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INTERNATIONAL JOURNAL OF PROFESSIONAL ENGINEERING STUDIES Volume VIII /Issue 2 / JAN 2017

IJPRES

FINITE ELEMENT ANALYSIS OF FATIGUE OVER DAMGED COMPOSITES UNDER CRACK PROPAGATION

1 MUTLOORI EZEKIEL, 2 D.V. SUBBARAO

1 Pg Scholar, Department of MECH, Malla Reddy College Of Engineering And Technology

(Sponsored by C.M.R. Educational society) Autonomous Institution – UGC, Govt. of India, NAAC-A Grade,

Accredited by AICTE-NBA, Affiliated to JNTUH Hyderabad, Maisammaguda, Dhulapally post, Secunderabad-

500100 2 Assistant Professor, Department of MECH, Malla Reddy College Of Engineering And Technology

(Sponsored by C.M.R. Educational society) Autonomous Institution – UGC, Govt. of India, NAAC-A Grade,

Accredited by AICTE-NBA, Affiliated to JNTUH Hyderabad, Maisammaguda, Dhulapally post, Secunderabad-

500100

ABSTRACT

Now a day’s composites are being widely used in

many applications. Composites are blend of two or

more materials forming a component. Fatigue life is

an important consideration used by the designers.

Each and every component designated with fatigue

number. Fatigue is a damage which is caused due to

continuous loading and unloading, certainly known

as cyclic loading. Continuous fiber reinforced metal

matrix composites are projected for use in high

temperature stiffness critical part that will be

subjected to cyclic loadings. This project presents a

general overview of the fatigue behavior of

composite fiber such as carbon fiber and Glass fiber

and presents the Crack propagation in composite

fiber. A numerical study was performed to

investigate the dynamic crack propagation on carbon

fiber and glass fiber reinforced composite plates

utilizing the finite element method. A rectangular

plate of uniform thickness, which had a propagating

crack, was used for the study. The plate was a

unidirectional composite panel and the load was

applied in the longitudinal direction of the composite

plate. The objective of the study was to examine the

effect of different composite material properties,

different ply orientations and densities on the

fracture. The metal matrix composite are grouped in

four types 1) Matrix dominated 2) Fiber dominated 3)

Self-similar damage 4) Fiber/matrix interfacial

failure. These four types of damage will be discussed

and illustrated by example with the emphasis on the

fatigue of un-notched laminates. The Crack

propagation and Fatigue life of the composite fiber is

analyzed by using MSC Fatigue software.

The experiment was performed on one of the

aluminum plate in order to investigate the Dynamic

crack propagation on the rectangular plate by using

composite fiber such as carbon fiber and glass fiber

validate the results of numerical simulation.

Introduction To Composites

Composite material is a material composed of two

or more distinct phases matrix phase and dispersed

phase and having bulk properties different properties

from those of any of the constituents.

Classification Of Composites:

Composites are classified into 3 groups namely

1. Particle Reinforced Composites.

2. Fiber Reinforced Composites.

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3. Sandwichpanel composites.

1. Particle Reinforced Composites are sub divided

into:

a) Large Particle composites.

b) Dispersion strengthened Composites.

2. Fiber Reinforced Composites are sub divided into:

a) Continuous and Aligned fiber Composites

b) Discontinuous and Randomly Oriented fiber

Composites.

3. Structural Composites are sub divided into:

a) Laminar Composites.

b) Sandwich Panels.

Structural-Composites

A structural composite is normally composed of both

homogeneous and composite materials, the properties

of which depend not only the properties of the

constituent materials but also on the geometrical

design of the most common structural composites;

only a relatively superficial examination is offered

here for them.

Laminar Composites

A laminar composite is composed of two-

dimensional sheet or panels they have a preferred

high-strength direction such as I found in wood and

continuous and aligned fiber-reinforced plastics

varies with each successive layer.

Fig: Laminar Composite

Matrix Composites:

Polymeric Matrix Composites

Fig: Composites in the Boeing 787

Polymer-matrix composites (PMCs) consist of a

polymer resin 1 as the matrix, with fibers as the

reinforcement medium. These materials are used in

the greatest diversity of composite applications, as

well as in the largest quantities, in light of their room-

temperature properties, ease of fabrication, and cost.

