study of cracks on aircraft structures
TRANSCRIPT
Study of Cracks on Aircraft Structures
C N Yashaswini1
B.E Aeronautical Engineering
Dayananda Sagar College of Engineering
Bangalore, India
Muskan Rastogi3
B.E Aeronautical Engineering
Dayananda Sagar College of Engineering
Bangalore, India
Manjunath B G2
B.E Aeronautical Engineering
Dayananda Sagar College of Engineering
Bangalore, India
Shahid Adnan4
B.E Aeronautical Engineering
Dayananda Sagar College of Engineering
Bangalore, India
Srikanth Salyan5
Assistant Professor, Department of Aeronautical Engineering
Dayananda Sagar College of Engineering
Bangalore, India
Abstract— Fatigue plays a significant role in crack growth in
aircraft structures. Besides, the Structures may also suffer from
corrosion damage and wear defects. The proper maintenance
and scheduled test intervals can avoid sudden failure. Therefore,
the inspection interval has to become shortened. Nevertheless,
the young machines may also be suffering from unexpected skin
rupture. During the last decades, the fracture toughness, design,
and the new alloying element have been enhancing. This study
revival the analysis of cracks on different structures of an
aircraft and states the deformation of the structure at different
positions. Also, a series of analysis will be carried out to examine
the effectiveness of the composites on preventing fatigue crack
propagation and extending the fatigue life using ANSYS
workbench. The cracks are emanating from the rivets and the
holes under cyclic loading. The stress concentration around the
notch has an effective role under the impact of cyclic loading.
The cracks propagate toward the high stressed area, such as the
notches or other crack locations. Therefore, the service life of
the structure for different composite materials, amount of
damage caused, and fatigue crack growth for the structural
component under subjected conditions are calculated.
Keywords— Fatigue, crack propagation, Service life, Aircraft
structures, damage, Remaining Flights.
I. INTRODUCTION
Since the early days of the aviation industry, safety has
been one of the major concerns. Aircraft always have been
expected to last longer than automobiles. Several problems
arise from the fact that aircraft when is expected to last so
long. One of the major sources of the problem, which is the
purpose of this research, is the presence of fatigue cracks in
Aircraft structures. For many years, techniques have been
developing and are used to address the problem of fatigue
cracks.
Cracks are local material separations in a machine frame
or structure. Cracks can develop later in the course of service
loading or cyclic loading when Aircraft experience all
different types of fatigue loadings. Take-offs and landings are
very fundamental types of cyclic loadings on aircraft. Cabin
pressurization is a type of cyclic loading as the plane
pressurizes to accommodate passengers at higher altitudes.
Vibration is a major source of fatigue cracking in aircraft,
present due to atmospheric turbulence but also due to many
factors related to the engines, whether reciprocating or
turbofan. Such structures need to be inspected non-
destructively to detect hidden damage such as fatigue cracks
before they have reached a critical length and repaired before
they lead to catastrophic failure. Therefore, accurate and
reliable techniques must be carried out routinely to detect such
defects in aircraft. Fatigue cracks Inspections in an aircraft is
most important because, if left unchecked, these cracks
continue to grow. In fact, it's generally considered that over 80
percent of all service failures can be traced to mechanical
fatigue, whether in association with cyclic plasticity, sliding or
physical contact (fretting and rolling contact fatigue),
environmental damage (corrosion), or elevated temperatures.
II. MATERIAL SELECTION
The most common metals used in aircraft construction are
aluminum, magnesium, titanium, steel, and their alloys.
Aluminum alloys are widely used in modern aircraft
construction. The outstanding characteristic of aluminum is its
lightweight. So, in this case we have used Aluminium alloy.
A. Aluminium Alloy 2024
2024 Aluminium alloy is an alloy with copper as the
primary alloying element. It is used in applications requiring a
high strength-to-weight ratio, as well as good fatigue
resistance. It has poor corrosion resistance. It is mostly used to
make the aircraft’s structural parts such as wing and fuselage.
B. Aluminium Alloy 6061
6061 Aluminium alloy a precipitation hardened
Aluminium alloy, containing magnesium and silicon as its
major alloying elements. It has very good corrosion resistance
and very good weldability although reduced strength in the
weld zone. 6061 is commonly used for the following
construction of aircraft structures, such as wings and
fuselages, more commonly in homebuilt aircraft than
commercial or military aircraft.
