A method for acquiring distributed aerodynamic characteristics of a missile based on CFD

Abstract

This invention belongs to the field of projectile aerodynamics technology, specifically relating to a CFD-based method for obtaining the distributed aerodynamic characteristics of projectiles. When the projectile is in level flight without other motion forms, the pressure and viscous forces obtained from steady CFD calculations are processed to obtain the aerodynamic characteristics of the projectile surface at different cross-sectional positions along the longitudinal axis of the projectile. When the projectile has unsteady motion forms such as rotation, conical motion, and bending deformation, the pressure and viscous forces at different times within a motion cycle can be obtained through CFD calculations. First, the CFD calculation results are converted to values ​​in a quasi-projectile coordinate system through coordinate transformation. Second, using the aforementioned method for obtaining distributed aerodynamic characteristics under steady conditions, the distributed aerodynamic characteristics of the projectile at different times are obtained. Finally, the results at different times are fitted using Fourier series to obtain the time-averaged distributed aerodynamic characteristics. This invention solves the difficult problems of calculating the distributed aerodynamic characteristics of projectiles with complex configurations and complex motions.

Description

Technical Field

[0001] This invention belongs to the field of projectile aerodynamics technology, specifically relating to a CFD-based method for obtaining distributed aerodynamic characteristics of projectiles. Background Technology

[0002] Distributed aerodynamic characteristics play a crucial role in engineering research. On one hand, they help reveal the flow mechanisms of induced forces. In typical canard-body-tail fin missiles, severe aerodynamic interference occurs between components during angle-of-attack flight. Using the total aerodynamic coefficient of the entire missile or its components is insufficient to reveal the intrinsic flow mechanisms of induced aerodynamic interference forces. A comprehensive understanding of the distributed aerodynamic characteristics of the entire missile or its components from start to finish, combined with comparisons of flow structures at characteristic cross-sections, is necessary to better elucidate the flow mechanisms. On the other hand, they help verify structural strength and analyze dynamic responses. For missile bodies with high aspect ratios or wings with high aspect ratios, the static deformation of the structure under aerodynamic loads is not negligible. Obtaining distributed aerodynamic loads under typical flight conditions is crucial for analyzing structural load-bearing capacity; conversely, static deformation also leads to changes in aerodynamic loads, thus affecting the structural dynamic response.

[0003] Projectiles exhibit diverse motion patterns during actual flight, including rotation, nutation, precession, conical motion, and bending / torsional deformation. Nutation, precession, and conical motion are collectively referred to as multi-axis angular motion, which is the circular motion of the projectile's longitudinal axis around its velocity vector. These complex motions result in significant unsteady aerodynamic characteristics for projectiles. However, the aerodynamic, structural, and flight responses of projectiles are not synchronous. For example, for a rotating projectile, at high rotational speeds, it will not respond to the transient aerodynamic characteristics caused by rotation, but only to the periodic average aerodynamic force generated by one rotation. Only at lower rotational speeds will it be able to respond to transient aerodynamic characteristics. Therefore, both transient and time-averaged distributed aerodynamic characteristics are crucial for flow mechanism analysis, structural strength verification, and structural and flight dynamic responses.

[0004] To meet diverse operational needs, missiles and rockets exhibit complex and varied shapes, with key aerodynamic components including canards, the missile body, and tail fins. Canards and tail fins may have installation angles, and tail fins can be straight, folded, or curved. Obtaining the distributed aerodynamic characteristics of complex configurations with complex motions is not easy. Currently, engineering methods generally employ wind tunnel testing or numerical methods. However, wind tunnel pressure testing only focuses on individual discrete points under simple motion conditions; numerical calculations typically divide the mesh into segments at the beginning of the calculation, extracting forces segment by segment to form distributed loads. Existing methods are difficult to implement and cannot guarantee accuracy. Summary of the Invention

[0005] (a) Technical problems to be solved

[0006] The technical problem to be solved by this invention is: how to provide a method for obtaining the distributed aerodynamic characteristics of projectiles and rockets, solve the difficult problems of calculating the distributed aerodynamic characteristics of projectiles and rockets with complex configurations and complex motions, streamline the process of obtaining distributed aerodynamic characteristics, and provide a way for engineering applications.

