Aerodynamic characteristics estimation method, device, equipment, medium and product for bionic aircraft

Abstract

The embodiment of the present invention provides a method, device, equipment, medium and product for estimating the aerodynamic characteristics of a bionic aircraft, and relates to the field of data processing technology. The method includes: inputting initial parameters into the wing geometry model of the bionic aircraft, outputting the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment; determining the torsion angle of the chord-wise load-bearing rod of the wing around the leading edge of the wing according to the aerodynamic resultant force; if the current moment has not reached the termination moment, updating the wing geometry model according to the torsion angle, and outputting the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment based on the updated wing geometry model after the interval time interval step, and executing the step of determining the torsion angle of the chord-wise load-bearing rod of the wing around the leading edge of the wing according to the aerodynamic resultant force; if the current moment reaches the termination moment, decomposing the aerodynamic resultant force. The present invention can estimate the aerodynamic characteristics of the bionic aircraft more accurately and quickly.

Description

Technical Field

[0001] The embodiments of the present invention relate to the field of data processing technology. More specifically, the embodiments of the present invention relate to a method, device, equipment, medium and product for estimating aerodynamic characteristics of a bionic aircraft. Background Art

[0002] This section is intended to provide a background or context to embodiments of the invention that are recited in the claims. No description herein is admitted to be prior art by inclusion in this section.

[0003] With the rapid development of science and technology, the application of bionics in the field of aircraft design is becoming more and more extensive. Bionic aircraft, especially aircraft that imitate the flapping flight mechanism of insects or birds in nature, have become a research hotspot in the aerospace field due to their unique flight mode and potential application value. However, the analysis of the aerodynamic characteristics of bionic aircraft has always been one of the key technical problems restricting its development.

[0004] Traditional aircraft aerodynamic characteristics analysis methods, such as numerical simulation methods based on computational fluid dynamics (CFD), have high accuracy, but they are computationally intensive and time-consuming, making it difficult to meet the requirements of real-time and rapid performance. Especially in the field of bionic aircraft, due to its complex flight mode and the particularity of its flexible flapping wing structure, traditional analysis methods often have difficulty in accurately and quickly estimating its aerodynamic characteristics. Summary of the invention

[0005] In this context, embodiments of the present invention are intended to provide a method, device, equipment, medium and product for estimating aerodynamic characteristics of a bionic aircraft.

[0006] In a first aspect of an embodiment of the present invention, a method for estimating aerodynamic characteristics of a bionic aircraft is provided, comprising:

[0007] Based on the three-dimensional coordinates of the leading edge point of the wing and the chord length, the wing geometric model of the bionic aircraft is established;

[0008] Inputting initial parameters into the wing geometric model, and outputting the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment; wherein the initial parameters include the end time and the time interval step;

[0009] Determining a torsion angle of a chord-wise load-bearing rod of the wing around a leading edge of the wing according to the aerodynamic resultant force;

[0010] If the current moment has not reached the termination moment, the wing geometric model is updated according to the torsion angle, and after the time interval step, the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment is output based on the updated wing geometric model, and the step of determining the torsion angle of the chord-wise load-bearing rod of the wing around the leading edge of the wing according to the aerodynamic resultant force is performed;

[0011] If the current moment reaches the end moment, the aerodynamic resultant force is decomposed to obtain the flapping lift in the vertical direction and the horizontal resultant force of the thrust and the drag in the horizontal direction.

[0012] In one embodiment of this implementation, the wing geometric model of the bionic aircraft is established based on the three-dimensional coordinates and chord length of the wing leading edge point, including:

[0013] Based on a preset spanwise segmentation number, determining the three-dimensional coordinates of all wing leading edge points after segmentation; wherein there are multiple wing leading edge points;

[0014] Obtaining the chord length of each leading edge point of the wing at an initial moment;

[0015] Based on the chord length and the number of chord-wise divisions at each leading edge point of the wing, a wing geometric model of the bionic aircraft is established.

[0016] In an example of the present implementation, the initial parameters also include the flight speed of the bionic aircraft, the flapping speed of the wing, and the flapping posture description function.

[0017] In one embodiment of this implementation, the wing geometric model is divided into a plurality of vortex grids, and the output of the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment includes:

[0018] The incoming flow velocity is calculated based on the incoming flow velocity and the conversion matrix in the global coordinate system; wherein the conversion matrix is ​​a conversion matrix for converting the global coordinate system to the aerodynamic calculation coordinate system;

[0019] The relative incoming flow velocity is calculated based on the length of the vortex grid from the flapping axis of the wing, the flapping angular velocity, and the flapping angle;

[0020] According to the incoming flow velocity, the relative incoming flow velocity and the induced velocity of each vortex grid, the sum velocity of each vortex grid is calculated;

[0021] Based on the summed speeds of the vortex grids, the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment is determined.

[0022] In an embodiment of this implementation manner, determining the torsion angle of the chord-wise load-bearing rod of the wing around the leading edge of the wing according to the aerodynamic resultant force includes:

[0023] Determining the resultant force action point of the aerodynamic resultant force in the chord direction;

[0024] The deflection generated at the end of the cantilever beam is calculated based on the aerodynamic resultant force, the point of action of the resultant force, the length of the cantilever beam of the wing, the elastic modulus of the material of the chord-wise load-bearing rod, and the moment of inertia of the cantilever beam of the wing around the fixed support position;

[0025] The deflection is calculated based on inverse trigonometric functions to obtain a torsion angle of the chord-wise load-bearing rod of the wing around the leading edge of the wing.

