CN117828752A - Aeroelasticity-based aircraft structure optimization method - Google Patents

Aeroelasticity-based aircraft structure optimization method Download PDF

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Publication number
CN117828752A
CN117828752A CN202311764630.1A CN202311764630A CN117828752A CN 117828752 A CN117828752 A CN 117828752A CN 202311764630 A CN202311764630 A CN 202311764630A CN 117828752 A CN117828752 A CN 117828752A
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control surface
optimization
finite element
aircraft
pneumatic
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田海涛
苑凯华
程萌
付志超
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Beijing Electromechanical Engineering Research Institute
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Beijing Electromechanical Engineering Research Institute
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    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/10—Geometric CAD
    • G06F30/15—Vehicle, aircraft or watercraft design
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/20—Design optimisation, verification or simulation
    • G06F30/23—Design optimisation, verification or simulation using finite element methods [FEM] or finite difference methods [FDM]
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F2119/00—Details relating to the type or aim of the analysis or the optimisation
    • G06F2119/18—Manufacturability analysis or optimisation for manufacturability
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation

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Abstract

The invention provides an aircraft structure optimization method based on aeroelasticity, which comprises the following steps: s1, building a control surface structure finite element model according to a control surface model of an aircraft; s2, establishing a pneumatic surface model of the control surface structure according to key nodes of the finite element model of the control surface structure; constructing an interpolation node set according to the finite element model and the pneumatic surface model of the control surface structure; s3, setting upper and lower limits of optimization variables including control surface ribs, skin size variables and angles of the control surface ribs according to static pneumatic elasticity characteristics of the control surface, and setting an optimization target value of pneumatic elasticity influence quantity and constraint variables in an optimization analysis process; setting optimization iteration parameters; and S4, performing static and pneumatic elasticity calculation and analysis, and adjusting the finite element model and interpolation nodes of the control surface structure according to the result of the optimized variable to finish the structural optimization of the aircraft. By applying the technical scheme of the invention, the technical problem that the structural optimization of the aircraft in the prior art cannot synchronously meet the pneumatic and structural requirements can be solved.

