A method and system for calculating pneumatic load distribution

Abstract

A method and system for calculating the pneumatic load distribution according to the present invention, the method steps are as follows: 1) Divide the pneumatic load distribution into multiple segments along the axial direction; 2) According to the segmented positions and the pneumatic load distribution laws of each segment, establish a correlation function between the normal force coefficient distributions of each segment; 3) According to the pneumatic load distribution, integrate to obtain the corresponding normal force coefficient and center of pressure coefficient values of the aircraft, and establish an equation relationship with the normal force coefficient and center of pressure coefficient obtained from the overall aerodynamic force of the aircraft; 4) Substitute the correlation function into the established equation relationship to calculate and obtain the normal pneumatic load distribution of the aircraft.

Description

Technical Field

[0001] The present invention belongs to the fields of aerodynamic prediction and load design of aircraft, and mainly relates to the calculation of the aerodynamic load distribution of aircraft, and can be applied to the rapid calculation and determination of the aerodynamic load distribution of winged aircraft. Background Technique

[0002] Aerodynamic load is the aerodynamic force generated by the airflow and acting on the aircraft when the aircraft moves at high speed in the air. Due to the different degrees of influence and different action laws of the interaction between different parts of the aircraft and the airflow, there are huge differences in the aerodynamic loads received by different parts. The aerodynamic load distribution is the distribution of the aerodynamic load at different positions of the aircraft, and is usually expressed by the aerodynamic force or aerodynamic coefficient per unit area or unit length, or by the total aerodynamic force or aerodynamic coefficient within a certain length range or area range. In the structural design of the aircraft, it is necessary to accurately determine the aerodynamic load conditions of different parts, which is one of the important bases for structural strength design. Among them, the lateral (normal) aerodynamic load distribution (hereinafter referred to as the aerodynamic load distribution) is the most important part in the calculation of the lateral load (bending moment load) of the aircraft, and is also the most important design basis parameter in the structural strength design of the aircraft.

[0003] The aerodynamic force received by the whole aircraft is usually obtained by the computational fluid dynamics method (CFD) or the wind tunnel force measurement test method. In recent years, the artificial intelligence machine learning method based on the surrogate model has been gradually used to quickly obtain the aerodynamic force received by the whole aircraft, and good engineering application effects have been achieved. Due to the interaction between different parts of the aircraft surface and between the surface and the flow field, there are too many geometric parameters and oncoming flow parameters that affect the aerodynamic load distribution, making it difficult to quickly obtain the aerodynamic load distribution data by the artificial intelligence machine learning method based on the surrogate model, and it is also necessary to further solve the consistency problem between the result after integrating the aerodynamic load distribution data and the aerodynamic force received by the whole aircraft. The algorithm is too complex to quickly obtain the aerodynamic load distribution by the surrogate model method.

[0004] At the present stage, the aerodynamic load distribution data of the aircraft still needs to be obtained through CFD numerical simulation calculation or subsonic-transonic-supersonic pressure measurement wind tunnel test. In the preliminary design stage, the pressure or pressure coefficient distribution on the aircraft surface is obtained by CFD numerical simulation and iterative solution of the N-S equation. A typical CFD numerical calculation of the surface pressure distribution cloud map of a bundled rocket is shown in Figure 1 . Furthermore, the aerodynamic load distribution data of different parts is obtained by integrating according to the cross-section position of the aircraft: that is, the distribution of the normal force or normal force coefficient per unit length along the surface position of the aircraft. Since the iterative solution of the N-S equation takes a long time, it takes 40 hours to obtain the aerodynamic load distribution under typical design conditions of the aircraft by CFD numerical calculation.

[0005] In the formal design stage, through the subsonic and supersonic pressure measurement wind tunnel test, the pressure coefficient distribution on the surface of the scaled model is obtained, and then the aerodynamic load distribution data of the model is obtained by integrating according to the model cross-section position. During the test, a series of pressure measurement profiles and as many pressure measurement points as possible need to be set on the surface of the scaled model (the characteristic length of the model is about 1m). At the positions where the pressure change gradient is large, the number of profiles and pressure measurement points need to be increased to obtain more accurate pressure distribution data, as shown in Figure 2 .