TABLE I. MECHANICAL PROPERTIES OF ALUMINIUM
MECHANICAL PROPERTIES Al 2024 Al 6061
Ultimate Tensile Strength 469 MPa 241 MPa
Tensile Yield Strength 324 MPa 145 MPa
Shear Strength 283 MPa 207 MPa
Fatigue Strength 138 MPa 96.5 MPa
Modulus of Elasticity 73.1 GPa 68.9 GPa
Shear Modulus 28 GPa 26 GPa
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III. METHODOLOGY
A. FATIGUE LIFE
A crack in a part will grow under conditions of cyclic
applied loading, or under a steady load in a hostile chemical
environment. Crack growth due to cyclic loading is called
fatigue crack growth. The crack initially grows very slowly,
but the growth accelerates (i.e., da/dN increases) as the crack
size increases. The reason for this acceleration in growth is
that the growth rate is dependent on the stress intensity factor
at the crack tip, and the stress intensity factor is dependent on
the crack size, a. As the crack grows the stress intensity factor
increases, leading to faster growth. The crack grows until it
reaches a critical size and failure occurs. The fatigue crack
growth of a structure can be obtained by using the given
formulas. This equation (1) gives here relates crack growth
rate with stress intensity factor range.
…………... (1)
Where ‘da/dN’ is fatigue crack growth rate, ‘C’ is the
material constant for walker crack growth rate equation, and
‘m’ is the Walker constant.
ΔK = Kmax – Kmin …………... (2)
The other mode of failure is plastic collapse at the net section
between two advancing crack tips of the rivet holes of wing
skin; if net section stress is greater than the yield strength of
the material, then wing skin fails due to plastic collapse. Net
section stress is calculated by,
…………... (3)
where W is the pitch between two riveted holes, aeff is the
effective crack length, and t is the wing thickness. Failure
mechanism in cracked wing skin is obtained by comparing
SIF results with Fracture Toughness (KIC) of the material. If
SIF is greater than the KIC value, then wing skin fails due to
fracture. The other way to calculate ∆K is given by the
equation (4),
Smax(1-R)*1.08899+0.04369*(a/b)1.77302*(a/b)2
+9.212*(a/b)3-15.8683*(a/b)4+9.48718*(a/b)5*
(3.142*0.001*a). …………... (4)
With the help of these formulas, we can predict the growth of
the crack in the aircraft structures like wings, fuselage, and
other structures as well. The formulas give here help us in
plotting the graph between the fatigue crack growth and the
number of cycles.
B. SERVICE LIFE
The service life of aircraft Structural components
undergoing random stress cycling was analyzed by the
application of fracture mechanics using MATLAB. The Initial
crack sizes at the critical stress points for the fatigue-crack
growth analysis were established on the structure. The fatigue-
crack growth rates for random stress cycles were calculated
using the half-cycle method. The equation (5) was developed
for calculating the number of remaining flights remaining for
the structural components.
i. Conventional method
If ∆a is the amount of crack growth induced by the first
flight, then the conventional method predicts the number of
remaining flights F1 (service life) based on the following
equation (5),
………... (5)
Where acp and ac
0 are calculated respectively from equation (6)
& (7),
………... (6)
………... (7)
Where, σ* and fσ* (f<1) are respectively, the proof load
induced stress (limit stress) and the operational peak stress at
the critical stress point. A is the crack location parameter
(A=1.00 for the through crack, A=1.12 for the surface and the
edge crack). Mk is the flaw magnification factor (Mk=1.0 for
very shallow surface cracks, Mk=1.6 when the depth of the
crack approaches the thickness of the plate). KIC is the critical
stress intensity factor, and Q is the surface flow shape and
plasticity factor of a surface crack which is expressed as, Critical stress intensity factors for through crack from equation
(8),
………... (8) Critical stress intensity factor for surface & edge crack from
equation (8),
………... (9)
Here, Q can be expressed as,
………... (10)
Where, σ ͚ is the Uniaxial tensile stress,
σy is the Yield stress,
E(k) is the Elliptic function.
