Aerodynamic / structural / acoustic blast multidisciplinary rapid optimization design method for supersonic civil aircraft

Abstract

The application provides a kind of aerodynamic / structure / sound blast three-discipline rapid optimization design method for supersonic civil aircraft, comprising: determining aerodynamic design variables, structural design variables; determining the same sound blast design variables as the aerodynamic design variables; establishing the correlation between the coordinates of the aerodynamic shape surface grid points and the structural design variables; establishing the aerodynamic / structure / sound blast multidisciplinary optimization mathematical model.1.Compared with the traditional single-discipline design, the application refines the comprehensive design of the three disciplines of aerodynamics / structure / sound blast, which helps to coordinate the overall performance of the conceptual scheme of the design.2.The application correlates the aerodynamic shape variables and the structural shape variables through the aircraft surface grid points, reducing the redundancy of the design variables, thereby helping to improve the optimization efficiency.3.The application uses a multidisciplinary rapid analysis method, which can comprehensively improve the performance of the three disciplines of aerodynamics / structure / sound blast within a few hours, achieving rapid design of the supersonic civil aircraft conceptual scheme.

Description

TECHNICAL FIELD

[0001] The application belongs to the technical field of aircraft design, and particularly relates to a rapid optimization design method for aerodynamics / structure / sound blast three disciplines of supersonic civil aircraft. BACKGROUND

[0002] Quickness and efficiency are the biggest advantages of civil aviation transportation system compared with other transportation systems. However, the cruise speed of the current civil aircraft has been in a bottleneck. The cruise speed of supersonic civil aircraft can reach 2 times or even more than that of traditional subsonic civil aircraft, which can greatly improve the travel efficiency and bring more comfortable travel experience. Therefore, supersonic civil aircraft is one of the key development directions of the next generation of civil aircraft, and is also an important part of the upgrading of the aviation industry. Therefore, developing supersonic civil aircraft and related technologies has great application prospect and practical significance.

[0003] It is known that the design of an aircraft involves multiple research fields, such as aerodynamics, structure, acoustics, propulsion and control, and is a typical multi-disciplinary problem. Therefore, the design of an aircraft is generally not to pursue the optimization of single-discipline performance, but to achieve the trade-off between multiple disciplines. In the prior art, a single-discipline design method is used in the initial stage of aircraft conceptual design, so that it is difficult to quickly obtain a design concept scheme with relatively optimal overall performance. SUMMARY

[0004] In view of the defects in the prior art, the application provides a rapid optimization design method for aerodynamics / structure / sound blast three disciplines of supersonic civil aircraft, which can effectively solve the above problems.

[0005] The technical scheme adopted by the application is as follows:

[0006] The application provides a rapid optimization design method for aerodynamics / structure / sound blast three disciplines of supersonic civil aircraft, comprising the following steps:

[0007] Step 1, inputting a supersonic civil aircraft reference layout to be optimized; the supersonic civil aircraft reference layout comprises an aerodynamic layout shape digital model and a structural arrangement shape digital model;

[0008] Step 2, parameterizing the aerodynamic layout shape digital model to determine aerodynamic design variables; parameterizing the structural arrangement shape digital model to determine structural design variables; determining the same sound blast design variables as the aerodynamic design variables;

[0009] Step 3, reading the aerodynamic layout shape digital model to obtain the coordinates of the aerodynamic shape surface grid points; establishing the correlation between the coordinates of the aerodynamic shape surface grid points and the structural design variables;

[0010] Step 4, determining the value of the aerodynamic design variable; when the value of the aerodynamic design variable changes, since the sonic boom design variable is the same as the aerodynamic design variable, the value of the sonic boom design variable changes accordingly; when the value of the aerodynamic design variable and the value of the sonic boom design variable change, the FFD control frame is deformed, thereby causing the coordinates of the aerodynamic shape surface grid points to change, and new coordinates of the aerodynamic shape surface grid points are obtained; using the new coordinates of the aerodynamic shape surface grid points, according to the correlation between the coordinates of the aerodynamic shape surface grid points and the structural design variables determined in step 3, new values ​​of the structural design variables are obtained;

[0011] Step 5: Establishing an aerodynamic / structure / sonic boom multidisciplinary optimization mathematical model; the aerodynamic / structure / sonic boom multidisciplinary optimization mathematical model includes an objective function f and constraint conditions;

[0012]

[0013] in:

[0014] w1, w2, and w3 are weighted coefficients of aerodynamics, structure, and sonic boom disciplines, respectively; w1+w2+w3=1;

[0015] represents the aerodynamic performance function; where L / D represents the cruise lift-to-drag ratio of the design layout, L is the cruise lift of the design layout, and D is the cruise drag of the design layout; (L / D)0 represents the cruise lift-to-drag ratio of the baseline layout; is the normalized cruise lift-to-drag ratio, which is the aerodynamic performance function;

[0016] represents the structural discipline performance function; where W represents the structural weight of the designed layout; W0 represents the structural weight of the baseline layout; is the normalized structural weight, which is the structural subject performance function;

[0017] represents the sonic boom performance function; where PLdB represents the cruise ground sonic boom intensity of the designed layout; PLdB0 represents the cruise ground sonic boom intensity of the reference layout; is the normalized cruise ground sonic boom intensity, which is the sonic boom subject performance function;

[0018] σ≤[σ] max The meaning is: the structural stress σ of the design layout should be less than or equal to the maximum structural stress [σ] max , is the structural discipline constraint;

[0019] δ≤[σ] max The meaning is: the structural displacement δ of the design layout should be less than or equal to the maximum allowable displacement [σ] maxThe constraint condition of the structure discipline is that the structural weight of the design layout is less than or equal to the structural weight of the reference layout;

[0020] The meaning of (L / D)≥(L / D)0 is that the cruise lift-drag ratio of the design layout is greater than or equal to the cruise lift-drag ratio of the reference layout, which is the constraint condition of the aerodynamics discipline;

[0021] The meaning of W≤W0 is that the structural weight of the design layout is less than or equal to the structural weight of the reference layout, which is the constraint condition of the structure discipline;

[0022] The meaning of PLdB≤PLdB0 is that the cruise ground sound blast intensity PLdB of the design layout is less than or equal to the cruise ground sound blast intensity PLdB0 of the reference layout, which is the constraint condition of the sound blast discipline;

[0023] Step 6, forming the current design layout by using the values of the aerodynamic design variables, the values of the structural design variables and the values of the sound blast design variables obtained in step 4;

[0024] Step 7, calculating the value of the target function f and the constraint response values of the constraint conditions in step 5;

[0025] Step 7, calculating the value of the target function f and the constraint response values of the constraint conditions in step 5;

[0026] Step 8, using the optimizer based on the surrogate optimization algorithm to minimize the target function, and judging whether the optimization process converges or not, if not, executing step 9; if yes, executing step 10;

[0027] Step 9, changing the values of the aerodynamic design variables, and returning to step 4;

[0028] Step 10, outputting the final design layout.