[0006] Due to the systematic error of the integral calculation caused by test measurement and the discreteness of measurement points, the normal force coefficient, drag coefficient, center of pressure coefficient, etc. of the whole rocket obtained by integrating the pressure coefficient on the surface of the rocket body in the pressure measurement test will inevitably deviate from the normal force coefficient, drag coefficient, center of pressure coefficient, etc. of the whole rocket directly measured in the force measurement test, as shown in Figure 3 .

[0007] It is necessary to correct the deviation of the pressure measurement test data:

[0008] 1) Judge the correctness of the data at each pressure measurement point, and eliminate / correct the pressure data of the measurement bad points according to the symmetry, the pressure distribution law under adjacent pressure measurement profiles or adjacent Mach numbers;

[0009] 2) Integrate the pressure distribution to obtain the aerodynamic load distribution of the model, and correct the test pressure measurement data according to the characteristics and laws of the aerodynamic load distribution on the surface of the aircraft;

[0010] 3) According to the characteristics and laws of the pressure distribution on the surface of the aircraft, the distribution characteristics and laws of the aerodynamic load, the sources, characteristics and laws of the discrete deviation of the pressure measurement test, combined with the CFD calculation, continuously adjust the aerodynamic load distribution data of each profile until the aerodynamic load distribution data conforms to the characteristics of the specific aircraft aerodynamic shape and Ma, and the integral result is consistent with the normal force coefficient and center of pressure coefficient of the whole rocket measured in the force measurement test.

[0011] In the data processing process, steps 1) → 3) need to be repeatedly corrected many times, which is time-consuming and requires high professional ability of designers. They can accurately judge the correctness of the test data, the accuracy of the CFD calculation, master the characteristics of the test deviation, and have a strong grasp of the aerodynamic pressure distribution characteristics and aerodynamic load distribution characteristics of specific shapes and different Ma. Through the subsonic and supersonic pressure measurement wind tunnel test, it takes 4 to 8 months to obtain the aerodynamic load distribution of the typical aircraft shape, and the data processing takes 2 to 4 weeks.

[0012] It can be seen that although the design method of the aerodynamic load distribution is theoretically mature, the process is complex, cumbersome, time-consuming, requires high professional ability, is difficult to achieve rapid design, is not conducive to the rapid demonstration of the design scheme, and a calculation method that can quickly obtain the aerodynamic load distribution of the aircraft needs to be developed. Summary of the Invention

[0013] The technical problem solved by the present invention is: to overcome the deficiencies of the prior art and provide a method and system for quickly calculating the aerodynamic load distribution of an aircraft, and the calculation accuracy meets the engineering use requirements of aircraft project demonstration, multi-scheme comparison, and structural weight estimation.

[0014] The technical solution of the present invention is: a method for quickly calculating the aerodynamic load distribution, including:

[0015] 1) Divide the aerodynamic load distribution into multiple segments along the axial direction;

[0016] 2) According to the segmented positions and the aerodynamic load distribution laws of each segment, establish a correlation function between the normal force coefficient distributions of each segment;

[0017] 3) According to the aerodynamic load distribution, integrate to obtain the corresponding normal force coefficient and center of pressure coefficient values of the aircraft, and establish an equation relationship with the normal force coefficient and center of pressure coefficient obtained from the overall aerodynamic force of the aircraft;

[0018] 4) Substitute the correlation function into the established equation relationship to calculate and obtain the normal aerodynamic load distribution of the aircraft.

[0019] The aerodynamic load distribution is divided into nine segments along the axial direction, including: from the head x0 to the cone-cylinder interface or the double-cone interface x1; from the cone-cylinder interface x1 to the front end face of the inverted cone x2; from the front end face of the inverted cone x2 to the rear limit of the inverted cone influence area x3; from the rear limit of the inverted cone influence area x3 to the front limit of the positive cone influence area x4; from the front limit of the positive cone influence area x4 to the rear end face of the positive cone x5; from the rear end face of the positive cone x5 to the rear limit of the positive cone influence area x6; from the rear limit of the positive cone influence area x6 to the front limit of the tail wing influence area x7; from the front limit of the tail wing influence area x7 to the leading edge point of the tail wing tip x8; from the leading edge point of the tail wing tip x8 to the bottom of the core stage x9.