Description

Aeroelasticity-based aircraft structure optimization method
Technical Field
The invention relates to the technical field of aeroelasticity of aircrafts, in particular to an aircraft structure optimization method based on aeroelasticity.
Background
The wide Mach number variation range of the flight airspace of the high-speed aircraft is large, the flight environment is complex, and the analysis subject involved is numerous, so that the aircraft design is a typical multidisciplinary optimization problem. Unlike previous high speed aircraft, first, the aircraft needs to meet the requirements of low speed, high lift force, high speed and low resistance, aerodynamic layout is generally designed into an elongated body, a lift body layout and a full or partial wave-taking body layout, and such layout results in weak structural bearing characteristics; secondly, the high-speed aircraft has strict constraint on weight due to the use requirements of range and speed, and the design is widely implemented by adopting light materials and thin skin structure design to reduce the structural quality, so that the aircraft has low structural rigidity and low flutter boundary, and the allowance for adjusting the structural frequency by adopting additional quality is small; finally, the aircraft needs to solve the problem of control accuracy in the turning process, and is extremely sensitive to the change of aerodynamic efficiency after deformation in the flying process. Therefore, an optimal design method based on static pneumatic elasticity needs to be developed, a pneumatic structure multi-field coupling optimization flow is formed, pneumatic/structure cooperative optimization design based on pneumatic elasticity is realized, and a foundation is laid for further realizing fine design and comprehensive performance improvement of an aircraft.
Disclosure of Invention
The present invention aims to solve at least one of the technical problems existing in the prior art.
The invention provides an aircraft structure optimization method based on pneumatic elasticity, which comprises the following steps:
s1, building a control surface structure finite element model according to a control surface model of an aircraft;
s2, establishing a pneumatic surface model of the control surface structure according to key nodes of the finite element model of the control surface structure; constructing an interpolation node set according to the finite element model and the pneumatic surface model of the control surface structure;
s3, setting upper and lower limits of optimization variables including control surface ribs, skin size variables and angles of the control surface ribs according to static pneumatic elasticity characteristics of the control surface, and setting an optimization target value of pneumatic elasticity influence quantity and constraint variables in an optimization analysis process; setting optimization iteration parameters;
and S4, performing static and pneumatic elasticity calculation and analysis, and adjusting the finite element model and interpolation nodes of the control surface structure according to the result of the optimized variable to finish the structural optimization of the aircraft.
Further, in S1, according to the aircraft control surface model, a control surface structure finite element model is built in Hypermesh, static force solving parameters are set, and static force calculation solving is performed.
Further, the control surface structure finite element model realizes grid division and unit attribute assignment of the control surface finite element model through a TCL/TK secondary development program in the Hypermesh.
Further, in S2, the key nodes include root leading edge points, root trailing edge points, tip leading edge points, and tip trailing edge points.
Further, in S3, an optimization method written by Matlab is adopted to analyze and generate the optimization variable analysis.
Further, the optimization method adopts a genetic algorithm.
Further, the constraint variables in the optimization analysis process comprise the quality of the control surface model.
Further, in S4, the static aeroelastic analysis method selects a face method analysis.
Further, in the process of optimizing analysis, the aeroelasticity influence quantity is extracted and analyzed; updating the control surface finite element grid and the unit attribute value according to the optimized variable, controlling the generation of Hypermesh by using a TCL/TK secondary development program to realize the optimized design of the structural topological structure and the dimensional parameter.
By applying the technical scheme of the invention, the aircraft structure optimization method based on the aeroelasticity is provided, the aircraft structure optimization is carried out based on the aeroelasticity, a control surface structure finite element model and a pneumatic surface model are established, and the static aeroelasticity optimization analysis is completed by setting optimization parameters according to the static aeroelasticity characteristics of the control surface. The invention improves the approach of the pneumatic elasticity problem under the condition of not changing the pneumatic appearance, realizes the construction and the programming of the structural optimization flow of the component based on the pneumatic elasticity, expands the application range of the structural optimization of the aircraft and is convenient for engineering application. Compared with the prior art, the technical scheme of the invention can solve the technical problem that the structural optimization of the aircraft in the prior art cannot synchronously meet the pneumatic and structural requirements.
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The accompanying drawings, which are included to provide a further understanding of embodiments of the invention and are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and together with the description serve to explain the principles of the invention. It is evident that the drawings in the following description are only some embodiments of the present invention and that other drawings may be obtained from these drawings without inventive effort for a person of ordinary skill in the art.
Fig. 1 shows a flow diagram of an aircraft structure optimization method based on aeroelasticity, according to a specific embodiment of the invention.
Detailed Description
It should be noted that, in the case of no conflict, the embodiments and features in the embodiments may be combined with each other. The following description of the embodiments of the present invention will be made clearly and completely with reference to the accompanying drawings, in which it is apparent that the embodiments described are only some embodiments of the present invention, but not all embodiments. The following description of at least one exemplary embodiment is merely exemplary in nature and is in no way intended to limit the invention, its application, or uses. All other embodiments, which can be made by those skilled in the art based on the embodiments of the invention without making any inventive effort, are intended to be within the scope of the invention.
It is noted that the terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of example embodiments in accordance with the present application. As used herein, the singular is also intended to include the plural unless the context clearly indicates otherwise, and furthermore, it is to be understood that the terms "comprises" and/or "comprising" when used in this specification are taken to specify the presence of stated features, steps, operations, devices, components, and/or combinations thereof.
The relative arrangement of the components and steps, numerical expressions and numerical values set forth in these embodiments do not limit the scope of the present invention unless it is specifically stated otherwise. Meanwhile, it should be understood that the sizes of the respective parts shown in the drawings are not drawn in actual scale for convenience of description. Techniques, methods, and apparatus known to one of ordinary skill in the relevant art may not be discussed in detail, but are intended to be part of the specification where appropriate. In all examples shown and discussed herein, any specific values should be construed as merely illustrative, and not a limitation. Thus, other examples of the exemplary embodiments may have different values. It should be noted that: like reference numerals and letters denote like items in the following figures, and thus once an item is defined in one figure, no further discussion thereof is necessary in subsequent figures.
As shown in fig. 1, according to an embodiment of the present invention, there is provided a method for optimizing an aircraft structure based on aeroelasticity, the method comprising:
s1, building a control surface structure finite element model according to a control surface model of an aircraft;
