CN112380624A - Helicopter cabin skeleton rigidity optimization method - Google Patents
Helicopter cabin skeleton rigidity optimization method Download PDFInfo
- Publication number
- CN112380624A CN112380624A CN202011309583.8A CN202011309583A CN112380624A CN 112380624 A CN112380624 A CN 112380624A CN 202011309583 A CN202011309583 A CN 202011309583A CN 112380624 A CN112380624 A CN 112380624A
- Authority
- CN
- China
- Prior art keywords
- cabin
- cabin skeleton
- skeleton
- optimization
- scheme
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
Links
Images
Classifications
-
- 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]
Landscapes
- Engineering & Computer Science (AREA)
- Physics & Mathematics (AREA)
- Geometry (AREA)
- Theoretical Computer Science (AREA)
- General Physics & Mathematics (AREA)
- Computer Hardware Design (AREA)
- Evolutionary Computation (AREA)
- General Engineering & Computer Science (AREA)
- Automation & Control Theory (AREA)
- Aviation & Aerospace Engineering (AREA)
- Computational Mathematics (AREA)
- Mathematical Analysis (AREA)
- Mathematical Optimization (AREA)
- Pure & Applied Mathematics (AREA)
- Investigating Strength Of Materials By Application Of Mechanical Stress (AREA)
- Body Structure For Vehicles (AREA)
Abstract
The invention relates to the technical field of aircraft structural strength design and verification, in particular to a helicopter cockpit skeleton stiffness optimization simulation analysis method. The method comprises the following steps: s1: finite element modeling of a cabin skeleton; s2: optimizing, analyzing and modeling the rigidity of the cabin skeleton; s3: and evaluating an optimization result to obtain a cabin skeleton rigidity optimization scheme. According to the method, on the premise of not influencing the airtightness of the cabin door and the normal work of the locking mechanism, a simulation optimization analysis model is established, the maximum influence of unit weight gain on the overall rigidity is an optimization design target, and the optimal layering strengthening scheme of different areas is given.
Description
Technical Field
The invention relates to the technical field of aircraft structural strength design and verification, in particular to a helicopter cockpit skeleton stiffness optimization simulation analysis method.
Background
The cockpit is an important constituent structure positioned at the front section of the helicopter, not only provides an internal space for a pilot, but also is a mounting platform of helicopter control, avionics and other systems, and mainly comprises components such as a cockpit skeleton, a skin, a cabin door and the like. At present, the cockpit of a domestic helicopter is mainly made of light composite materials such as carbon fibers, and during structural design, the static strength of the cockpit is generally checked only, and the deformation requirement cannot be clearly given, so that the problems of poor cabin door sealing performance, difficult lock catch closing and the like caused by excessive deformation in the later use process are frequent, and the user experience and even the flight safety are seriously influenced. Therefore, the optimized design for improving the bearing capacity and rigidity of the cabin skeleton is a basic condition for ensuring the normal use of the helicopter.
In the traditional cabin skeleton simulation analysis method, a detailed modeling method is mostly adopted for simulation, the structure needs to be integrally reinforced again during analysis, and then a test is carried out for verification.
Due to the fact that the structural form of the framework and the overall layering scheme are changed, redesign and unnecessary weight increase of the composite material mold are often caused, huge test cost is generated, and the traditional method is long in analysis period and low in efficiency.
Disclosure of Invention
The purpose of the invention is as follows: the invention provides a helicopter cabin skeleton rigidity optimization method, which is characterized in that a simulation optimization analysis model is established on the premise of not influencing the airtightness of a cabin door and the normal work of a locking mechanism, the maximum influence of unit weight gain on the overall rigidity is an optimization design target, and optimal layering strengthening schemes in different areas are provided.