[0020] The establishment of the correlation function between the normal force coefficient distributions of each segment includes:

[0021] Take the normal force coefficients corresponding to the positions of the cone-cylinder interface x1 and the leading edge point of the tail wing tip x8, that is, w1 and w8, as the unknowns to be solved; establish the relationship between other parameters and w1, w8 according to the variation laws between each segment:

[0022] w0 = 0;

[0023] w2 = k 21 ·w1;

[0024] w3 = w2;

[0025] w4 = k 46 ·w6;

[0026] w5 = k 51 ·w1;

[0027] w6 = w7;

[0028] w7 = k 78 ·w8;

[0029] w9 = w8; (1)

[0030] where w1 - w9 are the normal force coefficients at x1 - x9; k 21 is the correlation function between w1 and w2, k 46 is the correlation function between w4 and w6, k 51 is the correlation function between w1 and w5, k 78 is the correlation function between w7 and w8.

[0031] The specific forms of each correlation function are as follows:

[0032]

[0033]

[0034]

[0035]

[0036] In the formula:

[0037] α z1 is the cone angle of the first cone of the vehicle's conical section;

[0038] α dz is the cone angle of the vehicle's inverted cone, which is negative;

[0039] α zz is the cone angle of the positive cone of the vehicle's core stage;

[0040] D zhui is the diameter of the cylindrical section between X3 and X4 of the vehicle;

[0041] D xin is the diameter of the cylindrical section between X6 and X7 of the vehicle;

[0042] L tip is the chord length of the wing tip of the tail wing;

[0043] L root is the chord length of the wing root of the tail wing.

[0044] Establishing an equation relationship with the normal force coefficient and center of pressure coefficient obtained from the overall aerodynamic force of the vehicle includes:

[0045]

[0046] In the formula:

[0047] Δxi is the integral segment length;

[0048] x i is the axial coordinate where the integral segment is located;

[0049] C N is the normal force coefficient of the aircraft;

[0050] Xcp is the center of pressure coefficient of the aircraft;

[0051] L ref is the reference length of the aircraft.

[0052] The calculation to obtain the normal aerodynamic load distribution of the aircraft includes:

[0053] Substituting Equation (1) into Equation (3) can obtain a set of non - linear binary equations. Solving the equations of Equation (3) by the Newton - Raphson method to obtain the solutions of w1 and w8; substituting the obtained w1 and w8 into Equation (1) to get the parameters at other positions, thereby obtaining the normal aerodynamic load distribution of the aircraft.

[0054] A fast calculation system for aerodynamic load distribution includes a segmentation module, a correlation function establishment module, an equation relationship establishment module, and a calculation module;

[0055] The segmentation module divides the aerodynamic load distribution into multiple segments along the axis;

[0056] The correlation function establishment module establishes the correlation function between the normal force coefficient distributions of each segment according to the segmentation positions and the aerodynamic load distribution rules of each segment;

[0057] The equation relationship establishment module integrates according to the aerodynamic load distribution to obtain the corresponding values of the normal force coefficient and the center of pressure coefficient of the aircraft, and establishes an equation relationship with the normal force coefficient and the center of pressure coefficient obtained from the overall aerodynamic force of the aircraft;

[0058] The calculation module substitutes the correlation function into the established equation relationship to calculate and obtain the normal aerodynamic load distribution of the aircraft.

[0059] The segmentation module divides the aerodynamic load distribution into nine segments along the axis, including: from the head x0 to the cone - cylinder interface or the double - cone interface x1; from the cone - cylinder interface x1 to the front end face of the inverted cone x2; from the front end face of the inverted cone x2 to the rear limit of the inverted cone influence area x3; from the rear limit of the inverted cone influence area x3 to the front limit of the positive cone influence area x4; from the front limit of the positive cone influence area x4 to the rear end face of the positive cone x5; from the rear end face of the positive cone x5 to the rear limit of the positive cone influence area x6; from the rear limit of the positive cone influence area x6 to the front limit of the tail wing influence area x7; from the front limit of the tail wing influence area x7 to the leading edge point of the tail wing tip x8; from the leading edge point of the tail wing tip x8 to the bottom of the core stage x9.