Before the flight, the actual amount of crack growth ∆a_
for the first flight is unknown. The way to estimate ∆a , before
the actual flight is to perform a Transient Dynamic Analysis
of the flight vehicle under specified severe maneuvers such as
landing, braking, the ground turns, flight in severe buffet and
turbulence, etc. Actual ground maneuvering of the aircraft can
be conducted and generate an actual loading spectrum for each
critical component for a short period. Then, the loading
spectrum is extrapolated to meet the duration of one flight. For
large flexible aircraft, the ground maneuver could produce a
more severe loading spectrum than that of the actual steady
flight. F0 predicts a sufficient number of flights available
based on ∆a, calculated from the ground maneuver.
ii. Calculation of crack growth
The crack growth generated by the random stress cycling
of the first flight may be calculated by using the half-cycle
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theory. The half-cycle theory states that the damage, or crack
growth caused by each half cycle (either increasing or
decreasing load) of the load spectrum is estimated to equal
one-half of the damage caused by a full-cycle of the constant-
amplitude load spectrum of the same loading magnitude.
Thus, the total damage done by the load spectrum will be the
summation of the micro-damages caused by the individual
half-waves of different loading magnitudes
Thus, the crack growth ∆a caused by the first flight may be
calculated from the equation (11),
………... (11)
Where, Kmax & R are maximum stress intensity factor and
Stress ratio.
………... (12)
………... (13)
Here, σmax & σmin are maximum and minimum stresses
constant amplitude stress cycles.
IV. MODELLING
In this fatigue analysis of five different structures with and
without a crack of is considered to determine the effect of
crack on life, damage and safety factor under fatigue loading
conditions using software’s like CATIA and ANSYS. The
different structures chosen are:
➢ Wing Skin.
➢ Integral Wing Skin and Rib Panel.
➢ Integral Wing Panel of Wing Skin and Stringers.
➢ Integral Fuselage Panel Without Cut-Out.
➢ Integral Fuselage Panel with Cut-Out.
A. WING SKIN
i. Geometry
Wing skin used for the current study has the following
dimensions i.e., 60mm for width, 120mm for height, and
1.5mm thickness of the skin. The wing skin is joined to the
frame with the help of rivets which is 4mm in diameter,
separated by 26mm. The 2 rivets here are represented as holes
on the skin.
The geometry of the wing skin is shown in the Fig.1.
Cracks over wing skin usually occur in the rivet hole edges
and through the skin, these types of cracks are categorized as
edge crack and through crack respectively.
Fig.1 Design of a Wing Skin with a crack between the rivet edges and at the
rivet edges.
TABLE II. GEOMETRIC PROPERTIES OF WING SKIN
GEOMETRY UNIT (mm)
DIAMETER OF RIVET HOLES 4
WIDTH 60
PITCH (RIVETS) 26
LENGTH 120
SKIN THICKNESS 1.5
ii. Meshing
The modeling of wing skin is done with two riveted holes
in it and three-dimensional four-node tetrahedral elements of
size 1mm in FEA Solver Software Ansys as shown in Fig.2.
The tetrahedral shaped mesh used here are essential for crack
propagation for the Ansys System. The mesh around the crack
tip or crack fronts is defined finer than others by utilizing the
sphere of influence mesh with element size of 0.5mm.
Fig.2 Mesh of Wing Skin.
iii. Loads and Boundary Conditions
During the flights, there is a lot of loads acting on the wing
box of the aircraft such as the change in atmospheric pressure
due to which the drag acts on the wing skin similarly various
types of tensile and compressive loads occur during the take-
off or landing over the wing box sections.
In this study, we are considering the drag forces as the
tensile loads acting over the wing skin. Structural Analyses is
done by varying the tensile load as in Fig.3.
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Fig.3 Boundary Conditions and Loads of Wing Skin.
B. INTEGRAL WING SKIN AND RIB PANEL
i. Geometry
The length and width of the integral wing skin-rib panel
used here are120 mm and 100 mm respectively, and the
thickness of the skin is 5mm. The ribs have a thickness of
4mm and a height of 30mm and are spaced with 60mm. The
crack type included is an edge-type crack over the skin of the
rib panel as shown in the Fig.4.
