International Journal of Mechanical Engineering and Technology (IJMET)
Volume 10, Issue 03, March 2019, pp. 110-126. Article ID: IJMET_10_03_011
Available online at http://www.iaeme.com/ijmet/issues.asp?JType=IJMET&VType=10&IType=3
ISSN Print: 0976-6340 and ISSN Online: 0976-6359
© IAEME Publication Scopus Indexed
Dr. Hatem Hadi Obeid
Faculty of Engineering/ Department of Mechanical Engineering
University of Babylon
ABSTRACT
In this paper, the technique of structure health monitoring was applied to detect damage occurrence in the stiffeners of wing structure. A wing structure of airfoil shape according to digital NACA 0015 was considered for modelling the aerodynamic wing shape. Finite element models of the wing structure were created included undamaged and damage wing structures using high order eight nodes and six nodes shell elements. The damaged wing models included damaged at the main spar at 20%,
40%,60% and 80% the distance than the root chord. A mathematical model was developed using high order shell elements and programmed via MATLAB. The load cases were predicted experimentally using wind tunnel test on a wood airfoil prototype such that the pressure distributions through upper and lower wing skins measured at different attack angles. The static solutions were achieved for each finite element model under action of maximum pressure distribution at attack angle of 10
0
.
Strains distributions and lateral deflections through lower skin were estimated. The results of strains through lower skin showed occurring of climax in strain values initiated in the distribution at the location of the damage for all the cases of damaged structures, which not appeared in the undamaged wing structure. The results of lateral deformations through lower skin showed indicating variations in the shape of the curves, such that the curve appeared to be gradually increased the region between the chord and the damage. Semi to rapid increasing in the lateral deformation occurred in region between the damage and the end tip of wing. The difference in the behavior of the induced strains and lateral deformation between undamaged and damaged wing structures is intended to a technique predicting of the occurrence of the damage in the wing structure. The present of a climax in the strain distribution lead to presence of damage near the climax value of strain. Also, the indicating of variations in the shape of the lateral deformations of skins semi or rapid increasing lead to presence of damage near the rejoin of that variation. It is concluded the possibility of using the strain and deflection analysis for the wing structure to investigate occurring damage
http://www.iaeme.com/IJMET/index.asp 110 editor@iaeme.com

Detection of Damage in Stiffeners of Aircraft Wing Structure Based On Induced Skin Strains and
Lateral Deflections in the structure due to variation in the strains and lateral deflections between different damage locations.
Key Words: Wing structure, Structural health monitoring, finite element method, structural analysis, airfoil pressure distributions, strain, stress, deflection, damage in wing structure.
Cite this Article : Dr. Hatem Hadi Obeid, Detection of Damage in Stiffeners of
Aircraft Wing Structure Based On Induced Skin Strains and Lateral Deflections,
International Journal of Mechanical Engineering and Technology, 10(3), 2019, pp.
110-126. http://www.iaeme.com/IJMET/issues.asp?JType=IJMET&VType=10&IType=3
The Structural Health Monitoring of failures in aircraft and other flying equipment is very important because its concerned with the safety of their users or the risks that may be occurred due to falling on the ground. This analysis is applicable in wide range of engineering fields spanning from mechanical, aerospace, military, building, to bridge and transportation applications. Needs for prediction of failure in the aerospace structures are rapidly increasing due to demands of enhance safety, reduce inspection time and cost but maintaining structural function and uses. This field is also providing smart prognosis and diagnosis of failure instantaneously.
The deployment and development of structural health monitoring system is required to ensuring the safe operation of any engineering structure or system. The execution of the structural health monitoring is mainly considering the following processes: [1]
- Using sensors and instrumentation devices for sensing and tracing the interpretation of an engineering system or structure under operational loads.
- Assessment the performance of the system or structure for any propagation of damages or defects by the analysis of the measured values analytically or numerically.
- Indicating a notification or an alarm when the desired system parameters are exceeded.
The base on parameters or criteria that the structural health monitoring system executed are varied depending on the nature of the system responses or performance. Although the importance of structural health monitoring but their researches are relatively young if compared to other engineering fields, thus most of the researches have been conducted at universities or research and development centers of companies.
The applications of structural health monitoring in engineering fields included different problems such detection of damage in steel structure, aerospace, concrete structures, precast concrete, box girders in railroads, composite materials, dikes and pipelines of water and wastewater. Also, evaluation bridge deflection; stress or strain analysis in petroleum pipes, dam structures, or suspension bridge cables. In addition to more complicated problems such as detection of impact effects in wind turbine blades and water leaks. The vibration of highrise reinforced structures such as towers can be evaluated and assessment.
In this paper the structural health monitoring is applied for detection failure at the stiffeners of aircraft wing structure. The performance of an aircraft structure can be characterized by various factors, essentially the response (i.e. strain mapping, shape determination etc.) due to applied loading that can be evaluated by health monitoring technique [2]. Finite element method is powerful to be used to modelling the wing structure for estimation the effects in the structures due to the applied loading. The most wing structural stiffness comes from the interior stiffeners. The stiffeners included main, auxiliary spars and ribs. Having stiffeners (spars and ribs) attaing structure increases the load resistance of the http://www.iaeme.com/IJMET/index.asp 111 editor@iaeme.com

