A fast parameter estimation method for high-speed aircraft based on pneumatic parameter experience

Abstract

The present invention provides a method for quickly estimating parameters of a high-speed aircraft based on empirical pneumatic parameters, which relates to the field of aerospace technology. The method includes: determining the aircraft state variables and aerodynamic shape parameters that affect the values of the pneumatic parameters to be estimated; performing aerodynamic calculations through computational fluid dynamics methods based on the aerodynamic shape parameters and overall parameters to obtain aerodynamic data; based on the aerodynamic data, gradually establishing a multivariate function relationship between the pneumatic parameters, the aerodynamic shape, and the aircraft state variables by using the control variable method to obtain a pneumatic parameter estimation model; and obtaining the pneumatic parameters corresponding to the preset aircraft state variables according to the pneumatic parameter estimation model. The present invention solves the problems in the prior art that in the calculation of pneumatic parameters, when using the traditional theoretical calculation method of aerodynamic characteristics, the numerical calculation method is complex, and the parameter calculation accuracy of the engineering estimation method is low.

Description

Technical Field

[0001] The present invention relates to the field of aerospace technology, and particularly to a method for quickly estimating parameters of a high-speed aircraft based on empirical aerodynamic parameters. Background Art

[0002] With the research and development of high-speed aircraft design technology, performing aerodynamic analysis on the aircraft and obtaining its accurate aerodynamic characteristics are important prerequisites for establishing the aircraft dynamics model and designing an excellent control system. Considering the design cycle and R & D cost, in the initial stage of aircraft design, especially in the overall design optimization stage, theoretical calculation methods are usually used to obtain the aerodynamic data of the studied aircraft and conduct aerodynamic analysis based on this. The theoretical calculation of aerodynamic characteristics includes two methods: engineering estimation and numerical analysis. Among them, the traditional engineering estimation method is based on relevant knowledge of aerodynamics theory, which can grasp the core and some laws of the problem while conducting theoretical analysis, but its quantitative analysis ability is insufficient, and it is sometimes powerless in the face of complex problems. Numerical analysis, also known as computational fluid dynamics (CFD), has created favorable conditions for the development of computational fluid dynamics with the rapid development of computer technology, enabling a large amount of flow details and macroscopic data to be obtained quantitatively during aerodynamic calculation.

[0003] In the overall scheme demonstration stage of high-speed aircraft, it is necessary to continuously optimize and adjust the specific numerical values of the aircraft aerodynamic shape. In this process, on the one hand, it is necessary to generally master the variation law of the aircraft aerodynamic characteristics under each shape, so as to have the ability of qualitative analysis of key problems in the design. On the other hand, some simulation experiments in the scheme demonstration require a large amount of relatively reliable aerodynamic data for each aerodynamic shape as support. For the calculation of aerodynamic parameters, using the traditional theoretical calculation of aerodynamic characteristics, the calculation is complex and the accuracy of parameter calculation is low. Summary of the Invention

[0004] In order to overcome the deficiencies of the prior art, the purpose of the present invention is to provide a method for quickly estimating parameters of a high-speed aircraft based on empirical aerodynamic parameters. The present invention solves the problems in the prior art that for the calculation of aerodynamic parameters, using the traditional theoretical calculation of aerodynamic characteristics, the calculation is complex and the accuracy of parameter calculation is low.

[0005] To achieve the above purpose, the present invention provides the following scheme:

[0006] A method for quickly estimating parameters of a high-speed aircraft based on empirical aerodynamic parameters, comprising:

[0007] Determine the aircraft state variables and aerodynamic shape parameters that affect the value of the aerodynamic parameter to be estimated; the aircraft state variables are altitude, sideslip angle, angle of attack, and Mach number, and the aerodynamic shape parameter is the half-cone angle;

[0008] According to the aerodynamic shape parameters and overall parameters, perform aerodynamic calculations through computational fluid dynamics methods to obtain aerodynamic data;

[0009] Based on the aerodynamic data, adopt the method of controlling variables to gradually establish a multivariate function relationship between the aerodynamic parameters, the aerodynamic shape, and the aircraft state variables, and obtain an aerodynamic parameter estimation model;

[0010] According to the aerodynamic parameter estimation model, obtain the aerodynamic parameters corresponding to the preset aircraft state variables.