Fig.4 Wing Skin-Rib Panel with a crack.
TABLE III. GEOMETRIC PROPERTIES OF INTEGRAL WING SKIN-RIB
PANEL
GEOMETRY UNIT (mm)
LENGTH 120
WIDTH 100
RIB SPACING 60
SKIN THICKNESS 5
RIB THICKNESS 4
RIB HEIGHT 30
ii. Meshing
Here also modeling of the structure is done in ANSYS
Software and meshed using tetrahedron-shaped elements. The
whole model is divided into crack propagation regions and
other parts before meshing as shown the Fig.5. The structural
division was adopted, as a strategy to refine the grid in the
crack growth locations, with the cell size of 0.5 mm using a
sphere of influence and the rest with tetrahedral element size
of 1 mm.
Fig.5 Mesh of Wing Skin-Rib panel.
iii. Loads and Boundary Conditions
During flight conditions, lot of internal forces act on the
aircraft wing box such as the shear, bending moment, and
torque. For the crack to open, which is present on the upper
integral wing rib panel, compression stress caused by the
bending moment acts as the main force for its crack
propagation. So, in this analysis, the stress intensity at the
crack tip is found by focusing that the panel is under bending
load only as shown in the Fig.6.
Fig.6 Boundary Conditions and Loads of Wing Skin-Rib panel.
C. INTEGRAL WING PANEL OF WING SKIN AND
STRINGERS
i. Geometry
The integral panel of wing skin and stringer used here is an
I-type, 3-stringer panel. An I-type stringer is known for its
good strength under tensile loads. The dimensions used here
for the panel are 6mm for skin thickness, 4mm for I-type
flange and web thickness, 40 mm for web height, 30mm as the
width of upper flange, and 60mm for the lower flange width.
The spacing between stringers is 140mm for a panel span of
1400mm. The crack type introduced on the structure here is a
semi-elliptical surface crack between the stringer and skin as
shown in the Fig.7.
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Fig.7 Design of Wing Skin-Stringers with Semi-Elliptical Crack in between.
TABLE IV. GEOMETRIC PROPERTIES OF WING PANEL OF WING SKIN-
STRINGERS
GEOMETRY UNIT (mm)
SKIN THICKNESS 6
FLANGE THICKNESS 4
WEB THICKNESS 4
WEB HEIGHT 40
UPPER FLANGE LENGTH 30
LOWER FLANGE LENGTH 60
SPACING BETWEEN STRINGERS 140
ii. Meshing
Modeling and analysis of the structure are done here with
the help of FEA Solver software Ansys. The structure is
modeled with the given dimensions along with semi-elliptical
surface crack. The meshing is done using tetrahedral-shaped
elements of size 10mm, over the surface crack regions refined
mesh is employed by giving 20 as the number of divisions.
Material properties of aluminum alloys are applied as shown
in the Fig.8.
Fig.8 Mesh at the Wing Skin-Stringers.
iii. Loads and Boundary Conditions
Among all the main forces that act on the wing during its
flight, here for present analysis of the stringer panel bending
load is taken into consideration. Thus, the crack opening in the
integral lower panel of the wing is due to the tensile stress
caused by bending. The boundary condition and varying
bending moment load is applied over the model as shown in
the Fig.9.
Fig.9 Boundary Conditions and Loads of Wing Skin-Stringers.
D. INTEGRAL FUSELAGE PANEL WITHOUT CUT-OUT.
i. Geometry
The integral fuselage panel is designed with frames and
stringers which support the skin of the fuselage. The panel
consists of 2 frames and 5 stringers equally space with the
distance of 475 mm and 150 mm respectively. Stringer type
modelled here is of simple type with 5mm as its thickness and
height as 20 mm. The frames design used of the cap type with
height and the width of 35 mm and 20 mm respectively with a
thickness of 1.5mm. Structures such as the frames and stringer
act as the load bearing structure for the components. The crack
type introduced on the structure here is a through crack at the
bay of the fuselage panel as shown in the Fig.10.
Fig.10 Fuselage without cutout panel with through crack between the
Frames.
