SHM using embedded electrical crack gauges 43

4.7.1

Damage evidence

Cracks due to fatigue are the most common damages that can occur in an aircraft during operation. The cracks can initiate from high stress concentration points or in areas where corrosion is apparent. In aviation industry and according to some manufacturers, small cracks are tolerant up to specific sizes. However, aircrafts should be 100% airworthy, they should have no flaws and should be in perfect structural health. Therefore, to prevent a rapid growth of a crack and to avoid a possibility of a large-scale damage, it is very important that the stress sensitive areas are regularly monitored. A recent example of fatigue damage was the appearance of multiple cracks in the brackets of the wings of a new Airbus A380 aircraft from Qantas Airways on January 2012.

Figure 4.16:Drawings of components of the wing of an Airbus A380 aircraft from Qantas Airways where cracks have appeared [136].

In this PhD study, the crack gauges are conductive stripes that are embedded into the coatings of the metallic part under investigation. The crack gauges installation was performed in cooperation with the manufacturer of the aircraft components,

ASCO Industries N.V, Belgium. The SHM system is a proposed integrated system that can operate either as an online or an offline system.

4.7.2

Theoretical background

4.7.2.1

Methodology

Crack monitoring by crack gauges is based on the weakening or interruption of their electrical conductivity. The crack gauges are made of a conductive material that finally breaks together with the crack. An electric current is running through the conductive material of the gauges with a specific level of electrical resistance which depends on the kind of material used. The operation is as follows: Due to fatigue, cracks appear in the main material of the aircraft component that is under investigation. The crack passes through the crack gauge that is glued on the material and breaks it. When this happens, the resistance values of the conductive material of the crack gauges are reaching very high or infinite numbers. This occurs e ause the gauge has ee ut a d the ele t i u e t s flow has been interrupted. As it can be seen from the equation below, the resistance becomes infinite when the electric current is zero.

I

V

R

, I =0, R  ∞ Eq. 4.8 By monitoring the resistance values, it can be confirmed that the crack has reached the critical size and has cut the crack gauges. The technology makes use of resistance measurements with a relatively simple data analysis.

4.7.2.2

Closed crack issue

The fuselage and other structural components of an aircraft are constantly pressurized and depressurized due to stresses and pressure differences during take-off and landing. A crack in a component is likely to appear from fatigue. The crack will constantly stay open and close due to those cyclic pressurizations. When the aircraft is landed, the structural parts are subjected to less fatigue and due to the lack of pressurisation, cracks in the fuselage and other parts remain closed. These cracks might be macroscopically closed but they are microscopically open since they have altered the microstructure of the material. That means that they are difficult to be monitored since the material has been pressurized back in its original form in an atmospheric pressure and macroscopically look homogeneous. It is necessary that the crack gauges can detect the cracks instantly the moment that crack makes its appearance. Therefore, the electrical crack gauge should not regain its conductivity after the crack has penetrated through it. This is a great challenge since the crack gauge is re-attached and there is a chance of a regain of conductivity. In chapter 6, there is a proposal for an online monitoring solution that could be used to overcome this issue.

4.7.2.3

Material behaviour

Plasticity is also another factor that influences not only the resistance measurements using the crack gauge but also can result in the misjudgement of the critical break point. The physical theory of plastic deformation is quite difficult to be explained due to the complex geometries involved as well as the analysis of the non linear processes that occur [137]. However, it is important that any behaviour, elastic and plastic, prior the break point needs to be investigated in order to assure a reliable detection performance of the crack gauge. The figure below shows the diffe e t phases of the ate ial s defo atio he it is su je ted to a gradually increasing tensile force.

Figure 4.17:Elastic and plastic behaviour of a material indicated in the force- displacement x curve [138].

When the main material of the aircraft component is subjected to fatigue and the material can pass the elastic limit, local plastic deformation can occur. This plastic deformation is accompanied with void formation which is preceding the crack formation. This void formation is very important and needs to be taken into account when designing the crack gauge. This is because the crack gauge could malfunction during this plastic deformation of the material investigated and overlook the crack growth. Therefore, the crack gauge should be designed in order to have exactly the same or even less plastic deformation than the material under inspection. In that way, the crack gauge will break exactly at the point of the pre- appearance of the crack in the aircraft component.