[0011] Preferably, the step of performing aerodynamic calculations through computational fluid dynamics methods according to the aerodynamic shape and overall parameters to obtain aerodynamic data includes:

[0012] Based on the overall parameters of the high-speed aircraft, within the range of values of the half-cone angle, construct a three-dimensional geometric model of the aircraft corresponding to the values of the half-cone angle to be optimized;

[0013] Based on the three-dimensional geometric model of the aircraft, use finite element analysis software to obtain the aircraft grid file corresponding to the value of the half-cone angle;

[0014] Calculate the calculation nodes of the aircraft state variables to obtain the aerodynamic calculation state nodes;

[0015] According to the aerodynamic calculation state nodes, construct the working conditions corresponding to the range of values of the half-cone angle;

[0016] According to the grid file and the working conditions, based on the two-equation turbulence model as the aerodynamic characteristic calculation model, obtain the aerodynamic data.

[0017] Preferably, the step of gradually establishing a multivariate function relationship between the aerodynamic parameters, the aerodynamic shape, and the aircraft state variables based on the aerodynamic data includes:

[0018] Based on the aerodynamic data, fit the polynomial function relationship between the aerodynamic parameters and the Mach number under the condition that the angle of attack remains unchanged to obtain the first relationship;

[0019] Based on the first relationship, fit the polynomial function relationship between the coefficients of each term in the first relationship and the angle of attack under the condition that the Mach number remains unchanged to obtain the second relationship;

[0020] Based on the first relationship and the second relationship corresponding to the aerodynamic data, fit the polynomial function relationship between the coefficients of each term in the second relationship and the half-cone angle to obtain the third relationship;

[0021] According to the first relationship, the second relationship, and the third relationship, obtain the aerodynamic parameter estimation model.

[0022] Preferably, the expression of the aerodynamic parameter estimation model is:

[0023]

[0024]

[0025]

[0026]

[0027] Among them, C x is the axial force coefficient of the aircraft, C y is the normal force coefficient of the aircraft, is the pitching moment coefficient of the aircraft, θ is the half-cone angle, α is the angle of attack, Ma is the Mach number, is the fitting coefficient of the polynomial function between the coefficients of each term in the second relational expression and the half-cone angle, k ij , p ij , q ij , (i = 5, 4, 3, 2, 1, 0; j = 1, 2, 3) are the fitting coefficients of the polynomial function between the coefficients of each term in the first relational expression and the angle of attack, K i , P i , Q i , (i = 5, 4, 3, 2, 1, 0) are the fitting coefficients of the polynomial function between the aerodynamic parameters and the Mach number.

[0028] According to the specific embodiments provided by the present invention, the following technical effects are disclosed by the present invention:

[0029] The present invention provides a method for quickly estimating parameters of a high-speed aircraft based on aerodynamic parameter experience, including: determining the aircraft state variables and aerodynamic shape parameters that affect the numerical values of the aerodynamic parameters to be estimated; performing aerodynamic calculations through computational fluid dynamics methods based on the aerodynamic shape parameters and overall parameters to obtain aerodynamic data; based on the aerodynamic data, gradually establishing a multivariate function relationship between the aerodynamic parameters, the aerodynamic shape, and the aircraft state variables by using the control variable method to obtain an aerodynamic parameter estimation model; and obtaining the aerodynamic parameters corresponding to the preset aircraft state variables according to the aerodynamic parameter estimation model. The present invention establishes a corresponding relationship between the optimizable variables and the aerodynamic parameters, so that when the optimizable variables of the aerodynamic shape take any values, a large amount of relatively reliable aerodynamic data that can be used for simulation experiments such as dynamic analysis and control design can be quickly obtained, improving the accuracy of aerodynamic parameter calculation. Description of the Drawings

[0030] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the accompanying drawings required in the embodiments. Obviously, the accompanying drawings in the following description are only some embodiments of the present invention. For those of ordinary skill in the art, without creative efforts, other accompanying drawings can also be obtained based on these drawings.