[0029] Preferably, the structural arrangement shape digital model comprises a fuselage frame, a wing spar, a wing rib and a skin structure.

[0030] Preferably, the parameterized aerodynamic layout shape digital model is determined by using the FFD parameterization method to parameterize the aerodynamic layout shape digital model, and the aerodynamic design variables are the size of each fuselage section controlled by the control box and the sweepback angle of the wing.

[0031] The structural design variables are determined by using the FFD parameterization method to parameterize the structural arrangement shape digital model, and the structural design variables comprise the size of the fuselage frame, the position and length of the wing spar, the position, length and thickness of the wing rib, and the skin thickness of different partitions of the upper and lower surfaces of the wing.

[0032] Preferably, the correlation between the coordinates of the aerodynamic shape surface grid points and the structural design variables is established, and the correlation is specifically that

[0033] The radius of the i-th fuselage bulkhead in the structural design variable is R i , the center coordinates of the i-th fuselage bulkhead are (x i,f ,y i,f ,z i,f );The grid coordinates corresponding to the upper surface of the aerodynamic shape of the i-th fuselage bulkhead are The grid coordinates corresponding to the lower surface of the i-th fuselage bulkhead are The formula (1) shows the correlation formula:

[0034]

[0035] The chord length c j of the j-th wing rib in the structural design variable and its leading edge point coordinates (x j,w ,y j,w ,z j,w );The grid coordinates of the j-th wing rib at the wing leading edge are The grid coordinates of the wing trailing edge are The formula (2) shows the correlation formula:

[0036]

[0037] The correlation between the coordinates of the aerodynamic shape surface grid points and the structural design variables is established. Preferably, w1, w2 and w3 are set to 0.5, 0.1 and 0.4 respectively.

[0038] Preferably, in step 7, the cruise lift-drag ratio L / D of the design layout is determined by the following method: step A1, the lift coefficient C L of the current design layout is obtained by solving the high-order panel method;

[0039] Step A2, the pressure drag coefficient C Dp of the current design layout is obtained by solving the high-order panel method;

[0040] Step A3, the Reynolds number factor C fi (Re L ) and the Mach number factor f(M) are obtained by formula (3):

[0041]

[0042] Where: M is the cruise Mach number, Re L is the Reynolds number of the aircraft at cruise;

[0043] Step A4, the viscous drag coefficient C Df is obtained by formula (4):

[0044]

[0045] Where: S wet is the wetted area of ​​the aircraft, S w is the reference area of ​​the aircraft;

[0046] Step A5: Use formula (5) to obtain the resistance coefficient C D :

[0047] C D =C Dp +C Df (5)

[0048] In step A6, use formula (6) to obtain the cruise lift-to-drag ratio L / D of the design layout:

[0049]

[0050] The cruise lift-to-drag ratio L / D of the designed layout is thus obtained.

[0051] Preferably, for the current design layout, the finite element method is used to calculate the structural weight W, structural stress σ and structural displacement δ of the design layout.

[0052] Preferably, the following method is used to obtain the cruise ground sonic boom intensity PLdB of the current design layout:

[0053] Step B1: Use a plane to intercept the volume equivalent cross-sectional area distribution A of the aircraft along the incoming flow direction e,v (η) and lift equivalent cross-sectional area distribution A e,L (η), then use formula (7) to get the equivalent cross-sectional area distribution A of the aircraft e (η) is:

[0054] A e (η) = A e,v (η)+A e,L (η) (7)

[0055] Wherein, η is the distance from the nose point in the incoming flow direction;

[0056] Step B2, solve the equivalent cross-sectional area distribution A e (η) With respect to the second-order derivative of η, we get the second-order derivative of the equivalent cross-sectional area A" e (η);

[0057] Step B3, use formula (8) to calculate the second-order derivative A" of the equivalent cross-sectional area e (η) is integrated with a variable upper limit to obtain the F function with the aircraft as the disturbance source, expressed as F(x):

[0058]

[0059] Where: x is the axial distance between the sonic boom calculation point and the nose point;

[0060] Step B4, using formula (9), obtain the near-field sonic boom signal

[0061]

[0062] Where: γ is the atmospheric specific heat ratio; M is the Mach number; r0 is the distance between the location where the sonic boom is calculated and the aircraft axis; is the Prandtl-Grout correction factor, Ma is the Mach number;

[0063] Step B5, using the generalized Burgers equation to calculate the ground sonic boom waveform, the specific steps are as follows:

[0064] Step B5.1, describe the ordinary differential equations for the propagation of sound rays in the atmosphere according to formula (10);

[0065]

[0066] Where: R = (x, y, z) is the three-dimensional sound path vector; t represents the time in the propagation of the sound ray; T represents the transpose sign of the matrix; c0 is the ambient atmospheric sound speed; n is the wave surface unit normal vector; w is the atmospheric wind speed; I is the unit matrix; is the Kronecker product symbol; stands for Laplace operator;

[0067] Step B5.2, use the fourth-order four-step Runge-Kutta method to solve the ordinary differential equations shown in formula (10). The solution can be terminated when the z coordinate in R reaches the ground height. As the initial value of the acoustic pressure input, the generalized Burgers equation is solved along the sound ray to obtain the ground sonic boom waveform, which is expressed as formula (11):

[0068]

[0069] in:

[0070] p' is the acoustic pressure; s is the length coordinate of the sonic boom propagation ray;