[0026] In one embodiment of this implementation manner, the calculation formula for calculating the deflection generated at the end of the cantilever beam based on the aerodynamic resultant force, the resultant force action point, the cantilever beam length of the wing, the elastic modulus of the chord-wise load-bearing rod material, and the moment of inertia of the cantilever beam of the wing around the fixed support position is:

[0027]

[0028] in, is the deflection, is the aerodynamic force, is the point of action of the resultant force, l is the length of the cantilever beam of the wing, E is the elastic modulus of the material of the chord-wise load-bearing rod, and I is the moment of inertia of the cantilever beam of the wing around the fixed support position.

[0029] In a second aspect of the embodiments of the present invention, there is provided a device for estimating aerodynamic characteristics of a bionic aircraft, comprising:

[0030] Establishing a model for establishing a wing geometry model of a bionic aircraft based on the three-dimensional coordinates of the leading edge point of the wing and the chord length;

[0031] An output model is used to input initial parameters into the wing geometric model and output the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment; wherein the initial parameters include the end time and the time interval step;

[0032] A determination model is used to determine the twisting angle of the chord-wise load-bearing rod of the wing around the leading edge of the wing according to the aerodynamic resultant force;

[0033] An updating model is used to update the wing geometry model according to the torsion angle if the current moment has not reached the termination moment, and output the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment based on the updated wing geometry model after the time interval step, and perform the step of determining the torsion angle of the chord-wise load-bearing rod of the wing around the leading edge of the wing according to the aerodynamic resultant force;

[0034] The decomposition model is used to decompose the aerodynamic resultant force if the current moment reaches the termination moment, so as to obtain the flapping wing lift in the vertical direction and the horizontal resultant force of the thrust and the drag in the horizontal direction.

[0035] In a third aspect of an embodiment of the present invention, a computing device is provided, comprising: at least one processor, a memory, and an input-output unit; wherein the memory is used to store a computer program, and the processor is used to call the computer program stored in the memory to execute any one of the methods described in the first aspect.

[0036] In a fourth aspect of the embodiments of the present invention, a computer-readable storage medium is provided, which includes instructions, and when the instructions are executed on a computer, the computer executes the method described in any one of the first aspects.

[0037] In a fifth aspect of the embodiments of the present invention, a computer program product is provided, comprising a computer program, which implements any one of the methods in the first aspect when executed by a processor.

[0038] According to the method, device, equipment, medium and product for estimating the aerodynamic characteristics of a bionic aircraft according to the embodiments of the present invention, a wing geometry model based on the three-dimensional coordinates of the leading edge point of the wing and the chord length can be constructed, and based on the wing geometry model, the aerodynamic characteristics results of the flapping-wing aircraft can be quickly obtained through the rapid aerodynamics and solid mechanics simplified estimation method, and the flapping-wing aerodynamic force calculation method of the unsteady vortex lattice method is adopted by introducing the reverse velocity of the flapping velocity as the incoming flow velocity, and the flapping-wing deformation is simplified and corrected by the cantilever beam bending theory, so as to more accurately and quickly estimate the aerodynamic characteristics of the bionic aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0039] The above and other objects, features and advantages of the exemplary embodiments of the present invention will become readily understood by reading the detailed description below with reference to the accompanying drawings. In the accompanying drawings, several embodiments of the present invention are shown in an exemplary and non-limiting manner, in which:

[0040] Figure 1 A schematic diagram of a flow chart of a method for estimating aerodynamic characteristics of a bionic aircraft provided in one embodiment of the present invention;

[0041] Figure 2 A schematic diagram of a wing geometric model of a bionic aircraft provided in one embodiment of the present invention;

[0042] Figure 3 A simplified schematic diagram of a wing load-bearing structure provided by an embodiment of the present invention;

[0043] Figure 4 A simplified schematic diagram of wing structure deformation based on cantilever beam bending theory provided by an embodiment of the present invention;

[0044] Figure 5 Schematic diagram of the comparison between the instantaneous lift curve at an angle of attack of 0° calculated by the aerodynamic characteristics estimation method of the bionic aircraft of the present invention and the CFD simulation result;

[0045] Figure 6 Schematic diagram of comparison between the instantaneous lift curve at an angle of attack of 6° calculated by the aerodynamic characteristics estimation method of the bionic aircraft of the present invention and the CFD simulation result;

[0046] Figure 7 A schematic diagram showing a comparison between an instantaneous lift curve at an angle of attack of 12° calculated by the aerodynamic characteristics estimation method of a bionic aircraft according to the present invention and a CFD simulation result;

[0047] Figure 8 A schematic diagram showing the comparison between the periodic average lift force varying with the angle of attack calculated by the bionic aircraft aerodynamic characteristics estimation method of the present invention and the CFD simulation result;

[0048] Fig. 9 A schematic diagram of the structure of a bionic aircraft aerodynamic characteristics estimation device provided by an embodiment of the present invention;

[0049] Fig.10 A schematic diagram of the structure of a medium according to an embodiment of the present invention is schematically shown;

[0050] Fig.11 A schematic diagram of the structure of a computing device according to an embodiment of the present invention is schematically shown.

