CN116384290B - Hypersonic aircraft dynamic derivative prediction method considering real gas effect - Google Patents
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Abstract
The invention discloses a hypersonic aircraft dynamic derivative prediction method considering a real gas effect, which comprises the following steps: step 1: generating a first calculation grid according to the incoming flow parameters aiming at the trajectory points of the aircraft; step 2: carrying out flow field steady static aerodynamic characteristic numerical simulation by adopting a thermochemical unbalanced model based on a first calculation grid to obtain flow field parameters of the first trajectory point aircraft; step 3: acquiring first flow field parameter information of a flow field unsteady calculation initial moment based on the first trajectory point aircraft flow field parameters; step 4: based on the first flow field parameter information, carrying out flow field unsteady calculation by adopting a double-time-method to obtain an aircraft dynamic derivative calculation result; the method and the device realize efficient and accurate prediction of the dynamic derivative of the hypersonic aircraft.
Description
Technical Field
The invention relates to the technical field of fluid dynamics, in particular to a hypersonic aircraft dynamic derivative prediction method considering a real gas effect.
Background
In recent years, the high-speed sound velocity technology is rapidly developed, the speed of a near space vehicle is faster and faster, and the flying speed of some gliding aircrafts and gliding warheads at 20 km-30 km is already over Mach 10. Too high a flying speed can cause two problems: (1) Due to the incoming flow density in this height rangeLarger when the flying speed is->When larger, the incoming flow pressure is +.>The attitude adjustment of the aircraft is more difficult, and under the condition of limited control surface control efficiency, the prediction precision of the dynamic characteristics (dynamic derivatives) of the aircraft directly determines the strategy of a control system and the flight safety, so that the higher flying speed needs higher-precision dynamic derivative prediction capability; (2) When the aircraft flies at hypersonic speed, shock waves can appear in the flow field of the head of the aircraft, the temperature in the shock wave layer rises sharply (in the ideal gas case, the temperature is in direct proportion to the square of Mach number of incoming flow), the temperature of the shock wave of the head corresponding to Mach 10 can reach 5000K, under the temperature condition, the air components undergo severe thermochemical reaction, the flow field structure and aerodynamic parameters of the surface of the aircraft are changed, namely the real gas effect, and the dynamic aerodynamic characteristics of the aircraft are seriously influenced.
The traditional method calculates the dynamic derivative by adopting a hydrodynamic three-dimensional N-S equation, and the gas model adopts a complete gas model, namely the air component is unchanged, and no thermochemical reaction occurs. When the speed of the aircraft is too high, the method obviously ignores the influence of the real gas effect, and the prediction of the dynamic derivative is inaccurate, so that great hidden danger is possibly brought to flight safety.
When the actual gas effect is considered, the gas model needs to adopt a thermochemical unbalanced model, and compared with a complete gas model, the most obvious difference is that the basic equation is more than the component equation, and the state equation is different. Wherein the added component equations will increase the computational effort, resulting in a decrease in computational efficiency; meanwhile, the rigidity problem of the coefficient matrix in the numerical calculation process is also brought, so that the calculation stability and the robustness are reduced. Taking an air 5-component 11 reaction model as an example, 4 component equations are added to 5 components, the basic equations are increased from 5 to 9, the numerical calculation discrete equation coefficient matrix is changed from 5*5 to 9*9, and the calculated amount is increased by about 4 times. Meanwhile, a double-time step method is adopted in unsteady flow field calculation, when the gas model is completely calculated, the iteration steps of a aerodynamic force calculation sub-step number is generally 30 steps, and when the thermochemical unbalanced model is calculated, the aerodynamic force convergence is slowed down due to the problem of coefficient matrix rigidity caused by an added component equation, and the sub-iteration steps number is generally increased to 100 steps. Therefore, the calculation amount of the dynamic derivative calculation considering the actual gas effect is increased by 4*3 =12 times compared to the conventional dynamic derivative calculation. And because of the rigidity problem, the stability of calculation is poor, the quality requirement on the calculation grid is high, so that the dynamic aerodynamic characteristics of the aircraft can be obtained in 2-3 days by the traditional method, and a period of 1 month is required after the actual gas effect is considered, which seriously affects the iteration of the design of the appearance optimization and the design of the control system of the aircraft, and restricts the overall aerodynamic design of the aircraft.
Disclosure of Invention
The aim of the invention is to achieve an efficient and accurate prediction of the dynamic derivative of hypersonic aircraft.