[0007] (II) Technical Solution

[0008] To address the aforementioned technical problems, this invention provides a CFD-based method for obtaining distributed aerodynamic characteristics of projectiles and rockets, the method comprising the following steps:

[0009] Step 1: Based on the projectile cross-sectional data, including the pressure and viscous force of the projectile cross-section, obtained by CFD numerical calculation, the resultant normal force and resultant lateral force on the cross-section are obtained by circumferential angular integration;

[0010] Step 2: Based on the wing / rudder section data including wing / rudder section pressure and viscous force obtained from CFD numerical calculation, directly integrate the pressure and viscous force along the wingspan direction to obtain the resultant normal force and resultant lateral force on the section.

[0011] Step 3: When the projectile has unsteady motion forms including rotation, conical motion, and bending deformation, first convert the CFD numerical calculation results into CFD numerical calculation results in the quasi-projectile coordinate system through coordinate transformation. Then, referring to Step 1 and Step 2, obtain the distributed aerodynamic characteristics at a specific moment from the CFD numerical calculation results in the quasi-projectile coordinate system. Finally, use Fourier series to fit the distributed aerodynamic characteristics at different moments to obtain the time-averaged distributed aerodynamic characteristics.

[0012] In step one, for the wing area of ​​the missile body, the presence of the tail fin causes a discontinuity in the missile body data along the circumferential angle. Therefore, the missile body cross-sectional data is divided into segments, and the resultant force is calculated by integrating each segment separately.

[0013] In step one, the projectile cross-section data obtained by CFD numerical calculation includes grid node coordinates, pressure, and viscous force.

[0014] When integrating along the circumferential angle, the projectile cross-section data are arranged in ascending order of circumferential angle;

[0015] If there are discontinuities in the wing section cross-section data along the circumferential angle, the cross-section data is divided into sections according to the number and position of the tail fins; the resultant force in the normal direction and the resultant force in the lateral direction are obtained by integrating the local circumferential angle to the normal and lateral directions.

[0016] In step two, since the tail fin can have straight, bent, or curved configurations, the data of its windward / leeward or concave / convex surfaces are separated and then superimposed to calculate the resultant force.

[0017] In step two, the wing / rudder section data obtained by CFD numerical calculation includes grid node coordinates, pressure, and viscous force.

[0018] When integrating along the wingspan, the wing / rudder section data are arranged in ascending order of wingspan, and the windward / leeward or concave / convex surface data are separated. Based on the local azimuth normal and lateral projections, the resultant normal force and lateral force are obtained by integration.

[0019] In step three, the CFD numerical calculation results are transformed into the results of the quasi-bullet system through coordinate transformation between the inertial frame and the quasi-bullet system.

[0020] Based on the aerodynamic characteristics and frequency changes of the missile body and wing / rudder sections at different times, a Fourier series of appropriate order is selected for fitting to obtain the time-averaged distributed aerodynamic characteristics.

[0021] Furthermore, this invention also provides a CFD-based method for obtaining distributed aerodynamic characteristics of projectiles and rockets, the method comprising the following steps:

[0022] Step 1: When the projectile is flying horizontally without other motion, the pressure and viscous forces obtained from steady-state CFD calculations are processed to obtain the aerodynamic characteristics of the projectile surface at different cross-sectional positions along the longitudinal axis of the projectile; Step 1 includes:

[0023] Step 11: At different projectile cross-sectional positions, project the pressure and viscous forces along the circumferential angle and integrate to obtain the resultant normal force and resultant lateral force on the cross-section;

[0024] Step 12: At different wing / rudder cross-section positions, project the pressure and viscous forces along the azimuth angle and integrate them along the wingspan direction to obtain the resultant normal force and resultant lateral force on the cross-section;

[0025] Step 13: Superimpose the forces on the missile body and wings / rudders to obtain the distributed aerodynamic characteristics along the longitudinal axis of the missile body;

[0026] Step 2: When the projectile has unsteady motion forms including rotation, conical motion, and bending deformation, the pressure and viscous force at different times within a motion cycle are obtained through CFD calculation;

[0027] The unsteady CFD values ​​obtained from CFD calculations are converted into results in the quasi-projectile coordinate system through coordinate transformation.

