comprising: An acquisition module is used to acquire geometric features of an aircraft; determine N components according to the geometric features of the aircraft, and use the N components to construct a three-dimensional assembly model of the aircraft, where N is an integer greater than 1; divide the surface grid of the three-dimensional assembly model; acquire incoming flow conditions, acquire the geometric relationship between the Mth component and the M-1th component, and calculate the geometric parameters of the mth grid point of the Mth component, where the geometric relationship includes the object surface inclination angle, and the geometric parameters include the equivalent angle and the streamline length, where 1<M≤N, and M is an integer; A calculation module, used for calculating the flow parameters of the mth grid point of the Mth component according to the incoming flow condition, the geometric relationship between the Mth component and the M-1th component, and the geometric parameters of the mth grid point of the Mth component; and calculating the aerodynamic thermal 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 is used 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, then 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 thermal environment parameters of each grid point of the N components.

[0013] In a third aspect, the present application provides a computing device, including a memory and a processor; 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 executes the method as described in any one of the first aspects.

[0014] In a fourth aspect, the present application provides a computer-readable storage medium, wherein the computer-readable storage medium is used to store a computer program, and the computer program is used to execute the method as described in any one of the first aspects.

[0015] In a fifth aspect, the present application provides a computer program product, wherein the computer program product comprises one or more computer instructions, and when the computer instructions are executed by a computer, the computer executes the method as described in any one of the first aspects.

[0016] It can be seen from the above technical solution that the present application has at least the following beneficial effects: In the present application, the processing device obtains the geometric features of the aircraft, and then determines N components based on the geometric features of the aircraft, uses the N components to construct a three-dimensional assembly model of the aircraft, and then divides the surface grid of the three-dimensional assembly model. The processing device then obtains the incoming flow condition, obtains the geometric relationship between the Mth component and the M-1th component, calculates the geometric parameters of the mth grid point of the Mth component, and calculates the mth grid point of the Mth component based on the incoming flow condition, the geometric relationship between the Mth component and the M-1th component, and the geometric parameters of the mth grid point of the Mth component. The flow parameters of the m-th grid point of the M-th component are then calculated according to the flow parameters of the m-th grid point of the M-th component. After calculating the aerodynamic thermal environment parameters of the m-th grid point of the M-th component, it is determined whether the m-th grid point is the last grid point of the M-th component. If the m-th grid point is the last grid point of the M-th component, it is determined whether the M-th component is the last component among the N components. If the M-th component is the last component among the N components, the aerodynamic thermal environment parameters of each grid point of the N components are output. In the existing scheme, the traditional engineering algorithm simplifies the aircraft into a single-point heat flow calculation on a simple component based on the position and geometric characteristics 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 next component, which makes it difficult to handle complex-shaped aircraft with multi-level compression and expansion. Moreover, the existing technology only calculates a single point and cannot obtain the three-dimensional distribution of thermal environment parameters on the aircraft surface. If you want to obtain data from multiple points, you must repeatedly select points manually and then simplify it to a single point calculation on a single component. The calculation process is cumbersome, time-consuming and labor-intensive. It can be seen that the present application takes into account the mutual influence between components, obtains the flow expansion or compression relationship between components, can handle complex-shaped aircraft with multi-level compression and expansion, and can obtain the three-dimensional distribution of pressure and thermal environment parameters on the aircraft surface.

[0017] 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 realized in any single embodiment. On the contrary, it is 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 the present embodiment can also be combined in any appropriate manner. Those skilled in the art will understand that the embodiment can be realized 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 a specific embodiment that does not embody all embodiments. BRIEF DESCRIPTION OF THE DRAWINGS

