Aircraft wing weight calculation method and device

Abstract

The present application provides a method and device for calculating the weight of an aircraft wing. The method includes: obtaining the design parameters of the aircraft wing; calculating the weight of each structural unit of the aircraft wing corresponding to the design parameters based on the weight calculation function of the aircraft wing structural unit; calculating the total weight of the aircraft wing according to the weight of each structural unit of the aircraft wing; and outputting the total weight of the aircraft wing. In this way, by linking the wing design parameters with the structural dimensions and structural weight and fully considering the actual load conditions and strength requirements, a "parameter-structure-weight" calculation framework is established, which can more conveniently, reliably and accurately estimate the wing weight.

Description

Technical Field

[0001] The present application relates to the technical field of aircraft design, and in particular to a method and device for calculating the weight of an aircraft wing. Background Art

[0002] Aircraft wing weight estimation is an essential and important part of the overall design process, and its accuracy determines the effectiveness and rationality of aircraft design. In the past, when conducting overall aircraft design, the wing weight calculation generally adopted the empirical formula method, engineering beam theory analysis method and structural finite element method. Among them, when using the empirical formula for calculation, it is necessary to have statistical data on the wing weight of aircraft with similar wing layouts, and the error of the estimated value is large. When using the finite element theory calculation method, a detailed structural model is required and the discrete finite elements are applied with loads and boundary conditions, so as to obtain the minimum weight of the structure that meets the rigidity and structural stability based on the structural analysis. However, in the initial design, there is generally no definite structural design data, and complex finite element meshing and analysis are required. Therefore, this method is only suitable for use in the detailed design and final design stages.

[0003] Therefore, it is necessary to provide a method and device for calculating the weight of an aircraft wing. In this way, by linking the wing design parameters with the structural dimensions and structural weight and taking into full consideration the actual load conditions and strength requirements, a "parameter-structure-weight" calculation framework is established, which can more conveniently, reliably and accurately estimate the wing weight. Summary of the invention

[0004] An embodiment of the present application provides a method for calculating the weight of an aircraft wing.

[0005] Specifically, a method for calculating the weight of an aircraft wing includes:

[0006] Obtain aircraft wing design parameters;

[0007] Calculating the weight of each structural unit of the aircraft wing corresponding to the design parameters based on the aircraft wing structural unit weight calculation function;

[0008] Calculating the total weight of the aircraft wing according to the weight of each structural unit of the aircraft wing;

[0009] The total weight of the aircraft wing is output.

[0010] Furthermore, the aircraft wing design parameters include at least one design parameter of the density of the wing design material, the length of the wing design root chord, the length of the wing design tip chord, the half span of the wing design, the leading edge sweep angle of the wing design, and the trailing edge sweep angle of the wing design.

[0011] Furthermore, before calculating the weight of each structural unit of the aircraft wing corresponding to the design parameters based on the aircraft wing structural unit weight calculation function, the method further includes:

[0012] The aerodynamic simulation model is used to calculate the spanwise lift coefficient distribution of the aircraft wing and the aircraft lift coefficient under simulated flight conditions.

[0013] Furthermore, the aircraft wing structure unit weight calculation function expression is as follows:

[0014] ,

[0015] in, is the shear weight of the spar structural unit, It is the bending weight of the wing spar structural unit;

[0016] ,

[0017] ,

[0018] In the formula, is the allowable shear stress of the material, is the spar material density, F s (y) is the shear force at any section of the wing along the span direction, is the equivalent cross-sectional area of ​​the spar.

[0019] Furthermore, the aircraft wing structure unit weight calculation function expression is as follows:

[0020] ,

[0021] In the formula, is the material weight density of the skin, and is the numerical integration length of the upper and lower curves of the airfoil, and are the average upper and lower skin thicknesses in the wing section, z is the structural span of the wing skin surface; is the structural span of the wing skin surface.

[0022] Furthermore, the aircraft wing structure unit weight calculation function expression is as follows:

[0023] ,

[0024] In the formula, is the empirical correction factor, is an empirical constant, is the statistical reference thickness of the rib, is the material density of the rib, is the cross-sectional reference area of ​​the i-th rib.

[0025] Furthermore, before calculating the total weight of the aircraft wing based on the weight of each structural unit of the aircraft wing, the method also includes confirming the weight of the aircraft wing accessories.

[0026] Further, according to the weight of each structural unit of the aircraft wing, calculating the total weight of the aircraft wing specifically includes:

[0027] Calculating the total weight of the aircraft wing structural units according to the weight of each structural unit of the aircraft wing;

[0028] Calculating the total weight of the aircraft wing according to the total weight of the aircraft wing structural unit and the weight of the auxiliary parts;

[0029] The calculation function of the total weight of the aircraft wing is expressed as follows:

[0030] ,

[0031] In the formula, is the total weight of the aircraft wing, is the weight of the aircraft wing structure unit, is the weight of the aircraft wing accessories.

[0032] An embodiment of the present application also provides a device for calculating the weight of an aircraft wing.

[0033] Specifically, an aircraft wing weight calculation device includes:

[0034] An acquisition module, used to acquire the design parameters of the aircraft wing;

[0035] A calculation module, used for calculating the weight of each structural unit of the aircraft wing corresponding to the design parameters based on the aircraft wing structural unit weight calculation function; wherein the calculation module is also used for calculating the total weight of the aircraft wing according to the weight of each structural unit of the aircraft wing;

[0036] The output module is used to output the total weight of the aircraft wing.