[0071] G is the geometric acoustic factor; Among them, c n is the wave propagation velocity, v ray =|c0n+w| is the ray velocity, A is the area of ​​the sound tube, ρ0 is the ambient atmospheric density; c0 is the ambient atmospheric sound speed;

[0072] β is a nonlinear factor, β = 1 + (γ - 1) / 2; γ is the specific heat ratio of air, γ = 1.4, that is, 1.4 for calorimetrically perfect gas;

[0073] t′ is the delay time of the waveform;

[0074] δ is the diffusion coefficient, δ=(μ / ρ0)[4 / 3+0.6+(γ-1) 2 κ / (γRμ)]; where μ is the viscosity coefficient, κ is the thermal conductivity coefficient, and R is the gas constant;

[0075] (Δc) j and τ j are the sound velocity increment and relaxation time, respectively, j represents nitrogen molecules and oxygen molecules;

[0076] Step B5.3: After obtaining the ground sonic boom waveform, use Stevens' Mark VII method to calculate the cruise ground sonic boom intensity PLdB value corresponding to the waveform, specifically:

[0077] First, the ground sonic boom waveform is Fourier transformed to obtain the spectrum curve of the sound pressure level; then the spectrum curve is converted to equal loudness, and the converted loudness curve is weighted to obtain the cruise ground sonic boom intensity PLdB corresponding to the ground sonic boom waveform.

[0078] The present invention provides a rapid optimization design method for aerodynamics, structure, and sonic boom of supersonic civil aircraft, which has the following advantages:

[0079] 1. Compared with traditional single-discipline design, this invention refines the three disciplines of aerodynamics, structure, and sonic boom, which are very important in the conceptual design stage of supersonic civil aircraft, and conducts integrated design, which helps to coordinate the overall performance of the conceptual design scheme.

[0080] 2. The present invention associates aerodynamic shape variables with structural shape variables through aircraft surface grid points, reducing the redundancy of design variables and thus helping to improve optimization efficiency.

[0081] 3. This invention uses a multidisciplinary rapid analysis method to complete the performance evaluation of three disciplines in a short period of time. When combined with an efficient proxy optimization algorithm, it can comprehensively improve the performance of the three disciplines of aerodynamics, structure, and sonic boom within a few hours, achieving the extremely fast design of supersonic civil aircraft concept schemes. BRIEF DESCRIPTION OF THE DRAWINGS

[0082] Figure 1 A schematic diagram of the process flow of a rapid optimization design method for aerodynamics, structure, and sonic boom of a supersonic civil aircraft provided by the present invention;

[0083] Figure 2 Schematic diagram of the parameterization method and variable association method of the present invention;

[0084] Figure 3 Schematic diagram of the design structure matrix of the three disciplines of aerodynamics, structure and sonic boom according to the present invention;

[0085] Figure 4 This is an aerodynamic diagram of the LARM model in a specific embodiment of the present invention;

[0086] Figure 5 This is a structural layout diagram of the LARM model in a specific embodiment of the present invention;

[0087] Figure 6 This is an optimization convergence curve diagram for the LARM model in a specific embodiment of the present invention;

[0088] Figure 7 This is a comparison diagram of the appearance of the LARM model before and after optimization design in a specific embodiment of the present invention;

[0089] Figure 8 This is a comparison diagram of stress and displacement results before and after the optimization design of the LARM model in a specific embodiment of the present invention;

[0090] Figure 9 This is a comparison of sonic boom performance before and after the LARM model optimization design in a specific embodiment of the present invention. DETAILED DESCRIPTION

[0091] In order to make the technical problems, technical solutions and beneficial effects solved by the present invention more clearly understood, the present invention is further described in detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present invention and are not intended to limit the present invention.

[0092] The present invention provides a rapid optimization design method for supersonic civil aircraft across the three disciplines of aerodynamics, structure, and sonic boom. This method conducts multidisciplinary optimization design in the early stages of aircraft conceptual design, comprehensively considering the impacts of each discipline and facilitating the development of a conceptual design with superior overall performance. Specifically, during the conceptual design phase of a supersonic civil aircraft, the inventors discovered that aerodynamics, structure, and sonic boom significantly impact overall performance. Aerodynamic performance in cruise mode determines the aircraft's range and fuel consumption, impacting its ability to meet mission requirements and economic efficiency. Research has shown that for every 0.0001 decrease in the drag coefficient of large civil aircraft, payload capacity can increase by 7.56%. Aircraft structural weight reduction is crucial; lighter structural weight means lower fuel consumption, reduced emissions, and increased payload capacity. Furthermore, sonic boom performance is particularly important for supersonic civil aircraft. A sonic boom is a unique acoustic phenomenon unique to supersonic aircraft. Its strong pressure pulsation, short duration, and wide impact range significantly impact normal life and work on the ground, making it one of the obstacles hindering the commercial operation of supersonic civil aircraft. Therefore, the design of supersonic civil aircraft requires improved aerodynamic performance, reduced structural weight and lower sonic boom levels.

[0093] In the initial stage of conceptual design, the present invention not only comprehensively considers the influence between various disciplines, but also associates the design variables of each discipline to effectively improve the optimization efficiency.

[0094] The present invention aims to propose a rapid aerodynamic / structure / sonic boom three-disciplinary optimization design method for supersonic civil aircraft, capable of comprehensively improving the multidisciplinary performance of supersonic civil aircraft concept schemes in the early stages of conceptual design. The present invention relates the aerodynamic / structure / sonic boom design variables to the aerodynamic and structural shapes of a parameterized baseline layout through grid points on the aerodynamic shape surface, thereby reducing the dimensionality of the design variables. The present invention constructs objective functions and constraints using a multidisciplinary design structure matrix (DSM), establishes a weighted multi-objective optimization mathematical model for the aerodynamic / structure / sonic boom three-disciplinary design, calculates the objective function value and constraint response value using a rapid evaluation algorithm, and minimizes the objective function using an optimizer based on a proxy optimization algorithm. Finally, the proposed layout scheme with the best overall performance is output. The proposed method can comprehensively improve the cruise lift-to-drag ratio of a supersonic civil aircraft layout, reduce its structural weight, and reduce the perceived sound pressure level of the sonic boom in a relatively short period of time, subject to the constraints of structural strength and deformation. This method can meet the requirements for improving the multidisciplinary performance of supersonic civil aircraft in the early stages of conceptual design.