Glass Fiber-Reinforced Polymer (Gfrp)

Composites:

Fig: Glass Fiber

Fiberglass is simply a composite consisting of glass

fibers, either continuous or discontinuous, contained

within a polymer matrix.

1. It is easily drawn into high-strength fibers

from the molten state.

2. It is readily available and may be fabricated

into a glass-reinforced plastic economically

using a wide variety of composite-

manufacturing techniques.

Carbon Fiber-Reinforced Polymer

(CFRP) Composites

Figure 6 Carbon Fiber

Carbon is high-performance fiber material that is the

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most commonly used reinforcement in advanced (i.e.,

nonfiberglass) polymer-matrix composites. The

reasons for this are as follows:

1. Carbon fiber has the highest specific

modulus and specific strength of all

reinforcing fiber materials.

2. They retain their high-tensile modulus and

high strength at elevated temperatures;

high-temperature oxidation, however, may

be a problem.

Aramid Fiber-Reinforced Polymer

Composites

Fig: Aramid composite

Aramid fibers are high-strength, high-modulus

materials that were introduced in the early 1970s.

They are especially desirable for their outstanding

strength-to weight ratios, which are superior to

metals.

Ceramic-Matrix Composites

Fig: Ceramic Matrix Composite

Ceramic matrix composites (CMCs) provide high

temperature oxidation stability relative to metals and

enhanced toughness relative to monolithic ceramics.

Ultra met offers durable, refractory CMCs that

survive the severe environments of propulsion and

thermal management systems. Because they are

fabricated through a rapid melt infiltration process,

Ultra met CMCs are far less costly than composites.

Small particles of partially stabilized zirconia are

dispersed within the matrix material, often A12O3 or

ZrO2 itself. Typically, CaO, MgO, Y2O3, and CeO

are used as stabilizers.

Properties:

Mechanics

The physical properties of composite materials are

generally not isotropic (independent of direction of

applied force) in nature, but rather are typically

anisotropic (different depending on the direction of

the applied force or load).

For the special case of orthogonal isotropy, there are

three different material property constants for each of

Young’s Modulus, Shear Modulus and the Poisson’s

ratio – a total of 9 constants to describe the

relationship between force/ moments and

strains/curvatures. Techniques that take advantage of

the anisotropic properties of the materials include

mortise and tenon joints (in natural composites such

as wood) Pi joints in synthetic composites.

Failure

Shock, impact, or repeated cyclic stresses can cause

the damage to the composites.

Resin Systems

Depending on the particular thermoplastic material

used, thermoplastic matrix components can, however,

be used over a wide range of temperature from 100°c

to 3000°c. The advantages of thermoplastic matrices

are their improved fracture toughness over the thermo

set matrix and their potential of much lower cost in

the manufacturing of finished composites. For the

purpose of a simple classification, we may divide the

thermo set polymers into five categories. Polyester

resin,

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(i) Polyester resins,

(ii) Epoxy resin,

(iii) Vinyl ester resin,

(iv) Phenolic resin and

(v) High performance resins.

Different Types Of Fuselage Skin:

The fuselage is an aircraft's main body section that

holds crew and passengers or cargo. In single-engine

aircraft it will usually contain an engine, although in

some amphibious aircraft the single engine is

mounted on a pylon attached to the fuselage which in

turn is used as a floating hull.

There are three types of fuselage structure:

Truss type structure:

This type of structure is still in use in many

lightweight aircraft using welded steel tube trusses. A

box truss fuselage structure can also be built out of

wood often covered with plywood as can be seen on

this Ibis canard fuselage.

Fig: Truss Type Structure

Monocoque structure:

Monocoque is a structural approach that supports

loads through an object's external skin, similar to an

egg shell. The term is also used to indicate a structure

in which the skin provides the main structural

support, although this is rare and is usually confused

with either semi-monocoque or a unibody. The word

monocoque comes from the Greek for single (mono)

and French for shell (coque).