[0018] Figure 1 A flowchart of a method for predicting the three-dimensional distribution of the aerodynamic thermal environment of an aircraft based on a combined body model provided in an embodiment of the present application; Figure 2 A schematic diagram of the positions of reference features of a missile-like rotating body aircraft provided in an embodiment of the present application; Figure 3 A schematic diagram of node numbering of a three-dimensional assembly model provided in an embodiment of the present application; Figure 4 A schematic diagram of surface mesh division of a three-dimensional assembly model provided in an embodiment of the present application; Figure 5 A schematic diagram of determining a compression relationship between adjacent components provided in an embodiment of the present application; Figure 6 A schematic diagram of determining the expansion relationship between adjacent components provided in an embodiment of the present application; Figure 7 A schematic diagram of the geometric dimensions of a blunt double cone test model provided in an embodiment of the present application; Figure 8 A schematic diagram of the three-dimensional distribution results of heat flux on the surface of a blunt double-cone experimental model provided in an embodiment of the present application; Fig. 9 A schematic diagram of comparison curves of the blunt double cone test model provided in the embodiment of the present application on different meridians; Fig.10 A schematic diagram of a device for predicting the three-dimensional distribution of the aerodynamic thermal environment of an aircraft based on a combined body model provided in an embodiment of the present application; Fig.11 A schematic diagram of a computing device provided in an embodiment of the present application. DETAILED DESCRIPTION

[0019] The terms "first", "second", "third", etc. in the specification of this application and the accompanying drawings are used to distinguish different objects rather than to limit a specific order.

[0020] In the embodiments of the present application, words such as "exemplary" or "for example" are used to indicate examples, illustrations or descriptions. Any embodiment or design described as "exemplary" or "for example" in the embodiments of the present application should not be interpreted as being more preferred or more advantageous than other embodiments or designs. Specifically, the use of words such as "exemplary" or "for example" is intended to present related concepts in a specific way.

[0021] At present, traditional engineering algorithms simplify the aircraft into a single-point heat flow calculation on a simple component based on the position and geometric characteristics 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 next component, which makes it difficult to handle complex-shaped aircraft with multiple levels of compression and expansion.

[0022] Moreover, the existing technology only calculates a single point and cannot obtain the three-dimensional distribution of thermal environment parameters on the surface of the aircraft. If data from multiple points are to be obtained, it is necessary to repeatedly select points manually and then simplify the calculation to a single point on a single component. The calculation process is cumbersome, time-consuming and labor-intensive.

[0023] In view of this, an embodiment of the present application provides a prediction method for the three-dimensional distribution of the aerodynamic thermal environment of an aircraft based on an assembly model. In this method, a processing device obtains geometric features of the aircraft, and then determines N components based on the geometric features of the aircraft, and uses the N components to construct a three-dimensional assembly model of the aircraft. Then, the surface grid of the three-dimensional assembly model is divided. The processing device then obtains the incoming flow condition, obtains the geometric relationship between the Mth component and the M-1th component, calculates the geometric parameters of the mth grid point of the Mth component, and calculates the geometric parameters of the mth grid point of the Mth component based on the incoming flow condition, the geometric relationship between the Mth component and the M-1th component, and the mth grid point of the Mth component. Grid point geometric parameters, calculate the flow parameters of the mth grid point of the Mth component, then calculate the aerodynamic thermal environment parameters of the mth grid point of the Mth component based on the flow parameters of the mth grid point of the Mth component, after calculating the aerodynamic thermal environment parameters of the mth grid point of the Mth component, determine 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 determine 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 thermal environment parameters of each grid point of the N components. It can be seen that the present application takes into account the mutual influence between components, obtains the flow expansion or compression relationship between components, can handle complex-shaped aircraft with multi-level compression and expansion, and can obtain the three-dimensional distribution of pressure and thermal environment parameters on the aircraft surface.

[0024] In order to make the technical solution of the present application clearer and easier to understand, the following describes a method for predicting the three-dimensional distribution of the aerodynamic thermal environment of an aircraft based on a combined model provided in an embodiment of the present application in conjunction with the accompanying drawings. Figure 1 As shown, this figure is a flow chart of a method for predicting the three-dimensional distribution of the aerodynamic thermal environment of an aircraft based on a combination model provided in an embodiment of the present application.

[0025] The method is applied to the processing equipment, and the prediction method of the three-dimensional distribution of the aircraft aerodynamic thermal environment of the combined model includes: S101. A processing device obtains geometric features of an aircraft.