The technical scheme of the invention is as follows: in order to achieve the aim, the helicopter cabin skeleton rigidity optimization method is characterized by comprising the following steps of:
s1: finite element modeling of a cabin skeleton;
s2: optimizing, analyzing and modeling the rigidity of the cabin skeleton;
s3: evaluating an optimization result to obtain a cabin skeleton rigidity optimization scheme; evaluating the structural scheme and the weight cost according to the result of optimization evaluation, if the structural scheme and the weight cost are not satisfied, repeating the step S2 to perform iterative optimization,
and finally obtaining the optimal cabin skeleton rigidity optimization scheme.
In a possible embodiment, the step S1 specifically includes the following steps:
s101: simulating a cabin skeleton by adopting a shell unit;
the locking point connection between the cabin door and the cabin skeleton is simulated by a rigid body element RBE3, the most severe working condition is considered during analysis, only the main stress part is considered during modeling, the non-main stress part is removed, and the connection between the cabin door and the cabin skeleton is simulated by the rigid body element;
s102: assigning material properties;
inputting a basic layering sequence of the composite material of the cabin skeleton according to actual layering information based on a classical lamination theory, and assigning corresponding material attributes to the shell units;
s103: carrying out boundary definition;
defining the boundary comprises determining a load working condition and determining a constraint condition; the load working condition is determined by taking the working condition with the most serious load as the load working condition; determining the constraint condition means that the whole cabin framework is fixedly connected with the middle fuselage of the helicopter, and 3 degrees of freedom of a connection area of the cabin framework and the middle fuselage of the helicopter are constrained.
In a possible embodiment, the step S2 specifically includes the following steps:
s201: on the basis of finite element modeling of the cabin framework, partitioning the cabin framework;
s202: respectively calculating to obtain initial maximum deformation s for each cabin skeleton partition1Longitudinal equivalent elastic modulus and transverse equivalent elastic modulus; ex、EyIs an optimization design variable related to the number of layers of composite material layering, the laying angle, the layering sequence and the layering material;
calculating the equivalent longitudinal elastic modulus E according to the following calculation formula Ix(ii) a Calculating the transverse equivalent elastic modulus E according to the following formula IIy：
Wherein E is11Is the longitudinal elastic modulus of the composite material;
E22is the transverse elastic modulus of the composite material;
G12is the in-plane shear modulus of the composite material;
u12the cedar ratio of the composite material in XY direction;
theta is an included angle between the X direction of the reference coordinate system and the longitudinal direction of the composite material fiber;
s203: respectively implementing a strengthening scheme on each cabin skeleton partition;
s204: respectively calculating and counting the longitudinal equivalent elastic modulus and the transverse equivalent elastic modulus of each cabin skeleton partition subjected to the implementation of each strengthening scheme, and the maximum deformation;
s205: calculating the deformation difference deltas of each cabin skeleton subarea before and after implementing each strengthening scheme according to the formula III,
Δs＝s2-s1formula III
Wherein the maximum deformation after implementation of the reinforcing embodiment is s2Initial maximum deflection of s1；
Calculating the mass increment Deltam of each cabin skeleton partition before and after the strengthening scheme is implemented according to the formula four,
Δm＝m2-m1formula IV
Wherein the mass after implementation of the reinforcing scheme is m2Initial mass m1；
Defining a parameter 'deformation difference/mass increment' as a deformation mass increment ratio Y, and calculating the deformation mass increment ratio Y according to the formula V;
carrying out statistics on the deformation quality increment ratio Y obtained by calculation before and after different strengthening schemes are implemented on each cabin skeleton partition;
s206: optimization design analysis with commercial software for equivalent modulus of elasticity E in the longitudinal directionxTransverse equivalent modulus of elasticity EyAs a design variable, the deformation mass increment ratio Y is used as a function value, and a response surface is formed by fitting;
if the response surface is a plane, the rigidity characteristic is a linear problem; if the response surface is a curved surface with a single peak, the rigidity characteristic is a unimodal nonlinear problem; if a plurality of peaks exist in the response surface, the rigidity characteristic is proved to be a multi-peak nonlinear problem.
