CN108563916A - Aircraft wing fuselage thin-wall construction original dimension optimum design method - Google Patents
Aircraft wing fuselage thin-wall construction original dimension optimum design method Download PDFInfo
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Abstract
Aircraft wing fuselage thin-wall construction original dimension optimum design method, this method are based on thin wall engineering beam theory, determine the wing at aircraft project initial stage and the thin wall reinforced siding scheme of fuselage.Thin wall reinforced siding original dimension design is challenging and judgement property a job, and whether the original dimension that conceptual design defines is reasonable, directly affects subsequent design work.Ideal wall panel structure will fully meet the design requirements such as Static Strength Design, fatigue damage tolerance design, manufacture and repair.In original dimension design process, simplifies any one in analysis method or structure the two to be analyzed, can all improve project cost benefit.It is further introduced into manufacture and maintenance requirement, more optimized original dimension can be obtained under the premise of meeting static strength requirement in the schematic design phase.In the case where intensity requirement meets, original dimension definition is carried out to structure by the design requirements such as manufacturing, repairing, not only conforms with structure lightened design requirement, and improve design efficiency.
Description
Technical field
The present invention relates at the beginning of the aircraft wings such as military aircraft, civil aircraft, space shuttle, rocket, fuselage thin-wall construction
Beginning sizing method can be used for general thin wall reinforced structure original dimension design, belong to thin-wall construction design method field.
Background technology
Since modern project is for the needs of safety, economy, thin-wall construction is come across extensively in various engineering structures,
Such as Aeronautics and Astronautics, bridge, shipbuilding, building field.Thin-wall construction no matter from intensity, rigidity, weight and economy etc. all
With significant superiority, lighter construction weight can be obtained under conditions of meeting strength and stiffness requirement.
Design, manufacture and the experimental technique of aircraft wing fuselage thin wall reinforced structure are full machine systematization development processes
One of key technology.For structure design under the premise of ensuring reliability, comfort, product light-weight design receives designer pass
Note.Aircraft development process be one from coarse to fine, progressive alternate, optimization process, on the basis of cost control, design it is excellent
The more early of beginning is melted, the design scheme theoretically obtained is more reasonable.
At the same time, Scheme design excessively relies on system engineering teacher, the experience of structural engineer and with reference to corresponding type
Number configuration.Due to being pressed for time, data it is few, intensity engineer intervenes not enough in solution formulation and design process, cannot provide
One or more meet the prioritization scheme of intensity requirement, and excessively rely on Patran＆Nastran finite element analysis softwares.Its
Main reason is that designer is not thorough enough to thin-wall construction rigidity and strength study, the assembly and repair of siding are lacked
Solution.
Invention content
In view of the above problems, the invention proposes a kind of Wing-Body Configurations thin-wall construction original dimension optimization design side
Method, can be widely used in the aircraft wings such as military aircraft, civil aircraft, space shuttle, rocket, fuselage thin-wall construction side
Case designs, and specification analysis method can improve working efficiency, shorten the design cycle, reduce R＆D costs.