[0026] The aircraft includes a plane-symmetrical aircraft and a rotational body aircraft.

[0027] For cruise-type plane-symmetric aircraft, the plane-symmetric aircraft is simplified into a two-dimensional cylinder, plate, and wedge combination; for missile-type rotational-body aircraft, the rotational-body aircraft is simplified into a stationary point, sphere, cone, and column combination.

[0028] The first component of the assembly of a plane-symmetric aircraft is a two-dimensional cylinder (semi-infinite length cylinder), and the subsequent components are flat plates or wedges; the first component of the assembly of a rotational aircraft is a spherical surface, and the subsequent components are conical surfaces or cylindrical surfaces.

[0029] S102: The processing device determines N components according to geometric features of the aircraft, and constructs a three-dimensional assembly model of the aircraft using the N components.

[0030] According to the geometric features of the aircraft, N sub-features in the geometric features of the aircraft are determined; then according to the mapping relationship between the sub-features and pre-set reference features and reference components, the corresponding component of each sub-feature in the N sub-features is determined, wherein N is an integer greater than 1.

[0031] In some embodiments, taking a missile-type rotating body aircraft as an example, a typical combination of a ball head, a cone, and a column is shown in Table 1.

[0032] Table 1:

[0033] Among them, the position relationship of the reference feature in the projectile-like rotating body aircraft is as follows: Figure 2 As shown, this figure is a schematic diagram of the positions of reference features of a missile-type rotating body aircraft provided in an embodiment of the present application.

[0034] According to the above steps, the corresponding components of each of the N sub-features are determined, and the corresponding components of each of the N sub-features are used to obtain a combination of the aircraft. The combinations are combined to obtain a three-dimensional combination model of the aircraft.

[0035] S103: The processing device divides the surface grid of the three-dimensional assembly model.

[0036] After obtaining the three-dimensional assembly model of the aircraft, the surface mesh of the three-dimensional assembly model of the aircraft is divided, and the grid points are defined using node numbers (i, j), where i increases along the axial direction and j increases clockwise along the circumference (front view); i=0 is located at the front end of each component, j=0 is located on the back side of the longitudinal symmetry plane, different i values ​​correspond to different axial positions, and different j values ​​correspond to different circumferential meridians, such as Figure 3As shown in FIG. 1 , this figure is a schematic diagram of node numbering for a three-dimensional assembly model provided in an embodiment of the present application. A typical ball head, cone, and column assembly model and its surface mesh are shown in FIG. Figure 4 As shown, this figure is a schematic diagram of the surface mesh division of a three-dimensional assembly model provided in an embodiment of the present application.

[0037] S104: The processing device obtains the incoming flow condition, obtains the geometric relationship between the Mth component and the M-1th component, and calculates the geometric parameters of the mth grid point of the Mth component.

[0038] The incoming flow conditions are divided into ballistic conditions and wind tunnel conditions. Ballistic incoming flow parameters include altitude, speed, etc., which are calculated based on the 1976 standard atmospheric model of the United States. Wind tunnel incoming flow parameters include total temperature, total pressure, total density, etc., which are converted based on the flow stagnation parameter relationship. The user inputs the parameters of the aircraft's state into the 1976 standard atmospheric model of the United States and the flow stagnation parameter relationship to calculate and obtain the aircraft's incoming flow conditions.

[0039] During the flight of the aircraft, in order to calculate the thermal environment of the aircraft at all times along the entire trajectory, this application uses t i The subsequent calculations are based on the incoming flow conditions at the time.

[0040] Get the geometric relationship between the Mth component and the M-1th component. The geometric relationship is the object plane inclination. Therefore, compare the object plane inclination of the Mth component. Inclination angle to the object plane of the M-1th component , obtain the flow state of the Mth component.

