Composite truss unmanned aerial vehicle fuselage structure design method, electronic equipment and medium

Abstract

The invention discloses a composite truss unmanned aerial vehicle fuselage structure design method, electronic equipment and a medium. The method comprises the following steps: determining a bending moment distributed along a fuselage, and further determining initial geometric parameters of a truss structure; calculating inertia moment and rod piece strain according to the initial geometric parameters of the truss structure; equivalent material parameters are calculated, and then a truss structure finite element model is established; establishing a full-size fuselage truss structure model, and calculating the stability of a single pressing rod; and calculating the overall stability of the truss according to the finite element model to complete truss fuselage design. The method is used for designing the small-load and large-deformation flexible unmanned aerial vehicle fuselage structure, the unmanned aerial vehicle fuselage structure is lightest by utilizing material performance to the maximum extent, and the problem of overweight caused by the small-load and large-deformation flexible structure in an existing fuselage design method is solved on the basis that the rigidity, strength and stability of the fuselage are guaranteed.

Description

Technical Field

[0001] The present invention relates to the field of unmanned aerial vehicle (UAV) structural design, and more specifically, to a composite material truss UAV fuselage structural design method, electronic equipment and medium. Background Art

[0002] Since UAVs have extremely stringent requirements on structural weight, it is necessary to fully utilize the strength of materials. Especially when designing large-size and large-deformation structures, truss structures can save materials compared to web beams, while reducing their own weight and increasing rigidity.

[0003] The traditional UAV fuselage structure is mainly a semi-monocoque frame, beam, and skin wall panel structure. This structure is mostly suitable for large stress and small deformation structures. However, for small load and large deformation flexible structures, this traditional design method is difficult to achieve design weight closure and often exceeds the given weight by a lot.

[0004] Therefore, it is necessary to develop a composite truss UAV fuselage structure design method, electronic equipment and medium.

[0005] The information disclosed in the background technology section of the present invention is only intended to deepen the understanding of the general background technology of the present invention, and should not be regarded as acknowledging or suggesting in any form that the information constitutes the prior art already known to those skilled in the art. Summary of the invention

[0006] The present invention proposes a composite material truss UAV fuselage structure design method, electronic equipment and medium, which can be used for the design of a small-load, large-deformation flexible UAV fuselage structure, and achieves maximum utilization of material properties to make the UAV fuselage structure as light as possible. On the basis of ensuring the rigidity, strength and stability of the fuselage, the problem of overweight caused by the existing fuselage design method in the small-load, large-deformation flexible structure is solved.

[0007] In a first aspect, an embodiment of the present disclosure provides a composite material truss UAV fuselage structure design method, comprising:

[0008] Determine the bending moment distributed along the fuselage and then determine the initial geometric parameters of the truss structure;

[0009] Calculating the moment of inertia and the member strain according to the initial geometric parameters of the truss structure;

[0010] Calculate the equivalent material parameters and then establish the finite element model of the truss structure;

[0011] Establish a full-scale fuselage truss structure model and calculate the stability of a single strut;

[0012] The overall stability of the truss is calculated based on the finite element model to complete the truss fuselage design.

[0013] Preferably, the initial geometric parameters of the truss structure include truss height, truss width, outer diameter of the longitudinal rod, and inner diameter of the longitudinal rod.

[0014] Preferably, the moment of inertia is:

[0015]

[0016] Among them, I x is the x-direction moment of inertia, I y is the moment of inertia in the y direction.

[0017] Preferably, the rod is deformed as follows:

[0018]

[0019] Among them, ε zx Calculate the strain in the x direction, ε zy is the calculated strain in the y direction, and E is the equivalent elastic modulus of the composite material.

[0020] Preferably, the equivalent material parameters are:

[0021]

[0022] in, A 66 =G LT , E L 、E T is the elastic modulus in the single layer direction, G LT is the shear modulus in the single layer direction, E x 、E y is the equivalent material parameter, A is the stiffness matrix parameter, h is the rod wall thickness, μ LT , μ TL is the Poisson's ratio of a single layer in different directions.

[0023] Preferably, the beam units and shell units at the nodes of the full-size fuselage truss structure model are connected using MPC units, and the truss nodes and main load-bearing members are established through the shell units.

[0024] Preferably, the stability of a single compression rod is calculated using the Euler formula:

[0025]

[0026] Preferably, calculating the overall stability of the truss according to the finite element model and completing the truss fuselage design includes:

[0027] If the overall stability of the truss meets the design requirements, the truss fuselage design is completed. If not, the finite element model is re-established for iteration.