[0028] Using the method for obtaining the distributed aerodynamic characteristics under steady conditions in step 1 above, the distributed aerodynamic characteristics of the projectile and rocket at different times are obtained.

[0029] Finally, based on the aerodynamic characteristics and frequency of the missile body and wing / rudder sections at different times, a Fourier series of appropriate order is selected for fitting, and the time-averaged distributed aerodynamic characteristics are obtained by fitting the results at different times.

[0030] The steady CFD calculation uses solvers such as Fluent, CFD++, and CFX. The calculated projectile cross-section data and wing / rudder cross-section data include grid node coordinates, pressure, and viscous forces. When obtaining distributed aerodynamic forces, data from characteristic cross-sections are selected for processing.

[0031] When the projectile cross-sectional data is projected and integrated along the circumferential angle, the data is arranged in ascending order of circumferential angle. For a wing-type projectile, the presence of tail fins causes discontinuities in the projectile cross-sectional data along the circumferential angle. Therefore, the projectile cross-sectional data is segmented according to the number and position of tail fins. Each segment is projected onto the normal and lateral directions according to the local circumferential angle, and then integrated to obtain the resultant force in the normal direction and the resultant force in the lateral direction.

[0032] When integrating the wing / rudder section data along the wingspan direction, the wing / rudder section data are arranged in ascending order of wingspan. For tail fins with different configurations, including straight, bent, and rolled tail fins, their windward / leeward or concave / convex surface data are separated. Based on the local azimuth angle normal and lateral projections, the normal resultant force and lateral resultant force are then integrated to obtain the resultant force.

[0033] (III) Beneficial Effects

[0034] The method proposed in this invention is universally applicable in obtaining the distributed aerodynamic characteristics of complex configurations and complex motion projectiles, and the more data from the sample cross-section, the higher the accuracy, thus providing a way for flow mechanism analysis and structural dynamic response analysis.

[0035] Compared with the prior art, the present invention has the following beneficial effects:

[0036] 1. Compared with the traditional method of obtaining distributed aerodynamic characteristics by dividing the grid into segments, this invention innovatively adopts a cross-sectional data processing method. That is, based on the CFD calculation results, the wall data in any cross-section of the projectile is obtained by interpolation. The data of the projectile body is integrated along the circumferential angle to obtain the resultant force, and the data of the wing / rudder part is integrated along the wingspan direction to obtain the resultant force. At the same time, the data processing is taken into account when there are discontinuities in the projectile body data or complex tail fin configurations.

[0037] 2. When obtaining the distributed aerodynamic characteristics of complex-moving projectiles, the CFD calculation results in the inertial frame are first transformed into results for a quasi-projectile system through coordinate transformation. Then, the transient distributed aerodynamic characteristics at different times are obtained by using the force treatment method of the projectile body and wing / rudder sections. Furthermore, a Fourier series of appropriate order is selected to fit the cross-sectional aerodynamic characteristics at different times, thereby obtaining the time-averaged distributed aerodynamic characteristics. This solves the problem of calculating the distributed aerodynamic characteristics of complex-configured and complex-moving projectiles, providing a means for flow mechanism analysis and structural dynamic response analysis. Attached Figure Description

[0038] Figure 1 Flowchart for obtaining distributed aerodynamic characteristics;

[0039] Figure 2 This is a diagram illustrating the data processing partitioning.