In a possible embodiment, the partitioning of the cabin skeleton in step S201 includes a front region 1, a vertical region 2, and a lateral region 3; the front area 1 is positioned at the front end of the cabin skeleton; the vertical area 2 is positioned at the lower cross beams at two sides of the cabin skeleton; the transverse areas 3 are positioned at the upper longitudinal beams on the two sides of the cabin framework.
In a possible embodiment, in step S203, the step of performing a reinforcement scheme on each cabin skeleton segment includes adding and/or changing a direction of a mat.
In a possible embodiment, the steps S203-S204 are repeated for each cabin skeleton partition, and multiple reinforcement schemes are adopted to obtain the longitudinal equivalent elastic modulus E under each reinforcement schemexTransverse equivalent modulus of elasticity EyAnd a statistical value of the deformation mass increment ratio Y.
In a possible embodiment, in step S3, according to the optimal design objective with the largest influence of the unit weight gain on the overall stiffness, and in combination with the response surface obtained by fitting in step S205, the optimal reinforcement scheme for each cabin framework partition is determined, so as to form a cabin framework stiffness optimization scheme; the smaller the amount of change in the amount of deformation at the same weight gain, the greater the stiffness per unit mass increase, and the greater the structural efficiency.
In one possible embodiment, simulation optimization analysis is performed using Hypermesh software.
The invention has the beneficial effects that: the invention provides a helicopter cockpit skeleton rigidity optimization method, which adopts a structure optimization and parameter optimization method under the condition of meeting the strength design requirement, obtains the influence rule of unit weight gain of different subareas on the overall rigidity, effectively reduces unnecessary weight gain, avoids large-scale detail calculation, and effectively improves the design efficiency.
Drawings
FIG. 1 is a flow chart of the method of the present invention
Figure 2 is a schematic view of the skeleton structure of the cockpit of the present invention
FIG. 3 is a schematic view of a cabin skeleton partition according to an embodiment of the present invention
FIG. 4 is a graph showing the response of 1 unit weight gain to overall stiffness in the frontal region of the cabin frame according to an embodiment of the present invention
FIG. 5 is a graph showing the effect of 2 unit weight gain on overall stiffness in a vertical region of a cabin frame according to an embodiment of the present invention
FIG. 6 is a response graph of the effect of 3 unit weight gain on overall stiffness in the transverse region of the cabin frame according to the embodiment of the invention
FIG. 7 is a response graph of the effect of unit weight gain of the cabin skeleton on the overall stiffness according to the embodiment of the invention
Wherein:
1-a front region; 2-vertical region; 3-transverse region
Detailed Description
In order to make the objects, technical solutions and advantages of the present invention more apparent, the present invention is described in further detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are merely illustrative of the invention and are not intended to limit the invention.
As shown in fig. 1, a helicopter cabin skeleton stiffness optimization method is characterized by comprising the following steps:
s1: finite element modeling of a cabin skeleton;
s2: optimizing, analyzing and modeling the rigidity of the cabin skeleton;
s3: evaluating an optimization result to obtain a cabin skeleton rigidity optimization scheme; and (4) evaluating the structural scheme and the weight cost according to the result of optimization evaluation, and if the structural scheme and the weight cost are not satisfactory, repeating the step S2 to perform iterative optimization to finally obtain the optimal cabin skeleton stiffness optimization scheme.
In a possible embodiment, the step S1 specifically includes the following steps:
s101: simulating a cabin skeleton by adopting a shell unit;
the locking point connection between the cabin door and the cabin skeleton is simulated by a rigid body element RBE3, the most severe working condition is considered during analysis, only the main stress part is considered during modeling, the non-main stress part is removed, and the connection between the cabin door and the cabin skeleton is simulated by the rigid body element;
s102: assigning material properties;
inputting a basic layering sequence of the composite material of the cabin skeleton according to actual layering information based on a classical lamination theory, and assigning corresponding material attributes to the shell units;
s103: carrying out boundary definition;
defining the boundary comprises determining a load working condition and determining a constraint condition; the load working condition is determined by taking the working condition with the most serious load as the load working condition; determining the constraint condition means that the whole cabin framework is fixedly connected with the middle fuselage of the helicopter, and 3 degrees of freedom of a connection area of the cabin framework and the middle fuselage of the helicopter are constrained.
