Double-trailing-vortex encounter effect simulation method for flight simulator

Abstract

The invention belongs to the technical field of flight simulators and simulation, particularly relates to a double-trailing-vortex encounter effect simulation method for a flight simulator, and aims to solve the problem that the real-time performance and high-precision simulation requirements of a flight training platform cannot be met. The method comprises the following steps: calculating trailing vortex indexes by acquiring preorder aircraft operation related parameters, dividing a wake flow dissipation stage, establishing a trailing vortex intensity model, determining a vortex position, calculating an induced velocity, comparing trailing vortex intensity, determining a relative distance between a front aircraft and a rear aircraft, selecting different modes to calculate an incremental velocity and an angular velocity, and integrating to obtain aircraft state parameters; a new increment aerodynamic coefficient can be calculated according to the airplane state parameters, then the airplane state is obtained according to the new increment aerodynamic coefficient, the calculated airplane state tends to be accurate through coefficient iteration, and the average operation time is shortened. According to the invention, simulation precision and real-time performance are effectively improved, and related requirements of a flight training platform can be fully satisfied.

Description

Technical Field

[0001] The present invention belongs to the technical fields of flight simulators and simulation technologies, and particularly relates to a simulation method for double wake encounter effects of a flight simulator. Background Art

[0002] An aircraft wake is a by-product of aircraft lift. When a following aircraft enters the wake region of a preceding aircraft, especially for small aircraft of other series such as the ARJ21, dangerous situations such as roll and bump will occur to the following aircraft, especially when the wake effect level generated by the preceding aircraft is higher than the response ability of the following aircraft.

[0003] Wake encounter effect simulation technology provides important support for flight safety and air traffic management by simulating the generation, evolution of wakes and their effects on following aircraft. Current methods mainly include numerical simulation methods and wake decay simulation models. Among them, the numerical simulation method has high modeling accuracy but poor real-time performance and is not suitable for real-time simulation operations on flight training platforms. The wake decay model can quickly evaluate the effects of a following aircraft encountering a wake, but the model is simplified and does not consider the aerodynamic interference mechanism. Existing models are difficult to accurately calculate the incremental aerodynamic force caused by the wake of the preceding aircraft, and the simulated aircraft wake response state is difficult to approximate the real scenario. Summary of the Invention

[0004] In order to solve the above problems in the prior art, that is, in the existing wake encounter effect simulation technology, the numerical simulation method has poor real-time performance, and the wake decay model cannot accurately calculate the incremental aerodynamic force due to the lack of consideration of the aerodynamic interference mechanism, and the simulated aircraft wake response state is difficult to approximate the real situation, and it cannot meet the real-time and high-precision simulation requirements of the flight training platform. In the first aspect of the present invention, a simulation method for double wake encounter effects of a flight simulator is proposed. The method includes the following steps: S1, based on the operation-related parameters of the preceding aircraft in the flight simulator training scenario, obtain wake vortex parameters through a pre-constructed wake vortex model; S2, according to the stage of wake dissipation of the preceding aircraft, establish wake vortex intensity models for each stage based on the wake vortex parameters and the dimensionless buoyancy frequency, and obtain the wake vortex intensity values corresponding to the stages; S3, determine the vortex position; calculate the distance from any point on the following aircraft encountering the wake to the vortex position, and calculate the induced velocity of a single vortex at any point based on this distance; S4, based on the wake vortex intensity value and the initial circulation of the wake vortex, calculate the ratio of the current wake vortex intensity, and compare it with a critical value to obtain the relative distance between the preceding aircraft and the following aircraft; S5. Select different methods to calculate the incremental speed and incremental angular velocity of the whole aircraft according to the relative distance between the preceding aircraft and the following aircraft; integrate the incremental speed and incremental angular velocity respectively to obtain the aircraft state parameters, where the aircraft state parameters include the speed, position, attitude angular velocity and attitude angle of the aircraft.