[0041] If the object plane inclination angle of the Mth component The object plane inclination angle is greater than the M-1th component , then the flow of the Mth component relative to the M-1th component is in a compression state, and the relative inclination angle between the components is the compression angle ,like Figure 5 As shown in FIG. 1 , this figure is a schematic diagram of judging the compression relationship between adjacent components provided by an embodiment of the present application. Represents the incoming flow velocity, that is, the velocity at infinity.

[0042] If the object plane inclination angle of the Mth component Less than the object plane inclination of the M-1th component , then the flow of the Mth component relative to the M-1th component is in an expansion state, and the relative inclination angle between the components is the expansion angle ,like Figure 6 As shown, this figure is a schematic diagram of determining the expansion relationship between adjacent components provided in an embodiment of the present application.

[0043] Then, the geometric parameters of the mth grid point of the Mth component are calculated, wherein the geometric parameters include equivalent angle and streamline length, 1<M≤N, and M is an integer. The calculation method of the equivalent angle and streamline length is described in detail below.

[0044] The original half-cone angle, incoming flow angle of attack and meridian angle of the mth grid point of the Mth component are obtained. The incoming flow angle of attack is input by the user in the incoming flow condition. 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, incoming flow angle of attack and meridian angle of the mth grid point of the Mth component, the equivalent angle of the mth grid point of the Mth component, that is, the equivalent half-cone angle, is calculated, as shown in formula (1): (1) in, is the equivalent half cone angle, is the original half cone angle, is the incoming flow angle of attack, It is the meridian angle.

[0045] For the windward and leeward sides, the above formula is simplified to formula (2): (2) That is, the equivalent half-cone angle on the windward side is the original half-cone angle plus the angle of attack, and the equivalent half-cone angle on the leeward side is the original half-cone angle minus the angle of attack.

[0046] The streamline length of the mth grid point of the Mth component is the sum of the distances from the 1st grid point of the 1st component to the mth grid point of the Mth component.

[0047] In some embodiments, Figure 2 As shown, the streamline length of the mth grid point of the component is .

[0048] S105. The processing device calculates the flow parameters of the mth grid point of the Mth component according to the incoming flow conditions, the geometric relationship between the Mth component and the M-1th component, and the geometric parameters of the mth grid point of the Mth component.

[0049] The flow parameters include shock wave angle, post-wave flow parameters, surface pressure coefficient and boundary outer edge parameters, and the calculation method of the flow parameters is introduced in detail.

[0050] Calculate the shock wave angle. For a stagnation point or a sphere, the flow field shock wave type is a normal shock wave, and the shock wave angle is 90 degrees. For a swept cylinder, the flow field shock wave type 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 a wedge, the shock wave angle is solved by the oblique shock wave relationship. For a cone, the shock wave angle is calculated by iteratively solving the cone flow equation.

[0051] Calculate the flow parameters after the wave, take the incoming flow parameters, shock wave angle, equivalent angles of each component, and compression / expansion parameters as input conditions, and calculate the flow parameters after the wave based on the normal shock wave, oblique shock wave, and Prandtl- Meyer expansion flow relationship. For the normal shock wave, when using calorimetric complete gas calculation, solve the normal shock wave equation, and obtain a series of simplified relationships with the wavefront Mach number as a function; for the oblique shock wave, when using calorimetric complete gas calculation, take the velocity component of the wavefront incoming flow perpendicular to the shock wave as the basic quantity, and obtain a simplified relationship with the wavefront normal Mach number as a function; for the Prandtl- Meyer expansion flow, when using calorimetric complete gas calculation, iteratively solve the PM relationship. In some embodiments using high-temperature gas calculation, for normal shock waves and oblique shock waves, simple formulas cannot be obtained, and the equation group must be numerically solved.

[0052] Calculate the surface pressure coefficient based on the modified Newton theory, the wedge cutting method, and the cone cutting method.

[0053] The modified Newtonian theory uses the Lees formula, as shown in formula (3): (3) in, is the surface pressure coefficient, is the angle between the incoming flow vector and the object surface normal vector, is the pressure coefficient of the stagnation point after the normal shock wave, and the calculation formula is as follows: (4) in, is the stagnation pressure after the shock wave, is the incoming flow pressure, is the incoming flow density, is the incoming flow velocity, is the incoming flow Mach number, is the specific heat ratio of gases.