[0051] In the drawings, the same or corresponding reference numerals represent the same or corresponding parts. DETAILED DESCRIPTION

[0052] The principles and spirit of the present invention will be described below with reference to several exemplary embodiments. It should be understood that these embodiments are provided only to enable those skilled in the art to better understand and implement the present invention, and are not intended to limit the scope of the present invention in any way. On the contrary, these embodiments are provided to make the present disclosure more thorough and complete, and to fully convey the scope of the present disclosure to those skilled in the art.

[0053] Those skilled in the art know that the embodiments of the present invention can be implemented as a system, device, apparatus, method or computer program product. Therefore, the present disclosure can be specifically implemented in the following forms, namely: complete hardware, complete software (including firmware, resident software, microcode, etc.), or a combination of hardware and software.

[0054] According to an embodiment of the present invention, a method, device, equipment, medium and product for estimating aerodynamic characteristics of a bionic aircraft are proposed.

[0055] It should be noted that any number of elements in the drawings is for illustration rather than limitation, and any naming is only for distinction and does not have any limiting meaning.

[0056] The principle and spirit of the present invention are explained in detail below with reference to several representative embodiments of the present invention.

[0057] Exemplary Methods

[0058] Reference below Figure 1 , Figure 1 This is a flow chart of a method for estimating aerodynamic characteristics of a bionic aircraft provided by an embodiment of the present invention. It should be noted that the embodiments of the present invention can be applied to any applicable scenario.

[0059] Figure 1 The process of the method for estimating aerodynamic characteristics of a bionic aircraft provided by an embodiment of the present invention includes:

[0060] Step S101, establishing a wing geometric model of the bionic aircraft based on the three-dimensional coordinates of the wing leading edge point and the chord length.

[0061] Please also read Figure 2 , Figure 2 A schematic diagram of a wing geometric model of a bionic aircraft provided in one embodiment of the present invention.

[0062] As an optional implementation, the method of establishing the wing geometric model of the bionic aircraft based on the three-dimensional coordinates and chord length of the wing leading edge point in step S101 may include the following steps:

[0063] Based on a preset spanwise segmentation number, determining the three-dimensional coordinates of all wing leading edge points after segmentation; wherein there are multiple wing leading edge points;

[0064] Obtaining the chord length of each leading edge point of the wing at an initial moment;

[0065] Based on the chord length and the number of chord-wise divisions at each leading edge point of the wing, a wing geometric model of the bionic aircraft is established.

[0066] In the embodiment of the present invention, the input conditions for establishing the wing geometric model of the bionic aircraft are the three-dimensional coordinates of i (i ≥ 2) wing leading edge points ( , , ), the chord length at each selected wing leading edge point and the spanwise division number and the chordwise division number between every two coordinate points. The spanwise division number and the chordwise division number can both be the number of division vortex grids.

[0067] Based on the number of spanwise divisions, the three-dimensional coordinates of all wing leading edge points after division are first defined. The definition method can be to use a two-variable first-order equation to perform linear fitting on the position points between each segment, and the spacing of all grid points between every two input three-dimensional coordinate points of the wing leading edge points is evenly distributed.

[0068] Then, the chord length at each leading edge point of the wing after the given segmentation at the initial moment is obtained.

[0069] Finally, based on the chord length data at each wing leading edge point and the number of grids that need to be divided along the chord direction, the chord-wise grid points are divided using the same method of fitting a binary linear equation.

[0070] Finally, the coordinates of all grid points on the wing surface are obtained, and thus the wing geometric model of the bionic aircraft is established using this method.

[0071] Step S102, inputting initial parameters into the wing geometric model, and outputting the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment.

[0072] In the embodiment of the present invention, the initial parameters may include the end time and the time interval step. The initial parameters may also include the flight speed of the bionic aircraft, the flapping speed of the wing and the flapping attitude description function.

[0073] Optionally, the wing geometric model is divided into a plurality of vortex grids, and the method of outputting the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment in step S102 may include the following steps:

[0074] The incoming flow velocity is calculated based on the incoming flow velocity and the conversion matrix in the global coordinate system; wherein the conversion matrix is ​​a conversion matrix for converting the global coordinate system to the aerodynamic calculation coordinate system;

[0075] The relative incoming flow velocity is calculated based on the length of the vortex grid from the flapping axis of the wing, the flapping angular velocity, and the flapping angle;

[0076] According to the incoming flow velocity, the relative incoming flow velocity and the induced velocity of each vortex grid, the sum velocity of each vortex grid is calculated;

[0077] Based on the summed speeds of the vortex grids, the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment is determined.

