Analysis method and device for aerodynamic characteristics of remote tail control artillery shell

Abstract

The application discloses a kind of remote tail control shell aerodynamic characteristic analysis method and device. Among them, the method includes: the physical modeling of the shell body of shell is carried out, and the shell model is obtained;The shell model is divided into unstructured grid and boundary conditions are set;Based on the grid and the set boundary conditions, the shell model is analyzed, and the variation law of the aerodynamic characteristics of the shell model with flight Mach number and angle of attack is obtained.The application solves the technical problem that the control difficulty caused by high rotation speed of high-rotation shell.

Description

Technical Field

[0001] This invention relates to the field of high dynamic navigation technology, and more specifically, to a method and apparatus for analyzing the aerodynamic characteristics of a long-range tail-controlled guided projectile. Background Technology

[0002] In modern warfare, artillery shells have become a trump card, playing a crucial role. Conventional guided artillery shells maintain a high rotational speed during flight, and how to reduce this speed and improve accuracy is a current hot research topic.

[0003] In existing technologies, the dynamic characteristics of the terminal trajectory of a controlled projectile parachute system have been studied using multi-rigid-body dynamics methods, ensuring that the terminal trajectory meets design requirements. Anti-spin blades can reduce the projectile's spin to an ideal value. Another existing technology proposes a stable rolling motion region and anti-spin strategy suitable for electromagnetically launched hypersonic guided projectiles. Active spin initiation and anti-spin control strategies provide a novel and feasible method for the flight control of electromagnetically launched hypersonic guided projectiles. However, these existing technologies still suffer from the problem of high control difficulty due to the high rotational speed of the hypersonic guided projectile.

[0004] There is currently no effective solution to the above problems. Summary of the Invention

[0005] This invention provides a method and apparatus for analyzing the aerodynamic characteristics of a long-range tail-controlled guided projectile, in order to at least solve the technical problem of high control difficulty caused by the high rotational speed of high-spin guided projectiles.

[0006] According to one aspect of the present invention, a method for analyzing the aerodynamic characteristics of a long-range tail-guided projectile is provided, comprising: physically modeling the projectile body to obtain a projectile model; dividing the projectile model into an unstructured mesh and setting boundary conditions; and performing aerodynamic characteristic analysis on the projectile model based on the divided mesh and the set boundary conditions to obtain the variation law of the aerodynamic characteristics of the projectile model with flight Mach number and angle of attack.

[0007] According to one aspect of the present invention, an analysis device for the aerodynamic characteristics of a long-range tail-controlled guided projectile is provided, comprising: a model building module for physically modeling the projectile body to obtain a projectile model; a setting module for performing unstructured mesh generation on the projectile model and setting boundary conditions; and an analysis module for performing aerodynamic characteristic analysis on the projectile model based on the generated mesh and the set boundary conditions to obtain the variation law of the aerodynamic characteristics of the projectile model with flight Mach number and angle of attack.

[0008] In this embodiment of the invention, the projectile body is physically modeled to obtain a projectile model; the projectile model is divided into unstructured meshes and boundary conditions are set; based on the divided meshes and the set boundary conditions, the aerodynamic characteristics of the projectile model are analyzed to obtain the variation law of the aerodynamic characteristics of the projectile model with flight Mach number and angle of attack, thereby solving the technical problem of high control difficulty caused by the high rotational speed of high-spin guided projectiles. Attached Figure Description

[0009] The accompanying drawings, which are included to provide a further understanding of the invention and form part of this application, illustrate exemplary embodiments of the invention and, together with their description, serve to explain the invention and do not constitute an undue limitation thereof. In the drawings:

[0010] Figure 1 This is a schematic diagram of a physical projectile model according to an embodiment of the present invention;

[0011] Figure 2 This is a flowchart of an analysis method for the aerodynamic characteristics of a long-range tail-controlled guided projectile according to an embodiment of the present invention;

[0012] Figure 3 This is a flowchart illustrating another method for analyzing the aerodynamic characteristics of a long-range tail-controlled guided projectile according to an embodiment of the present invention.

[0013] Figure 4 This is a schematic diagram of the computational domain and mesh partitioning according to an embodiment of the present invention;

[0014] Figure 5 This is a schematic diagram of the grid division around the projectile according to an embodiment of the present invention;

[0015] Figure 6 This is a flowchart of another method for analyzing the aerodynamic characteristics of a long-range tail-controlled guided projectile according to an embodiment of this application;

[0016] Figure 7 This is a graph showing the change of the drag coefficient as a function of the iterative process according to an embodiment of the present invention;

[0017] Figure 8 This is a graph showing the change of the lift coefficient with the iterative process according to an embodiment of the present invention;

[0018] Figure 9 This is a graph showing the change of the torque coefficient with the iterative process according to an embodiment of the present invention;

[0019] Figure 10A , 10B 10C and 10D are velocity distribution contour maps in the X direction at 1.5 Ma, 2 Ma, 2.5 Ma, and 3 Ma, respectively, at an angle of attack of 4°.