[0037] Furthermore, the calculation module is also used to calculate the spanwise lift coefficient distribution of the aircraft wing and the aircraft lift coefficient under simulated flight conditions through an aerodynamic simulation model.

[0038] The technical solution provided in the embodiments of the present application has at least the following beneficial effects:

[0039] By linking the wing design parameters with the structural dimensions and weight and taking into full consideration the actual load conditions and strength requirements, a "parameter-structure-weight" calculation framework is established, which can more conveniently, reliably and accurately estimate the wing weight. BRIEF DESCRIPTION OF THE DRAWINGS

[0040] The drawings described herein are used to provide a further understanding of the present application and constitute a part of the present application. The illustrative embodiments of the present application and their descriptions are used to explain the present application and do not constitute an improper limitation on the present application. In the drawings:

[0041] Figure 1 A flowchart of a method for calculating aircraft wing weight provided in an embodiment of the present application.

[0042] Figure 2 A schematic diagram of the structure of an aircraft wing weight calculation device provided in an embodiment of the present application.

[0043] 100 Aircraft wing weight calculation device

[0044] 11 Get Module

[0045] 12 Computing Module

[0046] 13 Output Module DETAILED DESCRIPTION

[0047] In order to make the purpose, technical solution and advantages of the present application clearer, the technical solution of the present application will be clearly and completely described below in combination with the specific embodiments of the present application and the corresponding drawings. Obviously, the described embodiments are only part of the embodiments of the present application, not all of the embodiments. Based on the embodiments in the present application, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of the present application.

[0048] Please refer to Figure 1 The present application discloses a method for calculating the weight of an aircraft wing, comprising:

[0049] S100: Obtaining aircraft wing design parameters.

[0050] It is understandable that the stable flight of an aircraft requires close cooperation of various components. As an indispensable component of an aircraft, the wing can generate enough lift to ensure the stable flight of the aircraft. However, under the same atmospheric conditions, the flight states of wings of different weights are different. The weight distribution of the wing has a great influence on the load and flight state of the aircraft, and its accuracy determines the effectiveness and rationality of the aircraft design. Therefore, the calculation of the weight of the aircraft wing is an indispensable and important link in the aircraft design process. If the weight ratio of the designed wing is not appropriate, the stable flight of the aircraft cannot be guaranteed. Therefore, when the aircraft is initially designed, it is necessary to focus on the weight ratio of the aircraft wing. When designing the wing of an aircraft, there will be a series of wing design parameters. For example, the initial structure, appearance shape, size, selected materials, and position relative to the fuselage of the aircraft. The parameters can directly affect the flight state of the designed aircraft. Therefore, when calculating the wing weight, it is necessary to obtain the design value of the parameter. It is understandable that the design parameters of the wing design described here, such as the initial structure, appearance shape, size, selected materials, and position relative to the aircraft fuselage, obviously do not constitute a limitation on the specific protection scope of this application.

[0051] Furthermore, in a preferred embodiment provided in the present application, the aircraft wing design parameters include at least one design parameter of the density of the wing design material, the length of the wing design root chord, the length of the wing design tip chord, the half span of the wing design, the leading edge sweep angle of the wing design, and the trailing edge sweep angle of the wing design.

[0052] It is understandable that when the aircraft is initially designed, it is necessary to fully consider the impact of the aircraft wing weight ratio on the flight state of the aircraft. Although there will be a series of design parameters corresponding to the design of the aircraft, there will be no detailed design structure data due to the lack of a determined design structure. At this time, the initial design parameters related to the wing that can be determined are limited. Through the determined initial design parameters, a data basis can be provided for the initial shape modeling of the wing. The determination of the initial shape of the wing is closely related to the length of the wing design root chord, the length of the wing design tip chord, the half span of the wing design, the leading edge sweep angle of the wing design, and the trailing edge sweep angle of the wing design. By determining the parameters, the established wing model can be made more accurate. Through the establishment of the initial wing model, the reference values ​​of the wing-related structures in the initial design scheme can be obtained, thereby providing a basis for the calculation of the wing weight. In addition, since different materials have different densities, the same design structure of the same volume has different weights if different materials are selected. Therefore, when designing the wing, it is also necessary to fully consider the problem of design material selection. The corresponding wing weights of wings of different design materials are also different. Under the condition that the wing design structure dimensions are the same, the lift generated by materials of different densities is also different. Therefore, in order to facilitate the accurate calculation of the wing weight, it is necessary to obtain at least one design parameter of the wing design material density, the length of the wing design root chord, the length of the wing design tip chord, the half span of the wing design, the leading edge sweep angle of the wing design, and the trailing edge sweep angle of the wing design. It is understandable that the specific values ​​of the wing design parameters obtained here obviously do not constitute a limitation on the specific protection scope of this application.

[0053] Further, in a preferred embodiment provided by the present application, before calculating the weight of each structural unit of the aircraft wing corresponding to the design parameters based on the aircraft wing structural unit weight calculation function, the method further includes:

[0054] The aerodynamic simulation model is used to calculate the spanwise lift coefficient distribution of the aircraft wing and the aircraft lift coefficient under simulated flight conditions.