[0078] In the embodiment of the present invention, compared with the traditional steady-state vortex grid method, the unsteady vortex grid method used in this study introduces the relative incoming flow velocity generated by the flapping speed when calculating the wing circulation. Based on the wing model with divided vortex grids, the sum velocity at the i-th vortex grid is for:

[0079]

[0080] in, is the incoming flow velocity, is the relative flow velocity caused by the flapping speed of the wing surface, is the induced velocity of other vortices at the i-th vortex:

[0081]

[0082]

[0083]

[0084] in, is the incoming flow velocity in the global coordinate system, The transformation matrix from the global coordinate system to the aerodynamic calculation coordinate system; is the distance between the i-th vortex control point and the flapping axis, ω is the flapping angular velocity, and φ is the flapping angle; is the induced velocity of the jth vortex at the i-th vortex ( is the induced velocity of the unit intensity circulation at the jth vortex at the i-th vortex, is the actual circulation of the jth vortex grid). According to the no-penetration boundary condition, the normal velocity of the airfoil is 0:

[0085]

[0086] Step S103, determining the torsion angle of the chord-wise load-bearing rod of the wing around the leading edge of the wing according to the aerodynamic resultant force.

[0087] Please also read Figure 3 and Figure 4 , Figure 3 A simplified schematic diagram of a wing load-bearing structure provided in one embodiment of the present invention. Figure 4 A simplified schematic diagram of wing structure deformation based on cantilever beam bending theory provided in one embodiment of the present invention.

[0088] As an optional implementation, the method of determining the torsion angle of the chord-wise load-bearing rod of the wing around the leading edge of the wing according to the aerodynamic resultant force in step S103 may include the following steps:

[0089] Determining the resultant force action point of the aerodynamic resultant force in the chord direction;

[0090] The deflection generated at the end of the cantilever beam is calculated based on the aerodynamic resultant force, the point of action of the resultant force, the length of the cantilever beam of the wing, the elastic modulus of the material of the chord-wise load-bearing rod, and the moment of inertia of the cantilever beam of the wing around the fixed support position;

[0091] The deflection is calculated based on inverse trigonometric functions to obtain a torsion angle of the chord-wise load-bearing rod of the wing around the leading edge of the wing.

[0092] In the embodiment of the present invention, the wing surface load-bearing structure is first simplified. Except for the leading edge, it is assumed that only the chord-wise load-bearing rod structure is provided at the bending position, such as Figure 3 As shown, the two leading edge rods are hinged at the bending position, and the chord-wise load-bearing rod is fixedly supported with the inner wing leading edge rod to form a cantilever beam structure along the chord direction of the airfoil.

[0093] When considering the bending deformation angle of the chord-wise load-bearing rod, we first need to calculate the aerodynamic force of the outward extension of the load-bearing rod (in this embodiment, the outer wing section of the flapping wing) perpendicular to the wing surface. and its resultant force point in the x direction (chord direction) :

[0094]

[0095]

[0096] Assuming that the skin does not bear any load, the wing skin is in a tensioned state at any time, and the aerodynamic resultant can fully act on the load-bearing rod structure. The aerodynamic resultant of the outer wing section acts on the chord-wise load-bearing rod structure as shown in the figure. Figure 4 shown.

[0097] According to the cantilever beam bending theory, the force is applied to the fixed support point (the leading edge of the wing) When the cantilever beam is at a point, the calculation formula for calculating the deflection generated at the end of the cantilever beam can be:

[0098]

[0099] in, is the deflection, is the aerodynamic force, is the point of action of the resultant force, l is the cantilever beam length of the wing, that is, Figure 4The total length of the mid-chord bearing rod; E is the elastic modulus of the material of the chord bearing rod, and I is the moment of inertia of the cantilever beam of the wing around the fixed support position.

[0100] Furthermore, due to the small deformation, the bending deformation can be simplified to a rotation around the leading edge, and the torsion angle of the chord-wise load-bearing rod around the leading edge of the wing can be obtained by the inverse trigonometric function:

[0101]

[0102] Step S104, if the current moment has not reached the termination moment, the wing geometry model is updated according to the torsion angle, and after the time interval step, the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment is output based on the updated wing geometry model, and step S103 is executed.

[0103] In the embodiment of the present invention, after the torsion angle at the load-bearing rod is obtained, since the tail of the wing root is connected to the rudder surface, the displacement of the skin at the wing root is limited, and the chord-wise length at the wing tip is zero, the torsion angles around the leading edge at the wing root and the wing tip are both zero. Since the skin is assumed to be in a tensioned state, the torsion angles at other span-wise positions from the chord-wise load-bearing rod position to the wing root and the wing tip can be obtained by linear interpolation. After the torsion angles at the span-wise positions of each control point are obtained, the coordinates of each control point can be updated, and aerodynamic calculations can be performed again.

[0104] Step S105, if the current moment reaches the end moment, the aerodynamic resultant force is decomposed to obtain the flapping lift in the vertical direction and the horizontal resultant force of the thrust and the drag in the horizontal direction.

[0105] Please also read Figure 5 , Figure 6 , Figure 7 as well as Figure 8 , Figure 5 Schematic diagram of comparison between the instantaneous lift curve at an angle of attack of 0° calculated by the aerodynamic characteristics estimation method of the bionic aircraft of the present invention and the CFD simulation result; Figure 6 Schematic diagram of the comparison between the instantaneous lift curve at an angle of attack of 6° calculated by the aerodynamic characteristics estimation method of the bionic aircraft of the present invention and the CFD simulation result; Figure 7 A schematic diagram showing a comparison between an instantaneous lift curve at an angle of attack of 12° calculated by the bionic aircraft aerodynamic characteristics estimation method of the present invention and a CFD simulation result; Figure 8 This is a schematic diagram comparing the variation of the period average lift with the angle of attack calculated by the method for estimating the aerodynamic characteristics of a bionic aircraft according to the present invention with the CFD simulation results; it can be seen that the lift value within the period of this embodiment and the period average value have a good degree of similarity.