[0020] Figure 11A , 11B The 11C and 11D are surface pressure distribution cloud maps of the entire projectile at 1.5 Ma, 2 Ma, 2.5 Ma, and 3.0 Ma, respectively, at an angle of attack of 4°.

[0021] Figure 12 This is a graph showing the drag variation with Mach number and angle of attack according to an embodiment of the present invention;

[0022] Figure 13 This is a graph showing the change of lift with Mach number and angle of attack according to an embodiment of the present invention;

[0023] Figure 14 This is a graph showing the change of torque with Mach number and angle of attack according to an embodiment of the present invention;

[0024] Figure 15 This is a schematic diagram of the structure of an analysis device for the aerodynamic characteristics of a long-range tail-controlled guided projectile according to an embodiment of the present invention. Detailed Implementation

[0025] To enable those skilled in the art to better understand the present invention, the technical solutions of the present invention will be clearly and completely described below with reference to the accompanying drawings of the embodiments. Obviously, the described embodiments are merely some, not all, of the embodiments of the present invention. All other embodiments obtained by those skilled in the art based on the embodiments of the present invention without creative effort should fall within the scope of protection of the present invention.

[0026] It should be noted that the terms "first," "second," etc., in the specification, claims, and accompanying drawings of this invention are used to distinguish similar objects and are not necessarily used to describe a specific order or sequence. It should be understood that such data can be interchanged where appropriate so that the embodiments of the invention described herein can be implemented in orders other than those illustrated or described herein. Furthermore, the terms "comprising" and "having," and any variations thereof, are intended to cover a non-exclusive inclusion; for example, a process, method, system, product, or apparatus that comprises a series of steps or units is not necessarily limited to those steps or units explicitly listed, but may include other steps or units not explicitly listed or inherent to such processes, methods, products, or apparatus.

[0027] Overview

[0028] This application addresses the problem of high spin-reduction guided projectiles being difficult to control due to their high rotation speed by providing a guided projectile shape that utilizes a stern device to reduce projectile spin. The spin-reduction stern is non-rigidly connected to the forward section via a bearing device, which ensures that the projectile achieves the spin-reduction effect after leaving the barrel.

[0029] Furthermore, this application employs 3D software for physical modeling of the projectile, uses GAMBIT for unstructured mesh generation, and utilizes FLUENT software to simulate the aerodynamic characteristics of a guided projectile with a non-rigidly connected anti-spin stern based on the Spalart-Allmaras model. The simulation yields the variations in lift coefficient, drag coefficient, and moment coefficient of the anti-spin guided projectile with Mach number and angle of attack, and analyzes the results. The simulation results demonstrate that the aerodynamic shape of the guided projectile with an anti-spin stern is reasonable, and the obtained aerodynamic parameters can provide a design basis for anti-spin guided projectiles.

[0030] Example

[0031] This application provides a physical projectile model. When creating the projectile model, it is first necessary to use 3D modeling software such as UG and SolidWorks to model the projectile, and then import the completed projectile model into GAMBIT.

[0032] The shell model provided in this application embodiment is as follows: Figure 1 As shown, it consists of a warhead 11, a warhead 12, an engine 13, a sealing ring 14, a stern deflector 15, and a snorkel cover 16.

[0033] The projectile model in this application utilizes a stern device to reduce the spin of the guided projectile. This spin-reducing stern employs a non-rigid connection between the bearing and the stern and front end of the projectile. The friction on the bearing drives the stern to rotate, ensuring that the projectile achieves a spin-reducing effect after leaving the barrel. This reduces the projectile's rotational speed, achieving a controllable outcome.

[0034] After establishing the projectile model, its aerodynamic characteristics are analyzed. This application provides a method for analyzing the aerodynamic characteristics of a long-range tail-controlled guided projectile, such as... Figure 2 As shown, the method includes the following steps:

[0035] Step S202: Perform physical modeling on the projectile body to obtain the projectile model;

[0036] In an exemplary embodiment, a forward section of the projectile model is constructed; then, a spin-reducing stern of the projectile model is constructed. The spin-reducing stern is non-rigidly connected to the forward section via a bearing device. The frictional force on the bearing device can drive the spin-reducing stern to rotate, so that the projectile body can reduce rotation after leaving the barrel.

[0037] Step S204: Perform unstructured mesh generation on the projectile model and set boundary conditions;

[0038] In an exemplary embodiment, the computational domain of the projectile model is modeled and meshed, wherein the mesh is denser closer to the projectile body and relatively sparser farther away from the projectile body, and the mesh is densest in the projectile head portion; when setting boundary conditions, a density-based coupled explicit solver is selected for solving, and far-field boundary conditions and viscous boundary conditions are selected for the incoming flow and the object surface, respectively.

[0039] In an exemplary embodiment, before performing unstructured mesh generation on the projectile model, the method further includes: determining the drag coefficient expression and lift coefficient expression of the projectile model using engineering calculation methods; determining the continuity equation and transport equation of the projectile model, and solving the energy equation of the projectile model; and determining the turbulence model of the projectile model.

[0040] Step S206: Based on the divided grid and the set boundary conditions, perform aerodynamic characteristic analysis on the projectile model to obtain the variation law of the aerodynamic characteristics of the projectile model with flight Mach number and angle of attack.