TABLE V. GEOMETRIC PROPERTIES OF FUSELAGE WITHOUT CUTOUT
PANEL
GEOMETRY UNIT (mm)
SKIN THICKNESS 1.5
STRINGER THICKNESS 5
STRINGER HEIGHT 20
FRAME HEIGHT 35
FRAME WIDTH 20
FRAME THICKNESS 1.5
SPACING BETWEEN STRINGERS 150
SPACING BETWEEN THE FRAMES 475
ii. Meshing
Modeling and analysis of the structure are done here with
the help of FEA Solver software Ansys. The structure is
modeled with initial crack over the surface of the fuselage
panel. The meshing is done using tetrahedral-shaped elements
of size 5 mm, over the crack regions mesh is refined with
smaller element size as shown int the Fig.11. Material
properties of aluminum alloys are applied to the model.
Fig.11 Meshed cracked fuselage without cutout panel model.
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iii. Loads and Boundary Conditions
Among all the main forces that act on the fuselage during
its flight, here for present analysis of the fuselage panel cabin
pressure is taken into consideration. The pressure acts over the
skin of the panel and the frame of the panel acts as fixed
supports. The boundary condition and varying pressure loads
applied over the model as shown in the Fig.12.
Fig.12 Boundary Conditions and Loads applied on the fuselage without
cutout
E. INTEGRAL FUSELAGE PANEL WITH CUT-OUT.
i. Geometry
The integral fuselage panel is designed with frames and
stringers which support the skin of the fuselage. The panel
consists of 2 frames and 5 stringers equally space with the
distance of 475 mm and 150 mm respectively. Stringer type
modelled here is of simple type with 5mm as its thickness and
height as 20 mm. The frames design used of the cap type with
height and the width of 35 mm and 20 mm respectively with a
thickness of 1.5mm. Cut out is created over the panel with
dimensions as 300 mm for height and 200 mm for width and
the filleted edges of the cut-out with a radius of 80 mm, this
cut-out is supported by an additional frame of thickness 2 mm.
The crack type introduced on the structure here is a through
crack at the bay of the fuselage panel as shown in the Fig.13.
Fig.13 Fuselage with cutout panel with a Through Crack near the cutout.
TABLE VI. GEOMETRIC PROPERTIES OF FUSELAGE WITH CUTOUT PANEL
GEOMETRY UNIT (mm)
SKIN THICKNESS 1.5
STRINGER THICKNESS 5
STRINGER HEIGHT 20
FRAME HEIGHT 35
FRAME WIDTH 20
FRAME THICKNESS 1.5
CUT OUT WIDTH 200
CUT OUT HEIGHT 300
SPACING BETWEEN STRINGERS 150
SPACING BETWEEN THE FRAMES 475
ii. Meshing
The structure is modeled with initial crack over the surface
of the fuselage panel. The meshing is done using tetrahedral-
shaped elements of size 5 mm, over the crack regions mesh is
refined with smaller element size as shown in the Fig.14.
Model is defined with material properties of Aluminum alloys.
Fig.14 Meshed cracked fuselage with cutout panel model.
iii. Loads and Boundary Conditions
Among all the main forces that act on the fuselage during
its flight such as the torsion tension and compression caused
due to the wing loadings, here for present analysis of the
fuselage panel cabin pressure is taken into consideration. The
pressure acts over the skin of the panel and the frame of the
panel acts as fixed supports. The boundary condition and
varying pressure load applied over the model as in the Fig(15).
Fig.15 Boundary Conditions and loads applied on fuselage with cutout
panel.
V. RESULTS AND DISCUSSION
The analysis of crack on different aircraft structures at
different positions using ANSYS software has been carried
out and the results are presented in this section.
A. WING SKIN
The analysis is required for finding the lift and drag
performance at various velocities inputted. The parameters for
the analysis of the airfoils were
Stress analysis was conducted over the wing panel, to
identify the maximum and minimum stress contour regions.
These regions with maximum stresses will always initiate the
crack growth. Using finite element tools stress intensity
factors are estimated over the crack tips/front.
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i. Wing skin If crack is present at Rivet edges
Fig.16 Analysis of crack at the rivets.