[0095] See Figure 1 The present invention provides a rapid optimization design method for aerodynamics, structure, and sonic boom of supersonic civil aircraft, comprising the following steps:

[0096] Step 1: Input a baseline supersonic civil aircraft layout to be optimized; the baseline supersonic civil aircraft layout includes a digital model of the aerodynamic layout shape and a digital model of the structural layout shape; the digital model of the aerodynamic layout shape needs to be converted into a three-dimensional model file of a surface structured grid; the digital model of the structural layout shape includes components such as the fuselage bulkhead, wing spars, ribs, and skin structure.

[0097] Step 2: parameterize the digital model of the aerodynamic layout and shape to determine the aerodynamic design variables; parameterize the digital model of the structural layout and shape to determine the structural design variables; and ensure that the sonic boom design variables are the same as the aerodynamic design variables;

[0098] In this step, the digital model of the aerodynamic layout is parameterized to determine the aerodynamic design variables. Specifically, the digital model of the aerodynamic layout is parameterized using the FFD parameterization method. The aerodynamic design variables determined are the size of each fuselage section controlled by the control frame and the sweep angle of the wing. When performing the FFD parameterization, the FFD control frame is as follows: Figure 2 shown.

[0099] The FFD parameterization method is used to parameterize the digital model of the structural layout and shape. The determined structural design variables include the size of the fuselage bulkhead, the position and length of the wing spars, the position, length and thickness of the wing ribs, and the skin thickness of different partitions on the upper and lower surfaces of the wing.

[0100] Step 3: Read the digital model of the aerodynamic layout shape to obtain the coordinates of the grid points on the aerodynamic shape surface; establish a correlation between the coordinates of the grid points on the aerodynamic shape surface and the structural design variables, thereby reducing the dimension of the design variables;

[0101] In this step, the relationship between the coordinates of the grid points on the aerodynamic shape surface and the structural design variables is established, specifically:

[0102] Among the structural design variables, the radius of the i-th fuselage frame is R i , the center coordinates of the i-th fuselage frame are (x i,f ,y i,f ,z i,f ); the grid coordinates corresponding to the upper surface of the aerodynamic shape of the i-th fuselage frame are The grid coordinates corresponding to the lower surface of the aerodynamic shape of the i-th fuselage frame are: Then the correlation formula is as shown in formula (1):

[0103]

[0104] Among the structural design variables, the chord length c of the jth rib is j The coordinates of its front edge point (x j,w ,y j,w ,z j,w); The grid coordinates of the wing leading edge at the position of the jth rib are The grid coordinates of the wing trailing edge are Then the correlation formula is as shown in formula (2):

[0105]

[0106] Thus, the relationship between the coordinates of the grid points on the aerodynamic shape surface and the structural design variables is established.

[0107] Step 4, determining the values ​​of the aerodynamic design variables; when the values ​​of the aerodynamic design variables change, since the sonic boom design variables are the same as the aerodynamic design variables, the values ​​of the sonic boom design variables also change; when the values ​​of the aerodynamic design variables and the sonic boom design variables change, the FFD control frame is deformed, thereby causing the coordinates of the aerodynamic shape surface grid points to change, and new coordinates of the aerodynamic shape surface grid points are obtained; using the new coordinates of the aerodynamic shape surface grid points, according to the correlation between the coordinates of the aerodynamic shape surface grid points and the structural design variables determined in step 3, new values ​​of the structural design variables are obtained; therefore, after the structural shapes of the fuselage and wings are generated, structural analysis and calculation can be performed, avoiding the increase in the number of variables caused by using structural variables alone to describe the structural shapes.

[0108] Step 5: Construct the objective function and constraints based on the multidisciplinary design structure matrix DSM, such as Figure 3 As shown, a weighted multi-objective optimization mathematical model of aerodynamics / structure / sonic boom is established;

[0109] Specifically, a pneumatic / structural / sonic boom multidisciplinary optimization mathematical model is established; the pneumatic / structural / sonic boom multidisciplinary optimization mathematical model includes an objective function f and constraint conditions;

[0110]

[0111] in:

[0112] w1, w2, w3 are weighting coefficients of aerodynamics, structure and sonic boom respectively; w1+w2+w3=1; for example, w1, w2, w3 are set to 0.5, 0.1 and 0.4 respectively.

[0113] represents the aerodynamic performance function; where L / D represents the cruise lift-to-drag ratio of the design layout, L is the cruise lift of the design layout, and D is the cruise drag of the design layout; (L / D)0 represents the cruise lift-to-drag ratio of the baseline layout; is the normalized cruise lift-to-drag ratio, which is the aerodynamic performance function;

[0114] represents the structural discipline performance function; where W represents the structural weight of the designed layout; W0 represents the structural weight of the baseline layout; is the normalized structural weight, which is the structural subject performance function;

[0115] represents the sonic boom performance function; where PLdB represents the cruise ground sonic boom intensity of the designed layout; PLdB0 represents the cruise ground sonic boom intensity of the reference layout; is the normalized cruise ground sonic boom intensity, which is the sonic boom subject performance function;

[0116] σ≤[σ] max The meaning is: the structural stress σ of the design layout should be less than or equal to the maximum structural stress [σ] max , which are structural discipline constraints;

[0117] δ≤[σ] max The meaning is: the structural displacement δ of the design layout should be less than or equal to the maximum allowable displacement [σ] max , which are structural discipline constraints;

[0118] (L / D)≥(L / D)0 means that the cruise lift-to-drag ratio of the design layout should be greater than or equal to the cruise lift-to-drag ratio of the reference layout, which is an aerodynamic constraint.

[0119] W≤W0 means that the structural weight of the design layout should be less than or equal to the structural weight of the baseline layout, which is a structural discipline constraint.

[0120] PLdB≤PLdB0 means that the cruise ground sonic boom intensity PLdB of the design layout should be less than or equal to the cruise ground sonic boom intensity PLdB0 of the reference layout, which is a sonic boom discipline constraint condition;

[0121] The specific ideas for this step include:

[0122] The multidisciplinary design structure matrix DSM is used to characterize the direction of data flow. Figure 3 As shown in Figure 1, the objective function is constructed in the form of weighted multi-objectives. The performance functions involved in the three disciplines are cruise lift-to-drag ratio L / D, structural weight W, and cruise ground sonic boom intensity PLdB. L is the cruise lift and D is the cruise drag.