[0060] The correlation function establishment module establishes the correlation function between the normal force coefficient distributions of each segment, including:

[0061] Take the normal force coefficients corresponding to the cone-cylinder interface x1 and the leading edge point x8 of the wing tip, namely w1 and w8, as the unknowns to be solved; establish the relationships between other parameters and w1, w8 according to the variation rules between each section:

[0062] w0 = 0;

[0063] w2 = k 21 ·w1;

[0064] w3 = w2;

[0065] w4 = k 46 ·w6;

[0066] w5 = k 51 ·w1;

[0067] w6 = w7;

[0068] w7 = k 78 ·w8;

[0069] w9 = w8;

[0070] Among them, w1 - w9 are the normal force coefficients at x1 - x9; k 21 is the correlation function between w1 and w2, k 46 is the correlation function between w4 and w6, k 51 is the correlation function between w1 and w5, k 78 is the correlation function between w7 and w8.

[0071] The equation relationship establishment module establishes equation relationships including:

[0072]

[0073]

[0074] In the formula:

[0075] Δx i is the integral section length;

[0076] x i is the axial coordinate where this integral section is located;

[0077] C N is the normal force coefficient of the aircraft;

[0078] Xcp is the center of pressure coefficient of the aircraft;

[0079] L ref is the reference length of the aircraft.

[0080] The advantages of the present invention compared with the prior art are as follows:

[0081] Although the conventional design method of aerodynamic load distribution is theoretically mature, the process is complex, time-consuming, requires high professional capabilities, and it is difficult to achieve rapid design. The conventional method is to obtain the aerodynamic load distribution by using numerical simulation or subsonic-transonic-supersonic pressure-measuring wind tunnel tests. Numerical simulation: build a grid model, iteratively solve the N-S equations to obtain the pressure distribution on the rocket body surface, and then integrate according to the section position to obtain the load distribution, with a characteristic time consumption of 3 - 5 days. For wind tunnel tests, hundreds of pressure-measuring holes are drilled on the surface of the scaled model, and the pressure coefficient distribution on the model surface is obtained by blowing air in the wind tunnel. Then, the load distribution of the model is obtained by integrating according to the section, and the actual rocket body load distribution is obtained according to the scaling criterion. Due to systematic errors such as measurement and measurement point discreteness, there must be deviations between the normal force of the whole rocket, the center of pressure coefficient, etc. obtained by integrating through the pressure-measuring test and the coefficients directly measured by the force-measuring test. It is necessary to combine numerical calculations, eliminate bad points, and continuously correct the pressure coefficient of the measurement points and the discrete section integration error according to the load distribution law of a specific shape and the characteristics of the test discreteness deviation until the load distribution conforms to the aerodynamic law and the integration result is consistent with the aerodynamic coefficients of the whole rocket measured by the force-measuring test. This data correction process requires multiple repetitions, consumes a huge amount of time, and has high requirements for professional capabilities. It can accurately judge the accuracy of numerical calculations and tests, master various deviation sources and characteristics, and have a strong grasp of the pressure and load distribution characteristics of a specific shape and each Mach number. It takes 4 - 8 months for a mature designer to obtain the load distribution through tests. This method is based on the statistical analysis of a large amount of test data, extracts the aerodynamic load distribution law of slender body aircraft. Select the normal force coefficient per unit length at the interface of the nose cone - cylinder and the tail fin as independent variables, and establish an algebraic equation between the normal force coefficient per unit length at other positions and the independent variables. According to the integration result of the normal aerodynamic load distribution (the normal force coefficient of the whole rocket, the center of pressure) being equal to the aerodynamic coefficients of the whole rocket, and on the premise of knowing the aerodynamic coefficients of the whole rocket, use Newton iteration to solve and determine the independent variable parameters at the two key positions, and then obtain the aerodynamic load distribution according to the algebraic equation. This method can achieve rapid calculation of the aerodynamic load distribution, with a characteristic time consumption of seconds, and can strongly support the rapid scheme demonstration and structural weight estimation during rocket design. Description of the Drawings

[0082] Figure 1 It is a schematic diagram of the surface pressure contour distribution of the aircraft.