s2, establishing a pneumatic surface model of the control surface structure according to key nodes of the finite element model of the control surface structure; constructing an interpolation node set according to the finite element model and the pneumatic surface model of the control surface structure;
s3, setting upper and lower limits of optimization variables including control surface ribs, skin size variables and angles of the control surface ribs according to static pneumatic elasticity characteristics of the control surface, and setting an optimization target value of pneumatic elasticity influence quantity and constraint variables in an optimization analysis process; setting optimization iteration parameters;
and S4, performing static and pneumatic elasticity calculation and analysis, and adjusting the finite element model and interpolation nodes of the control surface structure according to the result of the optimized variable to finish the structural optimization of the aircraft.
By applying the configuration mode, the aircraft structure optimization method based on the aeroelasticity is provided, the aircraft structure optimization is carried out based on the aeroelasticity, a control surface structure finite element model and a pneumatic surface model are established, and the static aeroelasticity optimization analysis is completed by setting optimization parameters according to the static aeroelasticity characteristics of the control surface. The invention improves the approach of the pneumatic elasticity problem under the condition of not changing the pneumatic appearance, realizes the construction and the programming of the structural optimization flow of the component based on the pneumatic elasticity, expands the application range of the structural optimization of the aircraft and is convenient for engineering application.
Further, in the present invention, S1 is first executed, and a control surface structure finite element model is built according to the aircraft control surface model.
As a specific embodiment of the invention, a control surface structure finite element model can be established in hypersash according to an aircraft control surface model, static force solving parameters are set, and static force calculation solving is carried out.
The control surface structure finite element model can realize grid division and unit attribute assignment of the control surface finite element model through a TCL/TK secondary development program in Hypermesh.
Further, in the invention, after the control surface structure finite element model is constructed, S2 is executed, and a pneumatic surface model of the control surface structure is built according to key nodes of the control surface structure finite element model; and constructing an interpolation node set according to the finite element model and the pneumatic surface model of the control surface structure. The interpolation node set changes along with the change of the control surface structure in the optimization process and is continuously updated according to the finite element model.
As one embodiment of the invention, the key nodes include root leading edge points, root trailing edge points, tip leading edge points, and tip trailing edge points.
Further, in the invention, after the interpolation node set is constructed, S3 is executed, the upper limit and the lower limit of the optimization variables including the control surface rib, the skin size variable and the angle of the control surface rib are set according to the static pneumatic elasticity characteristic of the control surface, and the optimization target value of the pneumatic elastic influence and the constraint variable in the optimization analysis process are set; setting optimization iteration parameters.
As a specific embodiment of the present invention, an optimization method written by Matlab may be used to analyze and generate the optimization variable analysis, wherein the optimization algorithm uses a genetic algorithm. Constraint variables in the optimization analysis process comprise the quality of the control surface model. The angle of the control surface rib refers to the included angle between the control surface inner rib and the bottom.
Further, in the invention, after the setting of the optimization parameters is completed, S4 is executed, static pneumatic elasticity calculation and analysis are carried out, and the control surface structure finite element model and interpolation nodes are adjusted according to the result of the optimization variables, so that the aircraft structure optimization is completed.
As a specific embodiment of the invention, the static aeroelastic analysis method selects a face element method for analysis. In the process of optimizing analysis, extracting and analyzing the aeroelasticity influence quantity; and updating the control surface finite element grid and the unit attribute value according to the optimized variable, wherein the updating of the finite element grid controls the generation of Hypermesh through a TCL/TK secondary development program, thereby realizing the optimized design of the structural topological structure and the dimensional parameter.
For a further understanding of the invention, the method for optimizing the structure of an aircraft based on aeroelasticity according to the invention is described in detail below with reference to fig. 1.
As shown in fig. 1, according to an embodiment of the present invention, there is provided a method for optimizing an aircraft structure based on aeroelasticity, the method specifically comprising the following steps.
Step one, building a control surface structure finite element model in hypersash according to an aircraft control surface model, setting static force solving parameters, and carrying out static force calculation solving. The control surface structure finite element model realizes grid division and unit attribute assignment of the control surface finite element model through a TCL/TK secondary development program in the Hypermesh.
Step two, establishing a pneumatic surface model of the control surface structure according to key nodes (root chord leading edge points and trailing edge points, tip chord leading edge points and trailing edge points) of the finite element model of the control surface structure; and establishing an interpolation node set according to the control surface structure finite element model and the pneumatic surface model.
Setting upper and lower limits of optimization variables such as control surface ribs, skin size variables, rib angles and the like according to the static pneumatic elasticity characteristics of the control surface, and setting an optimization target value of pneumatic elasticity influence quantity and a constraint variable (the quality of a control surface model) in the optimization analysis process; setting optimization iteration parameters.
Wherein, the genetic algorithm written by Matlab is used for analyzing and generating the optimized variable analysis.
Step four, after parameter setting is completed, static pneumatic elasticity analysis is carried out; and adjusting the finite element model and the interpolation node according to the optimized variable result. The static pneumatic elasticity analysis method selects a surface element method for analysis.
Wherein, the aeroelasticity influence quantity is extracted and analyzed; updating the control surface finite element grid and the unit attribute value according to the optimized variable, and controlling the generation of Hypermesh by the updating of the finite element grid through a TCL/TK secondary development program, thereby realizing the optimized design of the structural topological structure and the dimensional parameter
In summary, the invention provides an aircraft structure optimization method based on aeroelasticity, which optimizes the aircraft structure based on the aeroelasticity, establishes a control surface structure finite element model and a pneumatic surface model, sets optimization parameters according to the static aeroelasticity characteristics of the control surface, and completes static aeroelasticity optimization analysis. The invention improves the approach of the pneumatic elasticity problem under the condition of not changing the pneumatic appearance, realizes the construction and the programming of the structural optimization flow of the component based on the pneumatic elasticity, expands the application range of the structural optimization of the aircraft and is convenient for engineering application.
The above description is only of the preferred embodiments of the present invention and is not intended to limit the present invention, but various modifications and variations can be made to the present invention by those skilled in the art. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention should be included in the protection scope of the present invention.