Semi-Monocoque Structure:

This is the preferred method of constructing an all-

aluminum fuselage. First, a series of frames in the

shape of the fuselage cross sections are held in

position on a rigid fixture. These frames are then

joined with lightweight longitudinal elements called

stringers.

Fig: Semi Monocoque Structure

Different Types Of Rivets used:

Rivets

Before modern welding techniques came into

common use, riveting was one of the most common

methods for joining sheet metal. Since the advent of

the new welding techniques and modern machines

that form seams on sheet metal, riveted seams are not

so common in modern sheet metal work.

Fig: Rivets

Riveting

Riveting may be done by hand or by machine. When

the job is performed by hand, as is usually the case in

sheet metal work, it is done with a hammer and rivet

set.

Rivet Sizes

The size of rivets is determined by the

weight of 1,000 rivets i.e. 1N. Rivets weigh 1 N. per

thousand, 2 N. rivets weigh 2 N. per thousand.

Flathead rivets vary in diameter from 3/32" to 7/16"

in steps of 1/32". Other rivets vary in size with 1/8"

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and 3/16" snaphead rivets being the most popular in

the sheet metal shop.

Fig: Rivet Sizes

Types of Rivet and Rivet-Head

The way in which thin material is joined to thick

material with countersunk rivets.

Always use the correct rivet for a particular metal to

be riveted. When riveting aluminium, for example,

use aluminium rivets; and when riveting copper use

copper rivets.

Fig: Standard Types of Rivet Heads

Fig: Countersunk Riveting Of Thin Material To

Thick Material

Defects in Riveted Joint

When making joints with rivets, the following points

should be followed to prevent many common defects:

1. Use The correct allowance for edge clearance and

pitch when marking out.

2. Use the proper type of rivet as specified on the

drawing.

3. Use rivets of the correct length.

Properties Of Typical Composite Materials

Various types of composite materials are possible by

choosing different constituents mixed in different

proportions. Those composites have higher specific

modulus and higher specific strength than steel.

Fiber properties dictate the stiffness of

unidirectional composites. For higher strength of

unidirectional composites, the ultimate strain is

generally lower at failure. For some composites, the

increase of stiffness has resulted in lowering the

strength zth stress-strain curve for unidirectional

composites in general is linear.

Advantages Of Composites:

1) A higher performance for a give weight

leads to fuel savings.

2) Part count is reduced.

3) Production cost is reduced. Composites

may be made by a wide range of processes.

Disadvantages Of Composites:

1) Composites are more brittle than wrought

metals and thus are more easily damaged.

Current Applications Ofcompositesin Aerospace:

The application of composites in place of

metal requires different approaches to the design and

service of structural components

Military Aircraft:

Fig: Fighter Aircraft

Fighter aircraft often drive new technologies because

of their special requirements

Therefore, it is not surprising that many of the early

composite technology programs in the 1970s were

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part of fighter aircraft development projects such as

the F15, Alphajet,Tornado and Mirage.

Introduction to Project:

In materials science, fatigue is the progressive and

localized structural damage that occurs when a

material is subjected to cyclic loading. The nominal

maximum stress values are less than the ultimate

tensile stress limit, and may be below the yield stress

limit of the material.

Fatigue occurs when a material is subjected to

repeated loading and unloading. If the loads are

above a certain threshold, microscopic cracks will

begin to form at the stress concentrators such as the

surface, persistent slip bands (PSBs), and grain

interfaces. Eventually a crack will reach a critical

size, and the structure will suddenly fracture.

Fatigue failures often occur quite suddenly with

catastrophic (disastrous) results and although most

insidious for metals, polymers and ceramics (except

for glasses) are also susceptible to sudden fatigue

failures. Fatigue causes brittle like failures even in

normally ductile materials with little gross plastic

deformation occurring prior to fracture.

The simplest is completely reversed constant

amplitude where the alternating stress varies from a

maximum tensile stress to a minimum compressive

stress of equal magnitude.

Fig: Comparison of Metal and Composite

Stiffness Reduction

Importance of composites in Aero space

In the year 2005 a new damage accumulation model

for fatigue life prediction of composite laminates is

proposed has been studied. The model is constructed

on the ply level and uses a new multiaxial damage

index to consider the damage caused by different

stress components. The fatigue life is predicted based

on the proposed model and experimental results of

the unidirectional laminates. The numerical results

are compared with the experimental data for balanced

laminates.