[0054] Wedge cutting method, cone cutting method: at any point i on the surface of the object, define its local deflection angle is the angle between the local surface tangent and the incoming flow vector. Extending the tangent at this point can give a semi-wedge angle or semi-cone angle of The cut wedge method or cut cone method assumes that the pressure at point i is equal to the pressure on the corresponding equivalent wedge or equivalent cone.

[0055] Calculate the parameters of the outer edge of the boundary. The first-order boundary layer approximation allows the pressure at the outer edge of the boundary layer to be equal to the corresponding surface pressure. In the case of hypersonic speed, 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 wave.

[0056] S106: The processing device calculates the aerodynamic thermal 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.

[0057] According to the flow parameters of the mth grid point of the Mth component, the aerodynamic thermal environment parameters of the mth grid point of the Mth component are calculated, including: Get the component type of the Mth component, the component types include stationary point, sphere, plate, wedge and cone; For the stagnation point, the aerodynamic thermal environment parameters of the ball head stagnation point can be calculated by the Fay Riddell method. The calculation formula is shown in formula (5): (5) in, is the stagnation heat flow, 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.

[0058] In formula (5) It can be calculated by formula (6): It can be calculated by formula (7), formula (6) (7) is as follows: (6) (7) in, 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, For pressure, is the viscosity coefficient, is the Lewis number, is the dissociation enthalpy, is the radius.

[0059] For a spherical surface, where the spherical atmosphere is laminar and turbulent, for laminar flow, the Lees method is used to calculate the aerodynamic thermal environment parameters of the mth grid point of the Mth component. For turbulent flow, the Detra method is used to calculate the aerodynamic thermal environment parameters of the mth grid point of the Mth component.

[0060] The calculation formula of Lees method is shown in formula (8) (9): (8) (9) in, is the laminar heat transfer coefficient, is the stagnation point heat transfer coefficient, is the sphere center angle, is the intermediate variable expression.

[0061] The calculation formula of the Detra method is shown in formula (10): (10) in, is the turbulent heat transfer coefficient, is the streamline length at the calculation point, is the local tangent angle of the sphere, 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.

[0062] According to the calculation strategy corresponding to the component type of the Mth component, the aerodynamic thermal environment parameters of the mth grid point of the Mth component are calculated using the flow parameters of the mth grid point of the Mth component.

[0063] For plates, wedges and cones, the Eckert reference enthalpy method can be used to calculate the aerodynamic thermal environment parameters of the mth grid point of the Mth component.

[0064] The calculation formula of laminar flow is shown in formula (11): (11) The calculation formula for turbulence is shown in formula (12): (12) in, is the heat transfer coefficient, Take 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.

[0065] In some embodiments, the Eckert reference enthalpy method is used for the plate, wedge, and cone, but different Mangler factors are used to perform equivalent treatment on the streamline length. For the cone, the Mangler factor in the laminar state is 3.0, and the Mangler factor in the turbulent state is 2.0; for the plate and wedge, the Mangler factors in the laminar state and the turbulent state are both 1.0.

[0066] S107: The processing device determines whether the mth grid point is the last grid point of the Mth component.

[0067] The aerodynamic thermal environment parameters of the mth grid point of the Mth component are obtained through the above steps, and then it is determined whether the mth grid point is the last grid point of the Mth component.

[0068] If the mth grid point is the last grid point of the Mth component, execute S108; If the m-th grid point is not the last grid point of the M-th component, execute S110.

[0069] S108: The processing device determines whether the Mth component is the last component among the N components.

[0070] If the Mth component is the last component among the N components, execute S109; if the Mth component is not the last component among the N components, execute S111.

[0071] S109. The processing device outputs the aerodynamic thermal environment parameters of each grid point of the N components.

[0072] If the Mth component is the last component among the N components, then t i The calculation of aerodynamic thermal environment parameters at the moment is completed.

