Kinetic Equation Study on Momentum Conservation of Aerodynamics Incompressible Fluid of High-Altitude Aircraft

This paper introduces an integrated aerodynamic design method for aircraft based on numerical solution of hydrodynamics and electromagnetics equations. By using Pro/Engineer, ANSA, Fluent and MATLAB, the geometric shape, unstructured fluid mesh, hydrodynamic parameters and unpowered flight trajectory data of an aircraft were processed and calculated. In this paper, N-S equations are studied, and a test scheme for a hydrodynamic material aerodynamic model is established. Then, relevant parameters of the aircraft are tested and studied. The feasibility of this method for unsteady aerodynamics analysis of aircraft is verified.


Introduction
Aircraft integrated with inward and outward flow have gradually attracted the attention of many countries in the world because of their advantages such as strong penetration ability, long range and short time.Unlike other conventional aircraft, this new type of aircraft has greater propulsion efficiency, greater carrying capacity, longer launch distances and can save a lot of money [1].Some key links, such as inlet separation, inlet opening/closing and combustion chamber starting/closing, all have the problem of unstable flight state caused by interference torque.The inlet is the core of hypersonic aircraft, and the unsteady change of its flight state will lead to the deviation of the flow pattern in the inlet from the actual situation.Whether it can be effectively opened in this case, how the characteristics and characteristics of its flow field will be changed, and what will be the quality of the air delivered by the backward combustion system are all major scientific problems to be solved urgently in the development of the aircraft.At present, computer simulation technology is one of the aerodynamic characteristics of aircraft.By using this method, the flow structure around the aircraft can be obtained, and the hydraulic parameters of the aircraft can be obtained, so as to achieve the optimal aerodynamic performance of the aircraft.Secondly, the use of computer simulation technology can realize some problems that cannot be realized under the experimental conditions, so as to greatly shorten the design cycle, reduce the design cost, reduce the project risk.The introduction of CFD method into the aerodynamic simulation of aircraft cannot only help people better understand the aerodynamic characteristics of aircraft, but also make up for the shortcomings of traditional wind tunnel test methods, greatly reduce the time and development of aircraft development and reduce the cost of aircraft development.Secondly, it can solve the basic equation.It cannot only provide detailed flow information, but also easily identify, and explore the law of interaction between various physical processes [2].The design of the technical feasibility and the selection of several design schemes.Experiments at this time cost a lot of money and time, and could not be tested on a wide range of people.After analysis, large-scale selection of samples and then experimental selection cannot only ensure the quality of products, but also avoid technical risks.Taking "small sample simulation + Kriging response surface simulation" as the research idea, based on underwater acoustic simulation, combined with dimension theory, and using underwater acoustic simulation technology, the underwater acoustic coupling mathematical model suitable for complex environment is established.A non-reflective model suitable for dynamic boundary is established by using the one-dimensional Riemann-invariant non-reflective property.The application of high precision performance analysis method in optimization design is realized.

Solving method of NS equation on structural grid
For the NS equation on the structural grid:  includes laminar stress and Reynolds stress caused by turbulence.When different turbulence modes are selected, the Reynolds stress is calculated differently.In this paper, we use the algebraic turbulence model of boundary layer.For equation ( 1), we will use Finite Verification (FLASH) algorithm and approximate factorization algorithm to implement numerical simulation of equation ( 1), where Roe flow difference and Central difference will be used for numerical simulation and numerical simulation, respectively.The correctness of the algorithm is proved by an example analysis [3].The basic equations of electromagnetic theory include Faraday's law of electromagnetic induction, Ampere's law, Gauss's law of electric field, The difference formula is: Where , , , H E B D is the field quantity in the electromagnetic field., ,    is the parameter.The calculation is usually to solve the following two curl differential equations: Q is the independent , , F G and H are the flux vectors, depending on the independent variable, , Q S is the source term.
Since Maxwell system and Euler system both belong to hyperbolic PDE, the time-domain onedimensional frequency-shift (FVTD) system of Euler system is introduced into Euler system.FVTD is a conservation equation based on time domain, which can realize the numerical solution of discontinuity, and is suitable for the electromagnetic scattering analysis of complex objects such as multi-medium and multi-medium [4].
In this project, the upwind scheme of MUSCL is adopted, and the initial parameters are extracted to the edge of the grid points according to the position of the grid points.Then, the vector cracking of Steg warming flow is used to obtain the current near the grid points.Then, the Runge-Kutta algorithm is used to integrate the overall electromagnetic field points in time, so as to obtain the spatial distribution of the overall electromagnetic field points [5].The value of electromagnetic field is transformed from time domain to frequency domain by Fourier transform.Then, using the Stratton-Chu integral, the electromagnetic parameters of the nearby region are converted to the electromagnetic parameters of the far field, and the RCS of the object are obtained according to the following formula: Where is the electric and magnetic field intensity of is the intensity of electric and magnetic fields scattering electromagnetic waves.