[0040] Figure 3 Flowchart for force distribution processing in projectile body;

[0041] Figure 4 Flowchart for wing / rudder force distribution processing;

[0042] Figure 5 Flowchart for handling force distribution in complex motion projectiles;

[0043] Figure 6a and Figure 6b A curve showing the force distribution coefficient of a canard-body-tail fin type projectile;

[0044] in, Figure 6a The lateral force distribution coefficient is the coefficient of rotation when α = 4°. Figure 6b The lateral force coefficient is the coefficient of rotation when α = 12°.

[0045] Figure 7a and Figure 7b This is a curve showing the force distribution coefficient of a rotating spring in conical motion.

[0046] in, Figure 7a The time-averaged normal force coefficient, Figure 7b is the time-averaged lateral force coefficient. Detailed Implementation

[0047] To make the objectives, contents, and advantages of the present invention clearer, the specific embodiments of the present invention will be described in further detail below with reference to the accompanying drawings and examples.

[0048] To address the aforementioned technical problems, this invention provides a CFD-based method for obtaining distributed aerodynamic characteristics of projectiles and rockets, the method comprising the following steps:

[0049] Step 1: Based on the projectile cross-sectional data, including the pressure and viscous force of the projectile cross-section, obtained by CFD numerical calculation, the resultant normal force and resultant lateral force on the cross-section are obtained by circumferential angular integration;

[0050] Step 2: Based on the wing / rudder section data including wing / rudder section pressure and viscous force obtained from CFD numerical calculation, directly integrate the pressure and viscous force along the wingspan direction to obtain the resultant normal force and resultant lateral force on the section.

[0051] Step 3: When the projectile has unsteady motion forms including rotation, conical motion, and bending deformation, first convert the CFD numerical calculation results into CFD numerical calculation results in the quasi-projectile coordinate system through coordinate transformation. Then, referring to Step 1 and Step 2, obtain the distributed aerodynamic characteristics at a specific moment from the CFD numerical calculation results in the quasi-projectile coordinate system. Finally, use Fourier series to fit the distributed aerodynamic characteristics at different moments to obtain the time-averaged distributed aerodynamic characteristics.

[0052] In step one, for the wing area of ​​the missile body, the presence of the tail fin causes a discontinuity in the missile body data along the circumferential angle. Therefore, the missile body cross-sectional data is divided into segments, and the resultant force is calculated by integrating each segment separately.

[0053] In step one, the projectile cross-section data obtained by CFD numerical calculation includes grid node coordinates, pressure, and viscous force.

[0054] When integrating along the circumferential angle, the projectile cross-section data are arranged in ascending order of circumferential angle;

[0055] If there are discontinuities in the wing section cross-section data along the circumferential angle, the cross-section data is divided into sections according to the number and position of the tail fins; the resultant force in the normal direction and the resultant force in the lateral direction are obtained by integrating the local circumferential angle to the normal and lateral directions.

[0056] In step two, since the tail fin can have straight, bent, or curved configurations, the data of its windward / leeward or concave / convex surfaces are separated and then superimposed to calculate the resultant force.

[0057] In step two, the wing / rudder section data obtained by CFD numerical calculation includes grid node coordinates, pressure, and viscous force.

[0058] When integrating along the wingspan, the wing / rudder section data are arranged in ascending order of wingspan, and the windward / leeward or concave / convex surface data are separated. Based on the local azimuth normal and lateral projections, the resultant normal force and lateral force are obtained by integration.

[0059] In step three, the CFD numerical calculation results are transformed into the results of the quasi-bullet system through coordinate transformation between the inertial frame and the quasi-bullet system.

[0060] Based on the aerodynamic characteristics and frequency changes of the missile body and wing / rudder sections at different times, a Fourier series of appropriate order is selected for fitting to obtain the time-averaged distributed aerodynamic characteristics.