To achieve the goals above, the technical solution that the invention uses for：At the beginning of aircraft wing fuselage thin-wall construction
Beginning size optimal design method, which is characterized in that its step is：
1) it according to aircraft wing, the total arrangement of fuselage, plans the position of lateral reinforcer and longitudinal elements, calculates winged
Row device wing, the section load of fuselage thin-wall construction conceptual design and basic structure definition；
2) according to aircraft cross section definition and section load, calculate covering, longitudinal elements stress, identified sign horizontal distribution and
Peak value；
3) according to aircraft wing, the intensity of fuselage wallboard, manufacture and maintenance requirement, part of wall plate cross section parameter is determined
Minimum dimension, including：
3.1) skin thickness tp：It is determined according to the maintainability that the intensity of covering is connected with siding；
3.2) longitudinal elements outer edge strip width wsf：Define wsf1For the fastener axis of connection covering and longitudinal elements to longitudinal elements
The distance of web, then outer edge strip width meets requires as follows：
wsf=wsf1+wsf2+tw=3Rf+RB+tw+Δ
Wherein：
wsf2For the back gauge value of fastener, wsf2=2Rf+Δ；
RfFor diameter fastener；
twFor longitudinal elements web thickness；
RBFor longitudinal elements bending radius；
Δ is tolerance, can take 1mm；
3.3) longitudinal elements outer edge strip thickness tsf：Longitudinal elements outer edge strip thickness needs to meet following require：
0.7tp＜ tsf
Wherein：
tpFor skin thickness；
tsfFor longitudinal elements outer edge strip thickness；
After skin thickness determines, t is accordingly obtainedsf,min=0.7tp。
4) longitudinal elements section design variable is defined, by designing analysis 4.1) -4.2) in parameter, build constraints,
Determine the feasibility space of design variable；
4.1) longitudinal elements rigidity is checked：
4.1.1) support rigidity of the edge strip to longitudinal elements web in check longitudinal elements；
4.1.2) under compressive load effect, support rigidity of the longitudinal elements to covering is checked；
4.1.3) under shear loads, support rigidity of the longitudinal elements to covering is checked；
4.2) stability analysis is carried out to longitudinal elements unit, calculates critical jitter stress；
5) design variable and constraints of definition scheme design phase, the final feasibility space for building design variable.
If the design variable that conceptual design is chosen meets the design requirements such as intensity, rigidity, manufacture in feasibility spatial dimension, can
To be selected as structural design scheme；
Design variable：
X=[x1,x2,x3,x4]
Wherein：
x1For edge strip width in longitudinal elements；
x2For edge strip thickness in longitudinal elements；
x3For longitudinal elements height；
x4For longitudinal elements web thickness.
Constraints：
Wherein：
IrThe moment of inertia for longitudinal elements with respect to covering neutral axis；
Ir=Istr+Astr·(ycg+tp/2)2
ycgFor longitudinal elements centre of form height；
IstrFor longitudinal elements section moment of inertia；
AstrFor the longitudinal elements area of section；
Astr=x1x2+x3x4+wsftsf
W is adjacent longitudinal elements spacing；
τcrFor covering shear buckling limit stress；
L is adjacent transverse reinforcer spacing；
NzFor aircraft cross section position of form center axle power；
A is the gross area of aircraft cross section；
IxFor the moment of inertia of aircraft cross section；
MxFor aircraft cross section position of form center moment of flexure；
tsfFor longitudinal elements outer edge strip thickness；
wsfFor longitudinal elements outer edge strip width；
tpFor skin thickness；
R is aircraft cross section radius；
EstrFor the modulus of elasticity in comperssion of longitudinal elements material；
EskinFor the modulus of elasticity in comperssion of skin material；
L ' is the effective length of longitudinal elements, and the schematic design phase can take L'=L；
ρ is the longitudinal elements section radius of gyration,
In the step 1), according to the layout of lateral reinforcer and longitudinal elements, aircraft wing and fuselage section are defined
Longitudinal elements is laid out basic configuration.Incorporation engineering beam theory can calculate section position of form center moment Mx, torque Mz, axle power NzWith
Shear Qy；
In the step 3.1), the intensity of covering：
It is required that local buckling does not occur under α times of ultimate load for covering, then covering stress is：
Wherein：
τ is covering stress；
τMThe lower covering shear stress generated is acted on for section torque；
τQFor the covering shear stress generated under section shearing action；
α is the initial flexion limit loading coefficient of covering；
MzFor aircraft cross section position of form center torque；
tpFor skin thickness；
R is aircraft cross section radius；
QyFor aircraft cross section position of form center shearing；
θiFor longitudinal elements and aircraft cross section coordinate system angle；
N is aircraft cross section longitudinal elements sum；
N ' be aircraft cross section it is quiet away from maximum when longitudinal elements number；
The elastic instability limit stress of covering is calculated according to simply supported on four sides tablet buckling analysis method, total according to siding
Body is laid out, and it is L, cell width W that can obtain covering element length.Under shear load, flat plate shear Buckling Critical Load
It is calculated as：
Wherein：
τcrFor flat plate shear critical buckling stress；
ksFor flat plate shear buckling coefficient；
W is adjacent longitudinal elements spacing；
L is adjacent transverse reinforcer spacing；
D is the bending stiffness of tablet；
EskinFor the modulus of elasticity in comperssion of skin material；
υ is the Poisson's ratio of skin material.