To achieve the above object, the present invention provides a hypersonic aircraft dynamic derivative prediction method considering a true gas effect, the method comprising:
step 1: generating a first calculation grid according to the incoming flow parameters aiming at the trajectory points of the aircraft;
step 2: carrying out flow field steady static aerodynamic characteristic numerical simulation by adopting a thermochemical unbalanced model based on a first calculation grid to obtain flow field parameters of the first trajectory point aircraft;
step 3: acquiring first flow field parameter information of a flow field unsteady calculation initial moment based on the first trajectory point aircraft flow field parameters;
step 4: based on the first flow field parameter information, adopting a double-time-step method to start flow field unsteady calculation to obtain an aircraft dynamic derivative calculation result; the time step iteration in the double-time-step method calculation meets any one of the following two convergence conditions:
convergence condition 1: the average residual error of the sub-iteration is reduced by two orders of magnitude or the number of sub-iteration steps reaches a preset value;
convergence condition 2: when the residual ratioLess than or equal to the set point.
The calculation of the dynamic derivative in the prior art does not take the actual gas effect into consideration, so that the calculation accuracy is low, and because the calculation uses a thermochemical unbalanced model to take the actual gas effect into consideration, the calculation result accuracy is high, and the calculation needs to use a chemical unbalanced model to take the actual gas effect into consideration, but after using the model, a plurality of calculation problems such as stability problems and efficiency problems are brought, so that the calculation can be quickly converged, the calculation efficiency problem is solved, and the stability problem is solved by using a steady flow field to optimize.
Furthermore, the method obtains the first flow field parameter information of the flow field unsteady calculation initial moment based on the flow field parameters of the first trajectory point aircraft, so that the initial flow field optimization is realized, the problems of equation rigidity and calculation stability caused by overlarge flow field parameter gradient are effectively solved, the double-time-step implicit solution sub-iteration convergence criterion is improved, and the calculation efficiency of the method is improved.
Preferably, step 2 in the present invention specifically includes: based on the first calculation grid, a 5-component 11-reaction Park thermochemical unbalanced model and a Navier-Stokes equation are adopted to perform numerical simulation of flow field steady static aerodynamic characteristics, so as to obtain flow field parameters of the first ballistic point aircraft.
Preferably, the residual ratio in the convergence condition 2 in the present inventionThe calculation mode of (a) is as follows:
；
；
；
wherein ,represents +.1 in iteration step>Is>Representing the nth iteration stepIs>For a given constant +.>Implicit solution obtained for the nth sub-iteration step,/->And (5) obtaining an implicit solving result for the n+1th sub-iteration step number.
Preferably, step 1 of the present invention further includes: the aircraft motion pattern is set to be forced pitch oscillation.
Preferably, in the present invention, the movement mode of the aircraft in step 1 is forced pitch oscillation, and the form of the oscillation is defined as:, wherein ,a0 For the central angle of attack of the oscillation, A m The amplitude of oscillation, w is the circular frequency of oscillation, t is the real time, and a is the attack angle corresponding to the moment t.
Preferably, the first ballistic point aircraft flow field parameters in the present invention include: the number of the grid block where the grid point is located, the position number of the grid point in three-direction dimensions of the grid blocks i, j and k, the actual coordinate position (x, y, z) of the grid point under the calculated coordinate system, and the air density on the grid pointAir velocity in x-direction on grid points +.>Velocity of air in y-direction on grid pointsAir velocity in z-direction at grid points>Air pressure on grid points->Component mass fraction of air at grid pointsDistribution, h represents the h-th component in air.
Preferably, in the present inventionThe following formula was used for calculation:
； wherein ,/>，Air density for grid point for iteration step n+1, +.>Air density for the nth iteration step grid point;，/>for the n+1th iteration step grid air velocity in x-direction, +.>The velocity of air in the x direction on the grid point of the nth iteration step; />，/>For the n+1th iteration step grid air velocity in y-direction, +.>The velocity of air in the y direction on the grid point of the nth iteration step; />，/>For the n+1th iteration step grid air velocity in z direction, +.>The velocity of air in the z direction on the grid point of the nth iteration step;，/>air pressure on grid point for n+1th iteration step, +.>Air pressure at the nth iteration grid point; />，/>Component mass fraction of air on grid point for n+1 iteration step +.>Distribution of->Component mass fraction of air at grid point for nth iteration step +.>Distribution.
Preferably, in the present invention, the first calculation grid is a structural grid.
Preferably, in the present invention, the preset value is 100 and the set value is 0.01.