[0005] In some preferred embodiments, the operation-related parameters include the mass, wingspan, true airspeed, normal overload and air density of the preceding aircraft; the wake vortex parameters include the initial circulation of the wake vortex, vortex core radius, characteristic time, and wake vortex characteristic velocity. The initial circulation of the wake vortex: ; ; Where is the initial circulation of the vortex; is the normal overload of the aircraft; is the mass of the preceding aircraft; is the acceleration due to gravity; is the air density; is the true airspeed of the preceding aircraft; is the initial vortex spacing; Vortex core radius: ; is the vortex core radius; is the wingspan of the aircraft; is the initial vortex core radius; Wake vortex characteristic velocity: ; Characteristic time: ; Where is the characteristic time; is the wake vortex characteristic velocity.

[0006] In some preferred embodiments, the stages of the dissipation of the wake vortex of the preceding aircraft include the diffusion stage and the rapid decay stage. The wake vortex intensity model in the diffusion stage is: ; The wake vortex intensity model in the rapid decay stage is: ; Where is the value of the wake vortex intensity after attenuation in the diffusion stage; is the dimensionless buoyancy frequency; is the duration of the wake vortex effect; is the duration of the diffusion stage.

[0007] In some preferred embodiments, the method for determining the position of the eddy current is as follows: Select a cross-section perpendicular to the wake flow field, and establish a coordinate system with the midpoint of the connection line between the left and right vortex cores as the origin, the x-axis along the connection line direction, and the y-axis perpendicular to the connection line; The left and right eddy currents are respectively located at the positions of (0, B / 2) and (0, -B / 2) in the coordinate system.

[0008] In some preferred embodiments, the method for obtaining the induced velocity of the single eddy current at any point is as follows: Based on the positions of the left and right eddy currents, calculate the distance from any point on the aircraft encountering the wake to the position of the eddy current: ; Calculate the induced velocity of the single eddy current at any point according to the distance: ; Wherein, is the distance from any point on the subsequent aircraft to the origin; is the intensity of the wake of the previous aircraft encountered by the subsequent aircraft; x and y are the coordinates of any point on the aircraft.

[0009] In some preferred embodiments, the method for obtaining the relative distance between the previous aircraft and the subsequent aircraft is as follows: Calculate the ratio of the current wake intensity: ; Compare the ratio of the current wake intensity with the critical value to obtain the relative distance between the previous aircraft and the subsequent aircraft: When D is less than or equal to the critical value, it is determined that the relative distance between the previous aircraft and the subsequent aircraft is far; otherwise, it is determined that the relative distance between the previous aircraft and the subsequent aircraft is close.

[0010] In some preferred embodiments, according to the relative distance between the previous aircraft and the subsequent aircraft, different methods are selected to calculate the incremental velocity and incremental angular velocity of the whole aircraft, and the method is as follows: If the relative distance between the previous aircraft and the subsequent aircraft is close, split the subsequent aircraft into set force components and calculate the induced velocity of each force component. Based on the induced velocity, calculate the incremental aerodynamic force, and respectively obtain the incremental velocity and incremental angular velocity of the corresponding force component through integration. Then, superimpose the incremental velocity and incremental angular velocity of each force component to obtain the incremental velocity and incremental angular velocity of the whole aircraft; If the relative distance between the previous aircraft and the subsequent aircraft is far, calculate the incremental aerodynamic force according to the induced velocity. The incremental aerodynamic force and control parameters pass through the six-degree-of-freedom equation to obtain the incremental velocity and incremental angular velocity of the whole aircraft; The incremental velocity is u, v, w; the incremental angular velocity is p, q, r.

[0011] In some preferred embodiments, the incremental aerodynamic forces include lift, drag, side force, rolling moment, pitching moment, and yawing moment; the calculation methods for the parameters in the incremental aerodynamic forces are as follows: Lift: ; Drag: ; Side force: ; Rolling moment: ; Pitching moment: ; Yawing moment: ; Wherein, is the induced velocity; is the current air density; is the projected area of the entire aircraft; is the wingspan; is the mean aerodynamic chord of the wing; is the lift coefficient; is the drag coefficient; is the side force coefficient; is the rolling moment coefficient; is the pitching moment coefficient; is the yawing moment coefficient.

[0012] In some preferred embodiments, the six-degree-of-freedom equations: ; ; ; ; ; ; Wherein, is the total weight of the aircraft; are the incremental velocity change rates along the xyz axes respectively; is the thrust; is the angle of attack; is the sideslip angle; is the pitch angle; is the roll angle; are the moments of inertia about the xyz axes respectively; are the incremental angular velocities about the xyz axes respectively; these belong to the control parameters.

[0013] Advantages of the present invention: The simulation module / platform incorporating this technology can implement the arrangement and application of the vortex effect on a flight simulator, enriching the training scenarios; Based on the subsequent aircraft's takeoff time and distance, model and analyze the wake effect it encounters, integrate the aerodynamic model and the wake model, and finally obtain the incremental aerodynamic force and aircraft state caused by the wake of the preceding aircraft. The new incremental aerodynamic force coefficient can be calculated through the difference algorithm according to the aircraft state parameters, and the aircraft state obtained based on the new incremental aerodynamic force coefficient will tend to be consistent as the number of coefficient iteration increases. The results of the two are within the ±5% error band, improving the pilot's emergency response ability in this scenario, ensuring flight safety, and thus enhancing the training effect; Compare the running time of the simulation module incorporating this technology: On a 100Hz simulation platform, that is, the sum of the running cycles of all aircraft modules should be less than 10ms. The average running time of the vortex simulation is approximately 200μs (30s is a test cycle, taking the average of 6 cycles), improving the real-time performance. Using the numerical simulation method would be more time-consuming. BRIEF DESCRIPTION OF THE DRAWINGS

[0014] Other features, objectives, and advantages of the present application will become more apparent by reading the detailed description of the non-limiting embodiments with reference to the following drawings: Figure 1 It is a flowchart of the steps of a method for simulating the double-vortex encounter effect for a flight simulator according to the present invention. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0015] The following further elaborates on the present application in conjunction with the drawings and embodiments. It can be understood that the specific embodiments described herein are only for explaining the relevant invention and not for limiting the invention. Additionally, it should be noted that only parts related to the relevant invention are shown in the drawings for ease of description.

[0016] It should be noted that, without conflict, the embodiments in the present application and the features in the embodiments can be combined with each other. The present application will be described in detail below with reference to the drawings and embodiments.

[0017] To more clearly illustrate a method for simulating the double-vortex encounter effect for a flight simulator according to the present invention, the following combines Figure 1 elaborates on each step in the embodiments of the present invention in detail.

[0018] The present invention provides a method for simulating the double-vortex encounter effect for a flight simulator. Refer to Figure 1 , this method includes the following steps: S1, based on the operating-related parameters of the preceding aircraft in the flight simulator training scenario, obtain the wake parameters through a pre-constructed wake model; In this embodiment, the operation-related parameters include the mass, wingspan, true airspeed, normal overload, and air density of the preceding aircraft; The wake vortex parameters include the initial circulation of the wake vortex, initial vortex spacing, vortex core radius, characteristic time, and wake vortex characteristic velocity; The initial circulation of the wake vortex: ; ; where is the initial circulation of the vortex; is the normal overload of the aircraft; is the mass of the preceding aircraft; is the acceleration due to gravity; is the air density; is the true airspeed of the preceding aircraft; is the initial vortex spacing; Vortex core radius: ; where is the vortex core radius; is the wingspan of the aircraft; is the initial vortex core radius, and simulation is carried out under the condition of generating the strongest moment in the wake vortex velocity field. The vortex core radius is approximately equal to the initial vortex core radius ; Wake vortex characteristic velocity: ; Characteristic time: ; is the characteristic time; Wake vortex characteristic velocity; S2. According to the stages of the dissipation of the wake of the preceding aircraft, based on the wake vortex parameters and the dimensionless buoyancy frequency, establish the wake vortex intensity models for each stage to obtain the wake vortex intensity values corresponding to each stage; In this embodiment, the stages of the dissipation of the wake of the preceding aircraft include the diffusion stage and the rapid decay stage; The wake vortex intensity model for the diffusion stage (near-field vortex) is: ; The wake vortex intensity model for the rapid decay stage (far-field vortex) is: ; where is the wake vortex intensity value after decay in the diffusion stage; is the dimensionless buoyancy frequency; is the duration of the wake vortex effect; is the duration of the diffusion stage; S3. Determine the position of the vortex; calculate the distance from any point on the subsequent aircraft encountering the wake to the position of the vortex, and calculate the induced velocity of a single vortex at any point based on this distance; In this embodiment, to determine the position of the vortex, the method is as follows: Select a cross-section perpendicular to the wake field, and establish a coordinate system with the midpoint of the line connecting the left and right vortex cores as the origin, the x-axis along the connection direction, and the y-axis perpendicular to the connection; The left and right vortices are respectively located at the positions of (0, B / 2) and (0, -B / 2) in the coordinate system; The method for obtaining the induced velocity of a single vortex at any point is as follows: Based on the positions of the left and right vortices, calculate the distance from any point on the aircraft encountering the wake to the position of the vortex: ; Calculate the induced velocity of a single vortex at any point according to the distance: ; Wherein, is the distance from any point on the subsequent aircraft to the origin; , respectively represent the distances of the right vortex and the left vortex. The formula for expressing the distance from any point on the aircraft to the double vortices can be substituted into the induced velocity formula to calculate the induced velocity; is the intensity of the wake of the previous aircraft encountered by the subsequent aircraft; x and y are the coordinates of any point on the aircraft; It should be noted here that when the relative distance between the previous aircraft and the subsequent aircraft is relatively far in the following text, the subsequent aircraft is regarded as a particle, and then the induced velocity is calculated with one distance; When the relative distance between the previous aircraft and the subsequent aircraft is relatively close in the following text, the non-uniformity of the velocity field encountered by the subsequent aircraft results in different distances for different aircraft components, and different induced velocities are calculated; Note: The influence on the aircraft is reflected in the area parameter S in the calculation of aerodynamic force / moment; S4. Based on the wake vortex strength value and the initial circulation of the wake vortex, calculate the ratio of the current wake vortex strength and compare it with the critical value to obtain the relative distance between the previous aircraft and the subsequent aircraft; In this embodiment, the geometric parameters, flight altitude, and current atmospheric parameters of the previous aircraft will all affect the attenuation intensity of the wake vortex. When the wake vortex strength reaches a certain critical value, the induced velocities caused by it on each component of the subsequent aircraft are approximately equal. Therefore, by judging whether the current wake vortex strength reaches the critical value, the calculation model is divided into two parts, and the judgment condition of the wake vortex strength is the ratio of the current wake vortex strength; The method for obtaining the relative distance between the previous aircraft and the subsequent aircraft is as follows: Calculate the ratio of the current wake vortex strength: ; Compare the ratio of the current wake intensity with a critical value to obtain the relative distance between the preceding aircraft and the following aircraft: When the D is less than or equal to the critical value, it is determined that the relative distance between the preceding aircraft and the following aircraft is far; otherwise, it is determined that the relative distance between the preceding aircraft and the following aircraft is close; where the critical value is preferably 0.1; S5. According to the relative distance between the preceding aircraft and the following aircraft, select different methods to calculate the incremental velocity and incremental angular velocity of the whole aircraft; integrate the incremental velocity and incremental angular velocity respectively to obtain the aircraft state parameters; In this embodiment, according to the relative distance between the preceding aircraft and the following aircraft, different methods are selected to calculate the incremental velocity and incremental angular velocity of the whole aircraft. The method is as follows: If the relative distance between the preceding aircraft and the following aircraft is close, since the effect caused by the distance is large, the induced velocities encountered by different components of the following aircraft are different. Therefore, the following aircraft is split into set force-bearing components (fuselage, engine, wing, and horizontal tail) and the induced velocity of each force-bearing component is calculated. Based on the induced velocity, the incremental aerodynamic force is calculated, and the incremental velocity and incremental angular velocity of the corresponding force-bearing component are obtained through integration respectively. Specifically, according to the current lift increment ΔF (the lift, drag, and side force of the incremental aerodynamic force are calculated below, and then the increments of the lift, drag, and side force are obtained), and Newton's second law ΔF = mΔa, the corresponding incremental velocity can be obtained through integration; according to the current moment change amount ΔM (the rolling moment, pitching moment, and yaw moment of the incremental aerodynamic force are calculated below, and then the increments of the rolling moment, pitching moment, and yaw moment are obtained), and the rotational law of a rigid body ΔM = IΔα, the corresponding incremental angular velocity can be obtained through integration), and the incremental velocities and incremental angular velocities of each force-bearing component are superimposed to obtain the incremental velocity and incremental angular velocity of the whole aircraft (specifically, the incremental velocities and incremental angular velocities encountered by different components of the aircraft are linearly superimposed to obtain the incremental velocities u, v, w and incremental angular velocities p, q, r of the whole aircraft); If the relative distance between the preceding aircraft and the following aircraft is far, the whole aircraft uses an induced velocity of the same magnitude, that is, the induced velocity of the whole aircraft, which is calculated using one distance. Therefore, according to the induced velocity, the incremental aerodynamic force is calculated, and the incremental aerodynamic force and control parameters pass through the six-degree-of-freedom equation to obtain the incremental velocity and incremental angular velocity of the whole aircraft; The incremental velocity is u, v, w; the incremental angular velocity is p, q, r; It should be noted here that the calculations in the close-distance and long-distance modes are two parallel calculation threads. Only one mode can be selected for one calculation, and there will be no conflict. There is no distinction between u, v, w; p, q, r; As the distance increases, it will change to the long-distance mode, but the letters and meanings of the iteration parameters remain unchanged, and there is no need to distinguish the parameters; The incremental aerodynamic forces include lift, drag, side force, rolling moment, pitching moment, and yawing moment. The calculation methods for the parameters in the incremental aerodynamic forces are as follows: Lift: ; Drag: ; Side force: ; Rolling moment: ; Pitching moment: ; Yawing moment: ; Among them, is the induced velocity; is the current air density; is the projected area of the entire aircraft; is the wingspan; is the mean aerodynamic chord of the wing; is the lift coefficient; is the drag coefficient; is the side force coefficient; is the rolling moment coefficient; is the pitching moment coefficient; is the yawing moment coefficient. The specific calculation processes for each coefficient are as follows: ; ; ; ; ; In the formula, (Basic rigid-airplane) represents the influence of the basic rigid aircraft on each coefficient (in each coefficient, respectively represent the influence of the basic rigid aircraft on the lift coefficient, drag coefficient, side force coefficient, rolling moment coefficient, pitching moment coefficient, and yawing moment coefficient, specifically the influence increment. The same applies to other parameters); (change in stabilizer incidence and elevator deflectionfrom zero degrees) represents the influence of the stabilizer incidence angle and elevator deflection on each coefficient; (ground effect) represents the influence of ground effect on the corresponding coefficient; (airframe deformation resulting from static aeroelasticity and inertia relief) represents the influence of airframe elastic deformation on the corresponding coefficient (airplane dynamics) represents the influence of airplane dynamics on the corresponding coefficient; (spoiler deflection) represents the influence of spoiler deflection on the corresponding coefficient; (aileron deflection) represents the influence of aileron deflection on the corresponding coefficient; (rudder deflection) represents the influence of rudder deflection on the corresponding coefficient; (sideslip) represents the influence of sideslip on the corresponding coefficient; (landing gear represents the influence of landing gear on the corresponding coefficient; (aerodynamic thrust effects) represents the influence of aerodynamic thrust effects on the corresponding coefficient; (high lift device failure) represents the influence of high lift device failure on the corresponding coefficient; (icing) represents the influence of icing on the corresponding coefficient; (windmilling engine) represents the influence of windmilling engine on the corresponding coefficient; (reverse thrust effects) represents the influence of reverse thrust effects on the corresponding coefficient; (shift of the center of gravity from aerodynamic reference point) represents the influence of the change in the position of the center of gravity of the airplane on the corresponding coefficient; (ground effect at non-zero bank angle) represents the influence of the ground effect at a non-zero bank angle on the corresponding coefficient; The six-degree-of-freedom equations: ; ; ; ; ; ; where is the total weight of the aircraft; are the incremental velocity change rates along the x, y, and z axes respectively; is the thrust; is the angle of attack; is the sideslip angle; is the pitch angle; is the roll angle; are the moments of inertia about the x, y, and z axes respectively; are the incremental angular velocities about the x, y, and z axes respectively; By integrating the aircraft speed and position at the current moment can be obtained. By integrating the aircraft attitude angular velocity and attitude angle at the current moment can be obtained, and the calculated aircraft state is continuously updated with the time step. Among them, the integration process is an existing technology and will not be elaborated here; S6, and according to the aircraft state parameters, a new incremental aerodynamic coefficient can be calculated using the difference algorithm (specifically, refer to the literature, the self-calibration method of aerodynamic coefficients and its application in flight simulation), or it can be obtained by referring to "CN117910145B Method for determining aerodynamic coefficients, determining device, and storage medium"; Based on the new incremental aerodynamic coefficient, the incremental aerodynamic force caused by the wake is calculated. Based on the incremental aerodynamic force, the incremental velocity and incremental angular velocity are calculated and integrated to obtain the updated aircraft state parameters (according to the above steps S1 - S5). As the number of coefficient iterations increases, it converges to the real flight parameters, making the results of the simulation and real flight state parameters within the ±5% error band, improving the pilot's emergency response ability in this scenario, ensuring flight safety, and thus improving the training effect.

[0019] In the above embodiments, although the various steps are described in the above order, those skilled in the art can understand that in order to achieve the effects of this embodiment, different steps do not have to be executed in such an order. They can be executed simultaneously (in parallel) or in reverse order, and these simple changes are all within the protection scope of the present invention.

[0020] In the second embodiment of the present invention, a double-vortex encounter effect simulation system for a flight simulator, based on the above-mentioned double-vortex encounter effect simulation method for a flight simulator, the system includes: Vortex parameter generation module: configured to obtain vortex parameters through a pre-constructed vortex model based on the operation-related parameters of the preceding aircraft in the flight simulator training scenario. Vortex intensity calculation module: configured to establish a vortex intensity model for each stage based on the vortex parameters and the dimensionless buoyancy frequency according to the stage of wake dissipation of the preceding aircraft, and obtain the vortex intensity value of the corresponding stage. Induced velocity calculation module: configured to determine the vortex position; calculate the distance from any point on the subsequent aircraft encountering the wake to the vortex position, and calculate the induced velocity of a single vortex at any point based on this distance. Relative distance evaluation module: configured to calculate the ratio of the current vortex intensity based on the vortex intensity value and the initial circulation of the vortex, and compare it with a critical value to obtain the relative distance between the preceding aircraft and the subsequent aircraft. State parameter calculation module: configured to calculate the incremental speed and incremental angular velocity of the whole aircraft in different ways according to the relative distance between the preceding aircraft and the subsequent aircraft; integrate the incremental speed and incremental angular velocity respectively to obtain the aircraft state parameters.

[0021] It should be noted that the above-mentioned double-vortex encounter effect simulation system for a flight simulator provided in the above embodiments is only illustrated by the above division of each functional module. In actual applications, the above functions can be allocated to different functional modules according to needs, that is, the modules or steps in the embodiments of the present invention can be further decomposed or combined. For example, the modules in the above embodiments can be combined into one module, or further split into multiple sub-modules to complete all or part of the functions described above. For the names of the modules and steps involved in the embodiments of the present invention, they are only used to distinguish each module or step and are not regarded as an improper limitation of the present invention.

[0022] Those skilled in the art should be able to realize that the modules and method steps of each example described in combination with the embodiments disclosed herein can be implemented by electronic hardware, computer software, or a combination of both. The programs corresponding to the software modules and method steps can be placed in a random access memory (RAM), internal memory, read-only memory (ROM), electrically programmable ROM, electrically erasable programmable ROM, registers, hard disk, removable disk, CD-ROM, or any other form of storage medium well-known in the technical field. To clearly illustrate the interchangeability of electronic hardware and software, the composition and steps of each example have been generally described according to functions in the above description. Whether these functions are executed in the form of electronic hardware or software depends on the specific application and design constraints of the technical solution. Those skilled in the art can use different methods to implement the described functions for each specific application, but such implementation should not be considered to exceed the scope of the present invention.

[0023] The terms "first", "second", etc. are used to distinguish similar objects, rather than to describe or represent a specific order or sequence.

[0024] The term "comprising" or any other similar term is intended to cover non-exclusive inclusion, so that a process, method, article, or device / equipment including a series of elements not only includes those elements, but also includes other elements not explicitly listed, or also includes the elements inherent in these processes, methods, articles, or devices / equipment.

[0025] So far, the technical solution of the present invention has been described in combination with the preferred embodiments shown in the drawings. However, it is easy for those skilled in the art to understand that the protection scope of the present invention is obviously not limited to these specific embodiments. Without departing from the principle of the present invention, those skilled in the art can make equivalent changes or substitutions to the relevant technical features, and the technical solutions after these changes or substitutions will fall within the protection scope of the present invention.

Claims

1. A simulation method for the double-vortex encounter effect of a flight simulator, characterized in that, The method includes the following steps: S1. Based on the operating related parameters of the preceding aircraft in the flight simulator training scenario, the wake vortex parameters are obtained through a pre-constructed wake vortex model; S2. According to the stages of the wake dissipation of the preceding aircraft, based on the wake vortex parameters and the dimensionless buoyancy frequency, wake vortex intensity models for each stage are established to obtain the wake vortex intensity values for the corresponding stages; S3. Determine the vortex position; calculate the distance from any point on the following aircraft encountering the wake to the vortex position, and calculate the induced velocity of a single vortex at any point based on this distance; S4. Based on the wake vortex intensity value and the initial circulation of the wake vortex, calculate the ratio of the current wake vortex intensity, and compare it with the critical value to obtain the relative distance between the preceding aircraft and the following aircraft; S5. According to the relative distance between the preceding aircraft and the following aircraft, select different methods to calculate the incremental velocity and incremental angular velocity of the whole aircraft; integrate the incremental velocity and incremental angular velocity respectively to obtain the aircraft state parameters, and the aircraft state parameters include the speed, position, attitude angular velocity and attitude angle of the aircraft.

2. The double-vortex encounter effect simulation method for a flight simulator according to claim 1, wherein The operating related parameters include the mass, wingspan, true airspeed, normal overload and air density of the preceding aircraft; the wake vortex parameters include the initial circulation of the wake vortex, vortex core radius, characteristic time, and wake vortex characteristic velocity; The initial circulation of the wake vortex: ； ； wherein, is the initial circulation of the vortex; is the normal overload of the aircraft; is the mass of the previous aircraft; is the acceleration of gravity; is the air density; is the true airspeed of the previous aircraft; is the initial vortex spacing; Vortex core radius: ； is the radius of the vortex core; is the wingspan of the aircraft; is the initial radius of the vortex core; Wake vortex characteristic velocity: ； Characteristic time: ； Among them, is the characteristic time; is the characteristic velocity of the wake vortex.