[0073] Next, the aerodynamic thermal environment parameters of each grid point of each component at the next moment are calculated, that is, t i+1 The aerodynamic thermal environment parameters of each grid point of each component at the moment are calculated as shown in S104 to S106, which will not be repeated here. Among them, each grid point data contains the data of all ballistic or test points. If a node number (i, j) is specified, only the grid point data is output. If a meridian number (j) is specified, all grid point data on the meridian are output.

[0074] S110: The processing device processes the next grid point of the Mth component.

[0075] The processing device calculates the aerodynamic thermal environment parameters of the m+1th grid point of the Mth component, and returns to S104 with m=m+1. The calculation method is shown in S104 to S106 and will not be repeated here.

[0076] S111, the processing device processes the next component.

[0077] The processing device calculates the aerodynamic thermal environment parameters of the first grid point of the M+1th component and returns to S104 with M=M+1, m=1. The calculation method is shown in S104 to S106 and will not be repeated here.

[0078] Based on the above content, the processing device obtains the geometric characteristics of the aircraft, and then determines N components according to the geometric characteristics of the aircraft, uses the N components to build a three-dimensional assembly model of the aircraft, and then divides the surface grid of the three-dimensional assembly model. The processing device then obtains the incoming flow condition, obtains the geometric relationship between the Mth component and the M-1th component, calculates the geometric parameters of the mth grid point of the Mth component, and calculates the mth grid point of the Mth component according to the incoming flow condition, the geometric relationship between the Mth component and the M-1th component, and the geometric parameters of the mth grid point of the Mth component. The flow parameters of the grid points are calculated, and then the aerodynamic thermal environment parameters of the mth grid point of the Mth component are calculated according to the flow parameters of the mth grid point of the Mth component. After calculating the aerodynamic thermal 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 thermal environment parameters of each grid point of the N components are output. It can be seen that the present application takes into account the mutual influence between components, obtains the flow expansion or compression relationship between components, can handle complex-shaped aircraft with multi-level compression and expansion, and can obtain the three-dimensional distribution of pressure and thermal environment parameters on the surface of the aircraft.

[0079] In some embodiments, the above method is verified by taking a blunt double cone test model as an example.

[0080] The blunt double cone test model is a ball-cone-cone combination. The specific size parameters are shown in Table 2. The geometric dimensions of the blunt double cone test model are shown in Figure 7 As shown, this figure is a schematic diagram of the geometric dimensions of a blunt double cone test model provided in an embodiment of the present application.

[0081] Table 2-Dimensional parameters of blunt double cone assembly

[0082] The incoming flow parameters are wind tunnel conditions: , , , , .

[0083] The calculation results of the method of the present invention are as follows Figure 8 , 9 shown. Figure 8 This is a schematic diagram of the three-dimensional distribution results of heat flux on the surface of the blunt double cone experimental model provided in the embodiment of the present application. Fig. 9 Schematic diagram of the comparison curves of the blunt double cone test model provided in the embodiment of the present application on different meridians. Wherein, exp is the wind tunnel test result, CFD is the numerical simulation result, and ZQheat is the calculation result of the present invention.

[0084] According to the above calculation results, it can be shown that the prediction method of the three-dimensional distribution of the aerodynamic thermal environment of an aircraft based on a combination model proposed in this application has a certain calculation accuracy, can handle 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.

[0085] Combination of the above Figures 1 to 9 A method for predicting the three-dimensional distribution of the aerodynamic thermal environment of an aircraft based on a combination model provided in an embodiment of the present application is introduced in detail. The device and equipment provided in the embodiment of the present application will be introduced in conjunction with the accompanying drawings.