Basic Parameters
The aircraft is a gliding underwater weapon based on shipborne launch; its overall parameters are as follows: total length: 3080mm; Total width: 2400mm; Machine type: NACA0012; Wing chord length: 350mm; Wing length: 1000mm.The geometric model of an aircraft is shown in Figure 1.

Grid Division
The author uses ANSA meshing tool.The quality of the surface mesh will directly affect the quality of the volume mesh.Before dividing the surface mesh, the geometric model should be cleaned, mainly the processing of acute angles.In nature and engineering, many fluid problems have large Reynolds numbers, such as the external air flow around aircraft or the internal flow of fluid machinery, which belong to large Re flow problems [6].At high Reynolds numbers, viscosity dominates fluid motion only in an extremely thin layer close to the surface, which is called the boundary layer.In the boundary layer, there is a large velocity gradient of the fluid.In order to accurately simulate the wall, it is required that the grid divided into the boundary layer is fine enough.The simulation results and the CPU processing time consumed.If the mesh quality is not good, it is difficult to converge: the mesh is too thin, cannot effectively capture some important information of the flow field, and may even lead to non-convergence.Too dense a grid can also dramatically increase CPU time consumption.The meshing of complex models takes a lot of time and effort.After repeated analysis and experiments, tetrahedral mesh cells were used in the pre-processing, and about 5.08 million unstructured fluid grids were divided into the model.The mesh quality acceptable by Fluent is skewness≤0.98. Figure 2 shows the volume grid divided after many attempts [7].

Case analysis
Figure 3 and 4 show the surface heat flow profile and deviation line obtained by various methods using Z=5 mm, Z=300 mm profile, 10-degree Angle of attack and Z=10-degree profile, respectively.At relatively low flight altitude, the conclusions obtained by the slip method and the non-slip method are similar.However, as the flight altitude increases, the difference between the conclusions obtained by the two methods will become larger and larger.From the quantization results of the error curve in FIG. 4, because the air on the windward side compresses and the air on the leeward side expands, the effect of slip will first occur on the leeward side.Therefore, in the absence of slip, the difference on the leeward side will be significantly smaller than that on the leeward side.Moreover, the rarefied effect in this case becomes more pronounced with increasing altitude, so the error between the heat flow results obtained in the slip and non-slip modes becomes larger and larger.As can be seen from the chart, the maximum deviation is mostly in the upwind expansion zone near the front of the wing.The deviation is about 5 to 15 percent over a wide area of the wing surface from 50 to 80 kilometers.Table 1 shows performance at different flight distances between taxiing mode and non-taxiing mode.With the increase of flight altitude, the coefficient of friction.The error analysis of the two calculation methods shows that the air flow becomes more and more sparse and its influence becomes more and more obvious with the takeoff of the aircraft, while the conventional non-slip NS method cannot accurately reflect the takeoff of the aircraft, and the calculation error of the two methods will become larger and larger [8].
Table 2 shows the relationship between peak heat and altitude obtained by various calculation methods.You can see clearly that at the nose of the aircraft, the lower the atmospheric concentration, the lower the maximum temperature, and at the nose of the aircraft, without slip, because the greater the slip of the strip leads to greater dilution of the air, so under the conventional continuous flow assumption, Calculations without slippage are not applicable.FIG. 5 shows the hysteresis loop for forced trim and heave of the total pressure recovery coefficient around, the pressurized gas volume in the inlet front area increases, the pressurized gas volume in the outer pressure area increases, shock wave interference between multiple wave systems will appear in the pressure area outside the inlet, and along with the thickening of the inlet wall layer, the inlet airflow will be intercepted and the outlet pressure will be restored.

Conclusion
This project intends to take the large aircraft with strong dynamic performance independently developed by China as the object.According to its complex and changeable characteristics under complex air conditions, combined with its special geometry under complex air flow conditions, the dynamic behavior under complex flow conditions will be analyzed, and combined with its practical application in practical engineering applications.A dynamic performance evaluation method for large aircraft is proposed.In view of the complex problems brought by the current high-precision numerical solution method, this project intends to conduct multi-scale solution for multi-parameters with multi-objectives and multi-parameters by combining "small sample solution + Kriging response surface simulation".The feasibility of this project is verified through the study of typical engineering examples.Applying it to engineering practice provides new ideas and ideas for engineering application.

Figure 3 .
Figure 3. Heat flow distribution results at different locations

Table 1 .
Results of aerodynamic characteristics under different methods

Table 2 .
Peak heat flow results under different methods