In a possible embodiment, the step S2 specifically includes the following steps:
s201: on the basis of finite element modeling of the cabin framework, partitioning the cabin framework;
s202: respectively calculating to obtain initial maximum deformation s for each cabin skeleton partition1Longitudinal equivalent elastic modulus and transverse equivalent elastic modulus;
calculating the equivalent longitudinal elastic modulus E according to the following calculation formula Ix(ii) a Calculating the transverse equivalent elastic modulus E according to the following formula IIy：
Wherein E is11Is the longitudinal elastic modulus of the composite material;
E22is the transverse elastic modulus of the composite material;
G12is the in-plane shear modulus of the composite material;
u12the cedar ratio of the composite material in XY direction;
theta is an included angle between the X direction of the reference coordinate system and the longitudinal direction of the composite material fiber;
s203: respectively implementing a strengthening scheme on each cabin skeleton partition;
s204: respectively calculating and counting the longitudinal equivalent elastic modulus and the transverse equivalent elastic modulus of each cabin skeleton partition subjected to the implementation of each strengthening scheme, and the maximum deformation; ex、EyIs an optimization design variable related to the number of layers of composite material layering, the laying angle, the layering sequence and the layering material;
s205: calculating the deformation difference deltas of each cabin skeleton subarea before and after implementing each strengthening scheme according to the formula III,
Δs＝s2-s1formula III
Wherein the maximum deformation after implementation of the reinforcing embodiment is s2Initial maximum deflection of s1；
The mass difference am of each cabin skeleton partition before and after implementing the reinforcement scheme is calculated according to equation four,
Δm＝m2-m1formula IV
Wherein the mass after implementation of the reinforcing scheme is m2Initial mass m1；
Defining a parameter 'deformation difference/mass increment' as a deformation mass increment ratio Y, and calculating the deformation mass increment ratio Y according to the formula V;
carrying out statistics on the deformation quality increment ratio Y obtained by calculation before and after the strengthening scheme is implemented on each cabin skeleton partition;
s206: optimization design analysis with commercial software for equivalent modulus of elasticity E in the longitudinal directionxTransverse equivalent modulus of elasticity EyAs a design variable, the deformation mass increment ratio Y is used as a function value, and a response surface is formed by fitting;
if the response surface is a plane, the rigidity characteristic is a linear problem; if the response surface is a curved surface with a single peak, the rigidity characteristic is a unimodal nonlinear problem; if a plurality of peaks exist in the response surface, the rigidity characteristic is proved to be a multi-peak nonlinear problem.
In one possible embodiment, the first step in step S2 is to partition the cabin skeleton, including front area 1, vertical area 2, and lateral area 3; the front area 1 is positioned at the front end of the cabin skeleton; the vertical area 2 is positioned at the lower cross beams at two sides of the cabin skeleton; the transverse areas 3 are positioned at the upper longitudinal beams on the two sides of the cabin framework.
In a possible embodiment, in step S203, the step of performing a reinforcement scheme on each cabin skeleton segment includes adding and/or changing a direction of a mat.
In a possible embodiment, the steps S203-S204 are repeated for each cabin skeleton partition, and multiple reinforcement schemes are adopted to obtain the longitudinal equivalent elastic modulus E under each reinforcement schemexTransverse equivalent modulus of elasticity EyAnd a statistical value of the deformation mass increment ratio Y.
