CN112699471A - Method and device for calculating effective width of skin under axial compression load of fuselage wallboard - Google Patents

Method and device for calculating effective width of skin under axial compression load of fuselage wallboard Download PDF

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CN112699471A
CN112699471A CN202011603089.2A CN202011603089A CN112699471A CN 112699471 A CN112699471 A CN 112699471A CN 202011603089 A CN202011603089 A CN 202011603089A CN 112699471 A CN112699471 A CN 112699471A
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skin
effective width
stringer
determining
case
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CN112699471B (en
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杜鹏良
张新全
刘彦杰
李明强
李禹�
冯军
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Xian Aircraft Design and Research Institute of AVIC
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    • 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
    • G06F2119/00—Details relating to the type or aim of the analysis or the optimisation
    • G06F2119/14—Force analysis or force optimisation, e.g. static or dynamic forces

Abstract

The invention belongs to the field of a method for calculating axial compression load of a fuselage wallboard, and relates to a method and a device for calculating effective width of skin under axial compression load of the fuselage wallboard. The method comprises the following steps: determining the average thickness t of the wallboard unit skin; determining a coefficient of influence c of a fastener on the effective width, wherein the fastener is a fastener for connecting the skin to the stringer; determining the rigidity A of the stringer and the rigidity B of the skin; determining the effective width B of the skin under the axial compression load of the fuselage panel according to the average thickness t of the skin, the influence coefficient c of the fastener on the effective width, the rigidity A of the stringer and the rigidity B of the skine. The method has the advantages that the effective skin width is calculated simply and accurately, the problem of conservation of the original calculation method is solved, engineering personnel can apply the method to airplane design conveniently, great value is brought to weight reduction and economic value creation of airplane structures, and the method is suitable for calculating the structural strength of various fuselage wall plates.