[0041] The drag coefficient, lift coefficient, and moment coefficient were calculated at different angles of attack and Mach numbers for dynamic monitoring window analysis. The variation laws of the lift coefficient, drag coefficient, and moment coefficient of the projectile model with flight Mach number and angle of attack were obtained: at a constant angle of attack, the drag coefficient increases continuously with increasing Mach number; at a constant angle of attack, the lift coefficient increases continuously with increasing Mach number; at a constant angle of attack, the moment coefficient increases continuously with increasing Mach number.

[0042] The X-direction velocity distribution cloud map and the projectile surface pressure distribution cloud map were calculated at different angles of attack and Mach numbers to perform a window analysis of airflow around the projectile. The variation law of the shock wave and projectile surface pressure of the projectile model with flight Mach number and angle of attack was obtained: at a constant angle of attack, the shock wave gradually decreases as the Mach number increases; at a constant angle of attack, the projectile surface pressure gradually increases as the Mach number increases, with the projectile head bearing the greatest pressure.

[0043] The drag, lift, and moment of the projectile model at final convergence are calculated to perform a force analysis on the projectile model, revealing the variation of the forces on the projectile model with flight Mach number and angle of attack:

[0044] 1) At a constant angle of attack, as the Mach number increases, the drag also increases continuously, and the rate of increase gradually changes from rapid to slow, but eventually shows an upward trend; at a constant Mach number, as the angle of attack increases, the drag first decreases and then increases to a peak value, and then decreases, but overall shows an upward trend.

[0045] 2) At a constant angle of attack, lift increases with increasing Mach number. At 0° angle of attack, the rate of climb is small, but the lift coefficient increases with increasing Mach number at other angles of attack. The larger the angle of attack, the faster the lift coefficient increases.

[0046] 3) When the angle of attack is constant, the torque coefficient increases with the increase of Mach number, and the difference in the increment of torque coefficient also gradually increases with the increase of angle of attack. The larger the angle of attack, the faster the increase of lift coefficient. When the Mach number is constant, the torque coefficient gradually increases with the increase of angle of attack, and the rate of increase gradually slows down.

[0047] After obtaining the variation law of the aerodynamic characteristics of the projectile model with flight Mach number and angle of attack, the method further includes: manufacturing a spin-reducing guided projectile based on the variation law.

[0048] This application provides another method for analyzing the aerodynamic characteristics of long-range tail-controlled guided projectiles, such as... Figure 3 As shown, the method includes the following steps:

[0049] Step S302: Determine the drag coefficient and lift coefficient of the projectile using engineering calculation methods.

[0050] Generally, engineering calculations are performed based on aerodynamic theory, with appropriate assumptions and simplifications made regarding the projectile structure and flow field, followed by final solution calculations. External ballistics research methods must meet three requirements: high calculation speed, simplicity, and high accuracy. Therefore, engineering calculation methods are widely used in external ballistics research. Compared to external shape analysis, engineering calculation methods can control the error between the calculated results and normal values ​​of coefficients such as drag, lift, and rollover moment to within 10%, or even less. Therefore, engineering calculation methods are more suitable for understanding the impact of changes in projectile structure on aerodynamic characteristics, and can obtain information such as the projectile's aerodynamic parameters through calculation.

[0051] Based on the derivation in existing technology, the expression for the drag coefficient of the projectile assembly is:

[0052]

[0053] The lift coefficient is:

[0054]

[0055] in, Cx represents the zero-lift drag coefficient of the projectile. iBw Cy represents the induced drag coefficient of the tail fin. BC represents the lift coefficient of the projectile. yw S represents the lift coefficient of the tail fin. w S represents the surface area of ​​the tail fin. M The surface area of ​​the projectile is represented by Ψ, and the logarithmic coefficient of the tail fin is represented by Ψ = 0.75 in the embodiment of this application. CyBw represents the interference factor correction coefficient between the missile body and the tail fin, and CyBw represents the lift coefficient.

[0056] Step S304: Determine the governing equations.

[0057] The continuity equation and momentum equation are shown below.

[0058]

[0059]

[0060] Where ρ represents density, μ represents viscosity coefficient, x represents velocity variable, t represents time, i represents fluctuating velocity constant, j represents time-averaged velocity constant, l represents instantaneous velocity constant, and μ′ i μ′ represents the pulsation velocity. j This indicates the average speed over time.

[0061] The Boussinesq assumption is typically used to solve the energy equation, which generally assumes that the Reynolds stress and the mean velocity gradient are positively correlated, as shown in the following expression.

[0062]

[0063]

[0064] Where k represents the thermal conductivity, δ ij Let E represent the stress tensor components, E represent the velocity vector components, T represent the static temperature, and k represent the static temperature. eff h represents the thermal conductivity. j′ J represents the diffusion flux of component h. j′ τ represents the diffusion flux of component j. ij The viscous dissipation term of the fluid, (τ) ij ) eff S represents the value of the viscous stress tensor at time-averaged velocity. h This represents the heat of chemical reaction and other user-defined volumetric heat sources.