In a possible embodiment, in step S3, according to the optimal design objective with the largest influence of the unit weight gain on the overall stiffness, and in combination with the response surface obtained by fitting in step S205, the optimal reinforcement scheme for each cabin framework partition is determined, so as to form a cabin framework stiffness optimization scheme; the smaller the amount of change in the amount of deformation at the same weight gain, the greater the stiffness per unit mass increase, and the greater the structural efficiency.
In one possible embodiment, simulation optimization analysis is performed using Hypermesh software.
Example 1
Taking the front area 1 as an example, simulation optimization analysis is carried out by adopting Hypermesh software, and initial ply [45C1/0C2/0C1/0C1/0C2/45C1 is given firstly]Calculating to obtain the initial longitudinal equivalent elastic modulus Ex11844.5MPa, equivalent modulus of elasticity E in transverse directiony7515.2MPa, initial maximum deformation 10.2 mm;
the embodiment of the reinforcement scheme is a build-up ply, and the reinforcement scheme is specifically [45C1/0C2/0C1/0C1/0C1/0C2/45C1]The equivalent modulus of elasticity E in the longitudinal direction after implementation of the reinforcing embodiment is calculatedx21873.8MPa, equivalent modulus of elasticity E in transverse directiony16625.1MPa, a deformation difference of 0.64mm from the initial maximum deformation, and a mass increment from the initial mass of 5.03E-04 t; as shown in Table 1, the equivalent modulus of elasticity E in the machine direction under various reinforcing schemes was calculated and countedxTransverse equivalent modulus of elasticity EyAnd a deformation mass increment ratio Y, and the other reinforcement scheme calculation methods are the same as the above method; with Ex，EyAs a design variable, the deformation mass increment ratio Y is taken as a function value, and the fitting function is
Constructing a corresponding response surface as shown in figure 4,
according to an optimal design target with the largest influence of unit weight gain on the overall rigidity and a response surface obtained by fitting, obtaining a reinforcement scheme corresponding to the Y of 1.26E +03 as an optimal reinforcement scheme of the front area 1;
the optimization process of other subareas of the cabin skeleton is analogized in turn, and the longitudinal equivalent elastic modulus ExTransverse equivalent modulus of elasticity EyAnd the deformation mass increment ratio Y statistics are shown in tables 2 and 3;
table 1 is a front region optimization scheme for the inventive cabin skeleton, with results corresponding to fig. 4.
TABLE 1 optimization of the frontal area of the cabin skeleton
Table 2 shows the vertical domain optimization of the inventive cabin skeleton, the results corresponding to fig. 5.
TABLE 2 vertical region optimization scheme for cabin skeleton
Table 3 is a lateral zone optimization scheme of the inventive cabin skeleton, the results corresponding to fig. 6.
TABLE 3 lateral zone optimization scheme for cabin skeleton
Table 4 shows the cabin skeleton optimization scheme of the present invention, with the results corresponding to fig. 7.
TABLE 4 optimization scheme for cabin skeleton
The above embodiments are only used for illustrating the technical solutions of the present application, and not for limiting the same; although the present application has been described in detail with reference to the foregoing embodiments, it should be understood by those of ordinary skill in the art that: the technical solutions described in the foregoing embodiments may still be modified, or some technical features may be equally replaced; such modifications and substitutions do not substantially depart from the spirit and scope of the embodiments of the present application, and are intended to be included within the scope of the present application.
Claims (8)
1. A helicopter cabin skeleton rigidity optimization method is characterized by comprising the following steps:
s1: finite element modeling of a cabin skeleton;
s2: optimizing, analyzing and modeling the rigidity of the cabin skeleton;
s3: evaluating an optimization result to obtain a cabin skeleton rigidity optimization scheme; and (4) evaluating the structural scheme and the weight cost according to the result of optimization evaluation, and if the structural scheme and the weight cost are not satisfactory, repeating the step S2 to perform iterative optimization to finally obtain the optimal cabin skeleton stiffness optimization scheme.