[0086] The present application also provides a device for predicting the three-dimensional distribution of the aerodynamic thermal environment of an aircraft based on a combined model. Fig.10 As shown, the figure is a schematic diagram of a prediction device for the three-dimensional distribution of the aerodynamic thermal environment of an aircraft based on a combined body model provided in an embodiment of the present application, the device comprising: an acquisition module 1001, a calculation module 1002 and a judgment module 1003; The acquisition module 1001 is used to acquire 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 assembly model of the aircraft, where N is an integer greater than 1; divide the surface grid of the three-dimensional assembly model; obtain the incoming flow condition, obtain the geometric relationship between the Mth component and the M-1th component, and calculate the geometric parameters of the mth grid point of the Mth component, where the geometric relationship includes the object surface inclination angle, and the geometric parameters include the equivalent angle and the streamline length, where 1<M≤N, and M is an integer; The calculation module 1002 is used to calculate the flow parameters of the mth grid point of the Mth component according to the incoming flow condition, the geometric relationship between the Mth component and the M-1th component, and the geometric parameters of the mth grid point of the Mth component; and calculate the aerodynamic thermal 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; The judgment module 1003 is used to judge whether the m-th grid point is the last grid point of the M-th component; if the m-th grid point is the last grid point of the M-th component, then judge whether the M-th component is the last component among the N components; if the M-th component is the last component among the N components, output the aerodynamic thermal environment parameters of each grid point of the N components.

[0087] In some possible implementations, the judgment module 1003 is further configured to calculate the aerodynamic thermal 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.

[0088] In some possible implementations, the judgment module 1003 is further configured to calculate the aerodynamic thermal environment parameter of the first grid point of the M+1th component if the Mth component is not the last component among the N components.

[0089] In some possible implementations, the flow parameters include: shock wave angle, post-wave flow parameters, surface pressure coefficient, and boundary outer edge parameters.

[0090] In some possible implementations, the calculation module 1002 is also used to obtain the component type of the Mth component; according to the calculation strategy corresponding to the component type of the Mth component, the aerodynamic thermal environment parameters of the mth grid point of the Mth component are calculated using the flow parameters of the mth grid point of the Mth component.

[0091] In some possible implementations, the acquisition module 1001 is further used to determine N sub-features in the geometric features of the aircraft; and determine the component corresponding to each sub-feature in the N sub-features based on the mapping relationship between the sub-features and pre-set reference features and reference components.

[0092] In some possible implementations, the aircraft includes a plane-symmetric aircraft and a rotational body aircraft.

[0093] The present application also provides a computing device. Fig.11 As shown, this figure is a schematic diagram of a computing device provided in an embodiment of the present application, and 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.

[0094] The bus 1101 may be a peripheral component interconnect (PCI) bus or an extended industry standard architecture (EISA) bus. The bus may be divided into an address bus, a data bus, a control bus, etc. For ease of representation, Fig.11 Only one thick line is used in the diagram, but this does not mean that there is only one bus or only one type of bus.

[0095] 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).

[0096] The communication interface 1103 is used for communicating with the outside.

[0097] The memory 1104 may include a volatile memory, such as a random access memory (RAM). The memory 1104 may also 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).

[0098] The memory 1104 stores executable codes, and the processor 1102 executes the executable codes to implement the aforementioned method for predicting the three-dimensional distribution of the aerodynamic thermal environment of an aircraft based on the combined body model.

[0099] The embodiment of the present application also provides a computer-readable storage medium. The computer-readable storage medium can be any available medium that can be stored by a computing device or a data storage device such as a data center containing one or more available media. The available medium can be a magnetic medium (such as a floppy disk, a hard disk, a tape), an optical medium (such as a DVD), or a semiconductor medium (such as a solid-state hard disk). The computer-readable storage medium includes instructions that instruct the computing device to execute the above-mentioned prediction method for the three-dimensional distribution of the aircraft's aerodynamic thermal environment based on the combined body model.

[0100] The embodiment of the present application further 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 process or function described in the embodiment of the present application is generated in whole or in part.

[0101] 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 via wired (e.g., coaxial cable, optical fiber) or wireless (e.g., infrared, wireless, microwave, etc.) means.