Dr. Hatem Hadi Obeid structure without further increasing in the weight. To further reduce the weight, a certain arrangement of stiffener on structure is being necessary to increase the rigidity with less material. The construction of both stiffeners and outer skin gives the overall wing structural stiffness. The response of wing structure into the loading depending mainly on the overall stiffness of wing. Then the present of damage in any location in the stiffeners will be affecting on the response of the wing to the applied loading.
In this work, a wing structure of a guided aircraft laboratory prototype was modelled via finite element method. Different finite element models were created each represent a case of damage in spar. The purpose of these finite element models is performing static analysis under action of static load and obtain strains induced in the skins and the lateral deformation at the tip of wing. Then the health monitoring technique is applied to estimate relations between damage of stiffeners and the static induced strains and deformations. The geometry configurations and dimensions are illustrated in figure (1), such that the wing consisting of upper and lower skins, spars and ribs. The wing stiffeners are constructed as eight spars and ten ribs connected for giving the wing structure the required stiffness. The four digits symmetrical NACA 0015 airfoil was used as an aerodynamic shape for the considered wing.
The (00) number of the NACA indicating that it has no camber. The (15) number of the
NACA indicates that the airfoil has a “thickness to chord length ratio” of 15% as shown in figure (2).
The equation describing the geometric shape of a NACA 0015 foil is [4,5]: x: is the location along chord from ranged (0 to 100%). y(x) is the half thickness at any location x . t: is maximum thickness of the wing as a fraction of the chord.
(1)
The chords of the wing structure were assigned as 1m and 0.4m at the root and tip respectively with length 1.75m. http://www.iaeme.com/IJMET/index.asp 112 editor@iaeme.com

Detection of Damage in Stiffeners of Aircraft Wing Structure Based On Induced Skin Strains and
Lateral Deflections
In order to study the effect of damage that occurred in the spars, finite elements models were created for each damage case. The damage was assumed to be as a cut in the connection of the main spar which is located at the maximum thickness of wing, such that it assumed that the damage occurred at five locations in the spar nearest to the maximum wing thickness. High order SHELL 8-node quadrilateral element and 6-nodes triangular element ware selected for the generation finite element model of the wing structures. Figure (3) is shown the high order
SHELL 8-Node element parameters and degrees of freedom. The upper and lower skins were discretized into ten divisions, to be generate one hundred shell elements at each skin. The spars and ribs were discretized into ten divisions in order to connected with both and skins element at the contracted nodes to maintain continuous element connections as shown in figure (4). The thickness of all of skins, spars and ribs were assigned to 0.4 mm. Aluminum
2024 alloy -T3 was used as the material of all parts their mechanical properties are shown in table (1). [7]
For the typical shell elements as shown in Figure (3), the external faces (surfaces) of the element are curved, while the thickness is generated by straight lines. Many coordinate sets employed in the formulation of degenerated shell element. These coordinate systems are described in the following sections.
It is the global Cartesian coordinate system, used to describe the wing geometry in the space.
The coordinates and deflections of nodes can be described using this coordinates system. In addition to describe the assembled global stiffness matrix and assembled force vector. The following notation is used:
X i
(I=1,3) and X1 =X , X2=Y , X3=Z ,
U i
(I=1,3) and U1=U, U2 =V, U3=W http://www.iaeme.com/IJMET/index.asp 113 editor@iaeme.com