[0061] Furthermore, this invention also provides a CFD-based method for obtaining distributed aerodynamic characteristics of projectiles and rockets, the method comprising the following steps:

[0062] Step 1: When the projectile is flying horizontally without other motion, the pressure and viscous forces obtained from steady-state CFD calculations are processed to obtain the aerodynamic characteristics of the projectile surface at different cross-sectional positions along the longitudinal axis of the projectile; Step 1 includes:

[0063] Step 11: At different projectile cross-sectional positions, project the pressure and viscous forces along the circumferential angle and integrate to obtain the resultant normal force and resultant lateral force on the cross-section;

[0064] Step 12: At different wing / rudder cross-section positions, project the pressure and viscous forces along the azimuth angle and integrate them along the wingspan direction to obtain the resultant normal force and resultant lateral force on the cross-section;

[0065] Step 13: Superimpose the forces on the missile body and wings / rudders to obtain the distributed aerodynamic characteristics along the longitudinal axis of the missile body;

[0066] Step 2: When the projectile has unsteady motion forms including rotation, conical motion, and bending deformation, the pressure and viscous force at different times within a motion cycle are obtained through CFD calculation;

[0067] The unsteady CFD values ​​obtained from CFD calculations are converted into results in the quasi-projectile coordinate system through coordinate transformation.

[0068] Using the method for obtaining the distributed aerodynamic characteristics under steady conditions in step 1 above, the distributed aerodynamic characteristics of the projectile and rocket at different times are obtained.

[0069] Finally, based on the aerodynamic characteristics and frequency of the missile body and wing / rudder sections at different times, a Fourier series of appropriate order is selected for fitting, and the time-averaged distributed aerodynamic characteristics are obtained by fitting the results at different times.

[0070] The steady CFD calculation uses solvers such as Fluent, CFD++, and CFX. The calculated projectile cross-section data and wing / rudder cross-section data include grid node coordinates, pressure, and viscous forces. When obtaining distributed aerodynamic forces, data from characteristic cross-sections are selected for processing.

[0071] When the projectile cross-sectional data is projected and integrated along the circumferential angle, the data is arranged in ascending order of circumferential angle. For a wing-type projectile, the presence of tail fins causes discontinuities in the projectile cross-sectional data along the circumferential angle. Therefore, the projectile cross-sectional data is segmented according to the number and position of tail fins. Each segment is projected onto the normal and lateral directions according to the local circumferential angle, and then integrated to obtain the resultant force in the normal direction and the resultant force in the lateral direction.

[0072] When integrating the wing / rudder section data along the wingspan direction, the wing / rudder section data are arranged in ascending order of wingspan. For tail fins with different configurations, including straight, bent, and rolled tail fins, their windward / leeward or concave / convex surface data are separated. Based on the local azimuth angle normal and lateral projections, the normal resultant force and lateral resultant force are then integrated to obtain the resultant force.

[0073] Example 1

[0074] In this embodiment, as Figure 1 As shown, the CFD-based method for obtaining the distributed aerodynamic characteristics of projectiles and rockets includes the following steps:

[0075] CFD result files can be calculated using solvers such as Fluent, CFD++, and CFX. They should include data such as mesh node coordinates, pressure, and viscous forces. Here, only the projectile wall data is needed; processing each section individually will yield the distributed aerodynamic characteristics. The wall data is divided into two categories: projectile body data and wing / rudder data, such as... Figure 2 As shown. Among them, the projectile body data can be further divided into spin-type projectile body data ( Figure 2 (shaded area in the middle diagonal) and wing area missile body data ( Figure 2 (Blank portion), Wing / Rudder data ( Figure 2 The solid black part in the middle has different treatment methods depending on the configuration.

