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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