Description

Method and device for calculating effective width of skin under axial compression load of fuselage wallboard
Technical Field
The invention belongs to the field of a method for calculating axial compression load of a fuselage wallboard, and relates to a method and a device for calculating effective width of skin under axial compression load of the fuselage wallboard.
Background
The wall plate is the main load transfer structure of the fuselage of the transport type airplane, and the bearing capacity under tensile compression, shearing, bending and torsion loads is the key point of airplane design research.
Because the skins are all thin plates and can be bent under the design load of the airplane, the stress of the part of the skin under the stringer after the skins are subjected to compressive bending can be continuously increased along with the increase of the load, and the part of the skin is called as an axial compressive load bearing effective skin. At present, the effective width of the skin under the stringer in engineering design is usually taken according to the thickness t of the skin by 30, the method is conservative, and the weight reduction and the economy of an airplane are not facilitated, so that a new effective skin width calculation method needs to be provided.
Disclosure of Invention
The purpose of the invention is as follows: the method and the device for calculating the effective width of the skin under the axial compression load of the fuselage panel with high accuracy are provided.
The technical scheme is as follows:
in a first aspect, a method for calculating an effective width of a skin under an axial compression load of a fuselage wallboard is provided, and comprises the following steps:
determining the average thickness t of the wallboard unit skin;
determining a coefficient of influence c of a fastener on the effective width, wherein the fastener is a fastener for connecting the skin to the stringer;
determining the rigidity A of the stringer and the rigidity B of the skin;
determining the effective width B of the skin under the axial compression load of the fuselage panel according to the average thickness t of the skin, the influence coefficient c of the fastener on the effective width, the rigidity A of the stringer and the rigidity B of the skine。
Further, effective width b of skin under axial compression load of fuselage paneleComprises the following steps:
in that
Figure BDA0002869825970000021
In the case of (a) in (b),
Figure BDA0002869825970000022
in that
Figure BDA0002869825970000023
In the case of (a) in (b),
Figure BDA0002869825970000024
in that
Figure BDA0002869825970000025
In the case of (A), (B) in the case of (A),
Figure BDA0002869825970000026
in that
Figure BDA0002869825970000027
In the case of (A), (B) in the case of (A),
Figure BDA0002869825970000028
in that
Figure BDA0002869825970000029
In the case of (a) in (b),
Figure BDA00028698259700000210
in that
Figure BDA00028698259700000211
In the case of (a) in (b),
Figure BDA00028698259700000212
further, determining an influence coefficient c of the fastener on the effective width, specifically:
and determining the influence coefficient of the bolt on the effective width and the influence coefficient of the rivet on the effective width according to the diameter of the bolt or the diameter of the rivet.
Further, the stiffness A of the stringer is the product of the modulus of elasticity of the stringer and the cross-sectional area of the stringer; the stiffness B of the skin is the product of the modulus of elasticity of the skin and the cross-sectional area of the skin.
Further, determining the average thickness t of the panel unit skin, specifically:
the average thickness t of the panel unit skin is the ratio of the sum of the cross-sectional area of the skin below the stringer, the cross-sectional area of the skin on the left side of the stringer in the panel unit, the cross-sectional area of the skin on the right side of the stringer in the panel unit to the width L of the panel unit.
In a second aspect, there is provided a device for calculating an effective width of a skin under an axial compressive load of a fuselage panel, comprising:
a first determination module for determining an average thickness t of a panel unit skin; determining a coefficient of influence c of a fastener on the effective width, wherein the fastener is a fastener for connecting the skin to the stringer; determining the rigidity A of the stringer and the rigidity B of the skin;
a second determination module for determining the effective width B of the skin under the axial compression load of the fuselage panel according to the average thickness t of the skin, the influence coefficient c of the fastener on the effective width, the rigidity A of the stringer and the rigidity B of the skine。
Further, the second determination module is particularly useful for determining the effective width b of the skin under axial compression loading of the fuselage paneleThe determination is as follows:
in that
Figure BDA00028698259700000213
In the case of (a) in (b),
Figure BDA00028698259700000214
in that
Figure BDA0002869825970000031
In the case of (a) in (b),
Figure BDA0002869825970000032
in that
Figure BDA0002869825970000033
In the case of (A), (B) in the case of (A),
Figure BDA0002869825970000034
in that
Figure BDA0002869825970000035
In the case of (A), (B) in the case of (A),
Figure BDA0002869825970000036
in that
Figure BDA0002869825970000037
In the case of (a) in (b),
Figure BDA0002869825970000038
in that
Figure BDA0002869825970000039
In the case of (a) in (b),
Figure BDA00028698259700000310
further, the first determining module is specifically configured to: and determining the influence coefficient of the bolt on the effective width and the influence coefficient of the rivet on the effective width according to the diameter of the bolt or the diameter of the rivet.
Has the advantages that: the method has the advantages that the effective skin width is calculated simply and accurately, the problem of conservation of the original calculation method is solved, engineering personnel can apply the method to airplane design conveniently, great value is brought to weight reduction and economic value creation of airplane structures, and the method is suitable for calculating the structural strength of various fuselage wall plates.
Drawings
FIG. 1 is a schematic view of an effective width of a panel skin;
fig. 2 is a schematic view of the effective width of the panel skin.
Detailed Description