Literature of composites in Aerospace

Composites are widely being used in many

applications like sporting goods, boating, building

construction, aviation industry etc, due to its high

strength and low weight advantages. Progressive loss

of stiffness during fatigue of a composite is a very

important characteristic.

Fig: Load curve for fatigue analysis

Fatigue Crack Propagation:

Crack growth rate da/dn is related to the cyclic stress

intensity factor range Δk, according to the power law

relationship formulated by paries and erdogan (1963).

da/dn = A (Δk)^m

The applied cyclic stress intensity range is given by:

Δk = YΔσ√πa

Y is a geometric factor, Δσ is the cyclic stress range,

and a is the crack length.

(S-N) Curve

S-NFrefers to a plot of Constant Amplitude Stress

Level (S) verses Number of Cycles to Failure (NF).

Fig: S-NF Curve

Introduction to Fatigue:

Fatigue, or metal fatigue, is the failure of a

component as a result of cyclic stress. The failure

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occurs in three phases: crack initiation, crack

propagation, and catastrophic overload failure.

Theoretical calculations:

Specimen’s diameter:

Sheet diameter = 1st raw material sheet length (l)

210+210=420

Sheet diameter= 2nd raw material sheet length (l)

210+150=360

Hole diameter =5 mm

Pitch diameter=12.

Width = 150

The mechanical properties =(0.2% yield stress =

291MPa,

Ultimate strength = 433 MPa,

Elongation = 13%

fatigue crack growth behavior of this material are

similar to those of the western Al 2024-T3 alloy

Force controlled riveting was applied using a

squeezer mounted in the grips of a MTS 810 fatigue

machine.The same machine was utilized in the

fatigue tests carried out under constant amplitude

loading at a stress ratio of 0.1.

Selecting the Correct Size of Rivet

To obtain the full strength of a riveted joint, a rivet of

the correct diameter and length must be used.

TD 25.1 (Urwin's formula);

Fatigue calculations

What is Finite Element Analysis?

FEA consists of a computer model of a material or

design that is stressed and analyzed for specific

results. It is used in new product design, and existing

product refinement.

Software’s used:

Solid-works: For CAD modeling of the lap jointed

fuselage skin and rivets

Hyper-mesh: For Meshing the model.

Radioss(solver): Radioss is used for solving the

structural analysis problems.

ANSYS Fatigue: Used for calculating the life span of

the fuselage skin under repeated loads.

Construction Of Fuselage Skin:

The fuselage skins are constructed by using lap joint

and by using nuts and bolts.In our case the fuselage

skin is joined together by using the lap joint.

Fig: Typical Fuselage Longitudinal Riveted Lap

Splice Joint

Figure: Geometry dimensions of the Fuselage

Geometry model:

Figure: Geometry of the fuselage skin with lap joint

Exploded view:

Figure: Exploded view of the geometry

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Meshing of the Geometrical model:

The fuselage skin is meshed by using the hyper-

mesh.Shell elements are used for fuselage skin and

beam elements are used for rivets modeling

Figure: Mesh model of the Fuselage skin

Modeling of rivets and rigid for coupling of rivets

and fuselage:

Figure: Mesh model with Bolts at hole locations

Figure: Exploded view of the bolt location

Enlarged view of the assembly showing rivets in pink

colour

Figure: Enlarged view of the Assembly

Material properties:

Table 1: Material properties

Load acting in our case:

Tensile load:

Figure: Load acting over the fuselage

Boundary conditions:

Figure: Constraint locations

Different load cases considering for analysis:

Loading Cases

Case 1: Material used as aluminum

Case 2: Composites (carbon fiber) with piles

orientation as(0/45/90/-45/0)

Case 3: Composites (carbon fiber) with piles

orientation as (0/90/0/90/0)

Case 4:Composites (carbon fiber) with piles

orientation as (0/45/90/-45/90/0/0)

Analysis results:

Case 1: Material–Aluminum

Displacement plot

Figure: Displacement contour plot

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Figure: Von Misses stress contour plot