[0055] It is understandable that the wing can provide stable lift for the flight of the aircraft. The calculation formula for the lift generated by the wing is: lift = 1 / 2 * air density * speed squared * wing area * wing lift coefficient. It can be seen that the lift generated by the wing is related to the wing lift coefficient in addition to the air density, flight speed and wing area. Among them, the wing area is determined by the external dimensions of the wing design. The wing lift coefficient is related to the shape of the wing cross section (airfoil), the angle between the airflow and the wing (angle of attack), etc. Under the same flight conditions, wings of different design schemes have different wing areas and airfoils. Therefore, the lift that can be generated by wings of different design schemes is different. In order to accurately calculate the weight of the wing, it is necessary to fully consider the impact of the wing design scheme on the flight state of the aircraft during actual flight. In order to obtain the impact of the wing on the flight state of the aircraft in the corresponding design scheme, it is necessary to simulate the actual flight state of the corresponding aircraft, that is, to perform aerodynamic simulation. In this way, the lift coefficient of the aircraft wing along the span direction and the lift coefficient of the aircraft as a whole can be obtained by calculation under simulated flight conditions.

[0056] Specifically, the aerodynamic load distribution function of the aircraft wing at any position along the wing span can be calculated by the following formula: ; In the formula, is the overall lift coefficient of the aircraft, which can be obtained from the aerodynamic analysis process; is any spanwise position of the wing relative to the wing root; is the chord length at any position in the span direction; is the geometric mean chord length of the wing; is the spanwise circulation distribution function. It can be obtained by the following formulas: ; ; ; In the formula, The spanwise lift coefficient distribution can be obtained by building a three-dimensional wing model and simulating it using computational fluid dynamics software under actual working conditions. is the root chord length of the wing; is the effective area of ​​the wing, which can be expressed as: . In this way, by performing aerodynamic simulation on the established three-dimensional model, the spanwise coefficient distribution of the wing under actual flight conditions can be obtained, thereby providing an accurate data basis for the calculation of the wing weight. It can be understood that the formulas described here are all calculation formulas for obtaining the spanwise lift coefficient distribution of the wing under actual conditions, and different methods can be used. The described embodiments are only part of the embodiments of the present application, not all of the embodiments. Based on the embodiments in the present application, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of this application.

[0057] S200: Calculating the weight of each structural unit of the aircraft wing corresponding to the design parameters based on a weight calculation function of the aircraft wing structural unit.

[0058] It is understandable that, by combining the design structure, design parameters, aerodynamic load and other factors affecting the wing weight of the aircraft wing, the weight of the aircraft wing can be calculated more accurately while avoiding the complex finite element analysis process. Among them, the aerodynamic load is obtained by performing aerodynamic simulation calculations on the designed wing and the aircraft. The design parameters can be obtained by obtaining at least one design parameter of the density of the wing design material, the length of the wing design root chord, the length of the wing design tip chord, the half span of the wing design, the leading edge sweep angle of the wing design, and the trailing edge sweep angle of the wing design in the initial design scheme. The design structure can be obtained by establishing a three-dimensional model of the wing. However, during the initial design, since there is no specific design structure determined, there are only simple main structures in the three-dimensional model of the wing. For example. A skin structure for transmitting aerodynamic force and aerodynamic loads, or a rib structure for maintaining the basic shape of the wing airfoil and transmitting the local aerodynamic loads on the skin to the wing beam, or a spar structure for maintaining the basic shape of the wing span. It can be seen from this that the wing is not a simple regular mechanical structure. Therefore, it is necessary to divide the wing structure into several structural units and then calculate them through corresponding calculation functions. That is, the wing structure is discretized into wing sections, and the weight of each wing section is calculated respectively. In this way, by calculating the weight of each wing section and then calculating the total weight of the wing, the weight of the wing can be made more accurate. At this time, the number of wing sections and the distance between each section can be determined according to parameters such as the efficiency factor and ultimate compressive strength of the design material. While the wing is discretized, the main structure of the wing is discretized into structural sections of corresponding number and size. It is understandable that the specific number and specific size of the wing sections after the wing discretization described here obviously do not constitute a limitation on the specific protection scope of this application.

[0059] Further, in a preferred embodiment provided in the present application, the aircraft wing structure unit weight calculation function expression is as follows:

[0060] ,

[0061] in, is the shear weight of the spar structural unit, It is the bending weight of the wing spar structural unit;

[0062] ,

[0063] ,

[0064] In the formula, is the allowable shear stress of the material, is the spar material density, F s (y) is the shear force at any section of the wing along the span direction, is the equivalent cross-sectional area of ​​the spar.

[0065] It is understandable that in order to make the wing weight more accurate, the wing needs to be discretized into corresponding wing sections, and the weight of each wing section needs to be calculated separately before calculating the total weight of the wing. When the initial design of the wing only includes the wing beam that makes up the wing, the weight of the wing structure unit needs to be calculated according to the calculation formula of the weight of the wing beam structure unit. Among them, since the wing beam needs to bear the bending moment and shear force of the wing at the same time during actual flight, when calculating the weight of the wing beam structure unit, its shear weight and bending weight need to be considered at the same time. At this time, the weight calculation function of the aircraft wing structure unit is: In the formula, is the shear weight of the spar structural unit, It is the bending weight of the wing spar structural unit.