[0076] Projectile data processing flow as follows Figure 3 As shown. First, the data for the characteristic cross-section needs to be selected, that is, the Y and Z coordinates, pressure, and viscous force data are selected by fixing the X coordinate of the cross-section. Next, the cross-sectional data are arranged in ascending order of circumferential angle, that is, the circumferential angle is determined and sorted according to the Y and Z coordinates, and the pressure and viscous force data are corresponding to the circumferential angle. If the cross-section of the projectile is located in the wing area of ​​the projectile body, the pressure and viscous force data will be discontinuous along the circumferential angle. It is necessary to segment the projectile body data according to the number and position of the tail fins, project the segments and integrate to obtain the cross-sectional normal force and lateral force coefficients. If the cross-section of the projectile body is located in the spinning body part, the pressure and viscous force are directly projected along the circumferential angle and integrated to obtain the cross-sectional normal force and lateral force coefficients. After completing the data processing of one characteristic cross-section, multiple characteristic cross-sections can be taken along the longitudinal axis of the projectile body, and the data can be processed one by one to obtain the distributed aerodynamic characteristics of the projectile body.

[0077] Wing / rudder data processing flow as follows Figure 4As shown. First, the data for the characteristic sections needs to be selected, that is, the Y and Z coordinates, pressure, and viscous force data are selected by fixing the X coordinate of the cross section; next, the cross section data are arranged from smallest to largest wingspan, that is, the wingspan length is determined and sorted according to the Y and Z coordinates, and the pressure and viscous force data are corresponding to the wingspan position; if the wing / rudder is a straight configuration, the windward and leeward data are separated with the position angle of the airfoil centerline as a reference, and the pressure and viscous force are integrated along the wingspan direction, then projected and the resultant force is calculated to obtain the cross section normal force and lateral force coefficients; if the tail fin For the curved configuration, the coordinates of the curvature center of the curved wing are first obtained, and then the distance between the coordinate point on the wing surface and the curvature center is obtained. Median filtering is then used to obtain the distance between the centerline of the curved wing section and the curvature center. Using this as a reference, the concave and convex surface data are separated. The pressure and viscous forces are integrated and projected along the wingspan direction in four quadrants, and the resultant force is obtained to obtain the normal force and lateral force coefficient of the section. After completing the data processing of one characteristic section, multiple characteristic sections can be taken along the longitudinal axis of the projectile, and the data can be processed one by one to obtain the distributed aerodynamic characteristics of the wing / rudder section.

[0078] When a projectile undergoes complex motion, the data processing flow is as follows: Figure 5 As shown. The case of a conical rotating projectile is used as an example. First, the projectile surface data in the inertial coordinate system obtained from CFD calculations is converted into quasi-projectile coordinate system results. The three-directional force components C in the inertial system obtained from CFD calculations can be expressed by the following formula. x C y C z Converted to the axial force C in a quasi-bullet system (subscript b is short for body). xb Normal force C yb Lateral force C zb Where α is both the angle of attack and the cone angle, and θ is the roll angle of the conical motion; secondly, the cross-sectional data is screened, and the normal force and lateral force coefficients of the characteristic cross-section are obtained by integration according to the data processing flow of the projectile body and wing / rudder; multiple characteristic cross-sections are taken along the longitudinal axis of the projectile, and the data is processed one by one to obtain the distributed aerodynamic characteristics of the projectile at the characteristic moment; further, according to the aerodynamic characteristic variation law and frequency of the same cross-section at different times, an appropriate order of Fourier series is selected to fit the cross-sectional aerodynamic characteristics to obtain the time-averaged distributed aerodynamic characteristics.

[0079]

[0080] Using the CFD-based method for acquiring the distributed aerodynamic characteristics of a projectile-launched rocket proposed in this invention, the time-averaged distributed aerodynamic characteristics of a rotating projectile-launched rocket were calculated and obtained. The effects of canard and conical motion on the distributed force characteristics were investigated. The numerical calculation conditions are shown in Table 1, with a total of 7 conditions. The numerical calculation used a time step Δt = 0.00001s, and the calculation stopped after one complete motion cycle. The CFD result files obtained at different times were used as the initial values ​​for calculating the time-averaged distributed aerodynamic characteristics.