SIFS (K1) Maximum (Pa mm0.5) vs Half-Crack Length in
mm at Wing skin for crack present at Rivet edge:
In this we can show that SIFS increases for both the material
as the crack length increases, and the material will fail as the
SIFS will increase beyond the fracture toughness of the
material.
Fig.17 Half Crack Length Vs SIFS of Wing skin at rivet edge – Al 2024.
Fig.18 Half Crack Length Vs SIFS of Wing skin at rivet edge – Al 6061.
ii. Wing skin If crack is present between the Rivet edges
Fig.19 Analysis of crack between rivet edges.
SIFS(K1) Maximum (Pa mm0.5) vs Half-Crack length in mm
at Wing skin for crack between the Rivet Edges:
If the crack is present between the Rivet edges, then we can
show that SIFS increases for both the material as the crack
length increases, and the material will fail as the SIFS will
increase beyond the fracture toughness of the material.
Fig.20 Half crack length Vs SIFS of wing skin between rivet edges – Al
2024.
Fig.21 Half crack length Vs SIFS of wing skin between rivet edges – Al
6061.
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B. INTEGRAL WING SKIN AND RIB PANEL
Here stress analysis of the structures is conducted by
varying bending moments, from which we find the maximum
and minimum stress and strain occurrence over the structure.
From the Von-Mises Stress contours, rib regions are more
stressed than others. At the crack tip or the crack front, the
stress intensity values are determined, from which these
values later utilized for the prediction of the number of life
cycles remaining in their service before their failure.
Fig.22 Analysis of crack of wing Skin-Rib panel.
SIFS (K1) Maximum (Pa mm0.5) vs Half-Crack Length in
mm for Integral Wing Skin and Rib Panel:
If the crack is present at Rib Panel, then we can show that
SIFS increases for both the material as the crack length
increases, and the material will fail as the SIFS will increase
beyond the fracture toughness of the material.
Fig.23 Half crack length Vs SIFS of Integral Wing Skin-Rib Panel – Al
2024.
Fig.24 Half crack length Vs SIFS of Integral Wing Skin-Rib Panel – Al
6061.
C. INTEGRAL WING PANEL OF WING SKIN AND STRINGERS
Here stress analysis of the structure is carried out by
varying its bending moments, from which the maximum and
minimum stress and strain values are estimated. Under the
current type of bending condition, the stress is more
accumulated at fixed regions over the upper flange. SIFs
values over the cracked surface are determined for each
loading conditions by varying crack length as well, to estimate
the service life left before the failure.
Fig.25 Analysis of crack of Wing Skin-Stringers.
SIFS (K1) Maximum (Pa mm0.5) vs Half-Crack Length in
mm for Integral Wing Panel of Wing Skin and Stringers.
If the crack is present at Wing Skin and Stringers, then we can
show that SIFS increases for both the material as the crack
length increases, and the material will fail as the SIFS will
increase beyond the fracture toughness of the material.
Fig.26 Half Crack Length Vs SIFS of Wing Skin-Stringers – Al 2024.
Fig.27 Half Crack Length Vs SIFS of Wing Skin-Stringers – Al 6061.
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D. INTEGRAL FUSELAGE PANEL WITHOUT CUT-OUT
Here stress analysis of the fuselage panel is carried out by
varying its pressures, from which the maximum and minimum
stress and strain values are estimated. Under the current load
condition of pressure acting over the skin of the fuselage, the
stress is distributed similarly over many regions and is more
over the free edges. SIFs values over the cracked surface are
determined for each loading conditions by varying crack
length as well, to estimate the service life left before the
failure.
Fig.28 Stress Analysis of cracked model of Fuselage without cutout panel.
SIFS (K1) Maximum (Pa mm0.5) vs Half-Crack Length in
mm for Fuselage Panel without cutout:
If the crack is present at Fuselage without cutout panel, then
we can show that SIFS increases for both the material as the
crack length increases, and the material will fail as the SIFS
will increase beyond the fracture toughness of the material.
Fig.29 Half crack length Vs SIFS of Fuselage without cutout panel – Al
2024.
Fig.30 Half crack length Vs SIFS of Fuselage without cutout panel – Al
6061.
E. INTEGRAL FUSELAGE PANEL WITH CUT-OUT
Here stress analysis of the structures is conducted by
varying pressure, from which we find the maximum and
minimum stress and strain occurrence over the structure. From
the Von-Mises Stress contours, stress distribution over the
skin is seemed to be equal over all skin except for the regions
near the supporting structure of the fuselage. At the crack tip
or the crack front, the stress intensity values are determined,
from which these values later utilized for the prediction of
the number of life cycles remaining in their service before
their failure.
Fig.31 Stress Analysis of cracked model of Fuselage with cutout panel.
SIFS (K1) Maximum (Pa mm0.5) vs Half-Crack Length in
mm for Fuselage Panel with cutout:
If the crack is present at Fuselage with cutout panel, then we
can show that SIFS increases for both the material as the crack
length increases, and the material will fail as the SIFS will
increase beyond the fracture toughness of the material.
Fig.32 Half crack length Vs SIFS of Fuselage with cutout panel – Al 2024.
Fig.33 Half crack length Vs SIFS of Fuselage with cutout panel – Al 6061.
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F. FATIGUE LIFE
Here the Fatigue life is calculated with the help of
MATLAB coding. Plotting a graph against the Number of
Cycles (N) versus the Fatigue Crack Length (a) for a Wing
Skin as shown in the Fig.34 and Fig.35.
i. MATERIAL:2024 ALUMINIUM ALLOY
Fig.34 Number of Cycles (N) Vs the Fatigue Crack Length (a) for a Wing
Skin – Al 2024.
ii. MATERIAL:6061 ALUMINIUM ALLOY
Fig.35 Number of Cycles (N) Vs the Fatigue Crack Length (a) for a Wing
Skin – Al 6061.
From the graph, we can conclude that the fatigue crack
growth is directly proportional to the number of cycles that is
the length of the crack is increasing as the number of cycles
increases.
G. SERVICE LIFE
The service life of aircraft Structural components
undergoing random stress cycling was analyzed by the
application of fracture mechanics using MATLAB coding.
Hence from the equations (5) the results obtained are the
number of Remaining Flights for the different structural
components having a different crack length.
TABLE VII. ESTIMATION OF SERVICE OF SKIN CRACK BETWEEN
RIVETS EDGES
SKIN CRACK BETWEEN RIVETS EDGES (THROUGH CRACK)
ALUMINIUM 2024 ALUMINUM 6061
Crack location parameter (A) = 1 Half-length of the crack (a) = 2.5
mm
Depth of the crack (c) = 1.5 mm operational peak stress factor (f)=
0.6
flaw magnification factor (Mk) = 1.6 uniaxial tensile stress (Si)= 66.67
MPa yield stress (Sy) = 334 MPa
min stress (Smin) = 15.02 MPa max stress (Smax) = 193.91Mpa
Crack location parameter (A) = 1 Half-length of the crack (c) = 2.5
mm
Depth of the crack (c) = 1.5mm operational peak stress factor (f)=
0.6
flaw magnification factor (Mk) = 1.6 uniaxial tensile stress (Si)= 66.67
MPa yield stress (Sy) = 288MPa
min stress (Smin) = 15.02 MPa max stress (Smax) = 193.91MPa
REMAINING SERVICE LIFE
122 Remaining flights 93 Remaining flights
In the above TABLE VII, after inputting all the
predefined values to estimate the Remaining Service life on
MATLAB software, it was found out to be 122 Remaining
flights for the Aluminium 2024 which was having higher
yield stress compared with the 93 Remaining flights of
Aluminium 6061.
TABLE VIII. ESTIMATION OF SERVICE OF SKIN CRACK BETWEEN
RIVETS EDGES
SKIN CRACK AT RIVETS EDGES (THROUGH CRACK)
ALUMINIUM 2024 ALUMINUM 6061
Crack location parameter (A) = 1 Half-length of the crack (a) = 2.5
mm
Depth of the crack (c) = 1.5 mm
operational peak stress factor (f)=
0.6
flaw magnification factor (Mk) = 1.6 uniaxial tensile stress (Si)= 66.67
MPa
yield stress (Sy) = 334 MPa
min stress (Smin) = 7.069 MPa max stress (Smax) = 241.77 MPa
Crack location parameter (A) = 1 Half-length of the crack (c) = 2.5mm
Depth of the crack (c) = 1.5mm
operational peak stress factor (f)=
0.6
flaw magnification factor (Mk) = 1.6
uniaxial tensile stress (Si)= 66.67
MPa yield stress (Sy) = 288MPa
min stress (Smin) = 7.069 MPa
max stress (Smax) = 241.77 MPa
REMAINING SERVICE LIFE
134 Remaining flights 103 Remaining flights
Similarly, as shown in the TABLE VIII it was carried out
for skin crack at rivets edges. It estimated 134 flights for
Aluminum 2024 and 103 flights for Aluminium 6061. This
process was carried out to find Remaining number of flights
for all the other 4 structures at a similar crack length of 5mm.
as shown in TABLE IX.
TABLE IX. ESTIMATION OF REMAINING FLIGHTS FOR DIFFERENT
STRUCTURES
STRUCTURES
REMAINING NUMBER OF FLIGHTS
ALUMINIUM
2024
ALUMINIUM
6061
INTEGRAL WING SKIN AND RIB PANEL
138 103
INTEGRAL WING PANEL
OF WING SKIN AND STRINGERS
218 171
FUSELAGE PANEL
WITHOUT CUTOUT 50 42
FUSELAGE PANEL WITH CUTOUT
183 139
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H. EXPERIMENTAL VALIDATION
i. CRACKED PLATE WITH THREE HOLES
Here in this a model of rectangular plate with dimensions
120 mm X 65 mm X 16 mm was created with two 13 mm
diameter holes near both ends, and a 20 mm hole at a distance
of 51 mm from the bottom of the plate as seen in Fig.36. Just
above the middle of the plate an initial edge crack of length
10mm is created. The size of the mesh element used is as 1
mm, creating a mesh consisting of 83448 nodes and 48024
elements of tetrahedral shape which is shown in the Fig.37.
Aluminium 7075-T6, with the material applied over the
model, and the fatigue load of P = 20 kN with a stress ratio R
= 0.1 is used.
Fig.36 Geometry of a cracked plate with three holes.
Fig.37 Initial Mesh of the Model.
The crack path growth was simulated with ANSYS
software and was compared with both experimental and
numerical results from ABAQUS software obtained by [22]
which showed that they have strong similarities in every
aspect. The distribution of the, the von Mises stress, and the
equivalent strain are shown in Fig.38 and Fig.39 respectively.
The predicted values of the two modes of stress intensity
factors, i.e., KI and KII. As shown in Fig.40, the crack starts
to grow in a straight direction, indicating the domination of KI
followed by a curved direction with an increasing negative
value of the second mode, KII, that results in the crack
growing toward the hole. Present work values of equivalent
stress intensity factor along with fracture toughness line in
Fig.41 indicates that the critical length or unstable cracks
growth occurs at an approximate crack length value of 21 mm
which is similar to the value obtained by prediction done by
Error! Reference source not found..
Fig.38 Equivalent Strain Distribution.
Fig.39 The equivalent von Mises stress distribution.
Fig.40 Predicted values of the first and second mode of stress intensity
factors.
Fig.41 Present work values of equivalent stress intensity factors along with
fracture toughness line.
The validation of the software results was revealed by
comparisons with the numerical results of crack propagation
by ANSYS and the experimental results.
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VI. CONCLUSION
With this thorough study of different types of cracks in
aircraft structures, it is shown that the main reason for aircraft
structural failure is fatigue failure by crack propagation. The
remaining service life of an aircraft structure can be estimated
by the process of Half-cycle method and fatigue crack growth
and the analysis of various structures. The FEM analysis of
crack on different aircraft structures using ANSYS software
has been carried out. Fracture mechanics is used for
predicting the propagation of the crack, and it is performed
for the most common failure mode of fracture mechanics. So,
it is concluded that the crack growth on aircraft structures
cannot be overlooked, and proper maintenance with
scheduled test intervals needs to be carried out for better
service life.
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International Journal of Engineering Research & Technology (IJERT)
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IJERTV10IS050328(This work is licensed under a Creative Commons Attribution 4.0 International License.)
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