[0031] Figure 1 Flowchart of a method for quickly estimating high-speed aircraft parameters based on pneumatic parameter experience provided by an embodiment of the present invention;

[0032] Figure 2 Flowchart of the principle of quickly estimating high-speed aircraft parameters based on pneumatic parameter experience provided by an embodiment of the present invention;

[0033] Figure 3 Flowchart of determining the state variables affecting the numerical values of pneumatic parameters provided by an embodiment of the present invention;

[0034] Figure 4 Flowchart of obtaining pneumatic data through CFD calculation provided by an embodiment of the present invention;

[0035] Figure 5 Flowchart of obtaining a method for quickly estimating high-speed pneumatic parameters provided by an embodiment of the present invention;

[0036] Figure 6 Three-view drawings of the three-dimensional geometric model of the high-speed aircraft provided by an embodiment of the present invention, where Figure 6 (a) is the front view, Figure 6 (b) is the left view, Figure 6 (c) is the top view;

[0037] Figure 7 Scatter plot and least squares fitting curve of the variation law of the pneumatic parameter (here the axial force coefficient C x ) with respect to a single state variable (here the Mach number) provided by an embodiment of the present invention, with other influencing state variables remaining unchanged;

[0038] Figure 8 Scatter plot and least squares fitting curve of the variation law of the pneumatic parameter (here the normal force coefficient C y ) with respect to a single state variable (here the Mach number) provided by an embodiment of the present invention, with other influencing state variables remaining unchanged;

[0039] Figure 9 Scatter plot and least squares fitting curve of the variation law of the pneumatic parameter (here the pitching moment coefficient C mz ) with respect to a single state variable (here the Mach number) provided by an embodiment of the present invention, with other influencing state variables remaining unchanged. Detailed implementation manners

[0040] The following will clearly and completely describe the technical solutions in the embodiments of the present invention with reference to the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are only a part of the embodiments of the present invention, rather than all of the embodiments. All other embodiments obtained by those of ordinary skill in the art based on the embodiments of the present invention without creative efforts shall fall within the protection scope of the present invention.

[0041] The object of the present invention is to provide a method for quickly estimating the parameters of a high-speed aircraft based on pneumatic parameter experience. The present invention solves the problems in the prior art that for the calculation of pneumatic parameters, using the traditional theoretical calculation of aerodynamic characteristics, the calculation is complex and the accuracy of parameter calculation is low.

[0042] To make the above objects, features, and advantages of the present invention more obvious and understandable, the present invention will be further described in detail below with reference to the accompanying drawings and specific embodiments.

[0043] As Figure 1 shown, the present invention provides a method for quickly estimating the parameters of a high-speed aircraft based on pneumatic parameter experience, including:

[0044] Step S100: Determine the aircraft state variables and aerodynamic shape parameters that affect the value of the pneumatic parameter to be estimated; the aircraft state variables are altitude, sideslip angle, angle of attack, and Mach number, and the aerodynamic shape parameter is the half-cone angle;

[0045] Step S200: Perform aerodynamic calculations through computational fluid dynamics methods based on the aerodynamic shape parameters and overall parameters to obtain aerodynamic data;

[0046] Step S300: Based on the aerodynamic data, gradually establish a multi-variable function relationship between the pneumatic parameter and the aerodynamic shape and aircraft state variables using the control variable method to obtain a pneumatic parameter estimation model;

[0047] Step S400: Obtain the pneumatic parameters corresponding to the preset aircraft state variables according to the pneumatic parameter estimation model.

[0048] Specifically, the overall parameters include the characteristic area and the reference length.