[0065] Step S306: Determine the turbulence model.

[0066] Turbulence, as a form of fluid flow, is typically disordered, multi-angled, and irregular. Generally, turbulence exhibits good diffusion characteristics. In aerodynamics, turbulence generally refers to wind speed fluctuations within a specific region and over a relatively short period. The physical model of turbulence consists of vortices of varying sizes, and the size, direction, and other characteristics of these vortices are disordered and cannot be represented by mathematical formulas.

[0067] In aerodynamics research, the Spalart-Allmaras model, k-ε model, k-ω model, and Reynolds stress model are frequently used. This application's embodiment employs the Spalart-Allmaras model. The Spalart-Allmaras model is a relatively simple single-equation model used in aerospace applications to solve problems related to wall-bounded flows, demonstrating excellent performance in addressing boundary layer problems with adverse pressure gradients. In aerodynamics, this model is generally used to solve flow field calculations around aircraft and other similar devices.

[0068]

[0069] in: Gv represents the viscosity at rest, Yv represents the viscosity at rest, and Vv represents the viscosity at rest. Gv represents the increase in turbulent viscosity, Yv represents the decrease in turbulent viscosity, and Vv represents the viscosity at rest. Indicates a user-defined source item, C b2 This represents a constant viscosity.

[0070] Step S308: Determine the mesh generation and boundary conditions.

[0071] To accurately obtain the calculated values ​​of drag coefficient, lift coefficient, and moment coefficient, a denser mesh needs to be placed near the projectile body when meshing in GAMBIT. Special attention is also required when modeling the projectile's computational domain; generally, two cylinders are selected: a large cylinder and a small cylinder. For the large cylinder, the length is 7.5 times the projectile's length, and the diameter is 40 times the projectile's diameter; the small cylinder, which is the internal denser region, has a length 2.5 times the projectile's length and a diameter 5 times the projectile's diameter. Figure 4 As shown. During mesh generation, the mesh closer to the projectile is denser, while the mesh further away can be relatively sparser. The denser the mesh in the projectile's region, the more accurate the calculation results. To speed up computer processing, the computational domain is typically divided into two parts, with only one half being simulated.

[0072] Model mesh generation model such as Figure 5As shown, the entire computational domain employs a mesh refinement method, resulting in more accurate calculations. The total number of meshes in the computational domain is approximately over one million. The meshed projectile flow field is then imported into FLUENT for calculation.

[0073] When setting boundary conditions, a density-based coupled explicit solver is used. Far-field boundary conditions and viscous boundary conditions are used for the incoming flow and the object surface, respectively, as shown in the following formulas:

[0074]

[0075] Among them, u i (i = 1, 2, 3) represents the velocity components in the coordinate direction, F represents the pressure, ρ represents the density, and n represents the normal direction.

[0076] Step S310: Analyze the calculation results and the aerodynamic characteristics of the projectile.

[0077] The analysis of the calculation results will be described in detail below, and will not be repeated here.

[0078] This application provides a method for analyzing the calculation results of the aerodynamic characteristics of a remotely tail-controlled projectile.

[0079] In this embodiment of the application, when performing aerodynamic simulation analysis of the anti-spin projectile using FLUENT software, it is necessary to calculate the operating conditions at different angles of attack and different Mach numbers. The Mach numbers selected are 1Ma, 1.5Ma, 2Ma, 2.5Ma, 3Ma, 3.5Ma, and 4Ma; the angles of attack α are 1°, 2°, 3°, and 4°; a total of 28 states. The corresponding lift coefficient, drag coefficient, moment coefficient, X-direction velocity distribution cloud map, and overall projectile surface pressure distribution cloud map are calculated and simulated as follows.

[0080] Figure 6 This is a flowchart illustrating a method for calculating and analyzing the aerodynamic characteristics of a long-range tail-controlled guided projectile according to an embodiment of this application. Figure 6 As shown, the method includes the following steps:

[0081] Step S602, dynamic monitoring window analysis.

[0082] Taking a 4° angle of attack and 2.5 Ma as an example, the following drag coefficient, lift coefficient, and moment coefficient were obtained through FLUENT calculations. Observation shows that after 3000 iterations, the coefficients do not change significantly, indicating that the system is in a convergence state.

[0083] After 500 iterations, the convergent values ​​of the drag coefficient at a 4° angle of attack and 1.5 Mach are approximately 0.1739719342; at a 4° angle of attack and 2.0 Mach, approximately 0.25507912; at a 4° angle of attack and 2.5 Mach, approximately 0.32457274; and at 3.0 Mach, approximately 0.46645228. At this point, the drag coefficient value remains essentially unchanged after iterations, indicating that the calculated drag coefficient has reached convergence. Comparing the drag coefficient values ​​at four different Mach numbers reveals that, with a fixed 4° angle of attack, the drag coefficient value continuously increases with increasing Mach number.

