 Open Access
 Authors : Manjunath B G , C N Yashaswini , Muskan Rastogi , Shahid Adnan, Srikanth Salyan
 Paper ID : IJERTV10IS050328
 Volume & Issue : Volume 10, Issue 05 (May 2021)
 Published (First Online): 02062021
 ISSN (Online) : 22780181
 Publisher Name : IJERT
 License: This work is licensed under a Creative Commons Attribution 4.0 International License
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
Srikanth Salyan5
Manjunath B G2
B.E Aeronautical Engineering Dayananda Sagar College of Engineering
Bangalore, India
Shahid Adnan4

Aeronautical Engineering Dayananda Sagar College of Engineering
Bangalore, India
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.

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. Takeoffs 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.

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.

Aluminium Alloy 2024
2024 Aluminium alloy is an alloy with copper as the primary alloying element. It is used in applications requiring a high strengthtoweight ratio, as well as good fatigue resistance. It has poor corrosion resistance. It is mostly used to make the aircrafts structural parts such as wing and fuselage.

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


METHODOLOGY

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
crack sizes at the critical stress points for the fatiguecrack growth analysis were established on the structure. The fatigue crack growth rates for random stress cycles were calculated using the halfcycle method. The equation (5) was developed for calculating the number of remaining flights remaining for the structural components.

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 a p and a 0 are calculated respectively from equation (6)
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.
c c
& (7),
… (6)
… (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.
K = 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 secton stress is greater than the yield strength of the material, then wing skin fails due to plastic collapse. Net section stress is calculated by,
… (7) Where, * and f* (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),
… (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
Here, Q can be expressed as,
Where, is the Uniaxial tensile stress, y is the Yield stress,
E(k) is the Elliptic function.
… (9)
… (10)
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(1R)*1.08899+0.04369*(a/b)1.77302*(a/b)2
+9.212*(a/b)315.8683*(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.


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
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 halfcycle
theory. The halfcycle 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 onehalf of the damage caused by a fullcycle 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 microdamages caused by the individual halfwaves 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, max & min are maximum and minimum stresses constant amplitude stress cycles.


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 softwares 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 CutOut.

Integral Fuselage Panel with CutOut.

WING SKIN

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

Meshing
The modeling of wing skin is done with two riveted holes in it and threedimensional fournode 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.

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.
Fig.3 Boundary Conditions and Loads of Wing Skin.


INTEGRAL WING SKIN AND RIB PANEL

Geometry
The length and width of the integral wing skinrib 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 edgetype crack over the skin of the rib panel as shown in the Fig.4.
Fig.4 Wing SkinRib Panel with a crack.
TABLE III. GEOMETRIC PROPERTIES OF INTEGRAL WING SKINRIB PANEL
GEOMETRY
UNIT (mm)
LENGTH
120
WIDTH
100
RIB SPACING
60
SKIN THICKNESS
5
RIB THICKNESS
4
RIB HEIGHT
30

Meshing
Here also modeling of the structure is done in ANSYS Software and meshed using tetrahedronshaped 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 SkinRib panel.

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 prsent 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 SkinRib panel.


INTEGRAL WING PANEL OF WING SKIN AND STRINGERS

Geometry
The integral panel of wing skin and stringer used here is an Itype, 3stringer panel. An Itype stringer is known for its good strength under tensile loads. The dimensions used here for the panel are 6mm for skin thickness, 4mm for Itype 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 semielliptical surface crack between the stringer and skin as shown in the Fig.7.
Fig.7 Design of Wing SkinStringers with SemiElliptical 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

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 semielliptical surface crack. The meshing is done using tetrahedralshaped 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 SkinStringers.

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 SkinStringers.


INTEGRAL FUSELAGE PANEL WITHOUT CUTOUT.

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

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 tetrahedralshaped 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.

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


INTEGRAL FUSELAGE PANEL WITH CUTOUT.

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 cutout with a radius of 80 mm, this cutout 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.
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
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
Fig.13 Fuselage with cutout panel with a Through Crack near the cutout. TABLE VI. GEOMETRIC PROPERTIES OF FUSELAGE WITH CUTOUT PANEL

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.

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.


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.

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.

Wing skin If crack is present at Rivet edges
Fig.16 Analysis of crack at the rivets.
SIFS (K1) Maximum (Pa mm0.5) vs HalfCrack 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.

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 HalfCrack 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.


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 VonMises 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 SkinRib panel.
SIFS (K1) Maximum (Pa mm0.5) vs HalfCrack 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 SkinRib Panel Al 2024.
Fig.24 Half crack length Vs SIFS of Integral Wing SkinRib Panel Al 6061.

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 SkinStringers.
SIFS (K1) Maximum (Pa mm0.5) vs HalfCrack 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 SkinStringers Al 2024.
Fig.27 Half Crack Length Vs SIFS of Wing SkinStringers Al 6061.

INTEGRAL FUSELAGE PANEL WITHOUT CUTOUT
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 HalfCrack 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.

INTEGRAL FUSELAGE PANEL WITH CUTOUT
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 VonMises 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 HalfCrack 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.

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.

MATERIAL:2024 ALUMINIUM ALLOY
Fig.34 Number of Cycles (N) Vs the Fatigue Crack Length (a) for a Wing Skin Al 2024.

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.


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 Halflength of the crack (a) = 2.5 mm
Depth of the crack () = 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 Halflength 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 Halflength 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 Halflength 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

EXPERIMENTAL VALIDATION

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 7075T6, 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.


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 Halfcycle 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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