2. The helicopter cabin skeleton stiffness optimization method of claim 1, wherein the step S1 specifically comprises the steps of:
s101: simulating a cabin skeleton by adopting a shell unit;
the locking point connection between the cabin door and the cabin skeleton is simulated by a rigid body element RBE3, the most severe working condition is considered during analysis, only the main stress part is considered during modeling, the non-main stress part is removed, and the connection between the cabin door and the cabin skeleton is simulated by the rigid body element;
s102: assigning material properties;
inputting a basic layering sequence of the composite material of the cabin skeleton according to actual layering information based on a classical lamination theory, and assigning corresponding material attributes to the shell units;
s103: carrying out boundary definition;
defining the boundary comprises determining a load working condition and determining a constraint condition; the load working condition is determined by taking the working condition with the most serious load as the load working condition; determining the constraint condition means that the whole cabin framework is fixedly connected with the middle fuselage of the helicopter, and 3 degrees of freedom of a connection area of the cabin framework and the middle fuselage of the helicopter are constrained.
3. The helicopter cabin skeleton stiffness optimization method according to claim 2, wherein the step S2 specifically comprises the steps of:
s201: on the basis of finite element modeling of the cabin framework, partitioning the cabin framework;
s202: respectively calculating to obtain initial maximum deformation, longitudinal equivalent elastic modulus and transverse equivalent elastic modulus aiming at each cabin skeleton partition;
calculating the equivalent longitudinal elastic modulus E according to the following calculation formula Ix(ii) a Calculating the transverse equivalent elastic modulus E according to the following formula IIy：
Wherein E is11Is the longitudinal elastic modulus of the composite material;
E22is the transverse elastic modulus of the composite material;
G12is the in-plane shear modulus of the composite material;
u12the cedar ratio of the composite material in XY direction;
theta is an included angle between the X direction of the reference coordinate system and the longitudinal direction of the composite material fiber;
s203: respectively implementing a strengthening scheme on each cabin skeleton partition;
s204: respectively calculating and counting the longitudinal equivalent elastic modulus and the transverse equivalent elastic modulus of each cabin skeleton partition subjected to the implementation of each strengthening scheme, and the maximum deformation; ex、EyIs an optimization design variable related to the number of layers of composite material layering, the laying angle, the layering sequence and the layering material;
s205: calculating the deformation difference deltas of each cabin skeleton subarea before and after implementing each strengthening scheme according to the formula III,
Δs＝s2-s1formula III
Wherein the maximum deformation after implementation of the reinforcing embodiment is s2Initial maximum deflection of s1；
The mass difference am of each cabin skeleton partition before and after implementing the reinforcement scheme is calculated according to equation four,
Δm＝m2-m1formula IV
Wherein the mass after implementation of the reinforcing scheme is m2Initial mass m1；
Defining a parameter 'deformation difference/mass increment' as a deformation mass increment ratio Y, and calculating the deformation mass increment ratio Y according to the formula V;
carrying out statistics on the deformation quality increment ratio Y obtained by calculation before and after the strengthening scheme is implemented on each cabin skeleton partition;
s206: optimization design analysis with commercial software for equivalent modulus of elasticity E in the longitudinal directionxTransverse equivalent modulus of elasticity EyAnd as a design variable, taking the deformation mass increment ratio Y as a function value, and fitting to form a response surface.
4. A helicopter cabin skeleton stiffness optimization method according to claim 3, characterized by that, the first step in step S2 is to divide the cabin skeleton into zones comprising a front zone (1), a vertical zone (2), and a lateral zone (3); the front area (1) is positioned at the front end of the cabin skeleton; the vertical regions (2) are positioned at the lower cross beams at two sides of the cabin skeleton; the transverse areas (3) are positioned at the upper longitudinal beams on the two sides of the cabin framework.