[0066] Specifically, the shear weight of the wing spar structural unit is The calculation formula is: ; In the formula, is the allowable shear stress of the material, is the spar material density, is the shear force at any section of the wing along the span direction. It can be calculated by the following formula: In the formula, is the structural length of the spar, is the chord length at any spanwise position, is the dynamic pressure. Assuming that the sweep angle of the elastic axis of the spar is equal to the sweep angle of the leading edge, the structural length of the spar can be expressed by the half span and the sweep angle as follows: The dynamic pressure calculation formula is: ; is the atmospheric density, is the actual flight speed. In addition, the bending weight of the wing beam structural unit The calculation formula is: In the formula, is the spar material density, is the equivalent cross-sectional area of ​​the wing beam. The equivalent cross-sectional area of ​​the wing beam can be calculated by the following formula: In the formula, is the efficiency factor, generally 0.81~0.84; effective depth It can be approximated as: , is the bending efficiency coefficient, generally ranging from 0.87 to 0.89. It is the maximum thickness of the wing at the current spanwise position, which can be calculated based on the upper and lower coordinate parameters of the airfoil; is the thickness of the spar at the current position, which can be calculated by the following formula: , is the safety factor, generally taken as 1.5; is the maximum bending stress that the material can withstand; is the actual lift of the wing, which can be approximately calculated as ; is the reference position of the wing's center of pressure, and the calculation formula is as follows: ; is the bending moment at any position in the span direction, and the calculation formula is as follows: . It is understandable that the formulas described here are all used to calculate the weight of the wing beam structural unit, and can be designed to different values ​​according to actual conditions. The described embodiments are only part of the embodiments of this application, not all of them. Based on the embodiments in this application, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of this application.

[0067] Further, in a preferred embodiment provided in the present application, the aircraft wing structure unit weight calculation function expression is as follows:

[0068] ,

[0069] In the formula, is the material weight density of the skin, and is the numerical integration length of the upper and lower curves of the airfoil, and are the average upper and lower skin thicknesses in the wing section, z is the structural span of the wing skin surface; is the structural span of the wing skin surface.

[0070] It is understandable that the wing weight can be more accurate if the wing is discretized into corresponding wing segments, the weight of each wing segment is calculated separately, and then the total weight of the wing is calculated. When the initial design of the wing only includes the skin that makes up the wing, it is necessary to calculate the weight of the wing structure unit according to the calculation formula of the skin structure unit weight. Among them, since the skin structure is subjected to compression force on the upper wing surface and tensile force on the lower wing surface during actual flight. Therefore, when calculating the weight of the skin structure unit, it is necessary to comprehensively consider the different stress conditions of the upper and lower wing skins of the wing. The equivalent upper and lower skin thickness can be determined based on the load-bearing conditions of the skin at different spanwise positions. At this time, the weight calculation function of the aircraft wing structure unit is: In the formula, is the material weight density of the skin, and is the numerical integration length of the upper and lower curves of the airfoil, and are the average upper and lower skin thicknesses in the wing section, is the structural span of the wing skin surface; is the structural span of the wing skin surface.

[0071] Specifically, the average upper skin thickness in the wing section is It can be calculated by the following formula: In the formula, is the bending moment distribution on the wing section i; is the allowable tensile stress; is the chord length of the upper skin, which is equal to The average lower skin thickness in the wing section It can be calculated by the following formula: In the formula, is the bending moment distribution on the wing section i; is the allowable extrusion stress; is the chord length of the lower skin, which is equal to The structural span of the wing skin surface can be calculated using the following formula: . It is understandable that the formulas described here are all used to calculate the weight of the wing skin structural unit, and can be designed to different values ​​according to actual conditions. The described embodiments are only part of the embodiments of this application, not all of the embodiments. Based on the embodiments in this application, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of this application.

[0072] Further, in a preferred embodiment provided in the present application, the aircraft wing structure unit weight calculation function expression is as follows:

[0073] ,

[0074] In the formula, is the empirical correction factor, is an empirical constant, is the statistical reference thickness of the rib, is the material density of the rib, is the cross-sectional reference area of ​​the i-th rib.

[0075] It is understandable that in order to make the wing weight more accurate, the wing needs to be discretized into corresponding wing sections, and the weight of each wing section needs to be calculated separately before calculating the total weight of the wing. When the initial design of the wing only includes the ribs that make up the wing, it is necessary to calculate the weight of the wing structure unit according to the calculation formula of the weight of the rib mechanism unit. The main function of the wing rib is to maintain the basic shape of the wing airfoil and transfer the local aerodynamic load on the skin to the wing beam. If you want to accurately calculate the weight of the wing rib, you need to understand the topological structure, rib cross-sectional area, material properties of the wing rib, and the distribution and number of the rib along the span of the wing. It can be seen that there are many factors affecting the weight calculation of the wing rib structure. In the initial design of the wing, the wing structure is relatively simple. Therefore, when calculating the weight of the wing rib, the modified form of the semi-empirical equation is used. At this time, the weight calculation function of the aircraft wing structure unit is: In the formula, is the empirical correction factor, is an empirical constant, is the statistical reference thickness of the rib, is the material density of the rib, is the cross-sectional reference area of ​​the i-th rib.

[0076] Specifically, the cross-sectional reference area of ​​the rib is It can be calculated by the following formula: In the formula, and are the normalized ordinate values ​​of the upper and lower curves of the airfoil respectively; the distance between two adjacent ribs It can be indirectly determined by the ultimate compressive strength of the wing material: .in, is the efficiency factor of the material, generally ranging from 0.8 to 0.9; is the material tangent modulus; is the ultimate compressive strength of the material; is the ultimate compressive load strength of the i-th section of the wing. In addition, according to the calculation results of the overall parameters of the wing and the spacing between the ribs of each unit, the number of ribs can be determined, and the ribs can be used as nodes as a reference for the segmentation of the wing. It can be understood that the formulas described here are all used to calculate the weight of the wing rib structure unit, and can be designed to different values ​​according to actual conditions. The described embodiments are only part of the embodiments of the present application, not all of the embodiments. Based on the embodiments in the present application, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of this application.

[0077] S300: Calculate the total weight of the aircraft wing according to the weight of each structural unit of the aircraft wing.