[0028] In a second aspect, an embodiment of the present disclosure further provides an electronic device, the electronic device comprising:

[0029] A memory storing executable instructions;

[0030] A processor runs the executable instructions in the memory to implement the composite material truss UAV fuselage structure design method.

[0031] In a third aspect, an embodiment of the present disclosure further provides a computer-readable storage medium storing a computer program, which, when executed by a processor, implements the composite material truss UAV fuselage structure design method.

[0032] Its beneficial effects are:

[0033] (1) The present invention combines engineering and finite element methods to solve the design and calculation problems of composite truss structures more efficiently and accurately.

[0034] (2) Since parameter equivalence is performed in the finite element modeling process, parametric modeling of the truss fuselage can be performed, which greatly improves the modeling efficiency and shortens the design iteration cycle.

[0035] (3) The present invention can make full use of the strength properties of the material, especially in the design of large-size and large-deformation structures, to minimize the deadweight and increase the rigidity of the fuselage.

[0036] The methods and apparatus of the present invention have other features and advantages that will be apparent from, or will be described in detail in, the accompanying drawings and subsequent detailed descriptions incorporated herein, which together serve to explain the specific principles of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0037] The above and other objects, features and advantages of the present invention will become more apparent through a more detailed description of exemplary embodiments of the present invention in conjunction with the accompanying drawings, wherein like reference numerals generally represent like components throughout the exemplary embodiments of the present invention.

[0038] Figure 1 A flow chart showing the steps of a composite material truss UAV fuselage structure design method according to an embodiment of the present invention.

[0039] Figure 2 A schematic diagram of an end surface of a truss fuselage structure according to an embodiment of the present invention is shown.

[0040] Figure 3 A schematic diagram of selecting two spans of a fuselage truss according to an embodiment of the present invention is shown.

[0041] Figure 4 A schematic diagram of a truss UAV fuselage structure finite element model according to an embodiment of the present invention is shown. DETAILED DESCRIPTION

[0042] The preferred embodiments of the present invention will be described in more detail below. Although the preferred embodiments of the present invention are described below, it should be understood that the present invention can be implemented in various forms and should not be limited to the embodiments set forth herein.

[0043] To facilitate understanding of the solutions and effects of the embodiments of the present invention, three specific application examples are given below. Those skilled in the art should understand that the examples are only for facilitating understanding of the present invention, and any specific details thereof are not intended to limit the present invention in any way.

[0044] Example 1

[0045] Figure 1 A flow chart showing the steps of a composite material truss UAV fuselage structure design method according to an embodiment of the present invention.

[0046] like Figure 1 As shown, the composite truss UAV fuselage structure design method includes:

[0047] Step 101, determining the bending moment distributed along the fuselage, and then determining the initial geometric parameters of the truss structure;

[0048] Step 102, calculating the moment of inertia and the member strain according to the initial geometric parameters of the truss structure;

[0049] Step 103, calculating equivalent material parameters, and then establishing a finite element model of the truss structure;

[0050] Step 104, establishing a full-scale fuselage truss structure model and calculating the stability of a single strut;

[0051] Step 105, calculating the overall stability of the truss according to the finite element model, and completing the truss fuselage design.

[0052] In one example, the initial geometric parameters of the truss structure include truss height, truss width, longitudinal member outer diameter, and longitudinal member inner diameter.

[0053] In one example, the moment of inertia is:

[0054]

[0055] Among them, I x is the x-direction moment of inertia, I y is the moment of inertia in the y direction.

[0056] In one example, the member strains become:

[0057]

[0058] Among them, ε zx Calculate the strain in the x direction, ε zy is the calculated strain in the y direction, and E is the equivalent elastic modulus of the composite material.

[0059] In one example, the equivalent material parameters are:

[0060]

[0061] in, A 66 =G LT , E L 、E T is the elastic modulus in the single layer direction, G LT is the shear modulus in the single layer direction, E x 、E y is the equivalent material parameter, A is the stiffness matrix parameter, h is the rod wall thickness, μ LT , μ TL is the Poisson’s ratio of a single layer in different directions.

[0062] In one example, the beam elements and shell elements at the nodes of the full-scale fuselage truss structure model are connected using MPC elements, and the truss nodes and main load-bearing members are established through the shell elements.

[0063] In an example, the stability of a single strut is calculated using Euler's formula:

[0064]

[0065] In one example, the overall stability of the truss is calculated based on the finite element model. The truss fuselage design includes:

[0066] If the overall stability of the truss meets the design requirements, the truss fuselage design is completed. If not, the finite element model is re-established for iteration.

[0067] Specifically, according to the load (horizontal tail and vertical tail load) and stiffness design requirements, the bending moment M distributed along the fuselage is calculated: x 、M y .

[0068] Figure 2 A schematic diagram of an end surface of a truss fuselage structure according to an embodiment of the present invention is shown.

[0069] According to the bending moment distribution of the fuselage, the initial geometric parameters of the truss structure are determined by engineering calculation methods, including the truss height H, truss width L, longitudinal bar outer diameter D, and longitudinal bar inner diameter d, such as Figure 2shown.

[0070] Calculate the moment of inertia in different directions as:

[0071]

[0072] Among them, I x is the x-direction moment of inertia, I y is the moment of inertia in the y direction.

[0073] The calculated member strain is:

[0074]

[0075] Among them, ε zx Calculate the strain in the x direction, ε zy is the calculated strain in the y direction, and E is the equivalent elastic modulus of the composite material. Since the allowable strain values ​​used are equal,

[0076] ε zy =ε zx =[ε]

[0077] Wherein, [ε] is the allowable strain value of the member, which is obtained based on the test of different material systems or the strength design requirements of engineering projects.

[0078] Truss nodes are generally nodes with a certain degree of rigidity. Due to the influence of node rigidity, rod bending and axial additional force occur. The additional force needs to take into account the axial deformation of the rod. The finite element model of the truss structure is established based on the preliminary obtained rod geometric dimensions and layup.

[0079] Composite rods are simulated using beam elements (bar elements), and equivalent material parameters are calculated based on material and ply information. The equivalent material parameters are:

[0080]

[0081] in, A 66 =G LT , E L 、E T is the elastic modulus in the single layer direction, G LT is the shear modulus in the single layer direction, E x 、E y is the equivalent material parameter, A is the stiffness matrix parameter, h is the rod wall thickness, μ LT , μ TL is the Poisson’s ratio of a single layer in different directions.

[0082] The equivalent shear modulus is

[0083] G xy =A 66 / h

[0084] The equivalent Poisson's ratio is

[0085]

[0086] Figure 3 A schematic diagram of selecting two spans of a fuselage truss according to an embodiment of the present invention is shown.

[0087] A full-scale fuselage truss structure model was established, and the static and stability analysis of the model was performed. The beam element (bar element) and the shell element (shell element) at the truss node were connected using MPC (RBE2) elements. Two typical truss structures were selected, such as Figure 3 As shown, shell elements are used to establish truss nodes and main load-bearing members.

[0088] Extract the beam element (bar element) force (FORCE) and calculate the stress and strain of the bar:

[0089]

[0090] The bar strain and critical pressure are calculated by beam unit force, composite material equivalent elastic modulus and effective length of the compression bar. The effective length of the compression bar can be given by experiment. The stability of a single compression bar is calculated using the Euler formula:

[0091]

[0092] Figure 4 A schematic diagram of a truss UAV fuselage structure finite element model according to an embodiment of the present invention is shown.

[0093] like Figure 4 As shown in the figure, the overall stability of the truss is obtained through the overall finite element model calculation to determine whether all calculation results meet the design requirements. If so, the truss fuselage design is completed. If not, the ply and geometric parameters are adjusted in a targeted manner according to the calculation results, and the finite element model is re-established for iteration to finally complete the design convergence.

[0094] The fuselage design of the UAV using this method has shown through ground tests that the weight reduction effect is obvious compared with the traditional design method, and the design iteration speed is increased by more than 30%.

[0095] Example 2

[0096] The present disclosure provides an electronic device, which includes: a memory storing executable instructions; and a processor running the executable instructions in the memory to implement the above-mentioned composite material truss UAV fuselage structure design method.

[0097] An electronic device according to an embodiment of the present disclosure includes a memory and a processor.

[0098] The memory is used to store non-temporary computer-readable instructions. Specifically, the memory may include one or more computer program products, which may include various forms of computer-readable storage media, such as volatile memory and / or non-volatile memory. The volatile memory may, for example, include random access memory (RAM) and / or cache memory (cache), etc. The non-volatile memory may, for example, include read-only memory (ROM), hard disk, flash memory, etc.