Stress plot

Case 2

Ply Orientation of Carbon-Fiber (0/45/90/-45/0)

Displacement plot


Stress plot

Figure: Von-Misses stress contour plot

Figure: Principal stress in Ply 1

Figure: Principal stresses in Ply 2

Figure: Principal stresses in ply 3

Figure: Principal stresses in ply 4

Figure: Principal stresses in ply

Table 2: Tabulated results for case 2

Case 3:

Ply Orientation of Carbon-Fiber(0/90/0/90/0)

Displacement plot

Figure: Displacement Contour plot

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Stress plot

Figure: Von-Misses stress contour plot

Figure: Principal stress in ply 1

Table 3: Tabulated results for case 3

Case 4

Ply Orientation of Carbon-Fiber (0/45/90/-

45/90/0/0)

Displacement plot


Stress plot

Figure: Von Misses stress plot

Figure: Principal stress in ply 1

Table 4: Tabulated results for the case 4

Figure: S -N Curve for the Aluminium 2024

Figure: Unit Load acting for Fatigue analysis

Figure 1: Total life of the component

Figure: Logarithmic life of the component

By performing the fatigue analysis over the

composite fuselage structure the number of life

cycles it can with stand is 7e12

Results And Discussions

Tabulated results

From the table 5 we conclude that 2 different material

having different properties.

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Where aluminum has applications requiring high

strength to weight ratio, as well as good fatigue

resistance.Where carbon fiber has good mechanical

properties and having good fatigue resistance then

aluminum.From the above experiment we can

conclude that the case 4 with 7 plies has more

stiffness when compared to other so we replace the

aluminium with the Carbon Fiber containing 7 plies.

Conclusion

Carbon fiber composite is capable of withstanding

higher amount of load when compared to that of

aluminium 2024. The Application of aluminium in

the construction of aircraft fuselage can be replaced

by the carbon fibre effectively. The advantage of

replacing the aluminium 2024 by the carbon fibre is

not only it can withstand load in its operating

condition but also we will be reducing the weight of

the total fuselage skin by half. By using the carbon-

fiber the weight is reduced with indirectly increases

the fuel efficiency of the airplane.since now a day’s

composites are playing a main role in increasing the

overall efficiency of the commercial aircrafts by

reducing the weight and to withstanding enough

loads in its manuvering, takeoff and landing

operations, the usage of carbon fiber will be very

useful in the aerospace industry.

Future scope:

1. Increasing the efficiency of composite single-shear

lap joints.

2. CFRP with aluminum alloy plates can be tested in

experimental.

3. CFRP with aluminum alloy plates can be tested in

experimentaland can find the strength and stress

distribution with differentcomposite materials.

4. The joint design can be made and optimization of

riveted jointcan be performed.

5. Stress and Failure Analysis of Laminated

Composite PinnedJoints can be performed. Carbon

fiber airframes mean that aircraft can be very

different in shape such a design could significantly

improve a plane's lift-to-drag ratio - making it much

more aerodynamically efficient, and also reduce its

overall weight.

References

1. www.callister –materials-sicence-an.html

2.www. Kawaius.com/carbon-fiber-pdf

3.Muller,R.P.G.(1995) Phd thesis ,Delft University of

Technology ,Netherlands.

4.www. Fiber materialsinc.com/cc comp.html

AUTHORS:

1. MUTLOORI EZEKIEL

Pg Scholar,

Department of MECH,

Malla Reddy College Of Engineering And

Technology (Sponsored by C.M.R. Educational

society) Autonomous Institution – UGC, Govt. of

India, NAAC-A Grade, Accredited by AICTE-

NBA, Affiliated to JNTUH Hyderabad,

Maisammaguda, Dhulapally post, Secunderabad-

500100

2. D.V. SUBBARAO

Assistant Professor

Department of MECH,

Malla Reddy College Of Engineering And

Technology (Sponsored by C.M.R. Educational

society) Autonomous Institution – UGC, Govt. of

India, NAAC-A Grade, Accredited by AICTE-NBA,

Affiliated to JNTUH Hyderabad, Maisammaguda,

Dhulapally post, Secunderabad-500100

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