[0078] It is understandable that the total weight of the aircraft wing can be calculated by the weight of the wing structure unit. In this way, when calculating the total weight of the wing, the wing design parameters, structural dimensions, and structural weight can be fully combined. In addition, after the aircraft wing is divided into wing structure units, the actual load conditions and strength requirements of the wing are fully combined in the calculation process of the structure unit weight. In this way, when the weight of the aircraft wing structure unit is calculated, the total weight of the aircraft wing can be calculated more accurately according to the weight of each structure unit of the aircraft wing.

[0079] Furthermore, in a preferred embodiment provided in the present application, before calculating the total weight of the aircraft wing based on the weight of each structural unit of the aircraft wing, it also includes confirming the weight of the aircraft wing accessories.

[0080] It is understandable that a complete wing cannot be assembled only through the main structural parts of the wing, such as the wing spars, skins, and wing ribs. In addition to the main structural parts, the wing also includes some secondary structures. That is, aircraft wing accessories. For example, secondary structures such as connectors, rivets, local opening reinforcement components, leading and trailing edge structures, glue and paint between wing structures. The weight of the accessories generally accounts for about 5% to 10% of the total weight of the wing structure, or about 5.26% to 11% of the weight of the main structure. Therefore, when calculating the total weight of the aircraft wing, the weight of the accessories cannot be ignored. It is understandable that the aircraft wing accessories described here are all parts used to assist the main structure of the wing to be assembled into a complete wing, which can be embodied in different forms. The aircraft wing accessories should not be understood as only including connectors, rivets, local opening reinforcement components, leading and trailing edge structures, glue and paint between wing structures. The described embodiments are only part of the embodiments of this application, not all of them. Based on the embodiments in this application, all other embodiments obtained by ordinary technicians in this field without making any creative work shall fall within the scope of protection of this application.

[0081] Further, in a preferred embodiment provided in the present application, calculating the total weight of the aircraft wing according to the weight of each structural unit of the aircraft wing specifically includes:

[0082] Calculating the total weight of the aircraft wing structural units according to the weight of each structural unit of the aircraft wing;

[0083] Calculating the total weight of the aircraft wing according to the total weight of the aircraft wing structural unit and the weight of the auxiliary parts;

[0084] The calculation function of the total weight of the aircraft wing is expressed as follows:

[0085] ,

[0086] In the formula, is the total weight of the aircraft wing, is the weight of the aircraft wing structure unit, is the weight of the aircraft wing accessories.

[0087] It is understandable that the total weight of the aircraft wing can be calculated by the weight of the wing structure unit. Among them, the wing structure unit is obtained by discretizing the wing. When the wing is initially designed, there are only relevant main structures, so the wing structure unit only includes the main structure unit in the wing structure unit. In this way, by calculating the weight of the wing structure unit, only the total weight of the main structure of the wing can be obtained. However, a complete wing cannot be assembled only by the main structural parts of the wing. In addition to the main structural parts, the wing also includes some secondary structures. That is, the auxiliary parts of the aircraft wing. For example, the connecting parts, rivets, local opening reinforcement parts, leading and trailing edge structures, glue and paint between the wing structures are secondary structures. The weight of the auxiliary parts generally accounts for about 5% to 10% of the total weight of the wing structure, or about 5.26% to 11% of the main structure. Therefore, when calculating the total weight of the aircraft wing, the weight of the auxiliary parts cannot be ignored. It can be seen that the total weight of the aircraft wing structure unit and the weight of the aircraft wing auxiliary parts must be fully considered when calculating the total weight of the aircraft wing. The calculation function of the total weight of the aircraft wing is expressed as: In the formula, is the total weight of the aircraft wing, is the weight of the aircraft wing structure unit, is the weight of the aircraft wing accessories.

[0088] Specifically, when the aircraft wing is designed and its initial design solution only includes the wing beams constituting the wing, the calculation function of the total weight of the aircraft wing can be expressed as: In the formula, is the shear weight of the wing spar structural unit: ; is the bending weight of the wing spar structural unit: .in, is the allowable shear stress of the material, is the spar material density, F s (y) is the shear force at any section of the wing along the span direction, is the equivalent cross-sectional area of ​​the spar.

[0089] When the aircraft wing is designed and its initial design only includes the skin that constitutes the wing, the calculation function of the total weight of the aircraft wing can be expressed as: In the formula, is the material weight density of the skin, and is the numerical integration length of the upper and lower curves of the airfoil, and are the average upper and lower skin thicknesses in the wing section, z is the structural span of the wing skin surface; is the structural span of the wing skin surface.

[0090] When the aircraft wing is designed and its initial design only includes the ribs that make up the wing, the calculation function of the total weight of the aircraft wing can be expressed as: In the formula, is the empirical correction factor, is an empirical constant, is the statistical reference thickness of the rib, is the material density of the rib, is the cross-sectional reference area of ​​the i-th rib.

[0091] When the aircraft wing is designed, and its initial design includes the spar and skin that make up the wing, the calculation function of the total weight of the aircraft wing can be expressed as: In the formula, is the shear weight of the wing spar structural unit: ; is the bending weight of the wing spar structural unit: ; is the material weight density of the skin, and is the numerical integration length of the upper and lower curves of the airfoil, and are the average upper and lower skin thicknesses in the wing section, z is the structural span of the wing skin surface; is the structural span of the wing skin surface. is the allowable shear stress of the material, is the spar material density, F s (y) is the shear force at any section of the wing along the span direction, is the equivalent cross-sectional area of ​​the spar.