Covering shear buckling safety margin is calculated as：
MS=0 is calculated by safety margin, obtains covering elastic instability critical thickness tbuckling。
In the step 3.1), the maintainability of siding connection：
From flight vehicle aerodynamic requirement, the fastener on covering, which uses, immerses oneself in form.And consider from repair angle, according to
The dependency structure size of the big level-one type design connection of fastener.If fastener immerses oneself in depth tcsk, tired in order to improve link position
Labor performance, it is desirable that skin thickness tpMeet following require：
tp≥1.5tcsk
Comprehensive strength and maintainability requirement, can obtain covering critical thickness is：
tp,min=max (1.5tcsk,tbuckling)
Wherein：
tcskDepth is immersed oneself in for fastener；
tbucklingFor covering elastic instability critical thickness.
The invention has the beneficial effect that：The above method analyzes covering, longitudinal elements knot based on thin wall engineering beam theory
Structure stress level is distributed, and determines covering in thin-wall construction, longitudinal elements peak stress.According to thin-wall construction manufacture and maintenance requirement,
Define covering, longitudinal elements partial cross section size.In the case where meeting intensity requirement and rigidity requirement, longitudinal elements section is cooked up
Size can design space, the selection and optimization of implementation.Wing-Body Configurations thin-wall construction conceptual design can be not only completed, and
And realize the lightweight target of design, manufacture, repair one.
Description of the drawings
Fig. 1：1 middle fuselage profile pattern schematic diagram of embodiment.
Fig. 2：Wing profile configuration schematic diagram in embodiment 1.
Fig. 3：Z-shaped longitudinal elements section defines schematic diagram in embodiment 1.
Fig. 4：1 middle fuselage section shearing flow distribution schematic diagram of embodiment.
Fig. 5：Longitudinal elements number schematic diagram in embodiment 1.
Fig. 6：Longitudinal elements fastener connection diagram in embodiment 1.
Fig. 7 a：Space-time function, works as x in embodiment 13Feasibility spatial contrast figure when being 24.
Fig. 7 b：Space-time function, works as x in embodiment 13Feasibility spatial contrast figure when being 26.
Fig. 8 a：In embodiment 1 under two-dimensional coordinate system, design variable x1For 6, x3Feasibility spatial contrast figure when being 26.
Fig. 8 b：In embodiment 1 under two-dimensional coordinate system, design variable x1For 8, x3Feasibility spatial contrast figure when being 26.
Specific implementation mode
Aircraft wing fuselage thin-wall construction original dimension optimum design method, is completed especially by following steps：
1) summarize Wing-Body Configurations thin-wall construction conceptual design section load and basic structure definition：
The total arrangement of known aircraft wing, fuselage plans lateral reinforcer (such as rib, frame) and longitudinal elements (such as stringer)
Position, it is determined that the spacing L of adjacent transverse reinforcer, the spacing W of adjacent longitudinal elements, longitudinal elements quantity n, longitudinal elements position
Deng.Incorporation engineering beam theory can calculate section position of form center moment Mx, torque Mz, axle power NzWith shearing Qy。
Aircraft wing and fuselage section longitudinal elements layout basic configuration are shown in Fig. 1 and Fig. 2, can obtain each indulging on section
To the occupy-place angle of part, the angle of i-th of longitudinal elements is θ in Fig. 1i。
2) according to aircraft cross section definition and section load, calculate covering, longitudinal elements stress, identified sign horizontal distribution and
Peak value.
Flight Vehicle Structure is a complicated thin-wall construction, needs to carry out necessary engineering simplification.In the work of conceptual design
In journey analysis, covering can be born in the ability conversion to the concentration area of the longitudinal elements such as beam, stringer of direct stress, formation is only held
Area in by the combination of sets of direct stress, it is assumed that covering is solely subjected to shear stress.
In the schematic design phase, need to assess longitudinal elements cross-sectional shape, basic size and skin thickness information.For engineering
Simplify and need, can according to etc. rigidity Designs, the z-shape longitudinal elements section of original hypothesis, dimension elements definition is such as Fig. 3.Longitudinal elements
Outer edge strip is connect with covering, and edge strip is inboard edge strip in longitudinal elements.
To thin wall engineering girder construction, section gross area A, the moment of inertia I are calculated separatelyx, as follows：
A=nAstr=n (wff·tff+h·tw+wsf·tsf)
Wherein：
wffFor edge strip width in longitudinal elements；
tffFor edge strip thickness in longitudinal elements；
H is longitudinal elements height；
twFor longitudinal elements web thickness；
wsfFor longitudinal elements outer edge strip width；
tsfFor longitudinal elements outer edge strip thickness；
R is aircraft cross section radius；
AstrFor the longitudinal elements area of section；
θiFor longitudinal elements and aircraft cross section coordinate system angle；
yiFor longitudinal elements and aircraft cross section coordinate system distance；
N is aircraft cross section longitudinal elements sum.
Assuming that section is along a certain coordinate axial symmetry, if Fig. 1 middle fuselage siding sections are symmetrical along Y-axis, load acts on
On line of symmetry, take the notch of single closed chamber section on symmetry axis at this time, then incision shear stress is zero, the shearing flow of single closed chamber section
Equal to section shearing flow is opened, fuselage opens the distribution of section shearing flow and sees Fig. 4.
It is quiet to be happened at thin-walled and X-axis intersection location away from maximum value, it defines n ' and is numbered for adjacent position longitudinal elements occupy-place, see figure
5.
Using shearing flow calculation formula of single closed chamber section under torque effect, i.e. Bredt formula calculate covering shearing flow, in turn
Calculate covering shear stress.
Wherein：
qMThe lower covering shearing flow generated is acted on for section torque；
τMThe lower covering shear stress generated is acted on for section torque；
Ω accumulates for aircraft cross section contour institute envelope surface；
tpFor skin thickness；
MzFor aircraft cross section position of form center torque；
R is aircraft cross section radius.
Q is sheared in sectionyUnder effect, covering shearing flow is calculated.
Wherein：
qQFor the covering shearing flow generated under section shearing action；
τQFor the covering shear stress generated under section shearing action.
Shear stress caused by shear stress caused by torque and shearing is overlapped, then obtaining maximum shear stress τ is：
By section axle power and moment of flexure, the extreme value of longitudinal elements axial stress on section is calculated separately：
Wherein：
σmaxFor longitudinal elements maximum stress；
σminFor longitudinal elements minimum stress；
NzFor aircraft cross section position of form center axle power；
A is the gross area of aircraft cross section；
IxFor the moment of inertia of aircraft cross section；
MxFor aircraft cross section position of form center moment of flexure；
R is aircraft cross section radius.
3) according to design requirements such as aircraft wing, the intensity of fuselage wallboard, manufacture, repairs, part of wall plate section is determined
The minimum dimension of parameter：
3.1) skin thickness tp
In the schematic design phase, the selection of the thickness specification of covering will consider following two factors：
3.1.1) the intensity of covering
In contemporary aircraft design, Wing-Body Configurations siding is designed as thin wall reinforced plate.Due to lateral reinforcer and indulge
Support to part to covering, for covering after elastic instability occurs, thin wall reinforced plate can also transmit load, it requires that covering exists
Local buckling does not occur under certain load, follow-up covering unstability is designed analysis according still further to tension field situation.
Assuming that design requirement is covering does not occur local buckling under α times of ultimate load, then covering stress is：
Wherein：
τ is covering stress；
τMThe lower covering shear stress generated is acted on for section torque；
τQFor the covering shear stress generated under section shearing action；
α is the initial flexion limit loading coefficient of covering；
MzFor aircraft cross section position of form center torque；
tpFor skin thickness；
R is aircraft cross section radius；
QyFor aircraft cross section position of form center shearing；
θiFor longitudinal elements and aircraft cross section coordinate system angle；
N is aircraft cross section longitudinal elements sum；
N ' be aircraft cross section it is quiet away from maximum when longitudinal elements number.
The elastic instability limit stress of covering can be calculated according to simply supported on four sides tablet buckling analysis method.According to wall
Plate total arrangement, it is L, cell width W that can obtain covering element length.Under shear load, flat plate shear buckling is critical
LOAD FOR is：
Wherein：
τcrFor flat plate shear critical buckling stress；
ksFor flat plate shear buckling coefficient；
W is adjacent longitudinal elements spacing；
L is adjacent transverse reinforcer spacing；
D is the bending stiffness of tablet；
EskinFor the modulus of elasticity in comperssion of skin material；
υ is the Poisson's ratio of skin material.
Covering shear buckling safety margin is calculated as：
MS=0 is calculated by safety margin, obtains critical thickness tbuckling。
3.1.2) the maintainability of siding connection
In order to keep the smoothness of aircraft wing fuselage outer surface, reach pneumatic property requirement, the fastener on covering
Using the form of immersing oneself in.And consider from repair angle, according to the dependency structure size of the big level-one type design connection of fastener.If
The connection of covering and longitudinal elements is connected using the fastener of 5/32 model, for subsequent maintenance, then uses the fastening of 6/32 model
Part carries out covering, longitudinal elements size design.
If fastener immerses oneself in depth tcsk, fastener immerses oneself in connect.In order to improve link position fatigue behaviour, it is desirable that covering
Thickness tpMeet following require：
tp≥1.5tcsk
Wherein：
tcskDepth is immersed oneself in for fastener；
tpFor skin thickness.
Comprehensive strength and maintainability requirement, can obtain covering critical thickness is：
tp,min=max (1.5tcsk,tbuckling)
Wherein：
tcskDepth is immersed oneself in for fastener；
tbucklingFor covering elastic instability critical thickness.
3.2) longitudinal elements outer edge strip width wsf；
With reference to figure 6 fastener connect configuration, empirically it is found that from connection covering and longitudinal elements fastener axis to
The distance w of longitudinal elements websf1Following requirement need to be met：
wsf1=Rf+RB
Wherein：
RfFor fastener radius；
RBFor bending radius.
Meanwhile also to meet the requirement of fastener back gauge.In most cases, using 5/32 fastener, it is contemplated that
It repairs, is all designed under normal circumstances using bigger fastener from now on, therefore using corresponding to 6/32 fastener
Back gauge value.Ordinary circumstance, the thickness for repairing angle section are identical as longitudinal elements web thickness.
To sum up, following result can be obtained：
wsf=wsf1+wsf2+tw=3Rf+RB+tw+Δ
Wherein：
wsf2For the back gauge value of fastener, wsf2=2Rf+Δ；
RfFor diameter fastener；
twFor longitudinal elements web thickness；
RBFor longitudinal elements bending radius；
Δ is tolerance, can take 1mm.
3.3) longitudinal elements outer edge strip thickness tsf
To avoid longitudinal elements from being forced to the generation of crushing failure, longitudinal elements outer edge strip thickness needs to meet following require：
0.7tp＜ tsf
Wherein：
tpFor skin thickness；
tsfFor longitudinal elements outer edge strip thickness.
After skin thickness determines, t is accordingly obtainedsf,min=0.7tp。
4) longitudinal elements section design variable X=(x are defined1,x2,x3,x4)=(wff,tff,h,tw).According to design requirement, really
Determine the constraints of structure size.
4.1) longitudinal elements rigidity is checked：
4.1.1) support rigidity of the edge strip to web in check longitudinal elements；
The interior edge strip of longitudinal elements, it is necessary to provide supporting function to longitudinal elements web, upper limb size will meet following requirement：
Wherein：
IffFor edge strip the moment of inertia in longitudinal elements；
AffFor edge strip area in longitudinal elements；
hwFor longitudinal elements web height；
twFor longitudinal elements web thickness.
It, can be by the following form of edge strip deflection constraint conditional definition in longitudinal elements by calculating edge strip the moment of inertia and area：
4.1.2) under compressive load effect, support rigidity of the longitudinal elements to covering is checked；
For covering under compressive load effect, longitudinal elements must play covering certain supporting function, then need to meet such as
Lower requirement：
Wherein：
W is adjacent longitudinal elements spacing；
tpFor skin thickness；
IrThe moment of inertia for longitudinal elements with respect to covering neutral axis；
Ir=Istr+Astr·(ycg+tp/2)2
ycgFor longitudinal elements section centre of form height；
tsfFor longitudinal elements outer edge strip thickness；
wsfFor longitudinal elements outer edge strip width；
IstrFor longitudinal elements section moment of inertia；
AstrFor the longitudinal elements area of section.
Astr=x1x2+x3x4+wsftsf
4.1.3) under shear loads, support rigidity of the longitudinal elements to covering is checked；
For covering under the action of shear-type load, longitudinal elements must play covering certain supporting function, then need to meet
It is following to require：
Wherein：
IrThe moment of inertia for longitudinal elements with respect to covering neutral axis；
W is adjacent longitudinal elements spacing；
L is adjacent transverse reinforcer spacing；
tpFor skin thickness；
EskinFor the modulus of elasticity in comperssion of skin material；
τcrFor flat plate shear critical buckling stress.
4.2) stability analysis is carried out to longitudinal elements unit, calculates critical jitter stress.
Squeeze, bend proximate matter axial compression load effect under be easy to happen general instability and local buckling.According to Euler
Formula carries out stable calculation.Compress limit stress σcrAccording to Calculation Using Euler Equations, method is as follows：
Wherein：
EstrFor the modulus of elasticity in comperssion of longitudinal elements material；
L ' is the effective length of longitudinal elements, and the schematic design phase can take L'=L；
ρ is the longitudinal elements section radius of gyration,
To meet Static Strength Design requirement, longitudinal elements general instability safety margin should be more than 0.
Wherein：
σcrLimit stress is compressed for longitudinal elements；
σminFor longitudinal elements minimum stress；
NzFor aircraft cross section position of form center axle power；
A is the gross area of aircraft cross section；
IxFor the moment of inertia of aircraft cross section；
MxFor aircraft cross section position of form center moment of flexure；
R is aircraft cross section radius.
5) design variable and constraints of definition scheme design phase, the final feasibility space for building design variable.
The value of design variable in feasibility space all meets the design requirements such as intensity, rigidity, manufacture, can be selected as structure design side
Case；Organization plan is then chosen as such as in feasibility spatial dimension for arbitrary longitudinal elements sectional dimension, it can be convenient and efficient
Ground judges the reasonability of scheme.
Design variable：
X=[x1,x2,x3,x4]
Constraints：
Wherein：
IrThe moment of inertia for longitudinal elements with respect to covering neutral axis；
Ir=Istr+Astr×(ycg+tp/2)2
ycgFor longitudinal elements centre of form height；
IstrFor longitudinal elements section moment of inertia；
AstrFor the longitudinal elements area of section；
Astr=x1x2+x3x4+wsftsf
W is adjacent longitudinal elements spacing；
τcrFor covering shear buckling limit stress；
L is adjacent transverse reinforcer spacing；
tpFor skin thickness；
tsfFor longitudinal elements outer edge strip thickness；
wsfFor longitudinal elements outer edge strip width；
NzFor aircraft cross section position of form center axle power；
A is the gross area of aircraft cross section；
IxFor the moment of inertia of aircraft cross section；
MxFor aircraft cross section position of form center moment of flexure；
R is aircraft cross section radius；
EstrFor the modulus of elasticity in comperssion of longitudinal elements material；
EskinFor the modulus of elasticity in comperssion of skin material；
L ' is the effective length of longitudinal elements, and the schematic design phase can take L'=L；
ρ is the longitudinal elements section radius of gyration,
Four constraintss can limit the value range of four design variables, in the four-dimensional feasibility function space of structure
Numerical value be all the longitudinal elements structure size that can be chosen the schematic design phase.Feasibility space according to aircraft cross section and can be cutd open
Face load change and change.Longitudinal elements height directly affects lateral reinforcer openings of sizes, it is possible to tentatively to longitudinal elements height
Degree, i.e. design variable x3It is defined.By Fig. 7 a and Fig. 7 b, a certain aircraft cross section is given under one group of load, when setting
Count variable x3The comparison in three-dimensional feasibility space when variation；And work as x3One timing, passes through x1The selection of variation can obtain not
It is selected for conceptual design with two dimension feasibility range, sees Fig. 8 a and Fig. 8 b.
Claims (4)
1. aircraft wing fuselage thin-wall construction original dimension optimum design method, which is characterized in that its step is：
1) according to aircraft wing, the total arrangement of fuselage, the position of lateral reinforcer and longitudinal elements is cooked up, calculates flight
Device wing, the section load of fuselage thin-wall construction conceptual design and basic structure definition；
2) according to aircraft cross section definition and section load, covering, longitudinal elements stress, identified sign horizontal distribution and peak are calculated
Value；
3) according to aircraft wing, the intensity of fuselage wallboard, manufacture and maintenance requirement, the minimum of part of wall plate cross section parameter is determined
Size, including：
3.1) skin thickness tp：It is determined according to the maintainability that the intensity of covering is connected with siding；
3.2) longitudinal elements outer edge strip width wsf：Define wsf1For the fastener axis of connection covering and longitudinal elements to longitudinal elements web
Distance, then outer edge strip width wsfMeet following require：
wsf=wsf1+wsf2+tw=3Rf+RB+tw+Δ
Wherein：
wsf2For the back gauge value of fastener, wsf2=2Rf+Δ；
RfFor diameter fastener；
twFor longitudinal elements web thickness；
RBFor longitudinal elements bending radius；
Δ is tolerance, can take 1mm；
3.3) longitudinal elements outer edge strip thickness tsf：Longitudinal elements outer edge strip thickness needs to meet following require：
0.7tp＜ tsf
Wherein：
tpFor skin thickness；
tsfFor longitudinal elements outer edge strip thickness.
After skin thickness determines, t is accordingly obtainedsf,min=0.7tp；
4) longitudinal elements section design variable is defined, by designing analysis 4.1) -4.2) in parameter, build constraints, determine
The feasibility space of design variable；
4.1) longitudinal elements rigidity is checked：
4.1.1) support rigidity of the edge strip to longitudinal elements web in check longitudinal elements；
4.1.2) under compressive load effect, support rigidity of the longitudinal elements to covering is checked；
4.1.3) under shear loads, support rigidity of the longitudinal elements to covering is checked；
4.2) stability analysis is carried out to longitudinal elements unit, calculates critical jitter stress；
5) design variable and constraints of definition scheme design phase, the final feasibility space for building design variable.If side
The design variable that case design is chosen then meets the design requirements such as intensity, rigidity, manufacture, Ke Yixuan in feasibility spatial dimension
For structural design scheme；
Design variable：
X=[x1,x2,x3,x4]
Wherein：
x1For edge strip width in longitudinal elements；
x2For edge strip thickness in longitudinal elements；
x3For longitudinal elements height；
x4For longitudinal elements web thickness.
Constraints：
Wherein：
IrThe moment of inertia for longitudinal elements with respect to covering neutral axis；
Ir=Istr+Astr·(ycg+tp/2)2
ycgFor longitudinal elements centre of form height；
IstrFor longitudinal elements section moment of inertia；
AstrFor the longitudinal elements area of section；
Astr=x1x2+x3x4+wsftsf
W is adjacent longitudinal elements spacing；
τcrFor covering shear buckling limit stress；
L is adjacent transverse reinforcer spacing；
NzFor aircraft cross section position of form center axle power；
tpFor skin thickness；
tsfFor longitudinal elements outer edge strip thickness；
wsfFor longitudinal elements outer edge strip width；
A is the gross area of aircraft cross section；
IxFor the moment of inertia of aircraft cross section；
MxFor aircraft cross section position of form center moment of flexure；
R is aircraft cross section radius；
EstrFor the modulus of elasticity in comperssion of longitudinal elements material；
EskinFor the modulus of elasticity in comperssion of skin material；
L ' is the effective length of longitudinal elements, and the schematic design phase can take L'=L；
ρ is the longitudinal elements section radius of gyration,
2. aircraft wing fuselage thin-wall construction original dimension optimum design method according to claim 1, feature exist
In：In the step 1), according to the layout of lateral reinforcer and longitudinal elements, aircraft wing and fuselage section longitudinal elements are defined
It is laid out basic configuration.Incorporation engineering beam theory can calculate section position of form center moment Mx, torque Mz, axle power NzAnd shearing
Qy。
3. aircraft wing fuselage thin-wall construction original dimension optimum design method according to claim 1, feature exist
In：In the step 3.1), the intensity of covering：
It is required that local buckling does not occur under α times of ultimate load for covering, then covering stress is：
Wherein：
τ is covering stress；
τMThe lower covering shear stress generated is acted on for section torque；
τQFor the covering shear stress generated under section shearing action；
α is the initial flexion limit loading coefficient of covering；
MzFor aircraft cross section position of form center torque；
QyFor aircraft cross section position of form center shearing；
tpFor skin thickness；
R is aircraft cross section radius；
θiFor longitudinal elements and aircraft cross section coordinate system angle；
N is aircraft cross section longitudinal elements sum；
N ' be aircraft cross section it is quiet away from maximum when longitudinal elements number；
The elastic instability limit stress of covering is calculated according to simply supported on four sides tablet buckling analysis method, according to siding totality cloth
Office, it is L, cell width W that can obtain covering element length.Under shear load, flat plate shear Buckling Critical Load calculates
For：
Wherein：
τcrFor flat plate shear critical buckling stress；
ksFor flat plate shear buckling coefficient；
W is adjacent longitudinal elements spacing；
L is adjacent transverse reinforcer spacing；
D is the bending stiffness of tablet；
EskinFor the modulus of elasticity in comperssion of skin material；
υ is the Poisson's ratio of skin material.
Covering shear buckling safety margin is calculated as：
MS=0 is calculated by safety margin, obtains critical thickness tbuckling。
4. aircraft wing fuselage thin-wall construction original dimension optimum design method according to claim 1, feature exist
In：In the step 3.1), the maintainability of siding connection：
From flight vehicle aerodynamic requirement, the fastener on covering, which uses, immerses oneself in form.And consider from repair angle, according to fastening
The dependency structure size of the big level-one type design connection of part.If fastener immerses oneself in depth tcsk, in order to improve link position fatigability
It can, it is desirable that skin thickness tpMeet following require：
tp≥1.5tcsk
Comprehensive strength and maintainability requirement, can obtain covering critical thickness is：
tp,min=max (1.5tcsk,tbuckling)
Wherein：
tcskDepth is immersed oneself in for fastener；
tbucklingFor covering elastic instability critical thickness.
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