[0083] Figure 2 It is a schematic diagram of the surface pressure measurement point distribution of the wind tunnel pressure-measuring test model.

[0084] Figure 3 It is a comparison between the integration result of the pressure-measuring test of a certain aircraft and the measurement result of the whole rocket force-measuring test.

[0085] Figure 4 It is a schematic diagram of the cross-section change position of the slender body aircraft.

[0086] Figure 5 It is a schematic diagram summarizing the distribution law of the normal force coefficient of a slender aircraft.

[0087] Figure 6 It is a schematic diagram showing the segmentation and parameter representation of the aerodynamic load distribution of a slender aircraft.

[0088] Figure 7 It is the aerodynamic shape of a typical winged aircraft.

[0089] Figure 8 It is a comparison of the calculation results of the normal force coefficient distribution.

[0090] Figure 9 It is a comparison of the calculation results of the aerodynamic bending moment load.

[0091] Figure 10 It is a comparison of the calculation results of the total bending moment load. Specific implementation manners

[0092] As Figure 4 shown, large change gradients of the surface pressure and pressure coefficient of the aircraft occur at the positions where the cross-section changes, thereby generating large changes in the normal force coefficient and the aerodynamic load distribution. Taking the slender aircraft as an example, the positions where the cross-section of the aircraft changes are sorted out, including the nose, conical section, inverse cone, normal cone, tail wing, etc.

[0093] Through statistical analysis, the law of the normal aerodynamic load is refined into a distribution curve as Figure 5 shown. The law of its segmented aerodynamic load distribution is as follows:

[0094] 1) For a blunt body, starting from the actual tip point, its aerodynamic load gradually increases; when reaching the cone-cylinder interface, due to the cross-section change causing air flow expansion, the aerodynamic load will decrease; if there is an inverse cone, its load will further decrease, usually becoming negative, and the influence distance of the air flow expansion in the inverse cone section is 1.5 times the length of the inverse cone section.

[0095] 2) For the straight section such as the core stage, its aerodynamic load can be considered to basically remain constant.

[0096] 3) For the normal cone section of the core stage, its aerodynamic load will increase to a certain extent. Since the normal cone angle is usually small, the increase in the aerodynamic load is usually small. When transitioning from the cone section to the straight section, the aerodynamic load will also decrease to a certain extent, and this decrease is usually also small. The maximum value of the aerodynamic load in this normal cone section is usually less than the peak value of the aerodynamic load at the fairing.

[0097] 4) For the position of the fin, its aerodynamic load will increase significantly. Due to the influence of flow field interference, the position where its aerodynamic load begins to increase will be before the vertex of the fin root chord, and the distance is taken as the chord length of the leading edge wedge section of the fin. It reaches the maximum near the axial position where the leading edge vertex of the fin tip is located and continues to the trailing edge of the fin.

[0098] A fast calculation method for aerodynamic load distribution of the present invention includes:

[0099] 1) Divide the aerodynamic load distribution into multiple segments along the axial direction;

[0100] 2) According to the segmented positions and the aerodynamic load distribution laws of each segment, establish the correlation function between the normal force coefficient distributions of each segment;

[0101] 3) According to the aerodynamic load distribution, integrate to obtain the corresponding normal force coefficient and center of pressure coefficient values of the aircraft, and establish an equation relationship with the normal force coefficient and center of pressure coefficient obtained from the overall aerodynamic force of the aircraft;

[0102] 4) Substitute the correlation function into the established equation relationship to calculate and obtain the normal aerodynamic load distribution of the aircraft.