[0049] Specifically, as Figure 2 shown, a principle for quickly estimating the parameters of a high-speed aircraft based on pneumatic parameter experience includes: S1: Determine the state variables that affect the pneumatic parameter; S2: Obtain aerodynamic data through CFD calculation; S3: Fit the function relationship between the state variables and the pneumatic parameter to establish a quick estimation formula.

[0050] Specifically, in combination with the trajectory characteristics of high-speed aircraft and engineering experience, the aircraft state variables that mainly affect the numerical values of aerodynamic parameters are determined. According to the aerodynamic shape and overall parameters of the high-speed aircraft, aerodynamic calculations are performed through relevant computational fluid dynamics (CFD) methods to obtain a large amount of aerodynamic data under different aircraft state variables. Based on the CFD calculation results, the control variable method is used to gradually establish a multivariate function relationship between the aerodynamic parameters and all aircraft state variables that mainly affect their values. According to this function, the numerical value of the aerodynamic parameter corresponding to a certain aircraft state variable can be calculated, thereby realizing the rapid estimation of high-speed aerodynamic parameters.

[0051] Specifically, as Figure 3 shown, S1.1: Determine the shape parameter affecting the aerodynamic parameter - the half-cone angle; S1.2 Determine the aircraft state variables affecting the aerodynamic parameter - Mach number, angle of attack, and sideslip angle.

[0052] Specifically, S101: In combination with the trajectory characteristics of high-speed aircraft and engineering experience, since the half-cone angle θ of the aircraft head is an optimizable shape parameter, different values of the half-cone angle θ will inevitably cause changes in the aerodynamic shape of the aircraft, resulting in changes in the numerical values of aerodynamic parameters. Therefore, the half-cone angle θ is one of the state variables affecting the numerical values of aerodynamic parameters. S102: In addition, the physical quantities that usually affect the numerical values of aerodynamic parameters include flight state variables such as Mach number Ma, altitude H, angle of attack α, and sideslip angle β. For this type of high-speed aircraft, the change in its aerodynamic parameters with altitude is largely due to the change in Mach number with the change in flight altitude, which causes changes in aerodynamic parameters. And considering its axisymmetric characteristics, only the influence of Mach number and angle of attack in the longitudinal channel on the axial force coefficient C x 、normal force coefficient C y and pitching moment coefficient is considered in the estimation. The influence laws of Mach number and sideslip angle in the lateral channel on the axial force coefficient C x 、lateral force coefficient C z and yaw moment coefficient are analogized with reference to the longitudinal channel, that is, Mach number, angle of attack / sideslip angle are the main state variables affecting the numerical values of aerodynamic parameters. In summary, the aircraft state variables that mainly affect the numerical values of aerodynamic parameters are determined.

[0053] Furthermore, the aerodynamic calculations are performed through computational fluid dynamics methods according to the aerodynamic shape and overall parameters to obtain aerodynamic data, including:

[0054] S201: Based on the overall parameters of the high-speed aircraft, according to the value range of the half-cone angle, construct a three-dimensional geometric model of the aircraft corresponding to the half-cone angle;

[0055] S202: Based on the three-dimensional geometric model of the aircraft, use finite element analysis software to obtain the mesh file corresponding to the value range of the half-cone angle;

[0056] A calculation node for calculating the state variables of the aircraft to obtain an aerodynamic calculation state node;

[0057] Construct working conditions corresponding to the value range of the half-cone angle according to the aerodynamic calculation state node;

[0058] S203: Based on the mesh file and the working conditions, using the two-equation turbulence model as the aerodynamic characteristic calculation model, obtain aerodynamic data.

[0059] Specifically, as Figure 4 shown, S2.1: Establish an aircraft geometric model corresponding to different half-cone angles; S2.2: Divide the CFD calculation area mesh based on the aircraft geometric model; S2.3: Perform CFD solution and process the calculation results to obtain aerodynamic data.

[0060] Specifically, based on the overall parameters of the high-speed aircraft, within the range of values of the head half-cone angle θ that can affect the aerodynamic shape of the aircraft and can be optimized, select respectively: (1) half-cone angle θ = 9.5°; (2) half-cone angle θ = 10°; (3) half-cone angle θ = 10.5°. Establish three-dimensional geometric models of high-speed aircraft under different aerodynamic shapes respectively. For the three-dimensional geometric model with half-cone angle θ = 10°, see Figure 6 .