When a fuselage panel unit is loaded in compression, the stringers carry the vast majority of the load, in addition to which the skin carries a non-negligible load, the load carrying capacity of the effective width skin (Be as shown in figure 1) being an important component of the panel load carrying capacity. At present, two methods exist for calculating the bearing capacity of the wallboard in engineering, wherein the most convenient calculation method is that the effective width of the skin under the stringer is 30 x the thickness t of the skin, but the method is conservative and is not beneficial to weight reduction and economy of an airplane; in addition, the wall plate strength calculation also has a mature empirical method, but the empirical method has complex calculation process and multiple steps, the effective width of the skin needs to be iterated for multiple times, and only a few aircraft design mechanisms hold due to copyright problem, so the applicability is not wide. The method solves the problem that the thickness of the skin is excessively conservative when the effective width of the skin is 30 times, greatly promotes the weight reduction of the aircraft structure, is convenient to calculate, and is very convenient for engineers to estimate and use in daily life.
As shown in fig. 1, a panel unit refers to a skin that is connected by stringers and stringers-rivets and supported by the stringers.
The invention provides a set of calculation method which is more convenient for calculating the effective width of the wall panel unit skin, the calculation method is simple, the problem of conservation of the original calculation method is solved, engineering personnel can be conveniently applied to the design of an airplane, great value is brought to the weight reduction and economic value creation of the airplane structure, and the calculation method is suitable for calculating the structural strength of various fuselage wall panels.
The method comprises the following steps:
(1) determining parameters of fuselage wall panel units affecting effective skin width
The fuselage panel of the conveyor consists of a Z-shaped stringer and a thin skin, and when the fuselage panel is designed, the ratio of the cross-sectional area of the stringer to the cross-sectional area of the skin is generally between 0.4 and 0.9 in consideration of the rigidity matching of the skin and the stringer. When the aircraft stiffened panel structure is subjected to compressive load, the thin skin structure can be subjected to buckling instability under low compressive load to reach the bearing limit, so that the main bearing capacity of the stiffened panel is provided by the stringers and the skin with effective width under the stringers (shown as be in fig. 1, sigma-deltacrIs the buckling stress of the skin, σCPFor buckling stringer-skin composite columnsStress), it can be seen that the value of the effective width of the skin under the stringer is an important factor for calculating the bearing capacity of the panel unit when the stiffened panel is pressed. And after the effective width of the skin is obtained, calculating the bearing capacity of the wallboard unit according to the parameters of the wallboard column unit and an Euler Johnson formula.
And analyzing the effective width of the skin by adopting a control variable method according to hundreds of groups of test data of the plurality of airplanes. Skin to stringer stiffness ratio, skin thickness, fastener diameter, and panel unit size were found to be factors that primarily affect the effective width.
(2) Determining the coefficient of influence of the fastener on the effective width
The test research shows that the larger the fastener for connecting the skin and the stringer is, the larger the bearing capacity of the wallboard unit is, and the stronger the bearing capacity of the skin under the stringer is.
Common skin to stringer tie fastener impact factors can be selected with reference to tables 1 and 2.
TABLE 1 bolt influence factor Table
Bolt diameter D/mm 4 6 8
Influencing factor c 1.18 1.21 1.24
TABLE 2 rivet influence factor Table
Rivet diameter D/mm 4 5 6
Influencing factor c 1.12 1.15 1.18
The method has the advantages that the influence of the fasteners on the effective width of the skin is given simply and visually, the use is convenient, the applicability is strong, and the calculation of the effective width of the skin stringer connecting fasteners of various types of airplanes at present is met.
(3) Calculating stiffness ratio of stringer to skin
According to analysis of multiple groups of test data, the relationships between the effective width and the skin thickness and the relationship between the stringer and the skin rigidity ratio are drawn, the E value and the sectional area of the stringer and the skin material are found, and the rigidity matching relationship of the formed wallboard has a large influence on the effective width. Wherein the content of the first and second substances,
defining stringer to skin stiffness ratio as:
Figure BDA0002869825970000051
the stringer to skin stiffness ratio can be calculated according to the above equation.
(4) Calculating the average thickness t of the wall plate unit skin;
referring to fig. 2, since the fuselage skin is generally a chemically milled skin, calculating the skin cross-sectional area needs to consider the skin thickness after chemically milling and the change of the skin cross-sectional area caused by the chemically milled boss.
Figure BDA0002869825970000052
SzcmpCross-sectional area of skin on left side of stringer of panel unit
SttIs the sectional area of the lower boss of the stringer of the wallboard unit
SycmpCross-sectional area of skin on right side of stringer of panel unit
(5) Calculating the effective width of the wall panel unit skin according to the effective width calculation formula
Figure BDA0002869825970000053
Figure BDA0002869825970000061
Figure BDA0002869825970000062
Figure BDA0002869825970000063
Figure BDA0002869825970000064
Figure BDA0002869825970000065
In the formula for calculating the effective width provided by the step, the rigidity ratio value of the stringer and the skin adopts a rounding method, the formula is simple and convenient to calculate, the calculation of the effective width of the skin of the wall plate of the fuselage of various conveyers is covered,
the invention provides a method for calculating the effective width of the wall panel unit skin more conveniently, the effective skin width is calculated simply and accurately, the problem of conservation of the original calculation method is solved, engineering personnel can be applied to the design of an airplane conveniently, great value is brought to the weight reduction and economic value creation of the airplane structure, and the method is suitable for calculating the structural strength of various fuselage wall panels.