Dr. Hatem Hadi Obeid
X i
(i=1,3) is a unit vector in the
X i
direction.
This system is used to define global stiffness matrix and applied force vector and in the structure geometry, as well as nodal coordinate and displacements are referred to this system.
This system type can be established at each nodal point as shown in Figure (3) with origin located at reference surface (shell mid-surface), V3f is a vector formulated from nodal coordinate of shell surfaces at node f,
V
3 f

X f top 
X f bot
,
V
3 f


X f
Y f
Z f

T
top


X f
Y f
Z f

T bot 2
The unit vector
V
3 f can normalize from the vectorV3f, thus:
V

3 f h f

V
3 f
 x
V
3 f y
V
3 f
V
3 z f

T h f


ΔX
1 f

2  
ΔX
2 f

2  
ΔX
3 f

2

1 / 2
ΔX if

X top if

X bot if
 i

1,2,3

3
4
5
The vector
V
3f is the direction of “normal” at the node f, which is not necessarily perpendicular to the mid-surface at f.
The vector V2f is normal to plane of vectors V1f and V3f: V2f = V1f

V3f .
The mathematical characteristic of vector V1f is perpendicular to V3f and parallel to the global X – Z plane, thus:
V x
1f
 z
V
3f
, V y
1f

0 .
0 , V z
1f

V
3f x
Or, if the vector is in the Y-direction

V
3f x 
V
3f z 
0 .
0

V
1f x   z
V
3f
, V
1f y 
0 .
0 , V
1f z 
x
V
3f
       
The description of this system is established by the definition of three coordinates, each one has its characteristic as shown in Figure (3). The
 coordinate is an intrinsic coordinate through thickness direction.

coordinate is defined as a function of V
3f
, the

-direction is considered normal to the shell mid-surface. It is assumed that the curvilinear coordinate varies between (-1 and +1), which represent the top and bottom surfaces of the element respectively.
The two other coordinates are
ξ, η
which are curvilinear coordinate in the middle plane of the shell element and are also assumed to vary between (-1 and +1) which represent the faces of the element.
X
' i or X
'
, Y
'
, Z
'
This system is a cartesian coordinate system defined at the sampling points wherein stresses and strains are to be calculated. The direction
'
Z is taken perpendicular to the mid-surface (
  constant
),
X
'
, Y
' and Z
'
X ' and Y ' tangent it as shown in Figure (3). The vectors
direction of mid-surface respectively, and it should be noted that,
V
1
'
,
V
2
'
,
V
3
' define http://www.iaeme.com/IJMET/index.asp 114 editor@iaeme.com

Detection of Damage in Stiffeners of Aircraft Wing Structure Based On Induced Skin Strains and
Lateral Deflections
ξ 













 x
ξ y
ξ z
ξ








,
η 













η z
 x
η y
η








,
V
1
  ξ 













 x
ξ y
ξ z
ξ








,
V
3
  ξ  η 













 x
ξ y
ξ z
ξ







 x












X


η
Y
η
Z
η








[T] =

V
1
'
,
V
2
'
,
V
3
'

V
1
'
, V
2
' , V
3
'
: unit vectors on X
'
, Y
' and Z
' .
6
7
To obtain the formulation of the coordinates of any point within the element, it is simpler to divide the general formulation of coordinate in two parts. The first part is to establish the term of formulation, which represents the intercept of the “normal” with the mid-surface. If there is a node (f) at the “normal” at  =0 surface and this “normal” has a top and bottom nodes at 
=+1 and

=-1 surfaces respectively, the coordinates of node (f) can be obtained as follows:
X ' f

1
2
   f top

  f bot

,





X
Y f
Z f f






1
2





X
Y f
Z f f




 top






X
Y
Z f f f 



 bot
8
Since node (f) is in the mid-surface (at function
N f


,



=0 surface), the two dimensional interpolation can be applied from Table (2) , to obtain relationship between the Cartesian point and the curvilinear coordinate as follows:
X i
 f n