[0103] As Figure 6 shown, divide the aerodynamic load distribution into nine segments along the axial direction as follows:

[0104] 1) From the nose to the cone-cylinder interface (single cone or ogive nose) or the double-cone interface (double cone) (x0~x1);

[0105] 2) From the cone-cylinder interface (or double-cone interface) to the front end face of the frustum cone (x1~x2);

[0106] 3) From the front end face of the frustum cone to the rear limit of the frustum cone influence area (x2~x3);

[0107] 4) From the rear limit of the frustum cone influence area to the front limit of the positive cone influence area (x3~x4);

[0108] 5) From the front limit of the positive cone influence area to the rear end face of the positive cone (x4~x5)

[0109] 6) From the rear end face of the positive cone to the rear limit of the positive cone influence area (x5~x6);

[0110] 7) From the rear limit of the positive cone influence area to the front limit of the fin influence area (x6~x7);

[0111] 8) From the front limit of the fin influence area to the leading edge point of the fin tip (x7~x8);

[0112] 9) From the leading edge point of the fin tip to the bottom of the core stage (x8~x9);

[0113] Table 1 Meanings of each parameter

[0114]

[0115]

[0116] According to the sectional position and the aerodynamic load distribution law of each section, establish the correlation function between the normal force coefficient distributions of each section. For the configuration with fins, the normal force coefficient distribution usually has two relatively large peaks, which appear at the cone-cylinder interface of the blunt head of the rocket body and at the fins respectively. Therefore, the Cnpm corresponding to the positions of X1 and X8 (i.e., w1 and w8) are taken as the unknowns to be solved. Establish the relationships between other parameters (w0, w2, w3, w4, w5, w6, w7, w9) and w1, w8 according to the variation law between each section as follows:

[0117] w0 = 0;

[0118] w2 = k 21 ·w1;

[0119] w3 = w2;

[0120] w4 = k 46 ·w6;

[0121] w5 = k 51 ·w1;

[0122] w6 = w7;

[0123] w7 = k 78 ·w8;

[0124] w9 = w8; (1)

[0125] Wherein, k 21 is the correlation function between w1 and w2, k 46 is the correlation function between w4 and w6, k 51 is the correlation function between w1 and w5, k 78 is the correlation function between w7 and w8. Specifically as follows:

[0126]

[0127]

[0128]

[0129]

[0130] In the formula:

[0131] α z1 is the cone angle of the first cone of the cone section of the aircraft;

[0132] α dz The cone angle of the inverted cone of the aircraft, which is negative;

[0133] α zz is the cone angle of the positive cone of the vehicle core stage;

[0134] D zhui is the diameter of the cylindrical section between X3 and X4 of the vehicle;

[0135] D xin is the diameter of the cylindrical section between X6 and X7 of the vehicle;

[0136] L tip is the chord length of the wingtip of the tail wing;

[0137] L root is the chord length of the wing root of the tail wing.

[0138] According to the above aerodynamic load distribution, the corresponding normal force coefficient and center of pressure coefficient values of the vehicle are obtained by integration, and an equation relationship is established with the normal force coefficient and center of pressure coefficient obtained from the overall aerodynamic force of the vehicle, that is

[0139]

[0140] In the formula:

[0141] Δx i is the integral segment length;

[0142] x i is the axial coordinate where the integral segment is located;

[0143] C N is the normal force coefficient of the vehicle;

[0144] Xcp is the center of pressure coefficient of the vehicle;

[0145] L ref is the reference length of the vehicle;

[0146] Substituting Equation (1) into Equation (3) can obtain a set of nonlinear binary equations. C N and Xcp are obtained by methods such as CFD and wind tunnel measurement tests in actual operation. C N and Xcp of conventional shapes can also be conveniently obtained by artificial intelligence machine learning. In this method, C N and Xcp are used as known terms that have been obtained. On this basis, the equations in Equation (3) are solved by the Newton iteration method to obtain the solutions of w1 and w8. Substituting the obtained w1 and w8 into Equation (1) can obtain the parameters (w0, w2, w3, w4, w5, w6, w7, w9) at other positions, thereby determining Figure 6 the curve shape and obtaining the normal aerodynamic load distribution of the vehicle.