Crack gauges can be made either of two or more constituent materials (composite) or of an isotropic material. Depending on the material that the crack gauge is made, it is very crucial that the gauge will break at the same moment as the component under investigation. In this study, in the test with the Airbus A320 slat- track, the attached crack gauge is made of aluminium and in the test with the aluminium 2024-T3 plates, the crack gauges were made of a conducting composite material. For the purpose of this study, a perfect bonding between the crack gauges and the main material is assumed.

4.7.2.4

Crack gauge made of an isotropic (uniform) material

In this study, only the tensile mode (opening mode I of the three fracture modes) has been taken into account since the fatigue tests performed were loaded in tension. The tensile mode distributes the stress in parallel to the length of the component. This stress distribution causes a crack which is perpendicular to the loading direction. Assuming that the load is applied uniformly on the material but also on the crack gauge then the tensile stress applied on the crack gauge is the following: g g

A

F





Eq. 4.9 where,



_g is the tensile stress applied on the crack gauge,

F

is the load, and

g

A

is the cross sectional area of the crack gauge. For isotropic materials such as aluminium, the yield strength is known and it can be directly compared to another isotropic material. The yield strength is preferably used here as a limitation factor. This is because if the ultimate tensile strength is reached, it will cause permanent deformation in the material and the whole structure will fail.

4.7.2.5

Crack gauge made of a composite material

Composite materials are composed out of two or more different components and therefore have a more complex structure than isotropic materials. To examine the cracking of an underlying material in this study, a simple unidirectional fibre- reinforced composite material is investigated. Unidirectional fibre-reinforced composites have all fibres in one and parallel direction inside the matrix. It is assumed that the fibres and the matrix have a linear behaviour during loading. In the case of a crack gauge that is made out of this composite, then similar to the previous equation, the load that is applied on the material is the following:

m m f

f

A

A

F



Eq. 4.10 where,



_fand



_m the tensile stress on the fibres and the matrix and

A

_f and

A

_m

A

is the total cross sectional area of the composite, then a volume fraction of the

fibres

A

A

V

_f



f can be defined. The elastic modulus of a material can be described from the following equation:







Y

Eq. 4.11

where

Y

is the ζou g s odulus of a ate ial,



the tensile stress and



the tensile strain. Hence, the maximum longitudinal tensile stress applied on the composite material is given by the following equation:

Y

C

V

f

Y

f

(1V

f

)Y

m Eq. 4.12

where

V

_f and

1V

_f are the volume fraction of the fibres and the remaining volume of the matrix respectively.

Y

_fand

Y

_m a e the ζou g s oduli of the fi es and the matrix respectively.

Y

_Cis lo gitudi al ζou g s odulus of the o posite. The equation above represents the parallel combination of the mixing rule of composites. This approach assumes that there is perfect bonding between the fibre and the matrix. By knowing the material properties of the composite individual components, the behaviour during cracking can be quantified.

In a crack gauge made of a composite, usually one part is conductive and the other non-conductive. In order to have a reliable crack gauge, the conductive part should be the first that brakes together with the main material since it carries all the information. Therefore, it is important that the fibres will be those that break first and their ultimate tensile strength has to be the same as the main material under inspection.

The design of the crack gauge can be optimized by the appropriate selection of the materials that will comprise the composite. Therefore, the maximum amount of stress that is applied to the crack gauge before it breaks defines the Young modulus of the materials that are selected for the design of the composite. This is very important since the stress limits of the composite crack gauge should be compared with the stress limits of the aircraft component under investigation. It is necessary that the fracture behaviour and the stress limits of the crack gauge are known in order to follow the fracture behaviour of the main material in the critical places of the aircraft component that is monitored. It has to be noted that the role of the crack gauge is purely for crack monitoring and does not strengthen the aircraft component.

In document Oplossingen voor Structural Health Monitoring van vliegtuigonderdelen (Page 63-68)