1
N f


,


X if
9
The second part is to establish the second term of the general formulation, which defines the position of the point along this “normal”. But to establish that, it is necessary to define arbitrary point (p) on the “normal” at node (f), therefore the vector V3f which represents the thickness of the shell hf, can be written ass: h f

V
3 f





X
Y
Z f f f



 top





X f
Y
Z f f



 bot 10
The distance between the arbitrary point (p) and the node f (at mid-surface) can be written as follows: h f
ζ
Unit Distance = 2 11 http://www.iaeme.com/IJMET/index.asp 115 editor@iaeme.com

Dr. Hatem Hadi Obeid
Since the
X i
 f n 

1 h f

V
3 i f means the coordinates of the arbitrary point, the equation (11) can be rewritten with the application of the two-dimensional interpolation function
N f
  from Table () so that:
N f
  ζ
2
V
3 i f
12
The general formulation of the coordinate defines the geometry of the shell element which represents the rotations between the coordinate (
ξ   η   ζ 
) and coordinates (X, Y, Z). It is obtained by the summation of the previous two parts (equation (8) & equation (12)) so that:
X i
 f n


1
N f
(

,

)
1

2
ζ   if top
 f n


1
N f
(

,

)
1

2
ζ   if bot
13

X



Y
Z


 f n


1
N f
(

,

)

X


Y
Z


 mid
 f n


1
N f
(

,

)
ζ h f
2




 
V
V
3 y f
V x
3 z
3 f f




 
X i
is the elemental Cartesian coordinate, (
X
1

X
,
X
2

Y
,
X
3

Z
), n: is the number of nodes per element, hf is the element thickness at node f, i.e. the respective “normal” length,
Xif is the Cartesian coordinate of nodal point f, and interpolation functions. (
ζ
=constant) at nodal point f.
N f
(

,

)
is the two dimensional
Each node of the high order element has five degrees of freedom representing the displacement along local coordinates. It is assumed that the strains in the directions to the mid-surface is assumed to be negligible. The deflections through the element mid surface can be defined by the three displacements (u,v,w), and two rotations of the nodal vector V3f about orthogonal directions normal to it. One of the two orthogonal directions is represented by unit vector V
1 f and the corresponding rotation (
α
1 f ). The other directions is
V
2 f and the corresponding rotation is (
α
2 f
), and it is the displacements (
δ
1 f
, δ
2 f ) of a point at unit distance
(h) from node f on the “normal” resulted from the two rotations (
α
1 f ,

2 f
) are calculated as follows:
α
1 f

δ
1 f h or
δ
1 f
 h
α
1 f
α
2 f

δ
2 f h or
δ
2 f
 h α
2 f
 h h where 2 f

,
δ
1 f
is the displacement in the direction of
V
1 f
14
,
δ
2 f is the displacement in the negative direction of
V
2 f . Equation (13) can be written as follows:
δ
δ
1
2 f f
 h 
α
α
1 f
2 f
The global displacement can be found from:
  i
 f n 

1
N f
 
οf mid
 
N f u
15
16 http://www.iaeme.com/IJMET/index.asp 116 editor@iaeme.com

Detection of Damage in Stiffeners of Aircraft Wing Structure Based On Induced Skin Strains and
Lateral Deflections u i : the nodal displacement through element thickness. u i
 f
Cartesian coordinate.
rotations

1 f and

2 f u
'
 i f
.
: nodal displacement through
:nodal displacement through “normal” of the cross product of
Since; u i
οf
 δ
1
' f
 δ '
2 f
17 where the corresponding displacement components (
 '
1 f follows:
,
 '
2 f
) of
' u  f can be calculated as
 '
1 f
 h f
2


1 f i
V
1 f
and
 '
1 f
 h f


1 f
V i
1 f
2 18
Since the global displacement can be found as shown before in equation (16) as the simulation of the mid-surface nodal displacement ( caused by the rotations of the normal ( u
' i
 f u i
 f
) and the relative displacements are
), then the element displacement field can be expressed by: u if
 u i
οf