[0123] First, the performance functions of the three disciplines are normalized to avoid the influence of the magnitude difference between the performance functions of the three disciplines on the objective function. Let the lift-to-drag ratio of the baseline layout be (L / D)0, the structural weight of the baseline layout be W0, and the circulating ground sonic boom intensity of the baseline layout be PLdB0. Then the normalized cruise lift-to-drag ratio is The normalized structure weight is The normalized cruise ground sonic boom intensity is

[0124] Secondly, the objective function is a weighted sum of the negative of the normalized cruise lift-to-drag ratio, the structural weight, and the cruise ground sonic boom intensity. The weighted coefficients for the three subjects are w1, w1, and w1, respectively, and the sum of the three coefficients satisfies w1+w2+w3=1.

[0125] Finally, the objective function for the aerodynamic / structure / sonic boom multidisciplinary optimization of supersonic civil aircraft is:

[0126]

[0127] Among them, the weighting coefficient can be selected according to the performance improvement of the subject.

[0128] The optimization mathematical model mainly includes five constraints:

[0129] Including the constraints of structural strength and stiffness: (1) The structural stress σ is less than the maximum stress of the structure [σ] max , that is, σ<[σ] max ; (2) The structural displacement δ is less than the given maximum allowable displacement [δ] max , that is, δ<[δ] max .

[0130] In addition, in order to ensure that the performance of the three disciplines of aerodynamics / structure / sonic boom is improved, the following conditions are added to ensure that the performance of the three disciplines is better than that of the baseline layout: (3) the lift-to-drag ratio L / D of the design layout is greater than or equal to the lift-to-drag ratio (L / D)0 of the baseline layout, that is, (L / D)≥(L / D)0; (4) the structural weight W of the design layout is less than or equal to W0, that is, W≤W0; (5) the cruise ground sonic boom intensity PLdB of the design layout should be less than or equal to the cruise ground sonic boom intensity PLdB0 of the baseline layout. In summary, the final optimization mathematical model is:

[0131]

[0132] Step 6, forming a current design layout using the values ​​of the aerodynamic design variables, the structural design variables, and the sonic boom design variables obtained in step 4;

[0133] Calculate the values ​​of the aerodynamic performance function, the structural performance function, and the sonic boom performance function of the current design layout;

[0134] Step 7, calculate the value of the objective function f in step 5 and the constraint response value of the constraint condition;

[0135] The specific calculation method is:

[0136] The cruise lift-to-drag ratio L / D of the design layout is determined using the high-order panel method + viscosity correction method. The specific method is as follows:

[0137] Step A1: Use the high-order panel method to solve the lift coefficient C of the current design layout. L ;

[0138] Lift coefficient C L The high-order panel method is used for solution, and the number of discrete panel elements of the aircraft surface is generally around 3000 to 5000.

[0139] Step A2: Use the high-order panel method to solve the pressure difference resistance coefficient C of the current design layout. Dp ;

[0140] Step A3: Use formula (3) to obtain the Reynolds number factor C fi (Re L ) and the Mach number factor f(M):

[0141]

[0142] Where: M is the cruise Mach number, Re L is the Reynolds number of the aircraft during cruising; , both are constants.

[0143] Step A4: Use formula (4) to obtain the viscous resistance coefficient C through the viscous resistance estimation formula. Df :

[0144]

[0145] Where: S wet is the wetted area of ​​the aircraft, S w is the reference area of ​​the aircraft;

[0146] Step A5: Use formula (5) to obtain the resistance coefficient C D :

[0147] C D =C Dp +C Df (5)

[0148] Therefore, the drag coefficient C D is the pressure difference resistance coefficient C Dp and the viscous drag coefficient C Df sum.

[0149] In step A6, use formula (6) to obtain the cruise lift-to-drag ratio L / D of the design layout:

[0150]

[0151] The cruise lift-to-drag ratio L / D of the designed layout is thus obtained.

[0152] The lift-to-drag ratio L / D, the objective function for performance evaluation in aerodynamics, can be calculated using the following formula:

[0153]

[0154] For the current design layout, the finite element method was used to calculate the structural weight W, structural stress σ, and structural displacement δ of the design layout. To improve the accuracy of the structural weight / stress / displacement calculations, a quadrilateral mesh was used, with as many meshes as possible maintaining an aspect ratio close to 1. The total number of meshes for the entire aircraft was approximately 30,000.

[0155] The cruise ground sonic boom intensity PLdB of the current design layout is obtained using the following method:

[0156] Step B1: Use a plane to intercept the volume equivalent cross-sectional area distribution A of the aircraft along the incoming flow direction e,v (η) and lift equivalent cross-sectional area distribution A e,L (η), then use formula (7) to get the equivalent cross-sectional area distribution A of the aircraft e (η) is:

[0157] A e (η) = A e,v (η)+A e,L (η) (7)

[0158] Wherein, η is the distance from the nose point in the incoming flow direction;

[0159] Step B2, solve the equivalent cross-sectional area distribution A e (η) With respect to the second-order derivative of η, we get the second-order derivative of the equivalent cross-sectional area A" e (η);

[0160] Step B3, use formula (8) to calculate the second-order derivative A" of the equivalent cross-sectional area e (η) is integrated with a variable upper limit to obtain the F function with the aircraft as the disturbance source, expressed as F(x):

[0161]

[0162] Where: x is the axial distance between the sonic boom calculation point and the nose point;

[0163] Step B4, using formula (9), obtain the near-field sonic boom signal

[0164]

[0165] Where: γ is the atmospheric specific heat ratio; M is the Mach number; r0 is the distance between the location where the sonic boom is calculated and the aircraft axis; is the Prandtl-Grout correction factor, Ma is the Mach number;

[0166] Step B5, using the generalized Burgers equation to calculate the ground sonic boom waveform, the specific steps are as follows:

[0167] Step B5.1, describe the ordinary differential equations for the propagation of sound rays in the atmosphere according to formula (10);

[0168]