[0106] The present invention can construct a wing geometry model based on the three-dimensional coordinates of the leading edge point of the wing and the chord length, and based on the wing geometry model, through fast aerodynamics and solid mechanics simplified estimation methods, quickly obtain the aerodynamic characteristics results of a flapping-wing aircraft, and use an unsteady vortex lattice method to calculate the aerodynamic force of flapping wings by introducing the reverse velocity of the flapping velocity as the incoming flow velocity, and a method for simplifying and correcting the flapping wing deformation using the cantilever beam bending theory, thereby more accurately and quickly estimating the aerodynamic characteristics of a bionic aircraft.

[0107] Exemplary Devices

[0108] After introducing the method of the exemplary embodiment of the present invention, next, refer to Fig. 9 A device for estimating aerodynamic characteristics of a bionic aircraft according to an exemplary embodiment of the present invention is described, and the device comprises:

[0109] Establishing a model 901 for establishing a wing geometric model of a bionic aircraft based on the three-dimensional coordinates of a wing leading edge point and a chord length;

[0110] Output model 902 is used to input initial parameters into the wing geometry model and output the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment; wherein the initial parameters include the end time and the time interval step; the initial parameters may also include the flight speed of the bionic aircraft, the flapping speed of the wing and the flapping posture description function.

[0111] Determine model 903, for determining the twisting angle of the chord-wise load-bearing rod of the wing around the leading edge of the wing according to the aerodynamic resultant force;

[0112] The updating model 904 is used to update the wing geometric model according to the torsion angle if the current moment has not reached the termination moment, and output the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment based on the updated wing geometric model after the time interval step, and execute the determination model 903 to determine the torsion angle of the chord-wise load-bearing rod of the wing around the leading edge of the wing according to the aerodynamic resultant force;

[0113] The decomposition model 905 is used to decompose the aerodynamic resultant force if the current moment reaches the termination moment, so as to obtain the flapping lift in the vertical direction and the horizontal resultant force of the thrust and the drag in the horizontal direction.

[0114] As an optional implementation, the model 901 is established based on the three-dimensional coordinates and chord length of the wing leading edge point, and the wing geometric model of the bionic aircraft can be established in the following manner:

[0115] Based on a preset spanwise segmentation number, determining the three-dimensional coordinates of all wing leading edge points after segmentation; wherein there are multiple wing leading edge points;

[0116] Obtaining the chord length of each leading edge point of the wing at an initial moment;

[0117] Based on the chord length and the number of chord-wise divisions at each leading edge point of the wing, a wing geometric model of the bionic aircraft is established.

[0118] As an optional implementation, the wing geometric model is divided into a plurality of vortex grids, and the output model 902 outputs the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment in a specific manner as follows:

[0119] The incoming flow velocity is calculated based on the incoming flow velocity and the conversion matrix in the global coordinate system; wherein the conversion matrix is ​​a conversion matrix for converting the global coordinate system to the aerodynamic calculation coordinate system;

[0120] The relative incoming flow velocity is calculated based on the length of the vortex grid from the flapping axis of the wing, the flapping angular velocity, and the flapping angle;

[0121] According to the incoming flow velocity, the relative incoming flow velocity and the induced velocity of each vortex grid, the sum velocity of each vortex grid is calculated;

[0122] Based on the summed speeds of the vortex grids, the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment is determined.

[0123] As an optional implementation manner, the determination model 903 determines the torsion angle of the chord-wise load-bearing rod of the wing around the leading edge of the wing according to the aerodynamic resultant force in a specific manner as follows:

[0124] Determining the resultant force action point of the aerodynamic resultant force in the chord direction;

[0125] The deflection generated at the end of the cantilever beam is calculated based on the aerodynamic resultant force, the point of action of the resultant force, the length of the cantilever beam of the wing, the elastic modulus of the material of the chord-wise load-bearing rod, and the moment of inertia of the cantilever beam of the wing around the fixed support position;

[0126] The deflection is calculated based on inverse trigonometric functions to obtain a torsion angle of the chord-wise load-bearing rod of the wing around the leading edge of the wing.

[0127] The calculation formula for calculating the deflection generated at the end of the cantilever beam is:

[0128]

[0129] in, is the deflection, is the aerodynamic force, is the point of action of the resultant force, l is the length of the cantilever beam of the wing, E is the elastic modulus of the material of the chord-wise load-bearing rod, and I is the moment of inertia of the cantilever beam of the wing around the fixed support position.