[0081] Table 1 Calculation Operating Condition Statistics

[0082]

[0083] When Ma = 2.5, the distribution curves of the average lateral force coefficient along the projectile axis for rotating projectiles with and without canards are as follows: Figure 6a and Figure 6b As shown in the figure, the curve and C z The area enclosed by (x / L) = 0 represents the total average lateral force coefficient of the entire missile, from which the yaw moment coefficient can also be easily obtained. The results curves show that the lateral force coefficients in the canard and tail fin areas are positive, while those in the middle of the missile body are negative. From the trailing edge of the canard to the leading edge of the tail fin, the flow near the missile body is disturbed by the canard, resulting in a significantly increased lateral force coefficient compared to the case without canards. In summary, after adding canards, the positive lateral force of the canards is limited, the negative lateral force coefficient of the missile body increases significantly, while the positive lateral force coefficient of the wing / rudder increases less, and the lateral force coefficient of the missile body dominates.

[0084] When Ma = 2.5, the distribution curves of the time-averaged aerodynamic force coefficients along the projectile axis for a spinning projectile without canards and a conical spinning projectile are as follows: Figure 7a and Figure 7b As shown in the figure, the curve and C n (x / L)=0, C z The area enclosed by (x / L) = 0 represents the total average normal force and total lateral force coefficient of the entire missile. The resulting curves show that conical motion has virtually no additional impact on the normal force distribution; the warhead and tail fins are the main sources of normal force, while the contribution of the middle body is very small. Conical motion induces significant changes in the lateral force distribution. When α = 4°, the lateral force distribution reverses direction in the middle and rear body. When α = 20.2°, the negative lateral force coefficient of the forebody and the positive lateral force coefficient of the tail fin both increase, and the resultant force approaches zero but forms a torque that generates a yaw moment.

[0085] 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 technical principles 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 obtaining distributed aerodynamic characteristics of projectiles and rockets based on CFD, characterized in that, The method includes the following steps: Step 1: Based on the projectile cross-sectional data, including the pressure and viscous force of the projectile cross-section, obtained by CFD numerical calculation, the resultant normal force and resultant lateral force on the cross-section are obtained by circumferential angular integration; Step 2: Based on the wing / rudder section data including wing / rudder section pressure and viscous force obtained from CFD numerical calculation, directly integrate the pressure and viscous force along the wingspan direction to obtain the resultant normal force and resultant lateral force on the section. Step 3: When the projectile has unsteady motion forms including rotation, conical motion, and bending deformation, first convert the CFD numerical calculation results into CFD numerical calculation results in the quasi-projectile coordinate system through coordinate transformation. Then, referring to Step 1 and Step 2, obtain the distributed aerodynamic characteristics at a specific moment from the CFD numerical calculation results in the quasi-projectile coordinate system. Finally, use Fourier series to fit the distributed aerodynamic characteristics at different moments to obtain the time-averaged distributed aerodynamic characteristics.

2. The method for obtaining distributed aerodynamic characteristics of projectiles and rockets based on CFD as described in claim 1, characterized in that, In step one, for the wing area of ​​the missile body, the presence of the tail fin causes a discontinuity in the missile body data along the circumferential angle. Therefore, the missile body cross-sectional data is divided into segments, each segment is integrated, and the resultant force is calculated.

3. The method for obtaining distributed aerodynamic characteristics of projectiles and rockets based on CFD as described in claim 2, characterized in that, In step one, the projectile cross-section data obtained by the CFD numerical calculation includes grid node coordinates, pressure, and viscous force. When integrating along the circumferential angle, the projectile cross-section data are arranged in ascending order of circumferential angle; If there are discontinuities in the wing section cross-section data along the circumferential angle, the cross-section data is divided into sections according to the number and position of the tail fins; the resultant force in the normal direction and the resultant force in the lateral direction are obtained by integrating the local circumferential angle to the normal and lateral directions.

4. The method for obtaining distributed aerodynamic characteristics of projectiles and rockets based on CFD as described in claim 3, characterized in that, In step two, since the tail fin may have straight, bent, or curved configurations, its windward / leeward or concave / convex surface data are separated and then superimposed to obtain the resultant force.

