1. INTRODUCTION
1.1 GOALS AND SIGNIFICANCE
The goal of this project is to simulate the wing box design process in industry and to
encourage team work. Integrally Stiffened Panel (Internal blade section) of wing box skin
was chosen for Casa NC212
To understand structural design criteria
•
To able to define the structure function,
•
To select the appropriate layout
•
To choose the right material and processes
•
To able to size and analyse the configuration
•
To draw the model using CAD
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•
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1.2 PROJECT ORGANIZATION
Firstly, choose material for each part of the wing box base on knowledge about structural
design criteria.
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Secondly, determine load distribution on the wing rely on every section, as Figure 1.
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Figure 1
Thirdly, use y, c, SF, BM as the initial parameter for designing wing box including
“rough” wing mass, with difference of stringer pitch and rib pitch in order to parametric study
for the last part.
Where, y: distance of each cross section from the central line
c: chord at each cross section
SF: shear force at each cross section
BM: bending moment at each cross section
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2. FEARTURE OF INTEGRAL STIFFENED PANELS
Present trends toward higher performance levels in machines and equipment continue to
place more exacting demands on the design of structural components. In aircraft, where
weight is always a critical problem, integrally stiffened structural sections have proved
particularly effective as a light weight, high-strength construction. Composed of skin and
stiffeners formed from the same unit of raw stock, these one-piece panel sections can be
produced by several different techniques. Size and load requirements are usually the
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important considerations in selecting the most feasible process.
Typical integral stiffened panels (planks)
For highly loaded long panels, extrusions or machined plates are most commonly
employed. Section discontinuities such as encountered in the region of cutouts can often be
produced more easily from machined plate. From a cost standpoint it is usually better to
machine a section from the extruded integrally stiffened structures than to machine a section
of the same size from a billet of plate.
From a structural standpoint, appreciable weight savings are possible through the
integral-section design which also develops high resistance to buckling loads. In addition, the
reduction in the number of basic assembly attachments gives a smooth exterior skin surface.
In aircraft applications, the most significant advantages of integrally stiffened structures over
comparable riveted panels has been :
● Reduction of amount of sealing material for pressurized fuel tank structures
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● Increase in allowable stiffener compression loads by elimination of attachment flanges.
● Increased joint efficiencies under tension loads through the use of integral doublers, etc.
● Improved performance through smoother exterior surfaces by reduction in number of
attachments and nonbuckling characteristics of skin
●
Light weight structures
Integrally stiffened structures have their greatest advantage in highly loaded applications
because of their minimum section size. Investigations have indicated that an integrally
stiffened section can attain an exceptionally high degree of structural efficiency. A weight
reduction of approximately 10-15% was realized by the use of an integrally stiffened
structure.
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It should be noted that in order to obtain the true weight difference, all non-optimum
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factors must by taken into account. The integrally stiffened design will have a relatively low
weight for the so-called non-optimum features. This is attributed to the machined local
padding and reinforcing material and permitted by integral cover construction. In contrast, the
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builtup type of design generally requires a relatively large non-optimum weight because of the
many chordwise splices for ease of tank sealing and fabrication, discrete doublers, etc.
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From the foregoing one concludes that the lightest cover panel design can be obtained with an
integrally stiffened cover structure supported by sheet metal ribs with a preference for a large
spacing. If the use of integral skin is prohibited for such reasons as exfoliation, etc, then
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special attention must be given to the non-optimum factor for using builtup skin-stringer
panel. End grain exposure and residual stresses (threshold stresses) due to fabrication “pullup” presented a stress corrosion risk. Airplane operator objections, due to experience with
integrally stiffened panels on some competitors’ airplane, led to a decision not to use them on
some of these commercial airplanes.
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3. DESIGN WING BOX IN DETAIL
3.1 MATERIAL SELECTION
As airframe design concepts and technology have become more sophisticated, materials’
requirements have accordingly become more demanding. The steps from wood to aluminum,
and then to titanium and other efficient high strength materials has involved some very
extensive development activities and the application of a wide range of disciplines.
Structure weight and therefore the use of light materials has always been important.
When a modern full-loaded subsonic transport takes off, only about 20% of its total weight is
payload. Of the remaining 80%, roughly half is empty weight and the other half is fuel.
Hence, any saving of structural weight can lead to a corresponding increase in payload.
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Alternatively, for a given payload, saving in aircraft weight means reduced power
requirements. Therefore, it is not surprising that the aircraft manufacturer is prepared to invest
heavily in weight reduction[1].
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Lower wing skin particularly prone to fatigue through the long-continued application and
relaxation of tension stress, so the standard material is aluminum alloy designated 2024-T3.
For upper wing skins, which have to withstand mainly compression stresses as the wing flexes
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upwards during flight, 7075-T6 is used.
With integrally Stiffened Panel (Internal blade section) of
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wing box skin . Just several material can be chosen because need
to pay attention to the thickness of form of material (Clad sheet,
bar and extrusions, sheet and strip,…) for machining.
In this design, following ESDU 76016, material can be chosen for upper skin is 2L 88-
T6 (Plate 40 mm< t < 63 mm) and material for lower skin is 2L 93-T651 (Plate 40 mm< t <
63 mm) or 2L 95-T651 (Plate 25 mm< t < 75 mm).
Anyway, in purpose to compare with another design, we choose the same material for
every part of wing box
Al 7075-T6 (DTD 5014) : Upper skin, spar web, rib web,…
Al 2024-T3 (DTD 5070B): Lower skin
ρ
fn
(MPa) (kg/m3)
Untimate tensile strength
(MPa)
Shear strength
(MPa)
E
(MPa)
c2
(MPa)
Al 7075-T6 (DTD 5014)
572
331
76000
487
22.2
444
2810
Al 2024-T3 (DTD 5070B)
483
283
73100
342
16.6
301
2780
Mat. Prop
m
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3.2 DETERMINE LOAD DISTRIBUTION ON THE WING
3.2.1 Specification of Casa NC212
Gross weight at crusing
7450 kg
One Engine weight
329.98 kg
Fuel weight (for half of wing)
800 kg
Air density at cruising altitude (2438 m):
ρ = 0.967 kg / m 3
= 40 m2
b
= 19 m
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Vcruise = 96.6 m/s
3.2.2 Wing model
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Taper ratio = 2
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outboard
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Base on real wing planeform, model of half-wing is illustrated as Figure 2, Fuel tank
Figure 2
3.2.3 Lift distribution
Using Shrenk’s method to estimate the lift distribution on the wing.
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Lift coefficient on each section
1⎛
4S
2y ⎞
c planformcl = ⎜⎜ c planform +
1 − ( ) 2 ⎟⎟
πb
2⎝
b ⎠
(1)
with
cplanform: chord length of section (m)
cl
: lift coefficient on section in case lift coefficient is 1.0
y
: position of section from central line (m)
S
: wing area (m2)
b
: wing span (m)
Since cl in equation (1) for lift coefficient is 1.0, real Cl is obtained by multiplying cl
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with CL at crusing condition.
L = Wcr
1
ρVcr2CL S
2
Wcr
7450 × 9.81
⇒ CL =
=
= 0.4
1
1
2
2
ρVcr S
× 0.967 × 97 × 40
2
2
(2)
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⇒ Wcr =
So lift on each section is
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ΔL =
1
ρV 2 cl ΔS
2
(3)
3.2.4 Shear force distribution
Using equilibrium condition on each section, calculate shear force for it
SF2 = ΔL1 + ΔL2 = SF1 + ΔSF2
General equation
SFi = SFi −1 + Δ Li
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3.2.5 Bending moment distribution
Similar to shear force, using equilibrium condition on each section, calculate bending
moment for it
General equation
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Δy ⎞
Δy
⎛
BM 2 = ΔL1 × ⎜ Δy2 + 1 ⎟ + ΔL2 × 2
2 ⎠
2
⎝
Δy
Δy
= ΔL1 × 1 + ΔL1 × Δy2 + ΔL2 × 2
2
2
SF − SF1 ⎞
⎛
= BM 1 + ⎜ SF1 + 2
⎟ × Δy2
2
⎠
⎝
⎛ SF + SF2 ⎞
= BM 1 + ⎜ 1
⎟ × Δy2
2
⎝
⎠
BM 2 = BM 1 + ΔBM
BM i = BM i −1 + ( SFi −1 + SFi )
Δy
2
3.2.6 Result
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= BM i −1 + ΔBM i
In this part, detail of load distribution in case Rib pitch (L) = 0.35 m and Stringer pitch
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(b) = 0.1 m
A/ Shear Force and Bending Moment distribution
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B/ Shear Force diagram (Figure 3)
Shear Force Diagram
SF (N) 30000
25000
20000
15000
10000
5000
0
0.0 0.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0 5.5 6.0 6.5 7.0 7.5 8.0 8.5 9.0 9.5
y (m)
Figure 3
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C/ Bending Moment diagram (Figure 4)
Bending Moment Diagram
BM (N.m) 100000
90000
70000
60000
50000
40000
30000
20000
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10000
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80000
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0.0 0.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0 5.5 6.0 6.5 7.0 7.5 8.0 8.5 9.0 9.5
y (m)
Figure 4
D/ Conclusion
Maximum Shear Force and Bending Moment occur at the wing center line (SFmax =
25,011 N; BMmax = 94,488 Nm). Choose wing root at the center line, so use these values to
design wing box at root.
3.3 WING BOX DESIGN
3.3.1 Wing box geometry selection
From airfoil NACA 653-218, choose shape of two-spars wing box showed in Figure 5.
• Wide of wing box (w) = 40% chord.
• Height of wing box (h) = 0.8 thickness = 14.4% chord
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Chord at each cross section of the wing is different so w and h are also different
Figure 5
%c (FS)
%c(RS)
%c max
20
60
d (max thickness at root) (m)
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0.45
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3.3.2 Design procedure
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y, c, SF, BM
Due to overall torsion moment
Checking buckling due to shear
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Sizing spar web (front and rear)
tw (front), tw (rear)
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Initial sizing upper skin panel
Intial tskin (upper)
Number of stringer
Distance of stringer
Actual sizing upper skin panel
tskin , tw , b , bw
Checking local buckling and crippling
Checking buckling for panel due to
compression and shear
Initial sizing lower skin panel
Intial tskin (lower)
Number of stringer
Distance of stringer
Actual sizing lower skin panel
tskin , tw , b , bw
Checking for tension (Mises Hensky
Criteria)
Checking buckling due to shear
Sizing rib webs
Rib flanges (thickness, width)
tRW
Due to concentrated loads
Due to shear loads
Due to crushing loads and checking
buckling due to compression and shear
The reason why sizing spar web was done first is getting shear flow on upper and lower
skin to check buckling due to shear when sizing upper and lower skin, illustrated in Figure 6
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3.3.4 Sizing Spar Webs ( Front Spar Web and Rear Spar Web)
Mat. Prop
Untimate tensile strength
(MPa)
Shear strength
(MPa)
E
(MPa)
c2
(MPa)
572
331
76000
487
Al 7075-T6 (DTD 5014)
m
22.2
ρ
fn
(MPa) (kg/m3)
444
2810
Spar webs shall be sized by using shear criteria :
τ all
≥ 1.0
τs
where
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τ s : applied shear stress
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τ all : allowable shear stress of the panel, which is the smallest value of:
skin shear local buckling stress
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allowable shear stress of the material used
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A. Due to Overall Torsion Moment
Estimate the enclosed area, A, of the primary structural box at representative sections
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across the span.
The corresponding shear flow is:
QT =
T
2A
with
Load factor: 2.5
Factor of Safety: 1.5
Shear flow due to torque
Figure 6
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Note that (from Figure 6), it’s evident shear flow in the rear spar web (+) larger than
shear flow in the front spar web (-), so rear spar web thickness will be thicker than rear spar
web thickness. This was proved in design results
Where T is now the applied distributed torsion, and QT will be nose up or nose down and
hence positive or negative depending on the sign convention
Select the allowable shear stress, f s as appropriate
The mean material thickness needed to react the torsion moment is then:
tq =
T
2 Af s
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Combined with vertical shear loads
The shear flow in the webs due to the shear force is then:
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hT
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QV =
where V is the applied vertical shear force
The net shear flow in the web is then approximately given by:
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Qw = QV ± 2
x
QT
w
(+) for rear spar web thickness and (-) for front spar web thickness
Where x is the chordwise location of particular web relative to the mid point of the box
•
The web thickness is then
tw =
Qw
fs
B. Checking due to shear
⎛t⎞
⎝b⎠
2
σ cr = K S E ⎜ ⎟
Figure 7
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Consider spar webs, skin (upper and lower), ribs as narrow panels with heavy skin, so
this panel will act as if it had hinged edges (panel 4 in Figure 7). Look up in Figure 5.4.6 ([1],
pp 139) with
Since
a
in line 4 to get value of K s .
b
a
always larger than 4, K s = 5 is used when checking buckling due to shear for
b
spar webs, skin (upper and lower) except ribs.
When choosing K s = 5 (minimum value), structure will be safe for all case but we pay
by make it heavier. Anyhow, the increasing weight for this carefulness is not too much, so it
is acceptable. In case sizing ribs web, buckling due to shear is the most dangerous, so try to
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make K s = 8 (maximum value) by using stiffener for ribs web.
Mat. Prop
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3.4.2 Sizing Upper Skin
Untimate tensile strength
(MPa)
Shear strength
(MPa)
E
(MPa)
c2
(MPa)
572
331
76000
487
22.2
444
2810
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Al 7075-T6 (DTD 5014)
ρ
fn
(MPa) (kg/m3)
m
A. Intial sizing Upper skin
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i. At various points across the span evaluate the idealised depth of the primary
structural box, h (Figure 8)
ii.
Figure 8
iii. Calculate the effective direct loads, P, in the top and bottom surfaces required to
react the appropriate bending moment, M, at each section from: P =
M
h
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iv. Evaluate the allowable stress, f b
When the concept is based on a distributed flange construction the allowable bending
stress at ultimate loading may be assumed to be the lesser of the 0.2% proof stress or f b ,
where f b may be approximately represented by: P 1/ 2
⎛
⎞
f b = A FB ⎜
⎟
⎝ wL ⎠
where
L: the local rib or frame spacing
w: the width of the box perpendicular to the bending axis
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P: the effective end load
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A : a function of the material
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FB: dependant upon the form of construction.
„
Note that the value of A are appropriate
to allowable stress and (P/wL) in MN/m2
units. In general the values of A give
conservative values for FB at stresses
below the limiting value.
„
Typical values for FB are also given.
With intergral blade stringer, get A = 138 and FB = 0.81
v. Evaluate the skin thickness is required at each section
• For a distributed flange assume initially a uniform effective thickness across the width,
w, to give
t=
M
hwf b
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• Typically this thickness will be made up of skin and stringer area. The effective stringer
are being about half of that of the skin area. Thus the actual skin thickness is about
t=
0.65M
hwf b
Follow optimization of value tw and bw in [1] pp 155-165 , so
from t, b get tw and bw
tw
= 2.25
te
bw
= 0.65
b
and
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B. Checking (compression and shear interaction criteria, buckling)
B1. Using compression and shear interaction criteria
τs
σc
τs
σ comp τ s2
+
≤ 1.0
σ cr τ cr2
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σc
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σ comp : applied compression stress
τs
: applied shear stress
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σ cr : critical compression stress (property of material)
τ cr
: critical shear stress (property of material)
Applied compression stress was calculated by
P
Aeff
σ comp =
where
P : the effective end load
Aeff : effective area (area of skin and stiffeners)
Applied shear stress was calculated by
τs =
tq
te (initial )
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where
tq
: shear flow due to torsion moment (from sizing spar webs)
te (initial ) : initial skin thickness (from initial sizing)
B1. Checking local buckling of skin and crippling stringer
due to
compression
Use ESDU-70003 to check local buckling of skin panels
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f b = ηf be
f c = (c2 f b )1/ 2
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f be
⎛t⎞
= KE ⎜ ⎟
⎝b⎠
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where
f be : average elastic compressive stress in panel at which local buckling first occurs
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K : buckling stress coefficent
f b : average compressive stress in panel at which local buckling first occurs
η : plasticity reduction factor defined by
fid
f b = ηf be
f c : crippling stress
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c2 : 0.2% compressive proof stress of strut material
B2. Checking global buckling of skin panel
⎛t ⎞
f cr , skin = KE ⎜ e ⎟
⎝b⎠
K = 3.62
2
K = 6.32
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In this design report, two kind of skin panels were checked for local buckling. However,
just need to check with hinged edge is enough
B3. Checking buckling due to shear
Do the same process with part B (B. Checking due to shear) in 2.3.2 Sizing Spar Webs
⎛t⎞
to get σ cr = K S E ⎜ ⎟
⎝b⎠
2
B. Actual sizing upper skin
From the initial skin thickness t, try to reduce it as thin as possible with condition that
Shear strength
(MPa)
E
(MPa)
c2
(MPa)
483
283
73100
342
m
ρ
fn
(MPa) (kg/m3)
16.6
301
2780
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Al 2024-T3 (DTD 5070B)
Untimate tensile strength
(MPa)
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Mat. Prop
σcomp < min(fbe, fc , fcr,skin,σcr )
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3.4.2 Sizing Lower Skin
and
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σ comp τ s2
+
≤ 1.0
σ cr τ cr2
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satisfy requirements below. Finally, skin thickness (t) at each section is obtained.
Do the same process of sizing upper skin but the difference here is design criteria.
Upper skin withstand mainly compression stresses as the wing flexes upwards during
flight, so almost unstable problem due to bucking, so checking buckling is the most
importance.
Beside that, lower skin particularly prone to fatigue through the long-continued
application and relaxation of tension stress, so checking damage tolerance is the most
importance. However, for checking damage tolerance take long time and need more detail
information about flight hour, and sizing without damage tolerance get value of skin thickness
(t) too much smaller than upper skin thickness. So get lower skin thickness equal to upper
skin thickness at each section.
Tension and shear interaction criteria for lower skin sizing is showed below
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Using tension and shear interaction criteria for lower skin sizing
σ all
≥ 1 .0
σ comb
where as material failure according to Von Mises :
σ comb =
σ all
σ comp + 3τ 2
: allowable tension stress of the material used (material property)
σ comp : applied tension stress
τ
l
: applied shear stress
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Note that Lower panel is also critical due to fatigue. The criteria have to be considered is :
Mat. Prop
Untimate tensile strength
(MPa)
Shear strength
(MPa)
E
(MPa)
c2
(MPa)
572
331
76000
487
m
22.2
ρ
fn
(MPa) (kg/m3)
444
2810
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Al 7075-T6 (DTD 5014)
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3.4.2 Sizing Ribs Web
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σ all −1G
−1 ≥ 0
σt
A. Intial sizing ribs web
i.
Due to concentrated loads (attachments: engine, flaps, etc):
Can be taken as a cantilever beam loaded by a vertical shear force equal to the hinge
reaction and a bending couple due to the offset of the hinge chordwise from the rear
spar location. The spar web will react most of the vertical shear, and in practice if the
hinge fitting is perpendicular to the rear spar, the rib flanges at the spar will be loaded
by direct forces given by:
R = ±V .
where
x
h
V: the hinge reaction
x: the offset of the hinge from the spar
h: the depth of the rib at the spar
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• Estimating lift on flap (V) at Take-off:
Assum that:
Take off velocity (Vtake _ off ) = 50 m/s
Cl (with alpha 80 )
= 2.25
ρ air (at sea level)
= 1.226 kg/m3
Lift on flap at each section (do the same as lift on wing at each section )
ΔV =
1
2
ρVtake
_ off ClΔS flap
2
• Estimating Flange area:
R
σ ys
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A flange =
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Vi = Vi −1 + Δ Vi
Due to shear loads :
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ii.
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Choose Flange thickness ⇒ Flange width
Shear flow on the ribs web
q3 =
R
rib (i+1)
rib (i)
rib (i-1)
Qz1
Qz2
R = Qz1 − Qz 2
R
(daN/mm) = shear flow
2. h
h=
hFS + hRS
2
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iii.
Due to crushing loads :
Crushing load act on the ribs web was given by:
σn =
2.σ 2 .t panel .L
E.h.trib _ web
σn
where
σupper
σn
: crushing stress at rib
σlower
σn
;
σ upper : normal stress at upper panel
σ lower : normal stress at lower panel
l
: rib height
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H
t panel
σ shear _ strength
abs (σ upperpanel ) + abs (σ lowerpanel )
2
= σ upperpanel
fid
σ=
: tskin actual
q3
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trib _ web : rib web thickness with trib _ web =
B. Checking buckling (due to compression and due to shear)
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Checking buckling due to compression
f cr , rib web
⎛t
⎞
= KE ⎜⎜ rib _ web ⎟⎟
⎝ w ⎠
2
Checking buckling due to shear
K = 3.62
Do the same process with part B (B. Checking due to shear) in 2.3.2 Sizing Spar Webs
2
to get σ cr , ribweb
⎛t
⎞
= K S E ⎜ rib _ web ⎟ . With KS = 8
⎜ w ⎟
3 ⎠
⎝
( )
σ cr , ribweb was compared by σ shear =
q3
trib _ web
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L = rib
If stiffener for ribs web was not use, rib webs thickness would be so thick ⇒ structure
would be very heavy. In this design, rib webs was added two stiffener ⇒ divide rib webs into
three part as figure below. Finally, get reasonable weight of ribs web
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C. Actual sizing ribs web
From the initial ribs web thickness trib _ web =
q3
σ shear _ strength
, try to reduce it as thin as
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possible with conditions that satisfy requirements below. Finally, skin thickness (t) at each
section is obtained.
σ shear < min(σ cr,ribweb,σ shear_ strength)
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3.4 RESULT
and
fid
σ n < fcr,ribweb
3.4.2 Data result
The main results of sizing wing box were shown in three tables below with three couple
value
(b) Stringer pitch (m) = 0.08 0.10 0.12
(L) Rib pitch (m) = 0.30 0.35 0.40
The purpose of given Stringer pitch in this report is to determine the number of stringers.
The number of stringer is still kept for all cross sections so stringer pitch of each cross section
will be different and decreased from root to tip. That’s why in the result stringer pitch is not
equal to the number above.
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3.4.2 Figure result
In this report, shape of cross section wing box at root was drawn from the
real scale for a general view. (Using Auto CAD to draw them)
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A. Wing box with Rib pitch L = 0.3 m and Stringer pitch = 0.08 m
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fid
Figure 9
Figure 10
26
B. Wing box with Rib pitch L = 0.35 m and Stringer pitch = 0.1 m
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Figure 11
Figure 12
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C. Wing box with Rib pitch L = 0.4 m and Stringer pitch = 0.12 m
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Figure 13
Figure 14
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4. PARAMETRIC STUDY
From the design results, draw mass of (upper skin, lower skin, spars, ribs)
vs stringer pitch and rib pitch.
(b) Stringer pitch
(m)
Upper skin mass
(kg)
Lower skin mass
(kg)
Spars mass
(kg)
Ribs mass
(kg)
Mass of hafl wing
(kg)
0.08
0.10
0.12
84
96
123
83
95
121
42
40
40
24
20
18
232
252
303
Parametric Study
mass (kg)
350
l
300
tia
250
W_upper skin
200
W_lower skin
W_spars
150
W_ribs
50
fid
0
0.07
0.08
0.09
0.10
C
on
(L) Rib pitch
(m)
W-total
en
100
0.11
0.12
Stringer pitch (m)
0.13
Upper skin mass
(kg)
Lower skin mass
(kg)
Spars mass
(kg)
Ribs mass
(kg)
Mass of hafl wing
(kg)
84
96
123
83
95
121
42
40
40
24
20
18
232
252
303
0.30
0.35
0.40
Parametric Study
mass (kg)
350
300
250
W_upper skin
200
W_lower skin
W_spars
150
W_ribs
W-total
100
50
0
0.25
Rib pitch (m)
0.27
0.29
0.31
0.33
0.35
0.37
0.39
0.41
29
It is evident that stringer pitch and rib pitch have the important role in mass
of half wing and also every part of wing box. Increasing stringer pitch and rib
pitch (reducing number of stringers, number of ribs) → mass of spars and ribs
reduce insignificantly, but the stringer have to be longer, thicker and skin also
→ mass of skins (including stiffener) increase promptly in order to keep
stability with the same loading (compression and shear).
In this report, lower skin thickness was chosen to be similar to upper skin
thickness, so mass of them is not different too much (just different from density
of material).
tia
l
From Figure 15, it is obvious to observe the change of wing box cross
section shape when L and b were changed.
en
The difference between wing box mass in case (L = 0.3m and b = 0.08m)
and in case (L = 0.35m and b = 0.1m) is not too much (232 kg → 252 kg) but
the difference between wing box mass in case (L = 0.35m and b = 0.1m) and in
fid
case (L = 0.4m and b = 0.12m) is significantly (252 kg → 303 kg). It mean that
case of (L = 0.4m and b = 0.12m) is not good and it will never be using for
C
on
design.
In case of (L = 0.3m and b = 0.08m), length of stringer still to long
(54mm).
In order to compare to anther design, this results was still kept.
Unfortunately, time is not enough for design other case, so it is kept for further
study.
Figure 15
30
5. FURTHER STUDY
• Reduce L and b to get more number of stringers and get less length of
stringer and less weight of skin (including stringer).
• Understand the influence of individual L and b.
From two suggestions above, withdrawn the optimization of wing mass and
C
on
fid
en
tia
l
configuration of the wing
31