[0169] Where: R = (x, y, z) is the three-dimensional sound path vector; t represents the time in the propagation of the sound ray; T represents the transpose sign of the matrix; c0 is the ambient atmospheric sound speed; n is the wave surface unit normal vector; w is the atmospheric wind speed; I is the unit matrix; is the Kronecker product symbol; ▽ represents the Laplace operator;

[0170] Step B5.2, use the fourth-order four-step Runge-Kutta method to solve the ordinary differential equations shown in formula (10). The solution can be terminated when the z coordinate in R reaches the ground height. As the initial value of the acoustic pressure input, the generalized Burgers equation is solved along the sound ray to obtain the ground sonic boom waveform, which is expressed as formula (11):

[0171]

[0172] in:

[0173] p' is the acoustic pressure; s is the length coordinate of the sonic boom propagation ray;

[0174] G is the geometric acoustic factor; Among them, c n is the wave propagation velocity, v ray =|c0n+w| is the ray velocity, A is the area of ​​the sound tube, ρ0 is the ambient atmospheric density; c0 is the ambient atmospheric sound speed;

[0175] β is a nonlinear factor, β = 1 + (γ - 1) / 2; γ is the specific heat ratio of air, γ = 1.4, that is, 1.4 for calorimetrically perfect gas;

[0176] t′ is the delay time of the waveform;

[0177] δ is the diffusion coefficient, δ=(μ / ρ0)[4 / 3+0.6+(γ-1) 2 κ / (γRμ)]; where μ is the viscosity coefficient, κ is the thermal conductivity coefficient, and R is the gas constant;

[0178] (Δc) j and τ j are the sound velocity increment and relaxation time, respectively, j represents nitrogen molecules and oxygen molecules;

[0179] Step B5.3: After obtaining the ground sonic boom waveform, use Stevens' Mark VII method to calculate the cruise ground sonic boom intensity PLdB value corresponding to the waveform, specifically:

[0180] First, the ground sonic boom waveform is Fourier transformed to obtain the spectrum curve of the sound pressure level; then the spectrum curve is converted to equal loudness, and the converted loudness curve is weighted to obtain the cruise ground sonic boom intensity PLdB corresponding to the ground sonic boom waveform.

[0181] Step 8: Use the optimizer based on the proxy optimization algorithm to minimize the objective function and determine whether the optimization process has converged. If not, execute step 9; if converged, execute step 10.

[0182] The proxy optimizer used is a proxy optimization algorithm based on the kriging interpolation model. Specifically, the parameters during optimization are set as follows:

[0183] The surrogate model adopts the Kriging model;

[0184] The initial sample points are generated by Latin hypercube sampling;

[0185] The initial sample point is 50;

[0186] The maximum number of sample points is 180;

[0187] The point addition criteria are “EI+PI+MSP+LCB”;

[0188] The correlation function is a cubic spline function;

[0189] The value range of the hyperparameter μ is 1×10 -6 ≤μ≤0.618.

[0190] Step 9, change the value of the aerodynamic design variable and return to step 4;

[0191] Step 10: Output the final design layout.

[0192] The final design layout is output, specifically: the sample point with the largest decrease in the objective function among all the sample points obtained by optimization. The design variable values ​​corresponding to this sample point are used to generate the optimal aerodynamic and structural shapes and output them.

[0193] A specific application example is used to further illustrate the supersonic civil aircraft aerodynamic / structure / sound blast three-discipline rapid optimization design method provided by the present application:

[0194] In this embodiment, the low sound blast research common model LARM published by the Sixth Aviation, Aerospace and Marine Industry Development Forum is selected, and aerodynamic / structure / sound blast three-discipline optimization design is carried out on the LARM model to verify the effectiveness of the method. The related parameters of the LARM model are shown in Table 1:

[0195] Table 1 Main parameters of LARM model

[0196]

[0197]

[0198] Under the design parameters of the baseline configuration, aerodynamic / structure / sound blast three-discipline optimization design is carried out on the LARM model.

[0199] The main steps of aerodynamic / structure / sound blast three-discipline rapid optimization design of the LARM model by the method of the present application are as follows:

[0200] (1) Input the baseline layout of the LARM model.

[0201] The LARM model adopts a large-sweeping arrow wing T-tail layout, and the digital model of the outer shape is shown in Figure 4 ; the internal structure arrangement adopts a truss beam fuselage and a multi-beam wing structure, and the materials are aluminum alloy and composite materials, and the digital model of the outer shape is shown in Figure 5 ;

[0202] (2) The aerodynamic shape of the LARM supersonic civil aircraft model is parameterized by using FFD parameterization, the size of the card is controlled by scaling the fuselage to describe the distribution change of the fuselage cross-sectional area, a total of 18 design variables; the wing sweep angle change is controlled by pulling the wing in parallel to describe the wing box, a total of 1 design variable; the above 19 design variables are associated with the structure variables through the fuselage surface grid points; in the structure discipline, the thickness of the skin is described by increasing the thickness of each of the three partitions on the upper and lower surfaces of the wing, a total of 6 design variables;

[0203] (3) The objective function and the constraint condition are constructed based on the multi-discipline design structure matrix (DSM), and a weighted multi-objective optimization mathematical model of aerodynamic / structure / sound blast three-discipline is established. In this embodiment, the weighted coefficients of the three disciplines are 0.5, 0.1 and 0.4, respectively;

[0204] (4) The cruise lift-to-drag ratio is calculated using the "high-order panel method + viscosity correction", the structural weight / stress / displacement is calculated using the finite element method, and the perceived sound pressure level of the cruise ground sonic boom is calculated using the generalized Burgers equation method. The three-disciplinary performance of the baseline configuration are: aerodynamic lift-to-drag ratio of 9.31, structural weight of 42517 kg, and sonic boom intensity of 95.77 PLdB;

[0205] (5) Calculate the objective function value and constraint response value;

[0206] (6) The weighted objective function is minimized using an optimizer based on a proxy optimization algorithm. The optimization settings are consistent with the above. The convergence curve is as follows: Figure 6 As shown, the optimal design result can be found after 5 hours and 39 minutes;

[0207] (7) The optimal solution after convergence is selected and brought into the parameterization program to generate the final design layout. Table 2 shows the performance comparison of the three disciplines before and after optimization. It can be seen that the performance of all three disciplines has improved. Figure 7 A comparison of the appearance before and after optimization is given; Figure 8 A comparison of stress and displacement distribution before and after optimization is given, showing that each region tends to be in a full stress state and material utilization is improved. Figure 9 A comparison of the sonic boom performance before and after optimization is given, and the rising peak of the head shock wave is significantly weakened.