[0102] When the computer program product is executed by a computer, the computer executes any of the aforementioned methods for predicting the three-dimensional distribution of the aerodynamic thermal environment of an aircraft based on a combined body model. The computer program product may be a software installation package, and when any of the aforementioned methods for predicting the three-dimensional distribution of the aerodynamic thermal environment of an aircraft based on a combined body model is needed, the computer program product may be downloaded and executed on a computer.

[0103] The descriptions of the processes or structures corresponding to the above-mentioned figures have different emphases. For parts that are not described in detail in a certain process or structure, please refer to the relevant descriptions of other processes or structures.

[0104] The above description is only a specific implementation 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 included in the protection scope of the present application.

Claims

1. A method for predicting the three-dimensional distribution of the aerodynamic thermal environment of an aircraft based on a combined model, characterized in that: The method comprises: Obtaining 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 assembly model of the aircraft, where N is an integer greater than 1; Dividing the surface grid of the three-dimensional assembly model; Obtaining the incoming flow condition, obtaining the geometric relationship between the Mth component and the M-1th component, and calculating the geometric parameters of the mth grid point of the Mth component, wherein the geometric relationship includes the object plane inclination angle, and the geometric parameters include the equivalent angle and the streamline length, 1<M≤N, and M is an integer; Calculate the flow parameter of the mth grid point of the Mth component according to the incoming flow condition, the geometric relationship between the Mth component and the M-1th component, and the geometric parameter of the mth grid point of the Mth component; Calculating the aerodynamic thermal environment parameter of the mth grid point of the Mth component according to the flow parameter of the mth grid point of the Mth component; Determine whether the m-th grid point is the last grid point of the M-th component; If the m-th grid point is the last grid point of the M-th component, determining whether the M-th component is the last component among the N components; If the Mth component is the last component among the N components, the aerodynamic thermal environment parameters of each grid point of the N components are output.

2. The method according to claim 1, characterized in that The method further comprises: If the mth grid point is not the last grid point of the Mth component, the aerodynamic thermal environment parameters of the m+1th grid point of the Mth component are calculated.

3. The method according to claim 1, characterized in that: The method further comprises: If the Mth component is not the last component among the N components, the aerodynamic thermal environment parameters of the first grid point of the M+1th component are calculated.

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

5. The method according to claim 1, characterized in that The step of calculating the aerodynamic thermal environment parameter of the mth grid point of the Mth component according to the flow parameter of the mth grid point of the Mth component comprises: Obtaining the component type of the Mth component; According to the calculation strategy corresponding to the component type of the Mth component, the aerodynamic thermal environment parameters of the mth grid point of the Mth component are calculated using the flow parameters of the mth grid point of the Mth component.

6. The method according to claim 1, characterized in that The step of determining N components according to the geometric features of the aircraft includes: Determining N sub-features of the geometric features of the aircraft; According to the sub-features and a mapping relationship between a preset reference feature and a reference component, a component corresponding to each sub-feature in the N sub-features is determined.

7. The method according to claim 1, characterized in that The aircraft includes a plane-symmetrical aircraft and a rotational body aircraft.

8. A prediction device for the three-dimensional distribution of the aerodynamic thermal environment of an aircraft based on a combined model, characterized in that: The device comprises: An acquisition module is used to acquire geometric features of an aircraft; determine N components according to the geometric features of the aircraft, and use the N components to construct a three-dimensional assembly model of the aircraft, where N is an integer greater than 1; divide the surface grid of the three-dimensional assembly model; acquire incoming flow conditions, acquire the geometric relationship between the Mth component and the M-1th component, and calculate the geometric parameters of the mth grid point of the Mth component, where the geometric relationship includes the object surface inclination angle, and the geometric parameters include the equivalent angle and the streamline length, where 1<M≤N, and M is an integer; A calculation module, used for calculating the flow parameters of the mth grid point of the Mth component according to the incoming flow condition, the geometric relationship between the Mth component and the M-1th component, and the geometric parameters of the mth grid point of the Mth component; and calculating the aerodynamic thermal 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 is used 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, then 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 thermal environment parameters of each grid point of the N components.

9. A computing device, characterized in that including memory and processor; 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 executes the method as claimed in 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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