[0147] In addition, it also relates to a rapid calculation system for aerodynamic load distribution, including a segmentation module, a correlation function establishment module, an equation relationship establishment module, and a calculation module;

[0148] The segmentation module divides the aerodynamic load distribution into multiple segments along the axial direction;

[0149] The correlation function establishment module establishes a correlation function between the normal force coefficient distributions of each segment according to the segmentation position and the aerodynamic load distribution law of each segment;

[0150] The equation relationship establishment module integrates according to the aerodynamic load distribution to obtain the corresponding normal force coefficient and center of pressure coefficient values of the aircraft, and establishes an equation relationship with the normal force coefficient and center of pressure coefficient obtained from the overall aerodynamic force of the aircraft;

[0151] The calculation module substitutes the correlation function into the established equation relationship to calculate and obtain the normal aerodynamic load distribution of the aircraft.

[0152] Taking Figure 7 the typical winged aircraft shown as an example, the comparison of the normal force coefficient distribution curves obtained by the CFD numerical calculation method and the rapid calculation method of this paper is shown in Figure 8 . Furthermore, taking the normal force coefficient distributions obtained by the two calculation methods as inputs, the bending moment load generated by the aerodynamic force and the total bending moment load of the entire rocket are calculated, and the comparison curve of the two is shown in Figures 8 to 10 . It can be seen from the comparison of the curves in the figure that the maximum local deviation of the aerodynamic bending moment load profile is 30% (Ma = 0.8) and 15% (Ma = 1.2), and the maximum deviation of the total bending moment load profile is 8% (Ma = 0.8) and 2% (Ma = 1.2), meeting the requirement that the total bending moment calculation deviation is not greater than 20%, and meeting the engineering use requirements of aircraft scheme demonstration and structural weight estimation for multi-scheme comparison.

[0153] Although the present invention has been disclosed above with preferred embodiments, it is not intended to limit the present invention. Any person skilled in the art can make possible changes and modifications to the technical solution of the present invention without departing from the spirit and scope of the present invention. Therefore, any simple modification, equivalent change, and modification made to the above embodiments based on the technical essence of the present invention without departing from the technical solution of the present invention shall fall within the protection scope of the technical solution of the present invention.

Claims

1. A rapid calculation method for pneumatic load distribution, characterized in that Including: Axially dividing the aerodynamic load distribution into multiple segments; According to the segmented positions and the aerodynamic load distribution laws of each segment, establishing a correlation function between the normal force coefficient distributions of each segment; According to the aerodynamic load distribution, integrating to obtain the corresponding normal force coefficient and center of pressure coefficient values of the aircraft, and establishing an equation relationship with the normal force coefficient and center of pressure coefficient obtained from the overall aerodynamic force of the aircraft; Substituting the correlation function into the established equation relationship to calculate and obtain the aerodynamic normal load distribution of the aircraft; The aerodynamic load distribution is axially divided into nine segments, including: from the nose x0 to the cone-cylinder interface or the double-cone interface x1; from the cone-cylinder interface x1 to the front end face of the inverted cone x2; from the front end face of the inverted cone x2 to the rear limit of the inverted cone influence area x3; from the rear limit of the inverted cone influence area x3 to the front limit of the positive cone influence area x4; from the front limit of the positive cone influence area x4 to the rear end face of the positive cone x5; from the rear end face of the positive cone x5 to the rear limit of the positive cone influence area x6; from the rear limit of the positive cone influence area x6 to the front limit of the tail wing influence area x7; from the front limit of the tail wing influence area x7 to the leading edge point of the tail wing tip x8; from the leading edge point of the tail wing tip x8 to the bottom of the core stage x9; The establishing of the correlation function between the normal force coefficient distributions of each segment includes: Taking the normal force coefficients corresponding to the positions of the cone-cylinder interface x1 and the leading edge point of the tail wing tip x8, i.e., w1 and w8, as the unknowns to be solved; establishing the relationships between other parameters and w1, w8 according to the variation laws between each segment: w0＝0； w2 = k 21 · w1; w3 = w2; w4 = k 46 ·w6; w5 = k 51 · w1; w6 = w7; w7 = k 78 · w8; w9 = w8; (1) Among them, w1 - w9 are the normal force coefficients at x1 - x9; k 21 is the correlation function between w1 and w2, k 46 is the correlation function between w4 and w6, k 51 is the correlation function between w1 and w5, k 78 is the correlation function between w7 and w8; The specific forms of each correlation function are as follows: Where: α z1 is the cone angle of the first cone of the cone section of the aircraft; α dz The inverted cone angle of the aircraft is negative; α zz is the cone angle of the forward cone of the vehicle core stage; D zhui is the diameter of the column section between aircraft X3 and X4; D xin is the diameter of the column section between aircraft X6 and X7; L tip Trailing edge wingtip chord length; L root Root chord length of the tail fin The establishing of the equation relationship with the normal force coefficient and center of pressure coefficient obtained from the overall aerodynamic force of the aircraft includes: Where: Δx i is the integral segment length; x i is the axial coordinate where the integral segment is located; C N is the normal force coefficient of the aircraft; Xcp is the center of pressure coefficient of the aircraft; L ref is the reference length of the aircraft; The calculating and obtaining of the aerodynamic normal load distribution of the aircraft includes: Substituting Equation (1) into Equation (3) to obtain a set of nonlinear binary equations, and solving the equations of Equation (3) by the Newton iteration method to obtain the solutions of w1 and w8; substituting the obtained w1 and w8 into Equation (1) to obtain the parameters at other positions, so as to obtain the aerodynamic normal load distribution of the aircraft.