Claims (8)

1. A method for calculating the effective width of a skin under the axial compression load of a fuselage wallboard is characterized by comprising the following steps:
determining the average thickness t of the wallboard unit skin;
determining a coefficient of influence c of a fastener on the effective width, wherein the fastener is a fastener for connecting the skin to the stringer;
determining the rigidity A of the stringer and the rigidity B of the skin;
determining the effective width B of the skin under the axial compression load of the fuselage panel according to the average thickness t of the skin, the influence coefficient c of the fastener on the effective width, the rigidity A of the stringer and the rigidity B of the skine。
2. The method of claim 1, wherein the effective width b of the skin under the fuselage panel axial compression load iseComprises the following steps:
in that
Figure FDA0002869825960000011
In the case of (a) in (b),
Figure FDA0002869825960000012
in that
Figure FDA0002869825960000013
In the case of (a) in (b),
Figure FDA0002869825960000014
in that
Figure FDA0002869825960000015
In the case of (A), (B) in the case of (A),
Figure FDA0002869825960000016
in that
Figure FDA0002869825960000017
In the case of (A), (B) in the case of (A),
Figure FDA0002869825960000018
in that
Figure FDA0002869825960000019
In the case of (a) in (b),
Figure FDA00028698259600000110
in that
Figure FDA00028698259600000111
In the case of (a) in (b),
Figure FDA00028698259600000112
3. the method according to claim 1, characterized in that the coefficient of influence c of the fastener on the effective width is determined, in particular:
and determining the influence coefficient of the bolt on the effective width and the influence coefficient of the rivet on the effective width according to the diameter of the bolt or the diameter of the rivet.
4. The method of claim 1, wherein the stringer stiffness, a, is the product of the modulus of elasticity of the stringer and the stringer cross-sectional area; the stiffness B of the skin is the product of the modulus of elasticity of the skin and the cross-sectional area of the skin.
5. The method according to claim 1, characterized in that the average thickness t of the panel unit skin is determined by:
the average thickness t of the panel unit skin is the ratio of the sum of the cross-sectional area of the skin below the stringer, the cross-sectional area of the skin on the left side of the stringer in the panel unit, the cross-sectional area of the skin on the right side of the stringer in the panel unit to the width L of the panel unit.
6. An effective width of skin calculation device under fuselage wallboard axle load, comprising:
a first determination module for determining an average thickness t of a panel unit skin; determining a coefficient of influence c of a fastener on the effective width, wherein the fastener is a fastener for connecting the skin to the stringer; determining the rigidity A of the stringer and the rigidity B of the skin;
a second determination module for determining the effective width B of the skin under the axial compression load of the fuselage panel according to the average thickness t of the skin, the influence coefficient c of the fastener on the effective width, the rigidity A of the stringer and the rigidity B of the skine。
7. The device according to claim 6, characterized in that the second determination module is particularly intended for determining the effective width b of the skin under axial compressive loading of a fuselage paneleThe determination is as follows:
in that
Figure FDA0002869825960000021
In the case of (a) in (b),
Figure FDA0002869825960000022
in that
Figure FDA0002869825960000023
In the case of (a) in (b),
Figure FDA0002869825960000024
in that
Figure FDA0002869825960000025
In the case of (A), (B) in the case of (A),
Figure FDA0002869825960000026
in that
Figure FDA0002869825960000027
In the case of (A), (B) in the case of (A),
Figure FDA0002869825960000028
in that
Figure FDA0002869825960000029
In the case of (a) in (b),
Figure FDA00028698259600000210
in that
Figure FDA00028698259600000211
In the case of (a) in (b),
Figure FDA00028698259600000212
8. the apparatus of claim 6, wherein the first determining module is specifically configured to: and determining the influence coefficient of the bolt on the effective width and the influence coefficient of the rivet on the effective width according to the diameter of the bolt or the diameter of the rivet.
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Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN103870613A (en) * 2012-12-10 2014-06-18 中国飞机强度研究所 Calculation method of bearing capacity of reinforced wall plate
CN111143941A (en) * 2019-12-24 2020-05-12 中国航空工业集团公司西安飞机设计研究所 Method for calculating axial compression bearing capacity of composite material reinforced wall plate

Patent Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN103870613A (en) * 2012-12-10 2014-06-18 中国飞机强度研究所 Calculation method of bearing capacity of reinforced wall plate
CN111143941A (en) * 2019-12-24 2020-05-12 中国航空工业集团公司西安飞机设计研究所 Method for calculating axial compression bearing capacity of composite material reinforced wall plate

Non-Patent Citations (3)

* Cited by examiner, † Cited by third party
Title
方阳等: "民用飞机壁板蒙皮及长桁布置结构优化设计", 《民用飞机设计与研究》 *
王一飞等: "长桁蒙皮截面积比对复合材料壁板压缩屈曲的影响", 《中国科技信息》 *
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