Claims (9)

1. The aeroelastic-based aircraft structure optimization method is characterized by comprising the following steps of:
s1, building a control surface structure finite element model according to a control surface model of an aircraft;
s2, establishing a pneumatic surface model of the control surface structure according to key nodes of the finite element model of the control surface structure; constructing an interpolation node set according to the finite element model and the pneumatic surface model of the control surface structure;
s3, setting upper and lower limits of optimization variables including control surface ribs, skin size variables and angles of the control surface ribs according to static pneumatic elasticity characteristics of the control surface, and setting an optimization target value of pneumatic elasticity influence quantity and constraint variables in an optimization analysis process; setting optimization iteration parameters;
and S4, performing static pneumatic elasticity calculation, and adjusting the finite element model of the control surface structure and interpolation nodes according to the result of the optimized variable to finish the structural optimization of the aircraft.
2. The aeroelastic-based aircraft structure optimization method according to claim 1, wherein in S1, according to an aircraft control surface model, a control surface structure finite element model is built in Hypermesh, static force solving parameters are set, and static force calculation solving is performed.
3. The aircraft structure optimization method based on aeroelasticity according to claim 1 or 2, wherein the control surface structure finite element model realizes grid division and unit attribute assignment of the control surface finite element model through a TCL/TK secondary development program in HyperMesh.
4. The aeroelastic-based aircraft structure optimization method according to claim 1, wherein in S2, the key nodes include root leading edge points, root trailing edge points, tip leading edge points, and tip trailing edge points.
5. The aeroelastic-based aircraft structure optimization method according to claim 4, wherein in S3, an optimization method written by Matlab is adopted to analyze and generate an optimization variable analysis.
6. The method for optimizing an aircraft structure based on aeroelasticity according to claim 5, wherein the optimization method uses a genetic algorithm.
7. The method of claim 5, wherein the constraint variables in the optimization analysis include the mass of the control surface model.
8. The aeroelastic-based aircraft structure optimization method according to any one of claims 1 to 7, wherein in S4, the static aeroelastic analysis method selects a bin method analysis.
9. The method for optimizing the structure of an aircraft based on aeroelasticity according to claim 8, wherein the aeroelasticity influence quantity is extracted and analyzed in the optimization analysis process; updating the control surface finite element grid and the unit attribute value according to the optimized variable, controlling the generation of Hypermesh by using a TCL/TK secondary development program to realize the optimized design of the structural topological structure and the dimensional parameter.
CN202311764630.1A 2023-12-20 2023-12-20 Aeroelasticity-based aircraft structure optimization method Pending CN117828752A (en)

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