2. A fast calculation system for pneumatic load distribution, characterized in that: Including a segmentation module, a correlation function establishing module, an equation relationship establishing module, and a calculation module; The segmentation module axially divides the aerodynamic load distribution into multiple segments; The correlation function establishing module establishes a correlation function between the normal force coefficient distributions of each segment according to the segmented positions and the aerodynamic load distribution laws of each segment; The equation relationship establishing module integrates according to the aerodynamic load distribution to obtain the corresponding normal force coefficient and center of pressure coefficient values of the aircraft, and establishes an equation relationship with the normal force coefficient and center of pressure coefficient obtained from the overall aerodynamic force of the aircraft; The calculation module substitutes the correlation function into the established equation relationship to calculate and obtain the aerodynamic normal load distribution of the aircraft; The segmented module divides the aerodynamic load distribution into nine segments along the axial direction, including: from the nose x0 to the cone-cylinder interface or the double-cone interface x1; from the cone-cylinder interface x1 to the front end face of the inverted cone x2; from the front end face of the inverted cone x2 to the rear limit of the inverted cone influence area x3; from the rear limit of the inverted cone influence area x3 to the front limit of the positive cone influence area x4; from the front limit of the positive cone influence area x4 to the rear end face of the positive cone x5; from the rear end face of the positive cone x5 to the rear limit of the positive cone influence area x6; from the rear limit of the positive cone influence area x6 to the front limit of the fin influence area x7; from the front limit of the fin influence area x7 to the leading edge point of the fin tip x8; from the leading edge point of the fin tip x8 to the bottom of the core stage x9; The correlation function establishment module establishes the correlation functions between the normal force coefficient distributions of each segment, including: Regarding the normal force coefficients corresponding to the positions of the cone-cylinder interface x1 and the leading edge point of the fin tip x8, namely w1 and w8, as the unknowns to be solved; establishing the relationships between other parameters and w1, w8 according to the variation rules between each segment: w0＝0； w2 = k 21 ·w1; w3 = w2; w4 = k 46 ·w6; w5 = k 51 · w1; w6 = w7; w7 = k 78 · w8; w9 = w8; Among them, w1 - w9 are the normal force coefficients at x1 - x9; k 21 is the correlation function between w1 and w2, k 46 is the correlation function between w4 and w6, k 51 is the correlation function between w1 and w5, k 78 is the correlation function between w7 and w8; The equation relationship establishment module establishes equation relationships, including: In the formula: Δx i is the integral segment length; x i is the axial coordinate where the integral segment is located; C N is the normal force coefficient of the aircraft; Xcp is the aerodynamic center coefficient of the aircraft; L ref is the reference length of the aircraft.

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