ζ
.
h f
2 u v w
 f n 

1
N f
( V i
1 f
α
1 f

V i
2 f
α
2 f
)

 u v f f w f


 f n 

1
N f
ζ h f
2



V x
1 f y
V
1 f z
V
1 f
V
2 x f
V
2
V z
2 y f f




α
1 f
α
2 f
19
20
The contribution to the global displacement from a given node f in the general form and for complete element is: u i
 f n 

1
N f

ξ, η, ζ

S f
21 u v w









N f
0
0
0
N f
0
0
0
N f
ζ h
N f
N f
ζ
2 h f f
N f
2
ζ h f
2
V
1 x f
V
1 y f
V
1 z f

N f

N f

N f
ζ h
ζ
2 h f f V
2 x f
V
2 y f
2
ζ h f V z
2 f
2













 u v w
α
1
 α
2 f f f f f





 
Nf is the shape function matrix of the degenerated shell element.
22
 u
 f
Sf is transformation matrix of the displacement vector at node f of shell element
, v
 f
, w
 f
,

1 f
,

2 f

T
.
The stress and strain components for the shell assumption of zero local stress through shell mid-surface along Z '
-direction (
 ' z

0
) and using Hooks law enables the stresses vector to be reduced to following five stress components,
ζ 





η


η
η
ζ
ζ x x x y '
'
' y
'
' z ' y ' z '








      
23
is the initial strain vector. http://www.iaeme.com/IJMET/index.asp 117 editor@iaeme.com

Dr. Hatem Hadi Obeid
 
is the strain vector.
[D] is the elasticity matrix given by,

1

E
ν 2 





1
ν
0
0
0
ν
1
0
0
0
0
0
G
0
0
K
0
0
0
1
G
0 K
0
0
0
0
2
G 





G: modulus rigidity, E: modulus of elasticity,

: Poisson’s ratio,
K1, K2 is the shear correction factors.
24
The normal strain in the Z
' -direction (
 ' z
) is neglected. Therefore, the general vector of green strains it will be reduced to the following five components,











ε
ε x x y x
' y
'
' y
' z '
' z
'
'























 u

 u y
' z v z
'
'
'


 u x v
'
'
'
'
'

 y '

 v

 x w


 x w
 y '
'
'
'
'
'












The local derivatives above of the displacement components coordinates system (
X '
1
) can be obtained as:
25 u
'
, v
' and w
' in the local












 u x u y u z
'
'
'
'
'
'


 v ' x v '
'


 z ' y
' v '
 w '

 x w
'
'

 y w '
'
 z '








 
T












 u x u y u z
 v



 x v y v
 z


 w

 y w
 x w z







[

] is the transformation matrix given by,












 u x u y u z
[

]













 x z
'
 x x y x '
'
 v

 x v

 y v
 z
 x

 y ' y

 y ' z
 y '


 x

 z z
'
 z y z '
'








 w

 x w

 y w
 z








 

1













 u
ξ u
η u
ζ
 v


ξ v


η v
 ζ
 w

 ξ w

 η w
 ζ








[J]














 x
ξ x
η x
ζ
 y


ξ y


η y
 ζ


 z
ξ z


η z
 ζ








[J] is Jacobian matrix
26
27
28
29 http://www.iaeme.com/IJMET/index.asp 118 editor@iaeme.com

Detection of Damage in Stiffeners of Aircraft Wing Structure Based On Induced Skin Strains and
Lateral Deflections
The transpose of the Jacobian matrix results from equation (12) can be expressed as follow:
[ J ] T
 f n


1




 
X
Y
Z f f f



N f ,

N f,

0

 h f
2

V


V
V x
3 f y
3 f z
3 f




0 0 N f

 h f

2

V


V
V x
3 f y
3 f z
3 f




N f ,

N f,

0 



  n is the number of node per element,
30 hf is the element thickness at node f, and
N f ,