5. A helicopter cabin skeleton stiffness optimization method according to claim 3, characterized in that in step S203, said reinforcement scheme for each cabin skeleton section comprises adding and/or changing the direction of the lay-up.
6. A helicopter cabin skeleton stiffness optimization method according to claim 3, characterized by being specific to each cabinThe steps S203-S204 are repeatedly carried out on the framework partition, and a plurality of reinforcement schemes are adopted to obtain the longitudinal equivalent elastic modulus E under each reinforcement schemexTransverse equivalent modulus of elasticity EyAnd a statistical value of the deformation mass increment ratio Y.
7. A helicopter cabin skeleton stiffness optimization method according to any one of claims 3-6, characterized in that in said step S3, according to the optimization design objective that the unit weight gain has the largest influence on the overall stiffness, the optimal reinforcement scheme for each cabin skeleton partition is determined, so as to form a cabin skeleton stiffness optimization scheme; the smaller the amount of change in the amount of deformation at the same weight gain, the greater the stiffness per unit mass increase, and the greater the structural efficiency.
8. The helicopter cabin skeleton stiffness optimization method of claim 1, wherein Hypermesh software is used for simulation optimization analysis.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011309583.8A CN112380624B (en) | 2020-11-20 | 2020-11-20 | Helicopter cabin skeleton rigidity optimization method |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011309583.8A CN112380624B (en) | 2020-11-20 | 2020-11-20 | Helicopter cabin skeleton rigidity optimization method |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN112380624A true CN112380624A (en) | 2021-02-19 |
| CN112380624B CN112380624B (en) | 2022-10-28 |
Family
ID=74584508
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202011309583.8A Active CN112380624B (en) | 2020-11-20 | 2020-11-20 | Helicopter cabin skeleton rigidity optimization method |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN112380624B (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113636098A (en) * | 2021-10-18 | 2021-11-12 | 成都飞机工业（集团）有限责任公司 | Design method of process stiffening piece for aircraft component |
Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20150170022A1 (en) * | 2013-12-13 | 2015-06-18 | King Fahd University Of Petroleum And Minerals | Method and apparatus for characterizing composite materials using an artificial neural network |
| CN107016182A (en) * | 2017-03-30 | 2017-08-04 | 中国航空工业集团公司西安飞机设计研究所 | A kind of composite hatch door optimization method |
| CN107563094A (en) * | 2017-09-21 | 2018-01-09 | 上海交通大学 | Three-dimensional woven carbon fibre composite fender optimization method |
| CN107844669A (en) * | 2017-12-01 | 2018-03-27 | 中国直升机设计研究所 | A kind of full machine static(al) FEM updating of helicopter and verification method |
| CN110704953A (en) * | 2019-09-30 | 2020-01-17 | 西北工业大学 | Analysis method for design sensitivity of static air elastic energy of high-aspect-ratio wing |
-
2020
- 2020-11-20 CN CN202011309583.8A patent/CN112380624B/en active Active
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20150170022A1 (en) * | 2013-12-13 | 2015-06-18 | King Fahd University Of Petroleum And Minerals | Method and apparatus for characterizing composite materials using an artificial neural network |
| CN107016182A (en) * | 2017-03-30 | 2017-08-04 | 中国航空工业集团公司西安飞机设计研究所 | A kind of composite hatch door optimization method |
| CN107563094A (en) * | 2017-09-21 | 2018-01-09 | 上海交通大学 | Three-dimensional woven carbon fibre composite fender optimization method |
| CN107844669A (en) * | 2017-12-01 | 2018-03-27 | 中国直升机设计研究所 | A kind of full machine static(al) FEM updating of helicopter and verification method |