The one or more technical schemes provided by the invention have at least the following technical effects or advantages:
according to the invention, a thermochemical unbalanced model is introduced into the unsteady flow field numerical solution, and the actual gas effect is considered, so that the numerical simulation is more similar to the actual flow environment of the hypersonic aircraft, and therefore, the calculation precision of the dynamic derivative is higher. By reasonably optimizing the unsteady calculation initial field, the calculated amount is reduced, and the calculation stability is increased; for a thermochemical unbalanced model, as a component equation is introduced, the convergence speed of the component is slower than the speed and the pressure, so that the calculation efficiency is reduced due to the increase of the number of sub-iteration steps under the traditional convergence criterion.
Drawings
The accompanying drawings, which are included to provide a further understanding of embodiments of the invention and are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and together with the description serve to explain the principles of the invention;
FIG. 1 is a flow chart of a hypersonic aircraft dynamic derivative prediction method taking into account the true gas effects in the present invention;
figure 2 is a schematic view of HBS profile;
FIG. 3 is a schematic view of an HBS computational grid;
FIG. 4 is a schematic diagram of the HBS steady calculation flow field shock wave structure;
FIG. 5 is a diagram showing the comparison of the calculated time for two convergence criteria;
FIG. 6 is a schematic diagram of the effect of real gas effects on dynamic derivatives.
Detailed Description
In order that the above-recited objects, features and advantages of the present invention will be more clearly understood, a more particular description of the invention will be rendered by reference to the appended drawings and appended detailed description. In addition, the embodiments of the present invention and the features in the embodiments may be combined with each other without collision.
In the following description, numerous specific details are set forth in order to provide a thorough understanding of the present invention, however, the present invention may be practiced in other ways than within the scope of the description, and the scope of the invention is therefore not limited to the specific embodiments disclosed below.
Example 1
Referring to fig. 1, fig. 1 is a flowchart of a hypersonic aircraft dynamic derivative prediction method considering real gas effects, the method comprising:
step 1: generating a first calculation grid according to the incoming flow parameters aiming at the trajectory points of the aircraft;
step 2: carrying out flow field steady static aerodynamic characteristic numerical simulation by adopting a thermochemical unbalanced model based on a first calculation grid to obtain flow field parameters of the first trajectory point aircraft;
step 3: acquiring first flow field parameter information of a flow field unsteady calculation initial moment based on the first trajectory point aircraft flow field parameters;
step 4: based on the first flow field parameter information, adopting a double-time-step method to start flow field unsteady calculation to obtain an aircraft dynamic derivative calculation result; the time step iteration in the double-time-step method calculation meets any one of the following two convergence conditions:
convergence condition 1: the average residual error of the sub-iteration is reduced by two orders of magnitude or the number of sub-iteration steps reaches a preset value;
convergence condition 2: when the residual ratioLess than or equal to the set point.
The method can more accurately and rapidly realize high-speed ultrasonic quick derivative prediction, and effectively relieves the problems of equation rigidity and calculation stability caused by overlarge flow field parameter gradient by optimizing an initial flow field, thereby obtaining the hypersonic speed aircraft dynamic derivative rapid prediction method considering the real gas effect.
The invention provides a new method for solving an unsteady flow field based on a thermochemical unbalanced model, improving an unsteady calculation double-time step iteration convergence criterion by optimizing the initial value of the aerodynamic parameters of the unsteady flow field, and realizing the rapid prediction of a high-speed ultrasonic quick derivative, which comprises the following specific implementation steps:
aiming at typical trajectory points of an aircraft, according to parameters such as Reynolds number, attack angle and the like of an incoming flow unit, grid generating software such as pointwise, gridgen is adopted to generate a set of multi-block structural grids which simultaneously meet the requirements of calculation precision and calculation efficiency. Based on the calculation grid, a 5-component 11-reaction Park thermochemical unbalanced model and a Navier-Stokes equation are adopted to carry out static aerodynamic characteristic numerical simulation of a steady flow field, and flow field parameters of the ballistic point aircraft are obtained, wherein the parameters comprise two parts of data: one part is information describing the position of a grid point in space around the aircraft, and the other part is aerodynamic aero-characteristics on the grid pointSex parameters. The information describing the positions of the grid points in space around the aircraft comprises: the number of the structural grid block where the grid point is located, the position number of the grid point in the i/j/k direction of three direction dimensions of the grid block, the actual coordinate position (x, y, z) of the grid point under the calculation coordinate system, and the aerodynamic characteristic parameters on the grid point comprise: air density at the grid pointsAir at the grid points is at a speed in the x-direction>Air velocity in y-direction on the grid points>The air at the grid points is at a speed in the z-direction>Air pressure on the grid points->Component mass fraction of air at the grid points +.>Distribution (subscript h represents the h-th component in air). The thermochemical unbalanced model can be used for calculating static aerodynamic characteristics of a steady flow field by referring to the following documents: liu Qingzong, dong Weizhong, ding Mingsong, et al numerical simulation study of the aerodynamic and thermal environment of Mars probe [ J ]]Aerodynamic journal, 2018, 36 (4): 642-650.