N f
 

,
; N f ,



N f
 
 
31 are the derivatives of the global displacements referred to the curvilinear coordinates are
 u obtained from equation (18), such as   . . .etc.
 u
 
 f n 

1

 N f ,
 u f
 h f

V
1 x f
N f ,


1f
2
 h f

2
V
2 x f
N f ,


2f


32
The strain matrix B can be formulated from equations (21), (29) and (31) as:

    where,
 

 u , v , w ,
α
1
,
α
2

T
33
34
[B]=







 a
0
0 c i i b i
0 c i
0 b i a i
0 b i a i
0
0 d i
V i
1 x
(e i
V e i
V i i
1 x

1 y d i
V i
(g i
V i
1 y
 e i
V i
1
1 y
) z
)
(d i
V i
1 z
 g i
V i
1 x
)
(-
(-
(e i
V g i
V d i
V
-
i i i d i
V i e i
V i
2
2 x
1 y
2 z



2 x x d i
V i e i
V g i
V i i
2
1
2 x y
) z
)
) 







35 a i

J *
11
N i
,
ζ 
J *
12
N i
,
η d i
 h i
2
(a i z ,

J *
31
N i
) b i

J *
21
N i
,
ζ 
J *
22
N i
,
η e i
 h i
2
(b i z ,

J *
23
N i
) c i

J *
31
N i
,
ζ 
J *
32
N i
,
η g i
 h i
2
(c i z ,

J *
23
N i
)
Consequently, it is used in the calculation of the stiffness matrix [K] using the midcoordinate rule. Hence [K] can be defined as follows:
[K]


1
  

1

1

1



1

1
[B] T [D][B] J

ξ, η, ζ
 dζ

 dξ d
η
Then K can be written as summing up the contribution of each layer at the gauss points,
36
[K]


1
  

1

1

1

 nL j

1
[B j
] T [D j
][B j
] J

ξ, η, ζ
 2 Δh h j

 dξ dη
[K] : stiffness matrix, [D] : elasticity matrix, [Bj]: strain-displacement matrix.
J (

,

,

)
is the determinant of the Jacobian matrix for layer ( j ).
 h j is the thickness of the jth layer. nL is the total number of layers.
In the same way, the internal force vector { f e
} can be determined as follows:
{ f e
}
  v
[ B ]
T
{

} J dV
37
38 http://www.iaeme.com/IJMET/index.asp 119 editor@iaeme.com

Dr. Hatem Hadi Obeid
{f e }



 j nL

1

1   

1

1

1
[B j
] T {ζ j
} J(ξ(
η, ζ)
2
Δh j h 

 dξ dη
{
 j
}
: is current stress vector,
{ f e
}
: is the internal force vector.
39
It should be noted that, it is essential in nonlinear analysis to determine the internal force vector (or equivalent nodal forces) at the end of each iteration. The transformation to the local coordinates system (
X
' 
Y
' ) using the following relation:
[D ' s
]

[S] T [D s
][S]
[S]








 l m
2
2
0
0
2 lm m
2 l 2
2 lm
0
0 l 2 lm
 lm
 m 2
0
0
0
0
0
 l m
0
0
0 m l







40
41
The wing is clamped through the root chord such that all the degrees of freedom of the regarded nodes are fixed to be zeros. While all other nodes are freely to be of three translations and three rotations along the corresponding local coordinates.
In order to predict the load that applied on the wing structure, a prototype for the airfoil was manufactured using wood material according to the digital naca number 0015. Then the prototype was amounted in the wind tunnel section as shown in figure (5). Its required to measure the pressure distribution through the airfoil, thus twelve manometers were used for this purpose. The wind tunnel is flowing air at 50 m/s over the airfoil. The airfoil was mounted with different attack angles (0 to 10 degrees). Figures (6),(7) showed the pressure distributions through upper and lower skins with different attack angles. The net pressure distribution is equal to the difference between the pressures through upper and lower skins.
The pressure distribution was measured for each attack angle which is represent the effect of aerodynamic lifting of the wing. http://www.iaeme.com/IJMET/index.asp 120 editor@iaeme.com

Detection of Damage in Stiffeners of Aircraft Wing Structure Based On Induced Skin Strains and
Lateral Deflections
Static analysis solution has been included the calculation of the effects of the applied load under steady loading conditions on a structure. The pressure distributions through upper and lower skins were applied as normal pressures on the elements. The static analysis is governing by the following equilibrium equation:
 