[0084] The curve of drag coefficient changing with the iterative process is as follows: Figure 7 As shown. Figure 7 In Figure (a), the drag coefficient curve is shown at a 4° angle of attack and 1.5 Ma. Figure 7 (b) is the drag coefficient curve at a 4° angle of attack and 2.0 Ma, (c) is the drag coefficient curve at a 4° angle of attack and 2.5 Ma, and (d) is the drag coefficient curve at a 4° angle of attack and 3.0 Ma.

[0085] After 500 iterations, the convergent values ​​of the lift coefficient at a 4° angle of attack and Mach 1.5 are approximately 0.066994468; at a 4° angle of attack and Mach 2.0, approximately 0.136173674; at a 4° angle of attack and Mach 2.5, approximately 0.23034896; and at Mach 3.0, approximately 0.40881864. At this point, the lift coefficient value remains essentially unchanged after iterations, indicating that the calculated lift coefficient has reached convergence. Comparing the lift coefficient values ​​at four different Mach numbers reveals that, with a fixed 4° angle of attack, the lift coefficient value continuously increases with increasing Mach number. The lift coefficient variation curve during the iteration process is shown below. Figure 8 As shown. Figure 8 (a) is the lift coefficient curve v at a 4° angle of attack and 1.5 Ma; (b) is the lift coefficient curve v at a 4° angle of attack and 2.0 Ma; (c) is the lift coefficient curve v at a 4° angle of attack and 2.5 Ma; and (d) is the lift coefficient curve v at a 4° angle of attack and 3.0 Ma.

[0086] After 500 iterations, the convergence value of the torque coefficient at a 4° angle of attack and 1.5 Mach is approximately 0.08667689; at a 4° angle of attack and 2.0 Mach, it is approximately 0.20028906; at a 4° angle of attack and 2.5 Mach, it is approximately 0.3550003; and at 3.0 Mach, it is approximately 0.64597743. At this point, the torque coefficient value remains essentially unchanged after iterations, indicating that the calculated torque coefficient has reached convergence. Comparing the torque coefficient values ​​at four different Mach numbers reveals that, with a fixed 4° angle of attack, the torque coefficient value continuously increases with increasing Mach number. The torque coefficient variation curve during the iteration process is shown below. Figure 9 As shown. Figure 9 (a) is the torque coefficient curve at a 4° angle of attack and 1.5 Ma; (b) is the torque coefficient curve at a 4° angle of attack and 2.0 Ma; (c) is the torque coefficient curve at a 4° angle of attack and 2.5 Ma; and (d) is the torque coefficient curve at a 4° angle of attack and 3.0 Ma.

[0087] Step S604: Window analysis of airflow around the projectile.

[0088] Figure 10A , 10B Figures 10C and 10D show the velocity distribution contours in the X direction at 1.5 Ma, 2 Ma, 2.5 Ma, and 3 Ma, respectively, at an angle of attack of 4°. When the Mach number Ma > 1, i.e., the velocity V is higher than the speed of sound, the projectile under high dynamic conditions is simultaneously subjected to friction, vortex drag, and shock wave effects. At this time, as the ambient pressure Pa, density ρ, and temperature T increase, the velocity on the surface of the projectile will decrease instantaneously. In FLUENT calculations, an ideal gas that satisfies the gas law is generally chosen. In mathematical theory, the shock wave effect of an ideal gas that satisfies the gas law has no thickness and is discontinuous. However, in the actual flight of a high-dynamic projectile, the effects of density and viscosity should also be considered, and the actual thermal conductivity rate cannot be ignored. Although the shock wave effect interferes with the high-dynamic flight of the projectile, the actual impact is very small. Therefore, the shock wave value obtained in numerical calculations is small, and the shock wave value gradually decreases as the Mach number increases. Projectiles flying at high Mach are affected by airflow compression and other factors, and the parameters calculated by the computer will also undergo abrupt changes. Careful observation of the images reveals that both the head and tail of the projectile exhibit shock waves, with the higher the Mach number, the more pronounced the shock wave effect.

[0089] Figure 11A , 11BThe four images, 11C and 11D, are surface pressure distribution cloud maps of the entire missile at angles of attack of 4° at Mach 1.5, Mach 2, Mach 2.5, and Mach 3.0. Pressure distribution on the aircraft surface is an important reference indicator for aircraft design. Figure 11D It can be seen that the projectile experiences the greatest pressure during flight. At Ma 1.5, the pressure is approximately 2.27e+0.5; at Ma 2, it is approximately 5.55e+0.5; at Ma 2.5, it is approximately 7.47e+0.5; and at Ma 3.0, it is approximately 9.82e+0.5. The pressure at the stern is relatively low, approximately 1.03e+0.5, 1.71e+0.5, 6.02e+0.4, and 2.42e+0.3 respectively.

[0090] Therefore, as the Mach number increases, the surface pressure of the projectile gradually increases, with the warhead bearing the greatest pressure. In terms of material processing, the warhead material must also undergo special treatment to ensure the stability of the projectile during flight.

[0091] Step S606: Force analysis of the projectile model.