The aerodynamic parameters of the flow field obtained by the method comprise the following information: grid block number, grid number information (i, j, k), grid coordinate information (x, y, z), air densityAir speed (U, V, W), air pressure +.>Air compositionScore of quantity->And assigning the information to a flow field unsteady calculation initial flow field to serve as a flow field parameter value of the flow field unsteady calculation initial time.
The unsteady flow field calculation generally adopts a double-time method, wherein two time steps are respectively that the real time is discretized into a plurality of time periods, namely the real physical time stepTheir set describes the history of the evolution of unsteady flows over real time; virtual time step +.>I.e. every real time step +.>In (i.e. sub-iteration in the two-step method, consider this moment as a steady flow field), the time advance course virtualized for solving the real flow characteristic at this moment is considered to be the real time +.>The flow field at this point. The core for determining the solving efficiency of the double-time-method is a sub-iteration solving method and a convergence criterion. In general, there are two options for the sub-iterative convergence criteria: (1) given a number of sub-iteration steps; (2) the flow field parameter average residual drops by two orders of magnitude. For a complete gas model, the sub-iteration step number 30 can meet the requirement that the average residual error is reduced by two orders of magnitude, while for a thermochemical unbalanced model, the convergence speed of the components is slower than the speed and the pressure due to the introduction of a component equation, and the iteration step number is increased in multiple. The following sub-iteration convergence criteria are therefore proposed: defining a new flow field parameter vector +.>It describes the aerodynamic characteristic information set of the flow field on a certain grid pointWherein->For the air density at grid points +.>For the air x-direction velocity on grid points, +.>For the velocity of air y direction on grid points, +.>For the air z-direction velocity on grid points, +.>For the pressure on the grid points,the mass fraction of air component on the grid points. The result obtained by the double-time step iteration in the implicit solving process is the difference of aerodynamic characteristic parameters of the flow field of the current virtual time and the flow field of the next virtual time, namely the implicit solving result obtained by the nth sub-iteration step number is +.>, wherein />Subscript n+1 represents the n+1th iteration step, subscript n represents the n-th iteration step,/->Representing the grid point air density for the n+1 iteration step, and so on. The implicit solving process of the two-time step iteration can be described as: to->As an input condition (flow field aerodynamic characteristic parameter of the nth sub-iteration step number), output +.>And then obtain(flow field aerodynamic characteristic parameter of n+1th sub-iteration step), then with +.>Output +.>And performing cyclic calculation according to the method.
The invention provides a new double-time-step iteration convergence termination condition suitable for considering a real gas effect component equation: definition of the definition, wherein ,/>Represents +.>Selecting the residual ratio as the convergence termination condition: />, wherein ,/>Represents +.1 in iteration step>Is>For a given constant, the value range is +.>When->When the value is smaller than a certain given value (0.01 is recommended), the sub-iteration is terminated; specific solving processes of the two-time method can be referred to as: zhao Huiyong, le Jialing double time stepApplication analysis of the method [ J]Calculated physical Vol.25, no.3.2008.
In order to obtain the dynamic derivative of the aircraft, the aircraft is required to move in a specific way, then the aerodynamic characteristics of the aircraft in the movement process are obtained by adopting unsteady calculation, and finally the dynamic derivative of the aircraft is identified. There are various ways to move the aircraft, the invention adopts a forced vibration method to identify the dynamic derivative, and the vibration form can be defined as:, wherein a0 For the central angle of attack of the oscillation, A m The amplitude of oscillation, w is the circular frequency of oscillation, t is the real time, and a is the attack angle corresponding to the moment t. The initial moment (t=0 moment) of unsteady calculation is initialized by the parameters of the steady flow field calculation, and the parameters comprise: grid block number, grid number information (i, j, k), grid coordinate information (x, y, z), air density +.>Three directional speeds (U, V, W) of air, air pressure +.>Mass fraction of air component>. The forced vibration method and the dynamic derivative identification method can be referred to as follows: yuan Xianxu, chen Qi, xiefei, chen Jianjiang. Related problems in dynamic derivative numerical prediction. Aeronautical journal 2016,37 (08); liu Xu hypersonic inside and outside flow Integrated aircraft dynamic Property research. National defense science and technology university. Shuoshi paper 2011.