[0130] Exemplary Media

[0131] After introducing the method and apparatus of the exemplary embodiment of the present invention, next, reference is made to Fig.10 For a description of a computer-readable storage medium according to an exemplary embodiment of the present invention, please refer to Fig.10 , the computer-readable storage medium shown is a CD 100, on which a computer program (i.e., a program product) is stored. When the computer program is executed by a processor, the steps recorded in the above method implementation are implemented, for example, a wing geometric model of a bionic aircraft is established based on the three-dimensional coordinates of the leading edge point of the wing and the chord length; initial parameters are input into the wing geometric model, and the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment is output; wherein the initial parameters include an end time and a time interval step; according to the aerodynamic resultant force, a torsion angle of the chord-wise load-bearing rod of the wing around the leading edge of the wing is determined; if the If the current moment has not reached the termination moment, the wing geometry model is updated according to the torsion angle, and after the time interval step, the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment is output based on the updated wing geometry model, and the step of determining the torsion angle of the chord-wise load-bearing rod of the wing around the leading edge of the wing according to the aerodynamic resultant force is executed; if the current moment reaches the termination moment, the aerodynamic resultant force is decomposed to obtain the flapping lift in the vertical direction and the horizontal resultant force of the thrust and drag in the horizontal direction; the specific implementation method of each step will not be repeated here.

[0132] It should be noted that examples of the computer-readable storage medium may also include, but are not limited to, phase change memory (PRAM), static random access memory (SRAM), dynamic random access memory (DRAM), other types of random access memory (RAM), read-only memory (ROM), electrically erasable programmable read-only memory (EEPROM), flash memory or other optical or magnetic storage media, which are not listed here one by one.

[0133] Exemplary Computing Devices

[0134] After introducing the method, apparatus and medium of the exemplary embodiments of the present invention, next, reference is made to Fig.11 A computing device for estimating aerodynamic characteristics of a bionic aircraft according to an exemplary embodiment of the present invention.

[0135] Fig.11 A block diagram of an exemplary computing device 110 suitable for implementing embodiments of the present invention is shown, and the computing device 110 may be a computer system or a server. Fig.11 The computing device 110 shown is only an example and should not bring any limitation to the functionality and scope of use of the embodiments of the present invention.

[0136] like Fig.11 As shown, the components of the computing device 110 may include, but are not limited to: one or more processors or processing units 1101 , a system memory 1102 , and a bus 1103 connecting various system components (including the system memory 1102 and the processing unit 1101 ).

[0137] The computing device 110 typically includes a variety of computer system readable media. These media can be any available media that can be accessed by the computing device 110, including volatile and nonvolatile media, removable and non-removable media.

[0138] System memory 1102 may include computer system readable media in the form of volatile memory, such as random access memory (RAM) 11021 and / or cache memory 11022. Computing device 110 may further include other removable / non-removable, volatile / non-volatile computer system storage media. By way of example only, ROM 11023 may be used to read and write non-removable, non-volatile magnetic media ( Fig.11 is not shown in the Fig.11 As shown in FIG. 1 , a disk drive for reading and writing a removable non-volatile disk (e.g., a "floppy disk") and an optical disk drive for reading and writing a removable non-volatile optical disk (e.g., a CD-ROM, a DVD-ROM, or other optical media) may be provided. In these cases, each drive may be connected to the bus 1103 via one or more data medium interfaces. The system memory 1102 may include at least one program product, which has a set (e.g., at least one) of program modules, which are configured to perform the functions of the various embodiments of the present invention.

[0139] A program / utility 11025 having a set (at least one) of program modules 11024 may be stored, for example, in system memory 1102, and such program modules 11024 include, but are not limited to: an operating system, one or more application programs, other program modules, and program data, each of which or some combination may include an implementation of a network environment. Program modules 11024 generally perform the functions and / or methods of the embodiments described herein.

[0140] The computing device 110 may also communicate with one or more external devices 1104 (e.g., a keyboard, a pointing device, a display, etc.). Such communication may be performed through an input / output (I / O) interface 1105. In addition, the computing device 110 may also communicate with one or more networks (e.g., a local area network (LAN), a wide area network (WAN), and / or a public network such as the Internet) through a network adapter 1106. Fig.11 As shown, the network adapter 1106 communicates with other modules (such as the processing unit 1101, etc.) of the computing device 110 via the bus 1103. It should be understood that although Fig.11 Not shown, other hardware and / or software modules may be used in conjunction with computing device 110 .

[0141] The processing unit 1101 executes various functional applications and data processing by running the program stored in the system memory 1102, for example, establishing a wing geometric model of the bionic aircraft based on the three-dimensional coordinates and chord length of the wing leading edge point; inputting initial parameters into the wing geometric model, and outputting the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment; wherein the initial parameters include the end time and the time interval step; according to the aerodynamic resultant force, determining the torsion angle of the chord-wise load-bearing rod of the wing around the wing leading edge; if the current moment If the termination moment is not reached, the wing geometry model is updated according to the torsion angle, and after the time interval step, the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment is output based on the updated wing geometry model, and the step of determining the torsion angle of the chord-wise load-bearing rod of the wing around the leading edge of the wing according to the aerodynamic resultant force is performed; if the current moment reaches the termination moment, the aerodynamic resultant force is decomposed to obtain the flapping lift in the vertical direction and the horizontal resultant force of the thrust and drag in the horizontal direction. The specific implementation method of each step is not repeated here. It should be noted that although several units / modules or sub-units / sub-modules of the bionic aircraft aerodynamic characteristics estimation device are mentioned in the above detailed description, this division is only exemplary and not mandatory. In fact, according to an embodiment of the present invention, the features and functions of two or more units / modules described above can be concretized in one unit / module. Conversely, the features and functions of one unit / module described above can be further divided into multiple units / modules to be concretized.