3. A simulation method for double-vortex encounter effect of a flight simulator according to claim 1, characterized in that, The stages of the wake dissipation of the preceding aircraft include the diffusion stage and the rapid decay stage; The wake vortex intensity model in the diffusion stage is as follows: ; The wake intensity model in the rapid decay stage is as follows: ; Among them, is the wake vortex intensity value after attenuation in the diffusion stage; is the dimensionless buoyancy frequency; is the wake vortex effect duration; is the diffusion stage duration.

4. A simulation method for double-vortex encounter effect of a flight simulator according to claim 1, characterized in that Determine the vortex position, and the method is as follows: Select a cross-section perpendicular to the wake flow field, and establish a coordinate system with the midpoint of the connection line of the left and right vortex cores as the origin, the x-axis along the connection line direction, and the y-axis perpendicular to the connection line; The left and right vortices are respectively located at the positions of (0, B / 2) and (0, -B / 2) in the coordinate system.

5. A double-vortex encounter effect simulation method for a flight simulator according to claim 1, characterized in that, The method for obtaining the induced velocity of a single vortex at any point is as follows: Based on the positions of the left and right vortices, calculate the distance from any point on the aircraft encountering the wake to the vortex position; ； Calculate the induced velocity of a single vortex at any point according to the distance; ； wherein, is the distance from any point of the subsequent aircraft to the origin; is the intensity of the subsequent aircraft encountering the wake of the preceding aircraft; x and y are the coordinates of any point on the aircraft.

6. A double-vortex encounter effect simulation method for a flight simulator according to claim 1, characterized in that, The method for obtaining the relative distance between the preceding aircraft and the following aircraft is as follows: Calculate the ratio of the current wake vortex intensity: ; Compare the ratio of the current wake vortex intensity with the critical value to obtain the relative distance between the preceding aircraft and the following aircraft: When the D is less than or equal to the critical value, it is determined that the relative distance between the preceding aircraft and the following aircraft is far, otherwise, it is determined that the relative distance between the preceding aircraft and the following aircraft is close.

7. A simulation method for double vortex encounter effect of a flight simulator according to claim 1, characterized in that According to the relative distance between the preceding aircraft and the following aircraft, select different methods to calculate the incremental velocity and incremental angular velocity of the whole aircraft, and the method is as follows: If the relative distance between the preceding aircraft and the following aircraft is close, split the following aircraft into set force components and calculate the induced velocity of each force component, calculate the incremental aerodynamic force based on the induced velocity, and respectively obtain the incremental velocity and incremental angular velocity of the corresponding force component through integration, and superimpose the incremental velocity and incremental angular velocity of each force component to obtain the incremental velocity and incremental angular velocity of the whole aircraft; If the relative distance between the preceding aircraft and the following aircraft is relatively far, an incremental aerodynamic force is calculated based on the induced velocity, and the incremental velocity and incremental angular velocity of the entire aircraft are obtained through the six-degree-of-freedom equation using the incremental aerodynamic force and control parameters; The incremental velocities are u, v, and w; the incremental angular velocities are p, q, and r.

8. A simulation method for double-vortex encounter effect of a flight simulator according to claim 7, characterized in that, The incremental aerodynamic force includes lift, drag, side force, rolling moment, pitching moment, and yawing moment; the calculation methods for the parameters in the incremental aerodynamic force are as follows: Lift: ; Drag: ; Lateral force: ; Rolling moment: ; Pitch moment: ; Course moment: ; Among them, is the induced velocity; is the current air density; is the projected area of the entire aircraft; is the wingspan; is the mean aerodynamic chord of the wing; is the lift coefficient; is the drag coefficient; is the side force coefficient; is the rolling moment coefficient; is the pitching moment coefficient; is the yawing moment coefficient.

9. A simulation method for double-vortex encounter effect of a flight simulator according to claim 7, characterized in that The six-degree-of-freedom equation: ； ； ； ； ； ； Among them, is the total weight of the aircraft; are the incremental velocity change rates along the x, y, and z axes respectively; is the thrust; is the angle of attack; is the sideslip angle; is the pitch angle; is the roll angle; are the moments of inertia about the x, y, and z axes respectively; are the incremental angular velocities about the x, y, and z axes respectively.

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