[0092] When the aircraft wing is designed, and its initial design includes the wing beam and wing ribs that make up the wing, the calculation function of the total weight of the aircraft wing can be expressed as: In the formula, is the shear weight of the wing spar structural unit: ; is the bending weight of the wing spar structural unit: ; is the empirical correction factor, is an empirical constant, is the statistical reference thickness of the rib, is the material density of the rib, is the cross-sectional reference area of ​​the i-th rib. is the allowable shear stress of the material, is the spar material density, F s (y) is the shear force at any section of the wing along the span direction, is the equivalent cross-sectional area of ​​the spar.

[0093] When the aircraft wing is designed, and its initial design includes the ribs and skin that make up the wing, the calculation function of the total weight of the aircraft wing can be expressed as: In the formula, is the material weight density of the skin, and is the numerical integration length of the upper and lower curves of the airfoil, and are the average upper and lower skin thicknesses in the wing section, z is the structural span of the wing skin surface; is the structural span of the wing skin surface; is the empirical correction factor, is an empirical constant, is the statistical reference thickness of the rib, is the material density of the rib, is the cross-sectional reference area of ​​the i-th rib.

[0094] When the aircraft wing is designed, and its initial design includes the wing beam, wing ribs and skin that make up the wing, the calculation function of the total weight of the aircraft wing can be expressed as: In the formula, is the shear weight of the wing spar structural unit: ;

[0095] is the bending weight of the wing spar structural unit: ; is the material weight density of the skin, and is the numerical integration length of the upper and lower curves of the airfoil, and are the average upper and lower skin thicknesses in the wing section, z is the structural span of the wing skin surface; is the structural span of the wing skin surface; is the empirical correction factor, is an empirical constant, is the statistical reference thickness of the rib,

[0096] is the material density of the rib, is the cross-sectional reference area of ​​the i-th rib. is the allowable shear stress of the material, is the spar material density, F s (y) is the shear force at any section of the wing along the span direction, is the equivalent cross-sectional area of ​​the spar.

[0097] It is understandable that the calculation formulas described here are all used to calculate the total weight of the wing, and can be designed to different values ​​according to actual conditions. The described embodiments are only part of the embodiments of this application, not all of them. Based on the embodiments in this application, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of this application.

[0098] S400: Output the total weight of the aircraft wing.

[0099] It is understandable that by dividing the aircraft wing into different wing structural units to calculate the total weight of the wing, a more accurate wing weight value can be obtained. When the total weight of the aircraft wing is calculated according to the total weight calculation function of the aircraft wing, the corresponding calculation result will be output. By outputting the wing weight calculation result, the designer can accurately know the weight of the wing under the current design scheme. In this way, through the comprehensive analysis of the wing weight corresponding to the design scheme and the weight of the remaining aircraft components, the designer can be assisted in judging whether the aircraft design corresponding to the current design scheme is suitable. After judgment, if the design is suitable, the next step of design can be carried out. If the design is not suitable, the design scheme can be modified in time.

[0100] Please refer to Figure 2 The present application discloses an aircraft wing weight calculation device 100, comprising:

[0101] The acquisition module 11 is used to acquire the design parameters of the aircraft wing.

[0102] It is understandable that the stable flight of an aircraft requires close cooperation of various components. As an indispensable component of an aircraft, the wing can generate enough lift to ensure the stable flight of the aircraft. However, under the same atmospheric conditions, the flight states of wings of different weights are different. The weight distribution of the wing has a great influence on the load and flight state of the aircraft, and its accuracy determines the effectiveness and rationality of the aircraft design. Therefore, the calculation of the weight of the aircraft wing is an indispensable and important link in the design process of the aircraft. If the weight ratio of the designed wing is not appropriate, the stable flight of the aircraft cannot be guaranteed. Therefore, when the aircraft is initially designed, it is necessary to focus on the problem of the weight ratio of the aircraft wing. When designing the wing of an aircraft, there will be a series of wing design parameters. For example, the initial structure, appearance shape, size, selected materials, position relative to the fuselage of the aircraft, etc. of the wing design. The parameters can directly affect the flight state of the designed aircraft. Therefore, when calculating the wing weight, it is necessary to obtain the design value of the parameter by the acquisition module 11. It is understandable that the design parameters of the wing design described here, such as the initial structure, appearance shape, size, selected materials, and position relative to the aircraft fuselage, obviously do not constitute a limitation on the specific protection scope of this application.

[0103] The calculation module 12 is used to calculate the weight of each structural unit of the aircraft wing corresponding to the design parameters based on the aircraft wing structural unit weight calculation function; wherein the calculation module is also used to calculate the total weight of the aircraft wing based on the weight of each structural unit of the aircraft wing.

[0104] It is understandable that, by combining the design structure, design parameters, aerodynamic load and other factors affecting the wing weight of the aircraft wing, the weight of the aircraft wing can be calculated more accurately while avoiding the complex finite element analysis process. Among them, the aerodynamic load is obtained by performing aerodynamic simulation calculations on the designed wing and the aircraft. The design parameters can be obtained by obtaining at least one design parameter of the density of the wing design material, the length of the wing design root chord, the length of the wing design tip chord, the half span of the wing design, the leading edge sweep angle of the wing design, and the trailing edge sweep angle of the wing design in the initial design scheme. The design structure can be obtained by establishing a three-dimensional model of the wing. However, during the initial design, since there is no specific design structure determined, there are only simple main structures in the three-dimensional model of the wing. For example. A skin structure for transmitting aerodynamic force and aerodynamic loads, or a rib structure for maintaining the basic shape of the wing airfoil and transmitting the local aerodynamic loads on the skin to the wing beam, or a spar structure for maintaining the basic shape of the wing span. It can be seen from this that the wing is not a simple regular mechanical structure. Therefore, it is necessary to divide the wing structure into several structural units and then calculate them through corresponding calculation functions. That is, the wing structure is discretized into wing sections, and the weight of each wing section is calculated respectively by the calculation module 12. In this way, the calculation module 12 calculates the total weight of the wing after calculating the weight of each wing section, which can make the wing weight more accurate. At this time, the calculation module 12 can determine the number of wing sections and the distance between each section according to parameters such as the efficiency factor and the ultimate compressive strength of the design material. While the wing is discretized, the main structure of the wing is discretized into structural sections of corresponding number and size. It can be understood that the specific number and specific size of the wing sections after the wing discretization described here obviously do not constitute a limitation to the specific protection scope of this application.