    
42
The above equilibrium equation is solved by jacobi iterative method to obtain the displacement [q] vector.
Naman Jain [11,12] was created a Finite element code to solve shell structure under action of static loads. In this work, a development was achieved for using Naman Jain code to generate stiffened shell structure using quadrilateral eight nodes shell element and triangular six nodes shell element. Jacobi iterative method was used to solve equation (42) to obtain the deformations along global coordinates. Then the deformations transformed using equation
(41) to the local coordinates. The local deformations were transformed into global coordinates x,y and z directions. The strains were estimated through local coordinates and substituted in the stress strain relations to obtain elemental stresses through local coordinates. The first run included the healthy structure that did not subjected into damage. Figure (8) showed the safe strains induced in the lower skin along the intersection of the skins and main spar. Its noted that the strains in the healthy wing were decreased gradually in the rejoins moved away than the wing chord. The same behavior was noted in the strains induced in the upper and lower skins. Figure (9) showed the safe lateral deformations induced in the lower skin along the intersection of the skins and main spar. The lateral deformations of both the lower and upper skins at the intersection of skins with main spar were approximately coincident. Also, the lateral deformations were increased gradually in the rejoins moved away than the wing chord.
A similar behavior was noted in the study of Salu Kumar Das and Sandipan Roy [13]. The second run included the static solution of the finite element model subjected to damage at the main spar at the location apart 20% than the chord. The damage simulated as a cut in the element connection of the spar and skins. The strain distributions are illustrated in figure
(10), where noted that a climax is initiated in the distribution at the location of the damage. In addition to variation in the shape of the distribution of the strains in the rejoin between the damage and the end tip of the wing. Figure (11) showed the distribution of the lateral http://www.iaeme.com/IJMET/index.asp 121 editor@iaeme.com

Dr. Hatem Hadi Obeid deformations of the lower skin at the intersection of the main spar with skins. The distribution indicated that there are two shapes, the first shape in the rejoin between the chord and the damage where the distribution was gradually increased. The second shape is between the damage and the end tip of the wing, where the distribution appeared to be semi gradually comparative with the first region. This is behavior due to the present of damage at the location
20% than the chord. In which the stiffness if reduced rapidly at that location. So, this behavior can be used as a key to detect the damage location. The other runs were executed for the finite element models that included the damages at the locations apart than the chord by 40%, 60% and 80%. The behavior of the strain distribution indicates a decreasing towards the end tip of the wing such that initiated a sudden climax at the location of the damage as shown in figures
(12, 14, and 16). The behavior of the lateral deformations indicated variations the shape of the curve, such that the curve appeared to be gradually increased the region between the chord and the damage. Semi to rapid increasing in the lateral deformation occurred in region between the damage and the end tip of wing as shown in figures (13, 15, and 17). http://www.iaeme.com/IJMET/index.asp 122 editor@iaeme.com

Detection of Damage in Stiffeners of Aircraft Wing Structure Based On Induced Skin Strains and
Lateral Deflections http://www.iaeme.com/IJMET/index.asp 123 editor@iaeme.com

Dr. Hatem Hadi Obeid http://www.iaeme.com/IJMET/index.asp 124 editor@iaeme.com

Detection of Damage in Stiffeners of Aircraft Wing Structure Based On Induced Skin Strains and
Lateral Deflections
Declare that there is no “conflict of interests” regarding the publication of this paper.
My thankful for the assistance from the staff of post graduate laboratory at Department of
Mechanical Engineering , College of Engineering, Babylon University ,Iraq.
All data concerned with the fixation and performing the test of pressure distributions through wing in wind tunnel are available in the author and can be obtained via mailing.
The codes of the finite element analysis can be obtained after publication of the research and can be contacting with author for this purpose.
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Rizzo, Yi Qing Ni, and Jinying Zhu Volume 2010, Article ID 165132. http://www.iaeme.com/IJMET/index.asp 125 editor@iaeme.com

Dr. Hatem Hadi Obeid
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Science and Engineering 402 (2018) 012077. http://www.iaeme.com/IJMET/index.asp 126 editor@iaeme.com