[0092] Simulations using FLUENT software, reading from cd-history, cl-history, and cm-history, yielded the final converged values ​​of drag, lift, and moment. To save computing time and reduce mesh generation difficulty, only half of the computational domain was calculated. This avoided numerical instability caused by reflections of pressure wave isobars in the flow field. Therefore, the actual drag, lift, and moment values ​​should be twice the calculated values.

[0093] from Figure 12 As can be seen from this, when the angle of attack is constant, as the Mach number increases, the drag value (Cd) also increases continuously, and the rate of increase gradually changes from rapid to slow, but eventually shows an upward trend; when the Mach number is constant, as the angle of attack increases, the drag value first decreases and then increases to a peak value, and then decreases, showing an overall upward trend.

[0094] from Figure 13 As can be seen, at a constant angle of attack, the lift coefficient (Cl) increases with increasing Mach number. Although the rate of ascent is small at 0° angle of attack, the lift coefficient increases continuously with increasing Mach number at other angles of attack, and the larger the angle of attack, the faster the lift coefficient increases. (The last sentence is a repetition of the previous one and can be omitted.)

[0095] from Figure 14As can be seen, at a constant angle of attack, the torque coefficient (Cm) increases continuously with the increase of the Mach number, and the difference in the increment of the torque coefficient also gradually increases with the increase of the angle of attack. The larger the angle of attack, the faster the increase in the lift coefficient. At a constant Mach number, the torque coefficient gradually increases with the increase of the angle of attack, and the rate of increase gradually slows down.

[0096] It should be noted that, for the sake of simplicity, the foregoing method embodiments are all described as a series of actions. However, those skilled in the art should understand that the present invention is not limited to the described order of actions, because according to the present invention, some steps can be performed in other orders or simultaneously. Furthermore, those skilled in the art should also understand that the embodiments described in the specification are preferred embodiments, and the actions and modules involved are not necessarily essential to the present invention.

[0097] Through the above description of the embodiments, those skilled in the art can clearly understand that the methods according to the above embodiments can be implemented by means of software plus necessary general-purpose hardware platforms. Of course, they can also be implemented by hardware, but in many cases the former is a better implementation method. Based on this understanding, the technical solution of the present invention, in essence, or the part that contributes to the prior art, can be embodied in the form of a software product. This computer software product is stored in a storage medium (such as ROM / RAM, magnetic disk, optical disk) and includes several instructions to cause a terminal device (which may be a mobile phone, computer, server, or network device, etc.) to execute the methods described in the various embodiments of the present invention.

[0098] This application also provides an analysis device for the aerodynamic characteristics of a long-range tail-controlled guided projectile, such as... Figure 15 As shown, the device includes: a model building module 152, a setting module 154, and an analysis module 156.

[0099] The model building module 152 is used to physically model the projectile body to obtain the projectile model; the setting module 154 is used to perform unstructured mesh generation on the projectile model and set boundary conditions; the analysis module 156 is used to perform aerodynamic characteristic analysis on the projectile model based on the divided mesh and the set boundary conditions to obtain the variation law of the aerodynamic characteristics of the projectile model with flight Mach number and angle of attack.

[0100] The model building module 152 is used to: build the front section of the shell model; build the anti-spin stern of the shell model, wherein the anti-spin stern is non-rigidly connected to the front section through a bearing device, and the friction on the bearing device can drive the anti-spin stern to rotate so that the shell body can reduce rotation after leaving the barrel.

[0101] The setting module 154 is used to: model the computational domain of the projectile model and perform mesh generation, wherein the mesh closer to the projectile body is denser, the mesh farther away from the projectile body is relatively sparser, and the mesh of the projectile head is the densest; when setting boundary conditions, a density-based coupled explicit solver is selected for solving, and far-field boundary conditions and viscous boundary conditions are selected for the incoming flow and the object surface, respectively.

[0102] The setting module 154 is also used to: determine the drag coefficient expression and lift coefficient expression of the projectile model through engineering calculation methods; determine the continuity equation and transport equation of the projectile model, and solve the energy equation of the projectile model; and determine the turbulence model of the projectile model.

[0103] The analysis module 156 is also used to: calculate the drag coefficient, lift coefficient, and moment coefficient at different angles of attack and different Mach numbers, so as to perform dynamic monitoring window analysis and obtain the variation law of the lift coefficient, drag coefficient, and moment coefficient of the projectile model with flight Mach number and angle of attack: when the angle of attack is constant, the drag coefficient also increases continuously with the increase of Mach number; when the angle of attack is constant, the lift coefficient also increases continuously with the increase of Mach number; when the angle of attack is constant, the moment coefficient also increases continuously with the increase of Mach number.

[0104] The analysis module 156 is also used to: calculate the velocity distribution cloud map in the X direction and the pressure distribution cloud map on the surface of the projectile at different angles of attack and different Mach numbers, so as to perform window analysis of the air flow around the projectile and obtain the variation law of the shock wave and the surface pressure of the projectile model with the flight Mach number and angle of attack: when the angle of attack is constant, the shock wave gradually decreases as the Mach number increases; when the angle of attack is constant, the surface pressure of the projectile gradually increases as the Mach number increases, among which the pressure borne by the projectile head is the greatest.