[0105] It is understandable that the calculation module 12 can calculate the total weight of the aircraft wing through the weight of the wing structure unit. In this way, when calculating the total weight of the wing, the wing design parameters, structural dimensions, and structural weight can be fully combined. In addition, after the aircraft wing is divided into wing structure units, the actual load conditions and strength requirements of the wing are fully combined in the calculation process of the structural unit weight. In this way, when the weight calculation of the aircraft wing structure unit is completed, the calculation module 12 can more accurately calculate the total weight of the aircraft wing according to the weight of each structural unit of the aircraft wing. However, only the main structural parts such as the wing spars, skins, and wing ribs of the wing cannot be assembled into a complete wing. In addition to the main structural parts, the composition of the wing also includes some secondary structures. That is, aircraft wing accessories. For example, secondary structures such as connectors, rivets, local opening reinforcement components, leading and trailing edge structures, and glue paint between the wing structures. The weight of the auxiliary parts generally accounts for about 5% to 10% of the total weight of the wing structure, or about 5.26% to 11% of the weight of the main structure. Therefore, when the calculation module 12 calculates the total weight of the aircraft wing, the weight of the auxiliary parts cannot be ignored. It can be understood that the aircraft wing auxiliary parts described here are all parts used to assist the main structure of the wing to be assembled into a complete wing, which can be embodied in different forms. The aircraft wing auxiliary parts should not be understood as only including connectors, rivets, local opening reinforcement parts, leading and trailing edge structures, glue and paint and other parts between the wing structures. The described embodiments are only part of the embodiments of this application, not all of the embodiments. Based on the embodiments in this application, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of this application.

[0106] The output module 13 is used to output the total weight of the aircraft wing.

[0107] It is understandable that by dividing the aircraft wing into different wing structural units to calculate the total weight of the wing, a more accurate wing weight value can be obtained. When the total weight of the aircraft wing is calculated according to the total weight calculation function of the aircraft wing, the output module 13 will output the corresponding calculation result. The wing weight calculation result is output by the output module 13, so that the designer can accurately know the weight of the wing under the current design scheme. In this way, through the comprehensive analysis of the wing weight corresponding to the design scheme and the weight of the remaining aircraft components, the designer can be assisted in judging whether the aircraft design corresponding to the current design scheme is suitable. After judgment, if the design is suitable, the next step of design can be carried out. If the design is not suitable, the design scheme can be modified in time.

[0108] Furthermore, in another preferred embodiment provided in the present application, the calculation module 12 is also used to calculate the spanwise lift coefficient distribution of the aircraft wing and the aircraft lift coefficient under simulated flight conditions through an aerodynamic simulation model.

[0109] It is understandable that the wing can provide stable lift for the flight of the aircraft. The calculation formula for the lift generated by the wing is: lift = 1 / 2 * air density * speed squared * wing area * wing lift coefficient. It can be seen that the lift generated by the wing is related to the wing lift coefficient in addition to the air density, flight speed and wing area. Among them, the wing area is determined by the external dimensions of the wing design. The wing lift coefficient is related to the shape of the wing cross section (airfoil), the angle between the airflow and the wing (angle of attack), etc. Under the same flight conditions, wings of different design schemes have different wing areas and airfoils. Therefore, the lift that can be generated by wings of different design schemes is different. In order to accurately calculate the weight of the wing, it is necessary to fully consider the impact of the wing design scheme on the flight state of the aircraft during actual flight. In order to obtain the impact of the wing on the flight state of the aircraft in the corresponding design scheme, it is necessary to simulate the actual flight state of the corresponding aircraft, that is, to perform aerodynamic simulation. In this way, the lift coefficient of the aircraft wing along the span direction and the lift coefficient of the entire aircraft under simulated flight conditions can be calculated by the calculation module 12.

[0110] Specifically, the aerodynamic load distribution function of the aircraft wing at any position along the wing span can be calculated by the following formula: ; In the formula, is the overall lift coefficient of the aircraft, which can be obtained from the aerodynamic analysis process; is any spanwise position of the wing relative to the wing root; is the chord length at any spanwise position; is the geometric mean chord length of the wing; is the spanwise circulation distribution function. They can be obtained by the following formulas: ; ; ; In the formula, The spanwise lift coefficient distribution can be obtained by building a three-dimensional wing model and simulating it using computational fluid dynamics software under actual working conditions. is the root chord length of the wing; is the effective area of ​​the wing, which can be expressed as: . In this way, by performing aerodynamic simulation on the established three-dimensional model, the spanwise coefficient distribution of the wing under actual flight conditions can be obtained, thereby providing an accurate data basis for the calculation of the wing weight. It can be understood that the formulas described here are all calculation formulas for obtaining the spanwise lift coefficient distribution of the wing under actual conditions, and different methods can be used. The described embodiments are only part of the embodiments of the present application, not all of the embodiments. Based on the embodiments in the present application, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of this application.