5. The method for obtaining distributed aerodynamic characteristics of projectiles and rockets based on CFD as described in claim 4, characterized in that, In step two, the wing / rudder section data obtained by CFD numerical calculation includes grid node coordinates, pressure, and viscous force. When integrating along the wingspan, the wing / rudder section data are arranged in ascending order of wingspan, and the windward / leeward or concave / convex surface data are separated. Based on the local azimuth normal and lateral projections, the resultant normal force and lateral force are obtained by integration.

6. The method for obtaining distributed aerodynamic characteristics of projectiles and rockets based on CFD as described in claim 5, characterized in that, In step three, the CFD numerical calculation results are transformed into the results of the quasi-bullet system through coordinate transformation between the inertial frame and the quasi-bullet system. Based on the aerodynamic characteristics and frequency changes of the missile body and wing / rudder sections at different times, a Fourier series of appropriate order is selected for fitting to obtain the time-averaged distributed aerodynamic characteristics.

7. A method for obtaining distributed aerodynamic characteristics of projectiles and rockets based on CFD, characterized in that, The method includes the following steps: Step 1: When the projectile is flying horizontally without other motion, the pressure and viscous forces obtained from steady-state CFD calculations are processed to obtain the aerodynamic characteristics of the projectile surface at different cross-sectional positions along the longitudinal axis of the projectile; Step 1 includes: Step 11: At different projectile cross-sectional positions, project the pressure and viscous forces along the circumferential angle and integrate to obtain the resultant normal force and resultant lateral force on the cross-section; Step 12: At different wing / rudder cross-section positions, project the pressure and viscous forces along the azimuth angle and integrate them along the wingspan direction to obtain the resultant normal force and resultant lateral force on the cross-section; Step 13: Superimpose the forces on the missile body and wings / rudders to obtain the distributed aerodynamic characteristics along the longitudinal axis of the missile body; Step 2: When the projectile has unsteady motion forms including rotation, conical motion, and bending deformation, the pressure and viscous force at different times within a motion cycle are obtained through CFD calculation; The unsteady CFD values ​​obtained from CFD calculations are converted into results in the quasi-projectile coordinate system through coordinate transformation. Using the method for obtaining the distributed aerodynamic characteristics under steady conditions in step 1 above, the distributed aerodynamic characteristics of the projectile and rocket at different times are obtained. Finally, based on the aerodynamic characteristics and frequency of the missile body and wing / rudder sections at different times, a Fourier series of appropriate order is selected for fitting, and the time-averaged distributed aerodynamic characteristics are obtained by fitting the results at different times.

8. The method for obtaining distributed aerodynamic characteristics of projectiles and rockets based on CFD as described in claim 7, characterized in that, The steady CFD calculation uses solvers such as Fluent, CFD++, and CFX. The calculated projectile cross-section data and wing / rudder cross-section data include grid node coordinates, pressure, and viscous forces. When obtaining distributed aerodynamic forces, data from characteristic cross-sections are selected for processing.

9. The method for obtaining distributed aerodynamic characteristics of projectiles and rockets based on CFD as described in claim 8, characterized in that, When the projectile cross-sectional data is projected and integrated along the circumferential angle, the data is arranged in ascending order of circumferential angle. For a wing-type projectile, the presence of tail fins causes discontinuities in the projectile cross-sectional data along the circumferential angle. Therefore, the projectile cross-sectional data is segmented according to the number and position of tail fins. Each segment is projected onto the normal and lateral directions according to the local circumferential angle, and then integrated to obtain the resultant force in the normal direction and the resultant force in the lateral direction.

10. The method for obtaining distributed aerodynamic characteristics of projectiles and rockets based on CFD as described in claim 8, characterized in that, When integrating the wing / rudder section data along the wingspan direction, the wing / rudder section data are arranged in ascending order of wingspan. For tail fins with different configurations, including straight, bent, and rolled tail fins, their windward / leeward or concave / convex surface data are separated. Based on the local azimuth angle normal and lateral projections, the normal resultant force and lateral resultant force are obtained by integration.

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