[0061] For the three-dimensional geometric models of high-speed aircraft corresponding to three different values of the half-cone angle θ, use finite element analysis software to select about 1 million meshes as the standard for the number of meshes when dividing the mesh. Considering the axisymmetric aerodynamic shape characteristics of the aircraft, set the flow field boundary conditions in the plane of any symmetry plane of the aircraft and generate the flow field mesh outside the aircraft contour line through automated mesh generation. Then generate the volume mesh by rotating the surface mesh to generate mesh files that can be used for CFD calculation in three cases of different values of the half-cone angle θ respectively.

[0062] According to the state variables of the aircraft that determine the main influencing aerodynamic parameter values, combined with the trajectory characteristics of the high-speed aircraft, select the aerodynamic calculation state of the aircraft. Considering that the aircraft has an axisymmetric aerodynamic shape, when selecting its aerodynamic calculation state, only study the influence of the change of the angle of attack α on the aerodynamic characteristics of the aircraft, and all sideslip angles are taken as 0. Then, combined with the approximate change range of the Mach number at different flight altitudes of the aircraft, the calculation nodes of the state variables such as the Mach number Ma and the angle of attack α for aircraft aerodynamic calculation are shown in Table 1. Table 1 is the aerodynamic calculation state node table, and Table 1 is as follows:

[0063] Table 1 Aerodynamic calculation state node table

[0064]

[0065]

[0066] Through the permutations and combinations of all the above computing nodes, a total of 688 groups of flight conditions for all the proposed calculations are obtained. The generated grid files are loaded into the CFD calculation software, and each aerodynamic configuration is solved by CFD under these 688 conditions. The two-equation turbulence model is selected as the aerodynamic characteristic calculation model. Then, the initial conditions are set according to the state quantity values corresponding to the conditions and the overall parameters of the aircraft. The Riemann solver is set to the default value for supersonic / hypersonic according to the Mach number. The boundary conditions such as the far field, the aircraft wall surface, and the grid symmetry plane in the calculation area are set respectively, and the number of iterative solution steps is set to 800 - 1000 steps, etc., and the CFD calculation is carried out. Usually, during the CFD iterative solution process, when the residual can converge stably within 1×10 -3 and is considered that the calculation result is credible and the iterative solution can be stopped. The numerical values of the aerodynamic parameters after the final iterative solution is stabilized are obtained from the result file.

[0067] Furthermore, based on the aerodynamic data, the control variable method is used to gradually establish the multivariate function relationship between the aerodynamic parameters, the aerodynamic configuration, and the aircraft state variables, including:

[0068] S301: Based on the aerodynamic data, fit the polynomial function relationship between the aerodynamic parameters and the Mach number under the condition that the angle of attack remains unchanged to obtain the first relationship;

[0069] S302: Based on the first relationship, fit the polynomial function relationship between the coefficients of each term in the first relationship and the angle of attack under the condition that the Mach number remains unchanged to obtain the second relationship;

[0070] S303: Based on the first relationship and the second relationship corresponding to the aerodynamic data, fit the polynomial function relationship with the half-cone angle to obtain the third relationship;

[0071] S304: According to the first relationship, the second relationship, and the third relationship, obtain the aerodynamic parameter estimation model.

[0072] Figure 5 where f i corresponds to the i-th relationship.

[0073] As Figure 5 shown, S3.1: Fix the half-cone angle and the angle of attack, and fit the polynomial function relationship f1 between the aerodynamic parameters and the Mach number; S3.2: Fix the Mach number, and fit the polynomial function relationship f2 between the coefficients of each term of f1 and the angle of attack; S3.3: According to f1 and f2 corresponding to different half-cone angles, fit the polynomial function relationship f3 with the half-cone angle; S3.4: Substitute the half-cone angle, the angle of attack, and the Mach number corresponding to the state to be estimated into f1, f2, and f3 in turn, and calculate the function relationship to quickly estimate the aerodynamic data of the current state.