[0208] Table 2 Comparison of performance of each subject before and after LARM model optimization

[0209] Performance indicators Baseline configuration Optimized configuration Absolute change Relative change lift-to-drag ratio 9.31 9.58 0.27 2.9％ Structural weight 42517kg 40581kg 1936kg 4.5％ Ground sonic boom intensity 95.76PLdB 94.74PLdB 1.02PLdB 1.1％

[0210] The present invention combines rapid analysis methods for aerodynamics, structure, and sonic boom with an efficient global proxy optimization algorithm to provide a rapid optimization design method for the three disciplines of aerodynamics / structure / sonic boom of supersonic civil aircraft in the early stages of conceptual design, which can comprehensively improve the multidisciplinary performance of supersonic civil aircraft concept schemes.

[0211] The above is only a preferred embodiment of the present invention. It should be pointed out that for ordinary technicians in this technical field, several improvements and modifications can be made without departing from the principles of the present invention. These improvements and modifications should also be considered within the scope of protection of the present invention.

Claims

1. A rapid optimization design method for aerodynamics, structure, and sonic boom of supersonic civil aircraft, characterized by: The following steps are involved: Step 1: inputting a supersonic civil aircraft reference layout to be optimized; the supersonic civil aircraft reference layout includes a digital model of an aerodynamic layout and a digital model of a structural layout; Step 2: parameterize the digital model of the aerodynamic layout and shape to determine the aerodynamic design variables; parameterize the digital model of the structural layout and shape to determine the structural design variables; and ensure that the sonic boom design variables are the same as the aerodynamic design variables; Step 3: Read the digital model of the aerodynamic layout shape to obtain the coordinates of the grid points on the aerodynamic shape surface; and establish a correlation between the coordinates of the grid points on the aerodynamic shape surface and the structural design variables. Step 4, determining the values ​​of aerodynamic design variables; When the value of the aerodynamic design variable changes, since the sonic boom design variable is the same as the aerodynamic design variable, the value of the sonic boom design variable also changes accordingly; when the value of the aerodynamic design variable and the value of the sonic boom design variable change, the FFD control frame is deformed, thereby causing the coordinates of the aerodynamic shape surface grid points to change, thereby obtaining new coordinates of the aerodynamic shape surface grid points; using the new coordinates of the aerodynamic shape surface grid points, according to the correlation between the coordinates of the aerodynamic shape surface grid points and the structural design variables determined in step 3, new values ​​of the structural design variables are obtained; Step 5: Establishing an aerodynamic / structure / sonic boom multidisciplinary optimization mathematical model; the aerodynamic / structure / sonic boom multidisciplinary optimization mathematical model includes an objective function f and constraint conditions; in: w1, w2, and w3 are weighted coefficients of aerodynamics, structure, and sonic boom disciplines, respectively; w1+w2+w3=1; represents the aerodynamic performance function; where L / D represents the cruise lift-to-drag ratio of the design layout, L is the cruise lift of the design layout, and D is the cruise drag of the design layout; (L / D)0 represents the cruise lift-to-drag ratio of the baseline layout; is the normalized cruise lift-to-drag ratio, which is the aerodynamic performance function; represents the structural discipline performance function; where W represents the structural weight of the designed layout; W0 represents the structural weight of the baseline layout; is the normalized structural weight, which is the structural subject performance function; represents the sonic boom performance function; where PLdB represents the cruise ground sonic boom intensity of the designed layout; PLdB0 represents the cruise ground sonic boom intensity of the reference layout; is the normalized cruise ground sonic boom intensity, which is the sonic boom subject performance function; σ≤[σ] max The meaning is: the structural stress σ of the design layout is less than or equal to the maximum structural stress [σ] max , which are structural discipline constraints; δ≤[δ] max The meaning is: the structural displacement δ of the designed layout is less than or equal to the maximum allowable displacement [δ] max , which are structural discipline constraints; (L / D)≥(L / D)0 means that the cruise lift-to-drag ratio of the designed layout is greater than or equal to the cruise lift-to-drag ratio of the reference layout, which is an aerodynamic constraint. W≤W0 means that the structural weight of the design layout is less than or equal to the structural weight of the baseline layout, which is a structural discipline constraint. PLdB≤PLdB0 means that the cruise ground sonic boom intensity PLdB of the design layout is less than or equal to the cruise ground sonic boom intensity PLdB0 of the reference layout, which is a sonic boom discipline constraint condition; Step 6, forming a current design layout using the values ​​of the aerodynamic design variables, the structural design variables, and the sonic boom design variables obtained in step 4; Calculate the values ​​of the aerodynamic performance function, the structural performance function, and the sonic boom performance function of the current design layout; Step 7, calculate the value of the objective function f in step 5 and the constraint response value of the constraint condition; Step 8: Use the optimizer based on the proxy optimization algorithm to minimize the objective function and determine whether the optimization process has converged. If not, execute step 9; if converged, execute step 10. Step 9, change the value of the aerodynamic design variable and return to step 4; Step 10: Output the final design layout.

2. The aerodynamic / structure / sonic boom three-discipline rapid optimization design method for supersonic civil aircraft according to claim 1 is characterized in that: The structural layout digital model includes fuselage bulkheads, wing spars, wing ribs and skin structures.

3. The aerodynamic / structure / sonic boom three-discipline rapid optimization design method for supersonic civil aircraft according to claim 1 is characterized in that: Parameterizing the digital model of the aerodynamic layout and shape to determine aerodynamic design variables. Specifically, the digital model of the aerodynamic layout and shape is parameterized using the FFD parameterization method. The aerodynamic design variables determined are the size of each fuselage section controlled by the control frame and the sweep angle of the wing. The FFD parameterization method is used to parameterize the digital model of the structural layout and shape. The determined structural design variables include the size of the fuselage bulkhead, the position and length of the wing spars, the position, length and thickness of the wing ribs, and the skin thickness of different partitions on the upper and lower surfaces of the wing.