Specific examples of the calculation of the dynamic derivative by the ballistic profile (HBS) unsteady flow field simulation are given below.
The HBS profile is shown in fig. 2, and the incoming flow condition is set as: flight altitude 25Km, flight Mach number 13, and flight initial angle of attack 4 °. The motion mode of the aircraft is forced pitching oscillation and the motion rule: the angle of attack is a function of time,，α 0 for an initial angle of attack of 4 DEG alpha m The amplitude is 0.5 DEG, the dimensionless oscillation frequency k is 0.2, and t is the real time of the motion. The computational grid is schematically shown in fig. 3, using a structural grid, a full field grid scale of 80 tens of thousands.
Firstly, adopting steady calculation and utilizing incoming flow conditions: the flight height is 25Km, the flight Mach number is 13, the initial attack angle is 4 degrees, and the initial attack angle state chemical unbalanced steady flow field is obtained, and the flow field structure is schematically shown in figure 4. Chemical non-equilibrium stationary flow field calculation reference: liu Qingzong, dong Weizhong, ding Mingsong numerical simulation of the catalytic properties of Mars detector surface materials [ J ]. Astronautics report Vol.39, no. 8.2018.
Then, the unsteady calculation is initialized by taking the unsteady flow field as the initial field, and the unsteady calculation is started. The specific steps are that the grid block number, grid number information (i, j, k), grid coordinate information (x, y, z) and air density of a steady flow field are addedAir speed (U, V, W), air pressure +.>Mass fraction of air component>These information are assigned to the 1 st iteration step at the time t=0 of the unsteady calculation as the initial value of the flow field calculation.
Meanwhile, a double-time-step method is adopted to start unsteady calculation, and two convergence criteria are adopted for time step iteration: (1) The convergence criterion 1 is a conventional method, the average residual error of the sub-iteration is reduced by two orders of magnitude or the number of sub-iteration steps reaches 100; (2) The convergence criterion 2 is the method after optimization herein,, wherein />Is common to experienceThe number (which is taken according to the extent and the range of influence of the actual gas effect in the actual problem), is generally taken +.>。
In the numerical calculation process, the convergence speed of the three quantities of pressure, density and speed is high, the convergence speed of the components is low, and the number of steps used for converging the components is about 3-5 times of the pressure density speed.
In the initial stage of sub-iteration, the pressure, density and speed in the flow field are greatly changed,the magnitude of these amounts is large, +.>The value of (2) depends entirely on these quantities. Sub-iterations continue with rapid convergence of pressure, density, speed, +.>The magnitude of these amounts will decrease rapidly, the component change valueAt->The contribution of (3) is becoming larger and larger. Through the structureSuch a convergence criterion, in effect, relaxes the convergence conditions of the components to a degree that is defined by the empirical constant +.>Given (I)>The smaller the relaxation degree the greater. When the flow field parameters meet the convergence condition, the precision of the flow field parameters can meet the requirement of accurate calculation of the dynamic derivative, and the calculation efficiency can be ensured. FIG. 5 shows the computational efficiency of two convergence criteriaBy comparing, the time for calculating the CPU given by the abscissa can be seen, the time for calculating the convergence criterion after optimization is obviously reduced.
Finally, for the initial attack angle of 4 degrees, 5 degrees, 6 degrees and 7 degrees, the pitching moment dynamic derivative of the HBS under the conditions of complete gas and real gas is calculated and obtained by adopting the same method, as shown in fig. 6, the influence of the real gas effect on the dynamic derivative is larger, and the effect reaches 60%. The calculation of the dynamic derivative under the complete gas condition belongs to the calculation of a conventional method, and the specific method can be referred to in the literature: liu Xu hypersonic inside and outside flow Integrated aircraft dynamic Property research. National defense science and technology university. Shuoshi paper 2011.
While preferred embodiments of the present invention have been described, additional variations and modifications in those embodiments may occur to those skilled in the art once they learn of the basic inventive concepts. It is therefore intended that the following claims be interpreted as including the preferred embodiments and all such alterations and modifications as fall within the scope of the invention.
It will be apparent to those skilled in the art that various modifications and variations can be made to the present invention without departing from the spirit or scope of the invention. Thus, it is intended that the present invention also include such modifications and alterations insofar as they come within the scope of the appended claims or the equivalents thereof.