[0099] The processor may be a central processing unit (CPU) or other forms of processing units having data processing capabilities and / or instruction execution capabilities, and may control other components in the electronic device to perform desired functions. In one embodiment of the present disclosure, the processor is used to run the computer-readable instructions stored in the memory.

[0100] Those skilled in the art should be able to understand that in order to solve the technical problem of how to obtain a good user experience, the present embodiment may also include well-known structures such as a communication bus and an interface, and these well-known structures should also be included in the protection scope of the present disclosure.

[0101] For detailed description of this embodiment, reference may be made to the corresponding descriptions in the aforementioned embodiments, which will not be repeated here.

[0102] Example 3

[0103] An embodiment of the present disclosure provides a computer-readable storage medium storing a computer program, which implements the composite material truss UAV fuselage structure design method when executed by a processor.

[0104] According to the computer-readable storage medium of the embodiment of the present disclosure, non-transitory computer-readable instructions are stored thereon. When the non-transitory computer-readable instructions are executed by a processor, all or part of the steps of the above-mentioned methods of each embodiment of the present disclosure are executed.

[0105] The above-mentioned computer-readable storage media include, but are not limited to: optical storage media (e.g., CD-ROM and DVD), magneto-optical storage media (e.g., MO), magnetic storage media (e.g., magnetic tape or mobile hard disk), media with built-in rewritable non-volatile memory (e.g., memory card) and media with built-in ROM (e.g., ROM box).

[0106] Those skilled in the art should understand that the purpose of the above description of the embodiments of the present invention is only to exemplarily illustrate the beneficial effects of the embodiments of the present invention, and is not intended to limit the embodiments of the present invention to any given examples.

[0107] The embodiments of the present invention have been described above, and the above description is exemplary, not exhaustive, and is not limited to the disclosed embodiments. Many modifications and changes will be apparent to those skilled in the art without departing from the scope and spirit of the described embodiments.

Claims

1. A composite material truss UAV fuselage structure design method, characterized in that: include: Determine the bending moment distributed along the fuselage and then determine the initial geometric parameters of the truss structure; Calculating the moment of inertia and the member strain according to the initial geometric parameters of the truss structure; Calculate the equivalent material parameters and then establish the finite element model of the truss structure; Establish a full-scale fuselage truss structure model and calculate the stability of a single strut; The overall stability of the truss is calculated based on the finite element model to complete the truss fuselage design.

2. The composite material truss UAV fuselage structure design method according to claim 1, wherein: The initial geometric parameters of the truss structure include truss height, truss width, longitudinal rod outer diameter, and longitudinal rod inner diameter.

3. The composite material truss UAV fuselage structure design method according to claim 1, wherein: The moment of inertia is: Among them, I x is the x-direction moment of inertia, I y is the moment of inertia in the y direction.

4. The composite material truss UAV fuselage structure design method according to claim 1, wherein: The member strain is: Among them, ε zx Calculate the strain in the x direction, ε zy is the calculated strain in the y direction, and E is the equivalent elastic modulus of the composite material.

5. The composite material truss UAV fuselage structure design method according to claim 1, wherein: The equivalent material parameters are: in, A 66 =G LT , E L 、E T is the elastic modulus in the single layer direction, G LT is the shear modulus in the single layer direction, E x 、E y is the equivalent material parameter, A is the stiffness matrix parameter, h is the rod wall thickness, μ LT , μ TL is the Poisson’s ratio of a single layer in different directions.

6. The composite material truss UAV fuselage structure design method according to claim 1, wherein: The beam units and shell units at the nodes of the full-size fuselage truss structure model are connected by MPC units, and the truss nodes and main load-bearing members are established through the shell units.

7. The composite material truss UAV fuselage structure design method according to claim 1, wherein: Euler's formula is used to calculate the stability of a single compression rod:

8. The composite material truss UAV fuselage structure design method according to claim 1, wherein: According to the finite element model, the overall stability of the truss is calculated, and the truss fuselage design is completed, including: If the overall stability of the truss meets the design requirements, the truss fuselage design is completed. If not, the finite element model is re-established for iteration.

9. An electronic device, characterized in that: The electronic device comprises: A memory storing executable instructions; A processor, wherein the processor runs the executable instructions in the memory to implement the composite material truss UAV fuselage structure design method according to any one of claims 1 to 8.

10. A computer-readable storage medium, characterized in that: The computer-readable storage medium stores a computer program, and when the computer program is executed by a processor, the composite material truss UAV fuselage structure design method according to any one of claims 1 to 8 is implemented.

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