[0142] In the description of the present invention, it should be noted that the terms “first”, “second” and “third” are only used for descriptive purposes and cannot be understood as indicating or implying relative importance.

[0143] Those skilled in the art can clearly understand that, for the convenience and brevity of description, the specific working processes of the systems, devices and units described above can refer to the corresponding processes in the aforementioned method embodiments and will not be repeated here.

[0144] In the several embodiments provided by the present invention, it should be understood that the disclosed systems, devices and methods can be implemented in other ways. The device embodiments described above are only schematic. For example, the division of the units is only a logical function division. There may be other division methods in actual implementation. For example, multiple units or components can be combined or integrated into another system, or some features can be ignored or not executed. Another point is that the mutual coupling or direct coupling or communication connection shown or discussed can be through some communication interfaces, and the indirect coupling or communication connection of devices or units can be electrical, mechanical or other forms.

[0145] The units described as separate components may or may not be physically separated, and the components shown as units may or may not be physical units, that is, they may be located in one place or distributed on multiple network units. Some or all of the units may be selected according to actual needs to achieve the purpose of the solution of this embodiment.

[0146] In addition, each functional unit in each embodiment of the present invention may be integrated into one processing unit, or each unit may exist physically separately, or two or more units may be integrated into one unit.

[0147] If the functions are implemented in the form of software functional units and sold or used as independent products, they can be stored in a non-volatile computer-readable storage medium that can be executed by a processor. Based on this understanding, the technical solution of the present invention, or the part that contributes to the prior art, or the part of the technical solution, can be embodied in the form of a software product. The computer software product is stored in a storage medium, including several instructions for enabling a computer device (which can be a personal computer, server, or network device, etc.) to perform all or part of the steps of the method described in each embodiment of the present invention. The aforementioned storage medium includes: U disk, mobile hard disk, read-only memory (ROM, Read-Only Memory), random access memory (RAM, Random Access Memory), disk or optical disk, etc., various media that can store program codes.

[0148] Finally, it should be noted that the above-described embodiments are only specific implementations of the present invention, which are used to illustrate the technical solutions of the present invention, rather than to limit them. The protection scope of the present invention is not limited thereto. Although the present invention is described in detail with reference to the above-described embodiments, ordinary technicians in the field should understand that any technician familiar with the technical field can still modify the technical solutions recorded in the above-described embodiments within the technical scope disclosed by the present invention, or can easily think of changes, or make equivalent replacements for some of the technical features therein; and these modifications, changes or replacements do not make the essence of the corresponding technical solutions deviate from the spirit and scope of the technical solutions of the embodiments of the present invention, and should be included in the protection scope of the present invention. Therefore, the protection scope of the present invention shall be subject to the protection scope of the claims.

[0149] In addition, although the operations of the method of the present invention are described in a specific order in the drawings, this does not require or imply that the operations must be performed in this specific order, or that all the operations shown must be performed to achieve the desired results. Additionally or alternatively, some steps may be omitted, multiple steps may be combined into one step, and / or one step may be decomposed into multiple steps.

[0150] In an exemplary embodiment, a computer program product is provided, including a computer program, and when the computer program is executed by a processor, the steps in the above method embodiments are implemented.

Claims

1. A method for estimating aerodynamic characteristics of a bionic aircraft, comprising: Based on the three-dimensional coordinates of the leading edge point of the wing and the chord length, the wing geometric model of the bionic aircraft is established; Inputting initial parameters into the wing geometric model, and outputting the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment; wherein the initial parameters include the end time and the time interval step; Determining a torsion angle of a chord-wise load-bearing rod of the wing around a leading edge of the wing according to the aerodynamic resultant force; If the current moment has not reached the termination moment, the wing geometric model is updated according to the torsion angle, and after the time interval step, the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment is output based on the updated wing geometric model, and the step of determining the torsion angle of the chord-wise load-bearing rod of the wing around the leading edge of the wing according to the aerodynamic resultant force is performed; If the current moment reaches the end moment, the aerodynamic resultant force is decomposed to obtain the flapping lift in the vertical direction and the horizontal resultant force of the thrust and the drag in the horizontal direction; The wing geometric model is divided into a plurality of vortex grids, and the output of the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment includes: The incoming flow velocity is calculated based on the incoming flow velocity and the conversion matrix in the global coordinate system; wherein the conversion matrix is ​​a conversion matrix for converting the global coordinate system to the aerodynamic calculation coordinate system; The relative incoming flow velocity is calculated based on the length of the vortex grid from the flapping axis of the wing, the flapping angular velocity, and the flapping angle; According to the incoming flow velocity, the relative incoming flow velocity and the induced velocity of each vortex grid, the sum velocity of each vortex grid is calculated; Based on the summed speeds of the vortex grids, determining the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment; And, determining the torsion angle of the chord-wise load-bearing rod of the wing around the leading edge of the wing comprises: Determining the resultant force action point of the aerodynamic resultant force in the chord direction; The deflection generated at the end of the cantilever beam is calculated based on the aerodynamic resultant force, the point of action of the resultant force, the length of the cantilever beam of the wing, the elastic modulus of the material of the chord-wise load-bearing rod, and the moment of inertia of the cantilever beam of the wing around the fixed support position; Calculating the deflection based on inverse trigonometric functions to obtain a torsion angle of the chord-wise load-bearing rod of the wing around the leading edge of the wing; And, the calculation formula for calculating the deflection generated at the end of the cantilever beam according to the aerodynamic resultant force, the resultant force action point, the cantilever beam length of the wing, the elastic modulus of the chord-wise load-bearing rod material, and the moment of inertia of the cantilever beam of the wing around the fixed support position is: in, is the deflection, is the aerodynamic force, is the point of action of the resultant force, l is the length of the cantilever beam of the wing, E is the elastic modulus of the material of the chord-wise load-bearing rod, and I is the moment of inertia of the cantilever beam of the wing around the fixed support position.