| CN110704953A (en) * | 2019-09-30 | 2020-01-17 | 西北工业大学 | Analysis method for design sensitivity of static air elastic energy of high-aspect-ratio wing |
Non-Patent Citations (6)
| Title |
|---|
| GAO, HP 等: "Feasibility analysis of high excess pressure cockpit using material substitution based on strain energy", 《SCIENCE CHINA-TECHNOLOGICAL SCIENCES》, vol. 57, no. 3, 31 March 2014 (2014-03-31), pages 598 - 606 * |
| SAN, XG 等: "Lightweight design of large scale theodolite"s turntable based on topology optimization of continuum structure", 《INTERNATIONAL CONFERENCE ON OPTOELECTRONICS AND MICROELECTRONICS (ICOM)》, 31 August 2012 (2012-08-31), pages 570 - 574, XP032244027, DOI: 10.1109/ICoOM.2012.6316340 * |
| 张文静等: "直升机金属平板入水冲击仿真及修正", 《机械设计与制造工程》, vol. 47, no. 06, 15 June 2018 (2018-06-15), pages 78 - 81 * |
| 杨留义等: "大型平面天线可展开支撑结构变形影响因素分析", 《遥测遥控》, vol. 40, no. 05, 15 September 2019 (2019-09-15), pages 42 - 46 * |
| 郭璐等: "基于SIMP拓扑优化的驾驶舱轻量化设计", 《计算机工程与设计》, vol. 36, no. 09, 16 September 2015 (2015-09-16), pages 2490 - 2495 * |
| 郭璐等: "多工况载荷下驾驶舱结构刚度优化设计与仿真", 《计算机仿真》, vol. 32, no. 10, 15 October 2015 (2015-10-15), pages 43 - 47 * |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113636098A (en) * | 2021-10-18 | 2021-11-12 | 成都飞机工业（集团）有限责任公司 | Design method of process stiffening piece for aircraft component |
Also Published As
| Publication number | Publication date |
|---|---|
| CN112380624B (en) | 2022-10-28 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN106874573B (en) | Design method of partitioned variable-thickness composite material laminated plate | |
| CN108427826B (en) | Whole-process analysis method for impact damage and residual strength of sewn composite material | |
| US7840386B2 (en) | Finite element modeling method utilizing mass distribution | |
| KR101200167B1 (en) | Analysis system and method for composite materials | |
| CN112380624B (en) | Helicopter cabin skeleton rigidity optimization method | |
| CN105447269A (en) | Non-structural mass trimming method for airplane in steady overload state | |
| CN107016182A (en) | A kind of composite hatch door optimization method | |
| Schwenke et al. | Optimization of load introduction points in sandwich structures with additively manufactured cores | |
| CN101612996A (en) | A kind of plate muscle construction design method | |
| Lu et al. | Fatigue damage modeling in laminated composite by using Rx-FEM and strength tracking method | |
| CN113987689B (en) | Comprehensive optimization design method for reinforced pressure frame of composite material grille based on shape modification | |
| CN105468826B (en) | The design method of composite material | |
| CN112131770B (en) | Functional gradient continuum structure lightweight design method considering reliability | |
| CN110334364B (en) | Composite structural strength checking system | |
| Rao et al. | Topology optimization of aircraft wing | |
| CN114912306B (en) | Modeling method of 2.5D woven composite material flat plate structure multi-scale model | |
| Chedrik | Two-level design optimization of aircraft structures under stress, buckling and aeroelasticity constraints | |
| Zink et al. | Robust design for aeroelastically tailored/active aeroelastic wing | |
| Özakça et al. | Buckling and post-buckling of sub-stiffened or locally tailored aluminium panels | |
| CN110659525B (en) | Optimized design method for composite structure | |
| CN112711801A (en) | Rectangular thin-wall machine body large-opening structure torsion load distribution calculation method | |
| CN113049360B (en) | Method for determining compression allowable strain value of aircraft composite material reinforced wallboard | |
| CN115906570B (en) | Data-driven composite material open-pore laminated plate progressive damage prediction method | |
| CN117669045B (en) | Aircraft crashworthiness reliability optimization design method | |
| Lionetti | MDO analysis of composite wing |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |