University of Liège
Aerospace & Mechanical Engineering
Aircraft Structures
Instabilities
Aircraft Structures - Instabilities
Ludovic Noels
Computational & Multiscale Mechanics of Materials β CM3
http://www.ltas-cm3.ulg.ac.be/
Chemin des Chevreuils 1, B4000 Liège
Elasticity
β’ Balance of body Bβ Momenta balance
β’ Linear
β’ Angular
β Boundary conditionsβ’ Neumann
β’ Dirichlet
β’ Small deformations with linear elastic, homogeneous & isotropic material
β (Small) Strain tensor , or
β Hookeβs law , or
with
β Inverse law
with
b
T
n
2ml = K - 2m/3
2013-2014 Aircraft Structures - Instabilities 2
β’ General expression for unsymmetrical beams
β Stress
With
β Curvature
β In the principal axes Iyz = 0
β’ Euler-Bernoulli equation in the principal axis
β for x in [0 L]
β BCs
β Similar equations for uy
Pure bending: linear elasticity summary
x
z f(x) TzMxx
uz =0
duz /dx =0 M>0
L
y
zq
Mxx
a
2013-2014 Aircraft Structures - Instabilities 3
β’ General relationships
β
β’ Two problems considered
β Thick symmetrical section
β’ Shear stresses are small compared to bending stresses if h/L << 1
β Thin-walled (unsymmetrical) sections
β’ Shear stresses are not small compared to bending stresses
β’ Deflection mainly results from bending stresses
β’ 2 cases
β Open thin-walled sections
Β» Shear = shearing through the shear center + torque
β Closed thin-walled sections
Β» Twist due to shear has the same expression as torsion
Beam shearing: linear elasticity summary
x
z f(x) TzMxx
uz =0
duz /dx =0 M>0
L
h
L
L
h t
2013-2014 Aircraft Structures - Instabilities 4
β’ Shearing of symmetrical thick-section beams
β Stress
β’ With
β’ Accurate only if h > b
β Energetically consistent averaged shear strain
β’ with
β’ Shear center on symmetry axes
β Timoshenko equations
β’ &
β’ On [0 L]:
Beam shearing: linear elasticity summary
ht
z
y
z
t
b(z)
A*
t
h
x
z
Tz
dx
Tz+ βxTz dx
gmax
gg dx
z
x
g
qy
qy
2013-2014 Aircraft Structures - Instabilities 5
β’ Shearing of open thin-walled section beams
β Shear flow
β’
β’ In the principal axes
β Shear center S
β’ On symmetry axes
β’ At walls intersection
β’ Determined by momentum balance
β Shear loads correspond to
β’ Shear loads passing through the shear center &
β’ Torque
Beam shearing: linear elasticity summary
x
z
y
Tz
Tz
Ty
Ty
y
z
S
Tz
TyC
q
s
y
t
t
h
b
z
C
t
S
2013-2014 Aircraft Structures - Instabilities 6
β’ Shearing of closed thin-walled section beams
β Shear flow
β’
β’ Open part (for anticlockwise of q, s)
β’ Constant twist part
β’ The q(0) is related to the closed part of the section,
but there is a qo(s) in the open part which should be
considered for the shear torque
Beam shearing: linear elasticity summary
y
z
T
Tz
Ty
C
q s
pds
dAh
y
z
T
Tz
Ty
C
q s
p
2013-2104 Aircraft Structures - Instabilities 7
β’ Shearing of closed thin-walled section beams
β Warping
β’
β’ With
β ux(0)=0 for symmetrical section if origin on
the symmetry axis
β Shear center S
β’ Compute q for shear passing thought S
β’ Use
Beam shearing: linear elasticity summary
y
z
T
Tz
Ty
C
q s
pds
dAh
y
z
S
Tz
C
q s
p ds
2013-2014 Aircraft Structures - Instabilities 8
β’ Torsion of symmetrical thick-section beams
β Circular section
β’
β’
β Rectangular section
β’
β’
β’ If h >> b
β &
β
β
Beam torsion: linear elasticity summary
t
z
y
C
Mx
r
h/b 1 1.5 2 4 β
a 0.208 0.231 0.246 0.282 1/3
b 0.141 0.196 0.229 0.281 1/3
z
y
C
tmaxMx
b
h
2013-2014 Aircraft Structures - Instabilities 9
β’ Torsion of open thin-walled section beams
β Approximated solution for twist rate
β’ Thin curved section
β
β
β’ Rectangles
β
β
β Warping of s-axis
β’
Beam torsion: linear elasticity summary
y
z
l2
t2
l1
t1
l3
t3
t
t
z
y
C
Mx
t
ns
t
R
pR
us
q
2013-2014 Aircraft Structures - Instabilities 10
β’ Torsion of closed thin-walled section beams
β Shear flow due to torsion
β Rate of twist
β’
β’ Torsion rigidity for constant m
β Warping due to torsion
β’
β’ ARp from twist center
Beam torsion: linear elasticity summary
y
z
C
q s
pds
dAh
Mx
y
z
R
C p
pRY
us
q
2013-2014 Aircraft Structures - Instabilities 11
β’ Panel idealization
β Boomsβ area depending on loading
β’ For linear direct stress distribution
Structure idealization summary
b
sxx1
sxx2
A1 A2
y
z
x
tD
b
y
z
xsxx
1sxx
2
2013-2014 Aircraft Structures - Instabilities 12
β’ Consequence on bending
β If Direct stress due to bending is carried by booms only
β’ The position of the neutral axis, and thus the second moments of area
β Refer to the direct stress carrying area only
β Depend on the loading case only
β’ Consequence on shearing
β Open part of the shear flux
β’ Shear flux for open sections
β’ Consequence on torsion
β If no axial constraint
β’ Torsion analysis does not involve axial stress
β’ So torsion is unaffected by the structural idealization
Structure idealization summary
Tz
y
z
x
dx
Ty
2013-2014 Aircraft Structures - Instabilities 13
β’ Virtual displacement
β In linear elasticity the general formula of virtual displacement reads
β’ s (1) is the stress distribution corresponding to a (unit) load P(1)
β’ DP is the energetically conjugated displacement to P in the direction of P(1) that
corresponds to the strain distribution e
β Example bending of semi cantilever beam
β’
β In the principal axes
β Example shearing of semi-cantilever beam
β’
Deflection of open and closed section beams summary
x
z Tz
uz =0
duz /dx =0 M>0
L
2013-2014 Aircraft Structures - Instabilities 14
β’ Torsion of a built-in end closed-section beam
β If warping is constrained (built-in end)
β’ Direct stresses are introduced
β’ Different shear stress distribution
β Example: square idealized section
β’ Warping
β’ Shear stress
Structural discontinuities summary
z
y
Mxx
L
b
h
dx
qb
qh
uxm
2013-2014 Aircraft Structures - Instabilities 15
β’ Shear lag of a built-in end closed-section beam
β Beam shearing
β’ Shear strain in cross-section
β’ Deformation of cross-section
β’ Elementary theory of bending
β For pure bending
β Not valid anymore because of the
cross section deformation
β Example
β’ 6-boom wing
β’ Deformation of top cover
Structural discontinuities summaryz
y
x
L
d
h
Tz/2d
A1 A2A1 qh
qd
Tz/2
d
y
x
d
2013-2014 Aircraft Structures - Instabilities 16
β’ Torsion of a built-end open-section beam
β If warping is constrained (built-in end)
β’ Direct stresses are introduced
β’ There is a bending contribution
to the torque
β Examples
β’ Equation for pure torque
with
β’ Equation for distributed torque
Structural discontinuities summary
Mx Mx
Mx
x
z
ymx
Mx+βxMx dx
dx
Mx
2013-2014 Aircraft Structures - Instabilities 17
β’ 2 kinds of buckling
β Primary buckling
β’ No changes in cross-section
β’ Wavelength of buckle ~ length of element
β’ Solid & thick-walled column
β Secondary (local) buckling
β’ Changes in cross-section
β’ Wavelength of buckle ~ cross-sectional dimensions
β’ Thin-walled column & stiffened panels
β Pictures:β D.H. Dove wing (max loading test)β Automotive beam β Local buckling
Column instabilities
2013-2014 Aircraft Structures - Instabilities 18
β’ Assumptions
β Perfectly symmetrical column (no imperfection)
β Axial load perfectly aligned along centroidal axis
β Linear elasticity
β’ Theoretically
β Deformed structure should remain symmetrical
β Solution is then a shortening of the column
β Buckling load PCR is defined as P such that if a
small lateral displacement is enforced by a
lateral force, once this force is removed
β’ If P = PCR, the lateral deformation is constant
(neutral stability)
β’ If P > PCR, this lateral displacement increases &
the column is unstable
β’ If P < PCR, this lateral displacement disappears &
the column is stable
β’ Practically
β The initial lateral displacement is due to
imperfections (geometrical or material)
Euler buckling
P
P P
F
P
P < PCR
P
P > PCR
2013-2014 Aircraft Structures - Instabilities 19
β’ Euler critical axial load
β Bending theory
β’
β Solution
β’ General form: with
β’ BCs at x = 0 & x = L imply C1 = 0 &
β’ Non trivial solution with k = 1, 2, 3, β¦
β’ In that case C2 is undetermined and can β β
β Euler critical load for pinned-pinned BCs
β’ with k = 1, 2, 3, β¦
Euler buckling
Puz
x
z
2013-2014 Aircraft Structures - Instabilities 20
β’ Euler critical axial load (2)
β Euler critical load for pinned-pinned BCs (2)
β’ with k = 1, 2, 3, β¦
β’ Buckling will occur for lowest PCR
β In the plane of lowest I
β For the lowest k k = 1
β In case modes 1, .. k-1 are prevented, critical load becomes the load k
Euler buckling
Puz
x
z
Px
z
Px
z
2013-2014 Aircraft Structures - Instabilities 21
β’ Euler critical axial load (3)
β For pinned-pinned BCs
β’
β’ (compression) with gyration radius
β General case
β’ Euler critical loads , (compressive)
β’ With le the effective length
Euler buckling
Puz
x
z
P
le = L/2
le = L/2
P
le = 2L
P
2013-2014 Aircraft Structures - Instabilities 22
β’ Practical case: initial imperfection
β Let us assume an initial small curvature of the beam
β’
β As this curvature is small the equation of bending
for straight beam can still be used, but with the
change of curvature being considered for the strain
β’
β’ The general form of the initial deflection satisfying the BCs is
the deflection equation becomes
β Solution
β’ With
Initial imperfection
uz0
x
z
Puz
x
z
2013-2014 Aircraft Structures - Instabilities 23
β’ Practical case: initial imperfection (2)
β Solution for an initial small curvature of the beam
β’
β’ With
β’ BCs at x = 0 & x = L imply C1 = C2 = 0, & as
β’ Clearly near buckling, so PβPCR, and the dominant term of the solution is for n = 1
Initial imperfection
uz0
x
z
Puz
x
z
2013-2014 Aircraft Structures - Instabilities 24
β’ Practical case: initial imperfection (3)
β Solution for an initial small curvature of the beam
β’ near buckling
β If central deflection is measured vs axial load
β’ As uz0(L/2) ~ A1
β’
β Southwell diagram
β’
β’ Allows measuring buckling loads without
breaking the columns
β’ Remark
β Critical Euler loading depends on BCs
Initial imperfection
uz0
x
z
Puz
x
z
D
D/P
A1
1/PCR
1
2013-2014 Aircraft Structures - Instabilities 25
β’ Thin-walled column under critical flexural loads
β Can twist without bending or
β Can twist and bend simultaneously
Flexural-torsional buckling of open thin-walled columns
2013-2014 Aircraft Structures - Instabilities 26
β’ Kinematics
β Consider
β’ A thin-walled section
β’ Centroid C
β’ Cyz principal axes
β’ Shear center S
β Section motion (CSRD)
β’ Translation
β Shear center is moved
β By uyS & uz
S
β To Sββ
β’ Rotation around shear (twist) center
β We assume shear center=twist center
β By q
β’ Centroid motion
β To Cβ after section translation
β To Cββ after rotation
β Resulting displacements uyC & uz
C
β Same decomposition for other
points of the section
Flexural-torsional buckling of open thin-walled columns
y
z
S
C
q
Cβ
uyS
uzS
uyS
uzS
Sββ
Cββ
a
a
uyC
uzC
y
z
S
C
q
Cβ
uyS
uzS
uyS
uzS
Sββ
Cββ
2013-2014 Aircraft Structures - Instabilities 27
β’ Kinematics (2)
β Relations
β’ Centroid
β’ Other points P of the section
β’ Considering axial loading
β If q remains small, the induced momentums are
β As we are in the principal axes (Iyz=0), and
as motion resulting from bending is uS
Flexural-torsional buckling of open thin-walled columns
Puz
x
z
Puy
x
y
y
z
S
C
q
Cβ
uyS
uzS
uyS
uzS
Sββ
Cββ
a
a
uyC
uzC
P
2013-2014 Aircraft Structures - Instabilities 28
β’ Torsion
β Any point P of the section
β As torsion results from axial loading,
this corresponds to a torque with
warping constraint
β’ See previous lecture
β Analogy between
beam bending/pin-ended column
β As
β The momentum at point P can be substituted by lateral loading
β Contributions on ds
Flexural-torsional buckling of open thin-walled columns
y
z
S
C
q
uyS
uzS
Sββ
CββP Pββ
ds
dfz
dfy
2013-2014 Aircraft Structures - Instabilities 29
β’ Torsion (2)
β Lateral loading analogy
β’ Contributions on ds
β’ As axial load leads to uniform
compressive stress on section
of area A
β’ Resulting distributed torque (per unit length) on ds
β
Flexural-torsional buckling of open thin-walled columns
y
z
S
C
q
Mxx
uzS
uyS
uzS
Sββ
CββP Pββ
ds
dfz
dfy
2013-2014 Aircraft Structures - Instabilities 30
β’ Distributed torque
β As
β As C is the centroid
Flexural-torsional buckling of open thin-walled columns
2013-2014 Aircraft Structures - Instabilities 31
β’ Distributed torque (2)
β The analogous torque by unit length resulting from the bending reads
β Polar second moment of area around S:
β For a built-in end open-section beam
β’ Warping is constrained
β Bending contribution to the torque
β’ New equation
β’ For a constant section
Flexural-torsional buckling of open thin-walled columns
Mx
x
z
ymx
Mx+βxMx dx
dx
Mx
2013-2014 Aircraft Structures - Instabilities 32
β’ Equations
β In the principal axes
β’
β’
β’
Flexural-torsional buckling of open thin-walled columns
2013-2014 Aircraft Structures - Instabilities 33
β’ Example
β Column with
β’ Deflection and rotation around x
constrained at both ends
β uy(0) = uy(L) = 0 & uz(0) = uz(L) = 0
β q(0) = 0 & q(L) = 0
β’ Warping and rotation around y & z allowed at both ends
β Twist center = shear center
β My(0) = My(L) = 0 uy,xx(0) = uy,xx(L) = 0
β Mz(0) = Mz(L) = 0 uz,xx(0) = uz,xx(L) = 0
β As warping is allowed
are equal to zero
q,xx(0) = 0 & q,xx(L) = 0
Flexural-torsional buckling of open thin-walled columns
x
z
y
P
L
2013-2014 Aircraft Structures - Instabilities 34
β’ Resolution
β Assuming the following fields satisfying the BCs
β’
β’
β’
β The system of equations
Flexural-torsional buckling of open thin-walled columns
x
z
y
P
L
2013-2014 Aircraft Structures - Instabilities 35
β’ Resolution (2)
β Non trivial solution leads to buckling load P
β’ Buckling load is the minimum root
Flexural-torsional buckling of open thin-walled columns
x
z
y
P
L
2013-2014 Aircraft Structures - Instabilities 36
β’ If shear center and centroid coincide
β System becomes
β This system is uncoupled and leads to 3 critical loads
β Buckling load is the minimum value
Flexural-torsional buckling of open thin-walled columns
x
z
y
P
L
2013-2014 Aircraft Structures - Instabilities 37
β’ Example
β Column
β’ Length: L = 1 m
β’ Young: E = 70 GPa
β’ Shear modulus: m = 30 GPa
β Buckling load?
β’ Deflection and rotation around x
constrained at both ends
β uy(0) = uy(L) = 0 & uz(0) = uz(L) = 0
β q(0) = 0 & q(L) = 0
β’ Warping and rotation around y & z allowed
at both ends
Flexural-torsional buckling of open thin-walled columns
y
t = 2 mm
h=
10
0 m
m
b = 100 mm
z
yβ
zβ
C
t = 2 mm
t = 2 mm
S(y,0 ) Oβ
2013-2014 Aircraft Structures - Instabilities 38
β’ Centroid position
β
β By symmetry on Oy
β’ Second moment of area
β
β
β By symmetry
Flexural-torsional buckling of open thin-walled columns
y
t = 2 mm
h=
10
0 m
m
b = 100 mm
z
yβ
zβ
C
t = 2 mm
t = 2 mm
S(yS,0 ) Oβ
2013-2014 Aircraft Structures - Instabilities 39
β’ Shear center
β On Cy by symmetry
β Consider shear force Tz
β’ As Iyz = 0
β’ Lower flange, considering frame Oβyβzβ
β’ Upper flange by symmetry
β’ As Tz passes through the shear center: no torsional flux
Flexural-torsional buckling of open thin-walled columns
y
t = 2 mm
h=
10
0 m
m
b = 100 mm
z
yβ
zβ
C
t = 2 mm
t = 2 mm
S(yS, 0) Oβ
s
Tz
q
q
q
MOβ
2013-2014 Aircraft Structures - Instabilities 40
β’ Uncoupled critical loads
β Using following definitions
β’ These values would be the critical loads
for an uncoupled system (if C = S)
?
Some values are missing
Flexural-torsional buckling of open thin-walled columns
y
t = 2 mm
h=
10
0 m
m
b = 100 mm
z
yβ
zβ
C
t = 2 mm
t = 2 mm
S(yS,0 ) Oβ
2013-2014 Aircraft Structures - Instabilities 41
β’ Uncoupled critical loads (2)
β
β’
β’
β’
β’
β Requires ARp(s)
Flexural-torsional buckling of open thin-walled columns
y
t = 2 mm
h=
10
0 m
m
b = 100 mm
z
yβ
zβ
C
t = 2 mm
t = 2 mm
S(yS, 0) Oβ
s
Tz
q
q
q
2013-2014 Aircraft Structures - Instabilities 42
β’ Uncoupled critical loads (3)
β Evaluation of the ARp(s)
β’ Lower flange
β’ Web
β’ Upper flange
Flexural-torsional buckling of open thin-walled columns
y
t = 2 mm
h=
10
0 m
m
b = 100 mm
z
yβ
zβ
C
t = 2 mm
t = 2 mm
S(yS, 0) Oβ
+
<0
+
<0>0
y
t = 2 mm
h=
10
0 m
m
b = 100 mm
z
yβ
zβ
C
t = 2 mm
t = 2 mm
S(yS, 0) Oβ
+
<0>0
y
t = 2 mm
h=
10
0 m
m
b = 100 mm
z
yβ
zβ
C
t = 2 mm
t = 2 mm
S(yS, 0) Oβ
<0
2013-2014 Aircraft Structures - Instabilities 43
β’ Uncoupled critical loads (4)
β
β’ Lower flange:
β
β
β
β
Flexural-torsional buckling of open thin-walled columns
y
t = 2 mm
h=
10
0 m
m
b = 100 mm
z
yβ
zβ
C
t = 2 mm
t = 2 mm
S(yS, 0) Oβ
s
Tz
q
q
q
2013-2014 Aircraft Structures - Instabilities 44
β’ Uncoupled critical loads (5)
β
β’ Web:
β
β
β
β
Flexural-torsional buckling of open thin-walled columns
y
t = 2 mm
h=
10
0 m
m
b = 100 mm
z
yβ
zβ
C
t = 2 mm
t = 2 mm
S(yS, 0) Oβ
s
Tz
q
q
q
2013-2014 Aircraft Structures - Instabilities 45
β’ Uncoupled critical loads (6)
β
β’ Upper flange:
β
Flexural-torsional buckling of open thin-walled columns
y
t = 2 mm
h=
10
0 m
m
b = 100 mm
z
yβ
zβ
C
t = 2 mm
t = 2 mm
S(yS, 0) Oβ
s
Tz
q
q
q
2013-2014 Aircraft Structures - Instabilities 46
β’ Uncoupled critical loads (7)
β
β’ Upper flange (2):
β
β
Flexural-torsional buckling of open thin-walled columns
y
t = 2 mm
h=
10
0 m
m
b = 100 mm
z
yβ
zβ
C
t = 2 mm
t = 2 mm
S(yS, 0) Oβ
s
Tz
q
q
q
2013-2014 Aircraft Structures - Instabilities 47
β’ Uncoupled critical loads (8)
β
β’ All contributions
β
β
β
Flexural-torsional buckling of open thin-walled columns
2013-2014 Aircraft Structures - Instabilities 48
β’ Uncoupled critical loads (9)
β
β’
β’
β’
β’
Flexural-torsional buckling of open thin-walled columns
y
t = 2 mm
h=
10
0 m
m
b = 100 mm
z
yβ
zβ
C
t = 2 mm
t = 2 mm
S(yS, 0) Oβ
s
Tz
q
q
q
2013-2014 Aircraft Structures - Instabilities 49
β’ Critical load β As the uncoupled critical loads read
, &
& as zS = 0, the coupled system is rewritten
Flexural-torsional buckling of open thin-walled columns
y
t = 2 mm
h=
10
0 m
m
b = 100 mm
z
yβ
zβ
C
t = 2 mm
t = 2 mm
S(yS, 0) Oβ
s
Tz
q
q
q
2013-2014 Aircraft Structures - Instabilities 50
β’ Critical load (2)
β Resolution
Flexural-torsional buckling of open thin-walled columns
>
2013-2014 Aircraft Structures - Instabilities 51
β’ Thin plates
β Are subject to primary buckling
β’ Wavelength of buckle ~
length of element
β So they are stiffened
Buckling of thin plates
2013-2014 Aircraft Structures - Instabilities 52
β’ Primary buckling of thin plates
β Plates without support
β’ Similar to column buckling
β Same shape
β Use D instead of EIzz
β Supported plates
β’ Other displacement buckling shapes
β’ Depend on BCs
Buckling of thin plates
p
b a
E1
E2
E3
A
f
f
b a
E1
E2
E3
A
f
f
0
0.5
1
0
0.5
10
2
4
x 10-3
x/ay/a
u3 D
/(p
0 a
4)
2013-2014 Aircraft Structures - Instabilities 53
β’ Kirchhoff-Love membrane mode
β On A:
β’ With
β Completed by appropriate BCs
β’ Dirichlet
β’ Neumann
Buckling of thin plates
βNA
n
n0 = na EaE1
E2
E3
AβDA
2013-2014 Aircraft Structures - Instabilities 54
β’ Kirchhoff-Love bending mode
β On A:
β’ With
β Completed by appropriate BCs
β’ Low order
β On βNA:
β On βDA:
β’ High order
β On βTA:
with
β On βMA:
Buckling of thin plates
βNA
T
n0 = na EaE1
E2
E3
AβDA
p
βMA
M
n0 = na EaE1
E2
E3
AβTA
p
D
2013-2014 Aircraft Structures - Instabilities 55
β’ Membrane-bending coupling
β The first order theory is uncoupled
β For second order theory
β’ On A:
β’ Tension increases the bending
stiffness of the plate
β’ Internal energy
β In case of small initial curvature (k >>)
β’ On A:
β’ Tension induces bending effect
Buckling of thin plates
Γ±11E1
E2
E3
A
Γ±22
Γ±12
Γ±21
E1
E2
E3
A
Γ±22
Γ±12
Γ±21
j03
u3
Γ±11
2013-2014 Aircraft Structures - Instabilities 56
β’ Primary buckling theory of thin plates
β Second order theory
β’ On A:
β Simply supported plate
with arbitrary pressure
β’ Pressure is written in a Fourier series
β’ Displacements with these BCs can also be written
with
β’ There is a buckling load Γ±11 leading to
infinite displacements for every couple (m, n)
β Lowest one?
Buckling of thin plates
Γ±11E1
E2
E3
A
Γ±22
Γ±12
Γ±21
p
b a
E1
E2
E3
A
Γ±11
Γ±11
2013-2014 Aircraft Structures - Instabilities 57
β’ Primary buckling theory of thin plates (2)
β Simply supported plate
β’ Displacements in terms of
β’ Buckling load Γ±11
β Minimal (in absolute value) for n=1
β Or again
with the buckling coefficient k
β Depends on ration a/b
Buckling of thin plates
p
b a
E1
E2
E3
A
Γ±11
Γ±11
2013-2014 Aircraft Structures - Instabilities 58
β’ Primary buckling theory of thin plates (3)
β Simply supported plate (2)
β’ Buckling coefficient k
β’ Mode of buckling depends on a/b
β’ k is minimal (=4) for a/b = 1, 2, 3, β¦
β’ Mode transition for
β’ For a/b > 3: k ~ 4
β This analysis depends on the BCs, but same behaviors for
β’ Other BCs
β’ Other loadings (bending, shearing) instead of compression
β’ Only the value of k is changing (tables)
Buckling of thin plates
2013-2014 Aircraft Structures - Instabilities 59
0 2 4 60
2
4
6
8
10
20.5
60.5
120.5
a/b
k
m=1
m=2
m=3
m=4
π/π
π
β’ Primary buckling theory of thin plates (4)
β Shape of the modes for
β’ Simply supported plate
β’ In compression
β’ n=1
Buckling of thin plates
0
0.5
1
0
0.5
1-1
0
1
x/ay/a
u3m
=1
0
0.5
1
0
0.5
1-1
0
1
x/ay/a
u3m
=2
2013-2014 Aircraft Structures - Instabilities 60
β’ Primary buckling theory of thin plates (5)
β Critical stress
β’ We found
β’ Or again
β’ This can be generalized to other loading
cases with k depending on the problem
β Picture for simply supported plate in
compression
β As
β’
β’ k ~ cst for a/b >3
We use stiffeners to reduce b
to increase sCR of the skin
β’ As long as sCR < sp0
Buckling of thin plates
b
2013-2014 Aircraft Structures - Instabilities 61
0 2 4 60
2
4
6
8
10
20.5
60.5
120.5
a/b
k
m=1
m=2
m=3
m=4
π/π
π
β’ Primary buckling theory of thin plates (6)
β What happens for other BCs?
β’ We cannot say anymore
β’ But buckling corresponds to a stationary point of the internal energy
(neutral equilibrium)
β’ So we can plug any Fourier series or displacement approximations in the form
and find the stationary point
Buckling of thin plates
2013-2014 Aircraft Structures - Instabilities 62
β’ Primary buckling theory of thin plates (7)
β Energy method
β’ Let us analyze the simply supported plate
β’ Internal energy
β’ First term
β As the cross-terms vanish
Buckling of thin plates
2013-2014 Aircraft Structures - Instabilities 63
β’ Primary buckling theory of thin plates (8)
β Energy method (2)
β’ Internal energy
β As
β And as cross-terms vanish
Buckling of thin plates
2013-2014 Aircraft Structures - Instabilities 64
β’ Primary buckling theory of thin plates (9)
β Energy method (3)
β’ As
Buckling of thin plates
2013-2014 Aircraft Structures - Instabilities 65
β’ Primary buckling theory of thin plates (10)
β Energy method (4)
β’ As
β’ At buckling we have at least for one couple (m, n)
β’ Most critical value for n=1
β’ In general
Buckling of thin plates
2013-2014 Aircraft Structures - Instabilities 66
β’ Experimental determination of critical load
β Avoid buckling Southwell diagram
β Plate with small initial curvature
β’
β Particular case of p = 0, tension Γ±11,
simply supported edges
β’ For
with
β’ When Γ±11 β
β Term bm1 is the dominant one in the solution
β Displacement takes the shape of buckling mode m (n=1)
Buckling of thin plates
E1
E2
E3
A
j03
u3
Γ±11
2013-2014 Aircraft Structures - Instabilities 67
β’ Experimental determination of critical load (2)
β Particular case of p=0, tension Γ±11 , simply supported edges (2)
β’ When Γ±11 β
β Term bm1 is the dominant one in the solution
β As
with
β’ Rearranging:
β m depends on ratio a/b
Buckling of thin plates
u3
-u3/Γ±11
bm1
-1/Γ±11CR
1
2013-2014 Aircraft Structures - Instabilities 68
β’ Primary to secondary buckling of columns
β Slenderness ratio le/r with
β’ le: effective length of the column
β Depends on BCs and mode
β’ r: radius of gyration
β High slenderness (le/r >80)
β’ Primary buckling
β Low slenderness (le/r <20)
β’ Secondary (local) buckling
β’ Usually in flanges
β In between slenderness
β’ Combination
Secondary buckling of columns
le = L/2
2013-2014 Aircraft Structures - Instabilities 69
β’ Example of secondary buckling
β Composite beam
β Design such that
β’ Load of primary buckling >
limit load >
web local buckling load
β Final year project
β’ Alice Salmon
β’ Realized by
β’ How to determine secondary buckling?
β Easy cases: particular sections
Secondary buckling of columns
2013-2014 Aircraft Structures - Instabilities 70
β’ Secondary buckling of a L-section
β Represent the beam as plates
β Take critical plate and evaluate
k from plate analysis
β’ k = 0.43, mode m=1
β Deduce buckling load
β’
β’ Check if lower than sp0
β This method is an approximation
β’ Experimental determination
Secondary buckling of columns
Loaded edges simply
supported
One unloaded edge free
one simply supported
(? Assumption)
0.43
b
a=3b
2013-2014 Aircraft Structures - Instabilities 71
β’ Primary buckling of thin plates
β We found
β’
β’ With k ~ constant for a/b >3
β In order to increase the buckling stress
β’ Increase h0/b ratio, or
β’ Use stiffeners to reduce effective b of skin
Buckling of stiffened panels
b
w
tst
tsk
bsk
bst
2013-2014 Aircraft Structures - Instabilities 72
β’ Buckling modes of stiffened panels
β Consider the section
β Different buckling possibilities
β’ High slenderness
β Euler column (primary) buckling with cross-section depicted
β’ Low slenderness and stiffeners with high degree of strength compared to skin
β Structure can be assumed to be flat plates
Β» Of width bsk
Β» Simply supported by the (rigid) stringers
β Structure too heavy
β’ More efficient structure if buckling occurs in stiffeners and skin at the same time
β Closely spaced stiffeners of comparable thickness to the skin
β Warning: both buckling modes could interact and reduces critical load
β Section should be considered as a whole unit
β Prediction of critical load relies on assumptions and semi-empirical methods
β Skin can also buckled between the rivets
Buckling of stiffened panels
w
tst
tsk
bsk
bst
2013-2014 Aircraft Structures - Instabilities 73
β’ A simple method to determine buckling
β First check Euler primary buckling:
β Buckling of a skin panel
β’ Simply supported on 4 edges
β’ Assumed to remain elastic
β Buckling of a stiffener
β’ Simply supported on 3 edges &
1 edge free
β’ Assumed to remain elastic
β Take lowest one (in absolute value)
Buckling of stiffened panels
w
tst
tsk
bsk
bst
0.43
2013-2014 Aircraft Structures - Instabilities 74
β’ Shearing instability
β Shearing
β’ Produces compression in the skin
β’ Leads to wrinkles
β The structure keeps some stiffness
β Picture: Wing of a Boeing stratocruiser
Buckling of stiffener/web constructions
2013-2014 Aircraft Structures - Instabilities 75
References
β’ Lecture notes
β Aircraft Structures for engineering students, T. H. G. Megson, Butterworth-
Heinemann, An imprint of Elsevier Science, 2003, ISBN 0 340 70588 4
β’ Other references
β Books
β’ MΓ©canique des matΓ©riaux, C. Massonet & S. Cescotto, De boek UniversitΓ©, 1994,
ISBN 2-8041-2021-X
2013-2014 Aircraft Structures - Instabilities 76
β’ Example
β Uniform transverse load fz
β Pinned-pinned BCs
β Maximum deflection?
β Maximum momentum?
Annex 1: Transversely loaded columns
P
x
z
fz
2013-2014 Aircraft Structures - Instabilities 77
β’ Equation
β Euler-Bernouilli
β’ This assumes deformed configuration ~ initial configuration
β’ But near buckling, due to the deflection, P is exerting a moment
β’ So we cannot apply superposition principle as the axial loading also produces a
deflection
β Going back to bending equation
β’
Annex 1: Transversely loaded columns
P
x
z
fz
P
Mxx
fzL/2
x
z
fz
2013-2014 Aircraft Structures - Instabilities 78
β’ Solution
β Going back to bending equation
β’
β General solution
β’ with
β’ BC at x = 0:
β’ BC at x = L:
β Deflection
β’ Deflection and momentum are maximum at x = L/2
Annex 1: Transversely loaded columns
P
x
z
fz
2013-2014 Aircraft Structures - Instabilities 79
β’ Maximum deflection
β Deflection is maximum at x = L/2
Annex 1: Transversely loaded columns
P
x
z
fz
2013-2014 Aircraft Structures - Instabilities 80
β’ Maximum deflection (2)
β Deflection is maximum at x = L/2 (2)
β’ From Euler-Bernoulli theory
β’ As for plates, compression induces bending
due to the deflection (second order theory)
Annex 1: Transversely loaded columns
P
x
z
fz
2013-2014 Aircraft Structures - Instabilities 81
10-2
10-1
100
100
101
102
P/PCR
uz/u
z(P=
0)
π/ππΆπ
π’π§/π’π§(π = 0)
β’ Maximum moment
β Maximum moment at x = L/2
β’
β’ With
Annex 1: Transversely loaded columns
P
x
z
fz
2013-2014 Aircraft Structures - Instabilities 82
β’ Maximum moment (3)
β Maximum moment at x = L/2 (2)
β’ Remark: for large deflections the bending equation which assumes linearity is no
longer correct as curvature becomes
Annex 1: Transversely loaded columns
P
x
z
fz
2013-2014 Aircraft Structures - Instabilities 83
10-2
10-1
100
100
101
102
P/PCR
Mxx
/Mxx
(P=
0)
π/ππΆπ
ππ₯π₯/ππ₯π₯(π = 0)
β’ Spar of wings
β Usually not a simple beam
β Assumptions before buckling:
β’ Flanges resist direct stress only
β’ Uniform shear stress in each web
β The shearing produces compression
in the web leading to a-inclined wrinkles
β Assumptions during buckling
β’ Due to the buckles the web can only
carry a tensile stress st in the wrinkle
direction
β’ This leads to a new distribution of
stress in the web
β sxx & szz
β Shearing t
Annex 2: Buckling of stiffener/web constructions
T
b
dA
B
CD
ta
Thickness t
a
C
st
A
B
C
D
F
t
a
szz
st
st
A
B
D
Ft
a
sxx
st
Ex
Ez
2013-2014 Aircraft Structures - Instabilities 84
β’ New stress distribution
β Use rotation tensor to compute in terms of st
Annex 2: Buckling of stiffener/web constructions
C
st
A
B
C
D
F
t
a
szz
st
st
A
B
D
Ft
a
sxx
st
Ex
Ez
2013-2014 Aircraft Structures - Instabilities 85
β’ New stress distribution (2)
β From
β Shearing by vertical equilibrium
β Loading in flanges
β’ Moment balance around bottom flange
β’ Horizontal equilibrium
Annex 2: Buckling of stiffener/web constructions
T
b
dA
B
CD
ta
Thickness t
a
PT
tsxx
Ex
Ez
PBL-x
2013-2014 Aircraft Structures - Instabilities 86
β’ New stress distribution (3)
β From
β Loading in stiffeners
β’ Assumption: each stiffener carries the loading
of half of the adjacent panels
β’ Stiffeners can be subject to Euler buckling if this load is too high
β Tests show that for these particular BCs, the equivalent length reads
Annex 2: Buckling of stiffener/web constructions
b
Ex
Ez
szz szz
P
2013-2014 Aircraft Structures - Instabilities 87
β’ New stress distribution (4)
β From
β Bending in flanges
β’ In addition to the flanges loading PB & PT
β’ Stress szz produces bending
β Stiffeners constraint rotation
β Maximum moment at stiffeners
β’ Using table for double cantilever beams
Annex 2: Buckling of stiffener/web constructions
b
Ex
Ez
szz szz
P
b
Ex
Ez
P
szz szz szzszz
Mmax
2013-2014 Aircraft Structures - Instabilities 88
β’ Wrinkles angle
β The angle is the one which minimizes
the deformation energy of
β’ Webs
β’ Flanges
β’ Stiffeners
β If flanges and stiffeners are rigid
β’ We should get back to a = 45Β°
β Because of the deformation of
flanges and stiffeners a < 45Β°
β’ Empirical formula for uniform material
β As non constant, a non constant
β’ Another empirical formula
Annex 2: Buckling of stiffener/web constructions
T
b
dA
B
CD
ta
Thickness t
a
Load in flange / Flange section
Load in stiffener / Stiffener section
2013-2014 Aircraft Structures - Instabilities 89
β’ Example
β Web/stiffener construction
β’ 2 similar flanges
β’ 5 similar stiffeners
β’ 4 similar webs
β’ Same material
β E = 70 GPa
β Stress state?
Annex 2: Buckling of stiffener/web constructions
T = 5 kN
b = 0.3 m
d=
40
0 m
m
ta
t = 2 mm
a
AS = 300 mm2
AF = 350 mm2
(Iyy/x) = 750 mm3
Ex
Ez
Ixx = 2000 mm4
2013-2014 Aircraft Structures - Instabilities 90
β’ Wrinkles orientation
β
Annex 2: Buckling of stiffener/web constructions
T = 5 kN
b = 0.3 m
d=
40
0 m
m
ta
t = 2 mm
a
AS = 300 mm2
AF = 350 mm2
(Iyy/x) = 750 mm3
Ex
Ez
Ixx = 2000 mm4
2013-2014 Aircraft Structures - Instabilities 91
β’ Stress in top flange (> than in bottom one)
β 1st contribution: uniform compression
β’ Maximum at x = 0
β 2nd contribution: bending
β’ Maximum at stiffener at x=0
β Maximum compressive stress
Annex 2: Buckling of stiffener/web constructions
T
b
dA
B
CD
ta
Thickness t
a
PT
tsxx
Ex
Ez
PB
2013-2014 Aircraft Structures - Instabilities 92
β’ Stiffeners
β Loading
β Buckling?
β’ As b < 3/2 d:
β’ Assuming centroid of stiffener lies in webβs plane
β We can use Euler critical load
β’ No buckling
Annex 2: Buckling of stiffener/web constructions
b
Ex
Ez
szz szz
P
2013-2014 Aircraft Structures - Instabilities 93
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