[0105] The analysis module 156 is also used to calculate the drag, lift and torque of the projectile model at the final convergence, so as to perform a force analysis of the projectile model and obtain the variation law of the force of the projectile model with the flight Mach number and angle of attack.

[0106] The aerodynamic characteristic analysis device for long-range tail-controlled guided projectiles provided in this embodiment can realize the aerodynamic characteristic analysis method for long-range tail-controlled guided projectiles provided in the above embodiment, therefore, it will not be described again here.

[0107] This embodiment has the following beneficial effects:

[0108] 1) Using FLUENT to simulate guided projectiles, it was confirmed that the aerodynamic shape of the non-rigid connection guided projectile anti-spin structure with bearing connecting the stern and the bow is reasonable, and the obtained aerodynamic parameters can provide a design basis for anti-spin guided projectiles.

[0109] 2) Simulation results show that the model has high prediction accuracy in solving lift, drag, and moment coefficient, providing new ideas for attitude measurement and guidance, and has certain engineering application value.

[0110] 3) By observing the pressure distribution cloud map of the entire missile surface, it can be seen that the pressure value borne by the warhead is the greatest. In terms of material processing, the material of the warhead must also be specially treated to meet the flight requirements.

[0111] Embodiments of the present invention also provide a storage medium. Optionally, in this embodiment, the storage medium is configured to store program code for performing the analysis method for the aerodynamic characteristics of the remotely controlled tail-guided projectile.

[0112] Optionally, in this embodiment, the storage medium may include, but is not limited to, various media capable of storing program code, such as USB flash drives, read-only memory (ROM), random access memory (RAM), portable hard drives, magnetic disks, or optical disks.

[0113] The sequence numbers of the above embodiments of the present invention are for descriptive purposes only and do not represent the superiority or inferiority of the embodiments.

[0114] If the integrated units in the above embodiments are implemented as software functional units and sold or used as independent products, they can be stored in the aforementioned computer-readable storage medium. Based on this understanding, the technical solution of the present invention, in essence, or the part that contributes to the prior art, or all or part of the technical solution, can be embodied in the form of a software product. This computer software product is stored in a storage medium and includes several instructions to cause one or more computer devices (which may be personal computers, servers, or network devices, etc.) to execute all or part of the steps of the methods described in the various embodiments of the present invention.

[0115] In the above embodiments of the present invention, the descriptions of each embodiment have different focuses. For parts not described in detail in a certain embodiment, please refer to the relevant descriptions of other embodiments.

[0116] In the several embodiments provided in this application, it should be understood that the disclosed client can be implemented in other ways. The device embodiments described above are merely illustrative; for example, the division of units is only a logical functional division, and in actual implementation, there may be other division methods. For example, multiple units or components may be combined or integrated into another system, or some features may be ignored or not executed. Furthermore, the displayed or discussed mutual couplings, direct couplings, or communication connections may be through some interfaces, indirect couplings, or communication connections between units or modules, and may be electrical or other forms.

[0117] The units described as separate components may or may not be physically separate. The components shown as units may or may not be physical units; that is, they may be located in one place or distributed across multiple network units. Some or all of the units can be selected to achieve the purpose of this embodiment according to actual needs.

[0118] Furthermore, the functional units in the various embodiments of the present invention can be integrated into one processing unit, or each unit can exist physically separately, or two or more units can be integrated into one unit. The integrated unit can be implemented in hardware or as a software functional unit.

[0119] The above description is only a preferred embodiment of the present invention. It should be noted that for those skilled in the art, several improvements and modifications can be made without departing from the principle of the present invention, and these improvements and modifications should also be considered within the scope of protection of the present invention.

Claims

1. A method for analyzing the aerodynamic characteristics of a long-range tail-controlled guided projectile, characterized in that, include: Physical modeling of the projectile body is performed to obtain the projectile model; The projectile model is divided into an unstructured mesh and boundary conditions are set. Based on the divided grid and the set boundary conditions, the aerodynamic characteristics of the projectile model are analyzed to obtain the variation law of the aerodynamic characteristics of the projectile model with flight Mach number and angle of attack. The process involves physically modeling the projectile body to obtain a projectile model, including: constructing the forward section of the projectile model; and constructing the anti-spin stern of the projectile model. The anti-spin stern is non-rigidly connected to the forward section via a bearing device. The friction on the bearing device can drive the anti-spin stern to rotate, so that the projectile body can reduce rotation after leaving the barrel. The process of unstructured mesh generation and boundary conditions setting for the projectile model includes: modeling the computational domain of the projectile body and meshing it, wherein the mesh is denser closer to the projectile body and relatively sparser farther away from the projectile body, with the densest mesh in the projectile head; when setting boundary conditions, a density-based coupled explicit solver is used for solving, and far-field boundary conditions and viscous boundary conditions are used for the incoming flow and the object surface, respectively. Before performing unstructured mesh generation on the projectile model, the method further includes: determining the drag coefficient expression and lift coefficient expression of the projectile model through engineering calculation methods; determining the continuity equation and transport equation of the projectile model, and solving the energy equation of the projectile model; and determining the turbulence model of the projectile model. The process includes: performing aerodynamic characteristic analysis on the projectile model to obtain the variation law of the aerodynamic characteristics of the projectile model with flight Mach number and angle of attack, including: calculating the drag coefficient, lift coefficient, and moment coefficient at different angles of attack and different Mach numbers for dynamic monitoring window analysis to obtain the variation law of the lift coefficient, drag coefficient, and moment coefficient of the projectile model with flight Mach number and angle of attack; calculating the X-direction velocity distribution cloud map and the projectile surface pressure distribution cloud map at different angles of attack and different Mach numbers for airflow window analysis around the projectile to obtain the variation law of the shock wave and projectile surface pressure of the projectile model with flight Mach number and angle of attack; and calculating the final convergence drag, lift, and moment of the projectile model to perform force analysis of the projectile model to obtain the variation law of the force of the projectile model with flight Mach number and angle of attack.

2. The method according to claim 1, characterized in that, The variation of the lift coefficient, drag coefficient, and moment coefficient of the projectile model with flight Mach number and angle of attack includes at least one of the following: With a fixed angle of attack, the drag coefficient increases as the Mach number increases. With a fixed angle of attack, the lift coefficient increases as the Mach number increases. With a fixed angle of attack, the torque coefficient increases as the Mach number increases.

3. The method according to claim 1, characterized in that, The variation of the shock wave and surface pressure of the projectile model with flight Mach number and angle of attack includes at least one of the following: With a fixed angle of attack, the shock wave gradually decreases as the Mach number increases; With a fixed angle of attack, as the Mach number increases, the surface pressure of the projectile gradually increases, with the projectile itself bearing the greatest pressure.

4. The method according to claim 1, characterized in that, The force distribution of the projectile model with respect to flight Mach number and angle of attack includes at least one of the following: With a fixed angle of attack, as the Mach number increases, the drag also increases continuously, and the rate of increase gradually slows down from rapid to slow, but eventually shows an upward trend. With a fixed Mach number, as the angle of attack increases, the drag first decreases, then increases to a peak value, and then decreases again, but overall shows an upward trend. At a constant angle of attack, lift increases with increasing Mach number. At 0° angle of attack, the rate of climb is small, but the lift coefficient increases with increasing Mach number at other angles of attack. The larger the angle of attack, the faster the lift coefficient increases. At a constant angle of attack, as the Mach number increases, the torque coefficient also increases continuously. Furthermore, as the angle of attack increases, the difference in the increment of the torque coefficient also gradually increases. The larger the angle of attack, the faster the lift coefficient increases. At a constant Mach number, as the angle of attack increases, the torque coefficient gradually shows an upward trend, and the rate of increase gradually slows down.

5. The method according to any one of claims 1 to 4, characterized in that, After obtaining the variation law of the aerodynamic characteristics of the projectile model with flight Mach number and angle of attack, the method further includes: manufacturing a spin-reducing guided projectile based on the variation law.

6. A device for analyzing the aerodynamic characteristics of a long-range tail-controlled guided projectile, characterized in that, include: The model building module is used to physically model the projectile body to obtain the projectile model; The setting module is used to perform unstructured mesh generation and set boundary conditions for the projectile model; The analysis module is used to perform aerodynamic characteristic analysis on the projectile model based on the divided grid and the set boundary conditions, and to obtain the variation law of the aerodynamic characteristics of the projectile model with flight Mach number and angle of attack. The model building module is further configured to: build the forward section of the shell model; build the anti-spin stern of the shell model, wherein the anti-spin stern is non-rigidly connected to the forward section through a bearing device, and the friction on the bearing device can drive the anti-spin stern to rotate so that the shell body can reduce rotation after leaving the barrel. The setting module is further configured to: model the computational domain of the projectile model and perform mesh generation, wherein the mesh closer to the projectile body is denser, the mesh farther away from the projectile body is relatively sparser, and the mesh of the projectile head is the densest; when setting boundary conditions, a density-based coupled explicit solver is selected for solving, and far-field boundary conditions and viscous boundary conditions are selected for the incoming flow and the object surface, respectively. Before performing unstructured mesh generation on the projectile model, the device is further configured to: determine the drag coefficient expression and lift coefficient expression of the projectile model through engineering calculation methods; determine the continuity equation and transport equation of the projectile model, and solve the energy equation of the projectile model; and determine the turbulence model of the projectile model. The analysis module is further configured to: calculate the drag coefficient, lift coefficient, and moment coefficient at different angles of attack and Mach numbers for dynamic monitoring window analysis, thereby obtaining the variation law of the lift coefficient, drag coefficient, and moment coefficient of the projectile model with flight Mach number and angle of attack; calculate the X-direction velocity distribution cloud map and the projectile surface pressure distribution cloud map at different angles of attack and Mach numbers for airflow window analysis around the projectile, thereby obtaining the variation law of the shock wave and projectile surface pressure of the projectile model with flight Mach number and angle of attack; and calculate the final convergence drag, lift, and moment of the projectile model for force analysis, thereby obtaining the variation law of the force of the projectile model with flight Mach number and angle of attack.

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