Claims (9)
1. A hypersonic aircraft dynamic derivative prediction method considering real gas effects, the method comprising:
step 1: generating a first calculation grid according to the incoming flow parameters aiming at the trajectory points of the aircraft;
step 2: carrying out flow field steady static aerodynamic characteristic numerical simulation by adopting a thermochemical unbalanced model based on a first calculation grid to obtain flow field parameters of the first trajectory point aircraft;
step 3: acquiring first flow field parameter information of a flow field unsteady calculation initial moment based on the first trajectory point aircraft flow field parameters;
step 4: based on the first flow field parameter information, adopting a double-time-step method to start flow field unsteady calculation to obtain an aircraft dynamic derivative calculation result; the time step iteration in the double-time-step method calculation meets any one of the following two convergence conditions:
convergence condition 1: the average residual error of the sub-iteration is reduced by two orders of magnitude or the number of sub-iteration steps reaches a preset value;
convergence condition 2: when the residual ratioLess than or equal to a set value;
residual ratio in Convergence Condition 2The calculation mode of (a) is as follows:
；/>；/>； wherein ,represents +.1 in iteration step>Is>Represents +.>Is>For a given constant +.>Implicit solution obtained for the nth sub-iteration step,/->And (5) obtaining an implicit solving result for the n+1th sub-iteration step number.
2. The hypersonic aircraft dynamic derivative prediction method considering the effect of real gas according to claim 1, wherein the step 2 specifically comprises: based on the first calculation grid, a 5-component 11-reaction Park thermochemical unbalanced model and a Navier-Stokes equation are adopted to perform numerical simulation of flow field steady static aerodynamic characteristics, so as to obtain flow field parameters of the first ballistic point aircraft.
3. A hypersonic aircraft dynamic derivative prediction method taking into account real gas effects as set forth in claim 1, wherein said step 1 further comprises: the aircraft motion pattern is set to be forced pitch oscillation.
4. A hypersonic aircraft dynamic derivative prediction method taking into account the effect of real gases as claimed in claim 3 wherein in step 1 the mode of motion of the aircraft is forced pitch oscillation, the form of oscillation being defined as:, wherein ,a0 For the central angle of attack of the oscillation, A m The amplitude of oscillation, w is the circular frequency of oscillation, t is the real time, and a is the attack angle corresponding to the moment t.
5. A hypersonic aircraft dynamic derivative prediction method considering real gas effects as claimed in claim 1 wherein the first ballistic point aircraft flow field parameters include: the number of the grid block where the grid point is located, the position number of the grid point in three-direction dimensions of the grid blocks i, j and k, the actual coordinate position (x, y, z) of the grid point under the calculated coordinate system, and the air density on the grid pointAir velocity in x-direction on grid points +.>Air velocity in y-direction on grid points +.>Air velocity in z-direction at grid points>Air pressure on grid points->Component mass fraction of air at grid points +.>Distribution, h represents the h-th component in air.
6. A hypersonic aircraft dynamic derivative prediction method taking into account real gas effects as set forth in claim 1, wherein,the following formula was used for calculation:
； wherein ,/>，/>Air density for grid point for iteration step n+1, +.>Is the nthIterating the air density of the step grid points; />，For the n+1th iteration step grid air velocity in x-direction, +.>The velocity of air in the x direction on the grid point of the nth iteration step; />，/>For the n+1th iteration step grid air velocity in y-direction, +.>The velocity of air in the y direction on the grid point of the nth iteration step; />，/>For the n+1th iteration step grid air velocity in z direction, +.>The velocity of air in the z direction on the grid point of the nth iteration step; />，/>Air pressure on grid point for n+1th iteration step, +.>For the nth iteration netAir pressure at the grid points; />，Component mass fraction of air on grid point for n+1 iteration step +.>Distribution of->Component mass fraction of air at grid point for nth iteration step +.>Distribution.
7. A hypersonic aircraft dynamic derivative prediction method taking into account the effect of real gases as claimed in claim 1 wherein the first computational grid is a structural grid.
8. The method for predicting the dynamic derivative of a hypersonic aircraft taking into account the effect of real gases according to claim 1 wherein the preset value is 100 and the set value is 0.01.
9. A hypersonic aircraft dynamic derivative prediction method taking into account real gas effects as claimed in claim 1 wherein the incoming flow parameters include: fly altitude, fly Mach number, and fly initial angle of attack.
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