2. The method for estimating aerodynamic characteristics of a bionic aircraft according to claim 1, wherein the wing geometric model of the bionic aircraft is established based on the three-dimensional coordinates of the wing leading edge point and the chord length, comprising: Based on a preset spanwise segmentation number, determining the three-dimensional coordinates of all wing leading edge points after segmentation; wherein there are multiple wing leading edge points; Obtaining the chord length of each leading edge point of the wing at an initial moment; Based on the chord length and the number of chord-wise divisions at each leading edge point of the wing, a wing geometric model of the bionic aircraft is established.

3. According to the method for estimating aerodynamic characteristics of a bionic aircraft according to claim 1, the initial parameters also include the flight speed of the bionic aircraft, the flapping speed of the wing and the flapping attitude description function.

4. A device for estimating aerodynamic characteristics of a bionic aircraft, comprising: Establishing a model for establishing a wing geometry model of a bionic aircraft based on the three-dimensional coordinates of the leading edge point of the wing and the chord length; An output model is used to input initial parameters into the wing geometric model and output the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment; wherein the initial parameters include the end time and the time interval step; A determination model is used to determine the twisting angle of the chord-wise load-bearing rod of the wing around the leading edge of the wing according to the aerodynamic resultant force; An updating model is used to update the wing geometry model according to the torsion angle if the current moment has not reached the termination moment, and output the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment based on the updated wing geometry model after the time interval step, and perform the step of determining the torsion angle of the chord-wise load-bearing rod of the wing around the leading edge of the wing according to the aerodynamic resultant force; A decomposition model is used to decompose the aerodynamic resultant force if the current moment reaches the termination moment, so as to obtain the flapping wing lift in the vertical direction and the horizontal resultant force of the thrust and the drag in the horizontal direction; The wing geometric model is divided into a plurality of vortex grids, and the output model outputs the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment in a specific manner: The incoming flow velocity is calculated based on the incoming flow velocity and the conversion matrix in the global coordinate system; wherein the conversion matrix is ​​a conversion matrix for converting the global coordinate system to the aerodynamic calculation coordinate system; The relative incoming flow velocity is calculated based on the length of the vortex grid from the flapping axis of the wing, the flapping angular velocity, and the flapping angle; According to the incoming flow velocity, the relative incoming flow velocity and the induced velocity of each vortex grid, the sum velocity of each vortex grid is calculated; Based on the summed speeds of the vortex grids, determining the aerodynamic resultant force perpendicular to the wing of the bionic aircraft at the current moment; And, the determination model determines the torsion angle of the chord-wise load-bearing rod of the wing around the leading edge of the wing in a specific manner: Determining the resultant force action point of the aerodynamic resultant force in the chord direction; The deflection generated at the end of the cantilever beam is calculated based on the aerodynamic resultant force, the point of action of the resultant force, the length of the cantilever beam of the wing, the elastic modulus of the material of the chord-wise load-bearing rod, and the moment of inertia of the cantilever beam of the wing around the fixed support position; Calculating the deflection based on inverse trigonometric functions to obtain a torsion angle of the chord-wise load-bearing rod of the wing around the leading edge of the wing; Furthermore, the determination model calculates the deflection generated at the end of the cantilever beam according to the aerodynamic resultant force, the resultant force action point, the cantilever beam length of the wing, the elastic modulus of the chord-direction load-bearing rod material, and the moment of inertia of the cantilever beam of the wing around the fixed support position, and the calculation formula is: in, is the deflection, is the aerodynamic force, is the point of action of the resultant force, l is the length of the cantilever beam of the wing, E is the elastic modulus of the material of the chord-wise load-bearing rod, and I is the moment of inertia of the cantilever beam of the wing around the fixed support position.

5. A computing device, comprising: at least one processor, memory, and input-output unit; The memory is used to store a computer program, and the processor is used to call the computer program stored in the memory to execute the method according to any one of claims 1 to 3.

6. A computer-readable storage medium comprising instructions, which, when executed on a computer, enables the computer to execute the method according to any one of claims 1 to 3.

7. A computer program product, comprising a computer program, characterized in that When the computer program is executed by a processor, the method according to any one of claims 1 to 3 is implemented.

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