4. The aerodynamic / structure / sonic boom three-discipline rapid optimization design method for supersonic civil aircraft according to claim 1 is characterized in that: Establish the relationship between the coordinates of the grid points on the aerodynamic shape surface and the structural design variables, specifically: Among the structural design variables, the radius of the i-th fuselage frame is R i , the center coordinates of the i-th fuselage frame are (x i,f ,y i,f ,z i,f ); the grid coordinates corresponding to the upper surface of the aerodynamic shape of the i-th fuselage frame are The grid coordinates corresponding to the lower surface of the aerodynamic shape of the i-th fuselage frame are: Then the correlation formula is as shown in formula (1): Among the structural design variables, the chord length c of the jth rib is j Its leading edge coordinates (x j,w ,y j,w ,z j,w ); The grid coordinates of the wing leading edge at the position of the jth rib are The grid coordinates of the wing trailing edge are Then the correlation formula is as shown in formula (2): Thus, the relationship between the coordinates of the grid points on the aerodynamic shape surface and the structural design variables is established.

5. The aerodynamic / structure / sonic boom three-discipline rapid optimization design method for supersonic civil aircraft according to claim 1 is characterized in that: w1, w2, and w3 are set to 0.5, 0.1, and 0.4, respectively.

6. The aerodynamic / structure / sonic boom three-discipline rapid optimization design method for supersonic civil aircraft according to claim 1 is characterized in that: In step 7, the cruise lift-to-drag ratio L / D of the design layout is determined using the following method: Step A1: Use the high-order panel method to solve the lift coefficient C of the current design layout. L ; Step A2: Use the high-order panel method to solve the pressure difference resistance coefficient C of the current design layout. Dp ; Step A3: Use formula (3) to obtain the Reynolds number factor C fi (Re L ) and the Mach number factor f(M): Where: M is the cruise Mach number, Re L is the Reynolds number of the aircraft during cruising; Step A4: Use formula (4) to obtain the viscous drag coefficient C Df : Where: S wet is the wetted area of ​​the aircraft, S w is the reference area of ​​the aircraft; Step A5: Use formula (5) to obtain the resistance coefficient C D : C D =C Dp +C Df (5) In step A6, use formula (6) to obtain the cruise lift-to-drag ratio L / D of the design layout: The cruise lift-to-drag ratio L / D of the designed layout is thus obtained.

7. The aerodynamic / structure / sonic boom three-discipline rapid optimization design method for supersonic civil aircraft according to claim 1 is characterized in that: For the current design layout, the finite element method is used to calculate the structural weight W, structural stress σ and structural displacement δ of the design layout.

8. The aerodynamic / structure / sonic boom three-discipline rapid optimization design method for supersonic civil aircraft according to claim 1 is characterized in that: The cruise ground sonic boom intensity PLdB of the current design layout is obtained using the following method: Step B1: Use a plane to intercept the volume equivalent cross-sectional area distribution A of the aircraft along the incoming flow direction e,v (η) and lift equivalent cross-sectional area distribution A e,L (η), then use formula (7) to get the equivalent cross-sectional area distribution A of the aircraft e (η) is: A e (h)=A e,v (h)+A e,L (h) (7) Wherein, η is the distance from the nose point in the incoming flow direction; Step B2, solve the equivalent cross-sectional area distribution A e (η) With respect to the second-order derivative of η, we get the second-order derivative of the equivalent cross-sectional area A″ e (η); Step B3, using formula (8), the second-order derivative A″ of the equivalent cross-sectional area e (η) is integrated with a variable upper limit to obtain the F function with the aircraft as the disturbance source, expressed as F(x): Where: x is the axial distance between the sonic boom calculation point and the nose point; Step B4, using formula (9), obtain the near-field sonic boom signal Where: γ is the atmospheric specific heat ratio; M is the Mach number; r0 is the distance between the location where the sonic boom is calculated and the aircraft axis; is the Prandtl-Grout correction factor, Ma is the Mach number; Step B5, using the generalized Burgers equation to calculate the ground sonic boom waveform, the specific steps are as follows: Step B5.1, describe the ordinary differential equations for the propagation of sound rays in the atmosphere according to formula (10); Where: R = (x, y, z) is the three-dimensional sound path vector; t represents the time in the propagation of the sound ray; T represents the transpose sign of the matrix; c0 is the ambient atmospheric sound speed; n is the wave surface unit normal vector; w is the atmospheric wind speed; I is the unit matrix; is the Kronecker product symbol; stands for Laplace operator; Step B5.2, use the fourth-order four-step Runge-Kutta method to solve the ordinary differential equations shown in formula (10). The solution can be terminated when the z coordinate in R reaches the ground height. As the initial value of the acoustic pressure input, the generalized Burgers equation is solved along the sound ray to obtain the ground sonic boom waveform, which is expressed as formula (11): in: p' is the acoustic pressure; s is the length coordinate of the sonic boom propagation ray; G is the geometric acoustic factor; Among them, c n is the wave propagation velocity, v ray =|c0n+w| is the ray velocity, A is the area of ​​the sound tube, ρ0 is the ambient atmospheric density; c0 is the ambient atmospheric sound speed; β is a nonlinear factor, β = 1 + (γ - 1) / 2; γ is the specific heat ratio of air, γ = 1.4, that is, 1.4 for calorimetrically perfect gas; t′ is the delay time of the waveform; δ is the diffusion coefficient, δ=(μ / ρ0)[4 / 3+0.6+(γ-1) 2 κ / (γRμ)]; where μ is the viscosity coefficient, κ is the thermal conductivity coefficient, and R is the gas constant; (Δc) j and τ j are the sound velocity increment and relaxation time, respectively, j represents nitrogen molecules and oxygen molecules; Step B5.3: After obtaining the ground sonic boom waveform, use Stevens' Mark VII method to calculate the cruise ground sonic boom intensity PLdB value corresponding to the waveform, specifically: First, the ground sonic boom waveform is Fourier transformed to obtain the spectrum curve of the sound pressure level; then the spectrum curve is converted to equal loudness, and the converted loudness curve is weighted to obtain the cruise ground sonic boom intensity PLdB corresponding to the ground sonic boom waveform.

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