[0074] Specifically, when the half-cone angle θ = 10°, the aerodynamic parameter data corresponding to each state of the aircraft at the angle of attack α = 10° and only changing the Mach number are screened from the corresponding CFD calculation data, and the scatter function values with the Mach number Ma as the independent variable and the axial force coefficient C x , the normal force coefficient C y and the pitch moment coefficient as the dependent variables respectively can be obtained. Then, the least squares method is used for fitting the N-degree polynomial function of one variable, where N is taken as 5 here. The scatter data graph and the fitted fifth-degree polynomial curve of one variable can be seen in Figures 7 - 9 . The function expressions (the first relational expressions) of each aerodynamic parameter with respect to the Mach number Ma are fitted as follows:

[0075] C x = K5Ma 5 + K4Ma 4 + K3Ma 3 + K2Ma 2 + K1Ma + K0;

[0076] C y = P5Ma 5 + P4Ma 4 + P3Ma 3 + P2Ma 2 + P1Ma + P0;

[0077] C mz = Q5Ma 5 + Q4Ma 4 + Q3Ma 3 + Q2Ma 2 + Q1Ma + Q0;

[0078] Among them, the coefficients K5, K4, K3, K2, K1, K0, P5, P4, P3, P2, P1, P0 and Q5, Q4, Q3, Q2, Q1, Q0 before each order term are all determined through least squares fitting.

[0079] After obtaining the fitting functions of various aerodynamic parameters and Mach number at several different angles of attack α, the relationship between aerodynamic parameters and the angle of attack needs to be obtained next. For the aerodynamic state at the same altitude and the same Mach number, the only factor affecting the aerodynamic parameters is the angle of attack. Therefore, the coefficients K5, K4, K3, K2, K1, K0, P5, P4, P3, P2, P1, P0, and Q5, Q4, Q3, Q2, Q1, Q0 before each order term are all functions of the angle of attack α. Combining the trajectory characteristics of high-speed aircraft and engineering experience, the angle of attack α basically remains within a very small range throughout the flight trajectory. Here, the CFD calculation data corresponding to when the angle of attack α varies within the set {10, 9, 8, 7, 5, 4, 3, 2, 1, 0} is taken for fitting. And according to both theoretical analysis and CFD calculation results, for an axisymmetric aircraft, as the independent variable angle of attack α changes, the axial force coefficient C x is an even function symmetric about the y-axis, and the normal force coefficient C y and the pitching moment coefficient are odd functions symmetric about the origin. And from theoretical analysis and CFD calculation results, it can be seen that when the angle of attack α = 0, regardless of the Mach number, the normal force coefficient and pitching moment coefficient of the aircraft are both 0. Therefore, the function fitting expression of K5, K4, K3, K2, K1, K0 with respect to the angle of attack α is a quadratic polynomial of one variable, the function fitting expression of P5, P4, P3, P2, P1, P0 with respect to the angle of attack α is a cubic polynomial of one variable without a constant term, and the function fitting expression of Q5, Q4, Q3, Q2, Q1, Q0 with respect to the angle of attack α is a cubic polynomial of one variable without a constant term (the second relation), specifically:

[0080]

[0081]

[0082]

[0083] Based on the comprehensive fitting results, when the aircraft state variables such as the Mach number and the angle of attack of the aircraft are the same, when the aerodynamic parameters change due to the change of the half-cone angle θ, that is, it is considered that in k i1 , k i2 , k i3 , p i1 , p i2 , p i3 , q i1 , q i2 , q i3 , (i = 5, 4, 3, 2, 1, 0) are functions of the half-cone angle θ, and because 3 groups of different values of the half-cone angle θ are selected in the CFD calculation, so k i1 , k i2 , k i3 、pi1 , p i2 , p i3 , q i1 , q i2 , q i3 , (i = 5, 4, 3, 2, 1, 0) is fitted to a quadratic polynomial function (the third relation) with respect to the half-cone angle θ, and the specific form is:

[0084]

[0085] Specifically, C x is the axial force coefficient of the aircraft, C y is the normal force coefficient of the aircraft, is the pitch moment coefficient of the aircraft, θ is the half-cone angle, α is the angle of attack, Ma is the Mach number, is the fitting coefficient of the polynomial function between the coefficients in the second relation and the half-cone angle, k ij , p ij , q ij , (i = 5, 4, 3, 2, 1, 0; j = 1, 2, 3) is the fitting coefficient of the polynomial function between the coefficients in the first relation and the angle of attack, K i , P i , Q i , (i = 5, 4, 3, 2, 1, 0) is the fitting coefficient of the polynomial function between the aerodynamic parameters and the Mach number.

[0086] Therefore, based on the CFD calculation results, the specific values of each coefficient can be obtained by fitting the aerodynamic data in the aerodynamic calculation state with the same Mach number and angle of attack, and thus the influence of the half-cone angle can be introduced into the estimation of the aerodynamic calculation results. The obtained specific values are shown in Table 2.

[0087] Table 2 is the fitting value table, and Table 2 is as follows:

[0088] Table 2 Fitting value table

[0089]

[0090] Table 3 is the fitting value table, and Table 3 is as follows:

[0091] Table 3 Fitting value table

[0092]

[0093] Table 4 is the fitting value table, and Table 4 is as follows:

[0094] Table 4 Fitting value table

[0095]

[0096] Table 5 is Fitting value table, Table 5 is as follows:

[0097] Table 5 Fitting value table

[0098]

[0099]

[0100] Table 6 is Fitting value table, Table 6 is as follows:

[0101] Table 6 Fitting value table

[0102]

[0103] Table 7 is Fitting value table, Table 7 is as follows:

[0104] Table 7 Fitting value table

[0105]

[0106] Table 8 is Fitting value table, Table 8 is as follows:

[0107] Table 8 Fitting value table

[0108]

[0109]

[0110] Table 9 is Fitting value table, Table 9 is as follows:

[0111] Table 9 Fitting value table

[0112]

[0113] Table 1 is Fitting value table, Table 10 is as follows:

[0114] Table 2 Fitting value table

[0115]

[0116] According to the fitting calculation formula, substitute the half-cone angle θ into it successively

[0117]

[0118] the coefficients k before each term in the polynomial function fitting expressions of the coefficients K5, K4, K3, K2, K1, K0, P5, P4, P3, P2, P1, P0 and Q5, Q4, Q3, Q2, Q1, Q0 with respect to the angle of attack α can be obtained i1 , k i2 , k i3 , p i1 , p i2 , p i3 , q i1 , q i2 , q i3 , (i = 5, 4, 3, 2, 1, 0), and then substitute the values of the angle of attack α and Mach number Ma corresponding to the aerodynamic state to be estimated. The axial force coefficient C x , normal force coefficient C y and pitching moment coefficient of the fast estimated values. Through the calculation of the above three groups of polynomial functions, the aerodynamic parameter values corresponding to the current state can be quickly estimated.

[0119] The beneficial effects of the present invention are as follows:

[0120] The present invention provides a method for quickly estimating the parameters of a high-speed aircraft based on aerodynamic parameter experience. Combining the engineering experience of aerodynamic estimation and the CFD calculation results, approximate aerodynamic data of the optimizable variables of the aerodynamic shape of a high-speed aircraft under different values can be quickly obtained, so as to qualitatively and quantitatively analyze the dynamic characteristics and control design difficulties of high-speed aircraft with different aerodynamic shapes and assist in the aerodynamic shape optimization design of the aircraft. The method for quickly estimating high-speed aerodynamic parameters proposed by the present invention fully combines the advantages of two methods of theoretical calculation of aerodynamic characteristics, combines the judgment of influencing factors of aerodynamic characteristics in engineering experience and relatively more accurate CFD calculation data, and is used to estimate aerodynamic parameters, improving the accuracy of aerodynamic parameter estimation. The formed method for quickly estimating aerodynamic parameters has the advantages of simple calculation and clear physical meaning of parameters, and is closely combined with engineering design, which is helpful for the application of the present invention in the subsequent aerodynamic shape optimization design of high-speed aircraft.