[0111] It should be noted that the terms "include", "comprises" or any other variations thereof are intended to cover non-exclusive inclusion, so that a process, method, commodity or device including a series of elements includes not only those elements, but also other elements not explicitly listed, or also includes elements inherent to such process, method, commodity or device. In the absence of further restrictions, an element defined by the phrase "comprises a ..." does not exclude the existence of other identical elements in the process, method, commodity or device including the element.

[0112] The above is only an embodiment of the present application and is not intended to limit the present application. For those skilled in the art, the present application may have various changes and variations. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present application should be included in the scope of the claims of the present application.

Claims

1. A method for calculating the weight of an aircraft wing, characterized in that, it includes: Obtain the design parameters of the aircraft wing; Based on the aircraft wing structural unit weight calculation function, calculate the weights of the respective structural units of the aircraft wing corresponding to the design parameters; The expression of the aircraft wing structural unit weight calculation function is as follows: ， in, is the shear weight of the spar structural unit, It is the bending weight of the wing spar structural unit; ， ， In the formula, is the allowable shear stress of the material, is the spar material density, F s (y) is the shear force at any section of the wing along the span direction, is the equivalent cross-sectional area of ​​the spar; ， In the formula, is the material weight density of the skin, and are the numerical integration lengths of the upper and lower curves of the airfoil, and are the average upper and lower skin thicknesses within the wing segment respectively, z is the structural span of the wing skin surface; is the structural span of the wing skin surface; ， In the formula, is the empirical correction factor, is an empirical constant, is the statistical reference thickness of the rib, is the material density of the rib, is the cross-sectional reference area of ​​the i-th rib; Calculate the total weight of the aircraft wing according to the weights of the respective structural units of the aircraft wing; Before calculating the total weight of the aircraft wing according to the weights of the respective structural units of the aircraft wing, it further includes confirming the weight of the aircraft wing accessories; Specifically, calculating the total weight of the aircraft wing according to the weights of the respective structural units of the aircraft wing includes: Calculate the total weight of the aircraft wing structural units according to the weights of the respective structural units of the aircraft wing; Calculate the total weight of the aircraft wing according to the total weight of the aircraft wing structural units and the weight of the accessories; Among them, the calculation function of the total weight of the aircraft wing is expressed as follows: ， In the formula, is the total weight of the aircraft wing, is the weight of the aircraft wing structure unit, is the weight of the aircraft wing accessories; Output the total weight of the aircraft wing.

2. The method for calculating the weight of an aircraft wing according to claim 1, characterized in that, The aircraft wing design parameters at least include at least one design parameter among the density of the wing design material, the length of the wing design root chord, the length of the wing design tip chord, the semi-span of the wing design, the leading edge sweep angle of the wing design, and the trailing edge sweep angle of the wing design.

3. The method for calculating the weight of an aircraft wing according to claim 1, characterized in that, Before calculating the weights of the respective structural units of the aircraft wing corresponding to the design parameters based on the aircraft wing structural unit weight calculation function, it further includes: Calculate the spanwise lift coefficient distribution of the aircraft wing and the aircraft lift coefficient under the simulated flight conditions through an aerodynamic simulation model.

4. An aircraft wing weight calculation device, characterized in that, it includes: An acquisition module for acquiring the design parameters of the aircraft wing; A calculation module for calculating the weights of the respective structural units of the aircraft wing corresponding to the design parameters based on the aircraft wing structural unit weight calculation function; wherein, the calculation module is further used to calculate the total weight of the aircraft wing according to the weights of the respective structural units of the aircraft wing; An output module for outputting the total weight of the aircraft wing, The expression of the aircraft wing structural unit weight calculation function is as follows: ， in, is the shear weight of the spar structural unit, It is the bending weight of the wing spar structural unit; ， ， In the formula, is the allowable shear stress of the material, is the spar material density, F s (y) is the shear force at any section of the wing along the span direction, is the equivalent cross-sectional area of ​​the spar; ， In the formula, is the material weight density of the skin, and is the numerical integration length of the upper and lower curves of the airfoil, and are the average upper and lower skin thicknesses in the wing section, z is the structural span of the wing skin surface; is the structural span of the wing skin surface; ， In the formula, is the empirical correction factor, is an empirical constant, is the statistical reference thickness of the rib, is the material density of the rib, is the cross-sectional reference area of ​​the i-th rib; Calculate the total weight of the aircraft wing according to the weights of the respective structural units of the aircraft wing; Before calculating the total weight of the aircraft wing according to the weights of the respective structural units of the aircraft wing, it further includes confirming the weight of the aircraft wing accessories; Specifically, calculating the total weight of the aircraft wing according to the weights of the respective structural units of the aircraft wing includes: Calculate the total weight of the aircraft wing structural units according to the weights of the respective structural units of the aircraft wing; Calculate the total weight of the aircraft wing according to the total weight of the aircraft wing structural units and the weight of the accessories; Among them, the calculation function of the total weight of the aircraft wing is expressed as follows: ， In the formula, is the total weight of the aircraft wing, is the weight of the aircraft wing structure unit, is the weight of the aircraft wing accessories.

5. The aircraft wing weight calculation device according to claim 4, characterized in that, The calculation module is also used to calculate the spanwise lift coefficient distribution of the aircraft wing and the aircraft lift coefficient under simulated flight conditions through an aerodynamic simulation model.

Images