[0121] Each embodiment in this specification is described in a progressive manner. The key point of each embodiment is to illustrate the differences from other embodiments. The same and similar parts among the embodiments can be referred to each other.

[0122] In this article, specific examples are used to illustrate the principles and implementation modes of the present invention. The descriptions of the above embodiments are only used to help understand the method of the present invention and its core idea. At the same time, for those of ordinary skill in the art, according to the idea of the present invention, there will be changes in the specific implementation modes and application scopes. To sum up, the content of this specification should not be construed as a limitation to the present invention.

Claims

1. A method for quickly estimating the parameters of a high-speed aircraft based on the experience of pneumatic parameters, characterized in that Including: Determine the aircraft state variables and aerodynamic shape parameters that affect the value of the aerodynamic parameter to be estimated; the aircraft state variables are altitude, sideslip angle, angle of attack, and Mach number, and the aerodynamic shape parameter is the half-cone angle; Based on the aerodynamic shape parameters and overall parameters, perform aerodynamic calculations using computational fluid dynamics methods to obtain aerodynamic data; Based on the aerodynamic data, use the method of controlling variables to gradually establish a multivariate function relationship between the aerodynamic parameter, the aerodynamic shape, and the aircraft state variables, and obtain an aerodynamic parameter estimation model; According to the aerodynamic parameter estimation model, obtain the aerodynamic parameters corresponding to the preset aircraft state variables; The performing aerodynamic calculations using computational fluid dynamics methods based on the aerodynamic shape and overall parameters to obtain aerodynamic data includes: Based on the overall parameters of the high-speed aircraft, within the value range of the half-cone angle, construct a three-dimensional geometric model of the aircraft corresponding to the value of the half-cone angle to be optimized; Based on the three-dimensional geometric model of the aircraft, use finite element analysis software to obtain the aircraft grid file corresponding to the value of the half-cone angle; Calculate the calculation nodes of the aircraft state variables to obtain the aerodynamic calculation state nodes; According to the aerodynamic calculation state nodes, construct the working conditions corresponding to the value range of the half-cone angle; According to the grid file and the working conditions, based on the two-equation turbulence model as the aerodynamic characteristic calculation model, obtain aerodynamic data; The using the method of controlling variables to gradually establish a multivariate function relationship between the aerodynamic parameter, the aerodynamic shape, and the aircraft state variables based on the aerodynamic data includes: Based on the aerodynamic data, fit the polynomial function relationship between the aerodynamic parameter and the Mach number under the condition that the angle of attack remains unchanged to obtain the first relationship; Based on the first relationship, fit the polynomial function relationship between the coefficients of each term in the first relationship and the angle of attack under the condition that the Mach number remains unchanged to obtain the second relationship; Based on the first relationship and the second relationship corresponding to the aerodynamic data, fit the polynomial function relationship between the coefficients of each term in the second relationship and the half-cone angle to obtain the third relationship; According to the first relationship, the second relationship, and the third relationship, obtain the aerodynamic parameter estimation model; The expression of the aerodynamic parameter estimation model is: Among them, C x is the axial force coefficient of the aircraft, C y is the normal force coefficient of the aircraft, is the pitching moment coefficient of the aircraft, θ is the half-cone angle, α is the angle of attack, Ma is the Mach number, are the fitting coefficients of the polynomial function between the coefficients of each term in the second relational expression and the half-cone angle, k ij , p ij , q ij are the fitting coefficients of the polynomial function between the coefficients of each term in the first relational expression and the angle of attack, K i , P i , Q i are the fitting coefficients of the polynomial function between the aerodynamic parameters and the Mach number, where i = 5, 4, 3, 2, 1, 0; j = 1, 2, 3.

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