Continuing Demand For High Performance Materials Biology Essay

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Materials technology impacts on the quality of our lives. The continuing demand for high performance materials in structural applications has posed a great challenge for scientists and engineers. Materials technology enables us to design durable components in military that serve several generations by understanding the characteristics of material that we need in different operating conditions. The world we live in has finite resources and because of that, materials engineers have duty of taking care for the research, design, specification and materials construction for future technologies. As example, in military armored vehicles, the material they use not only strong enough to protect against impact from projectiles of conventional bullet shape, but also light to reduce the weight of the vehicles.

Many of the products nowadays are produced from monolithic materials. Majority of composites have two components materials: a binder or matrix, and reinforcement. The reinforcement is typically stronger and stiffer than the matrix. The matrix holds the reinforcements in an arranged design. Because the reinforcements are regularly discontinuous, the matrix as well helps to transform load among the reinforcements.

Composites are classify into three various categories, there are; particle-reinforced, structural and fiber-reinforced. Sandwich panel is one example from structural composites. It has two face sheets at the outer, coated by a core which has lower in stiffness also in strength inside. In addition, the core also can be change with a honeycomb structure.

Honeycomb sandwich materials are one of such high performance materials. Sandwich, by definition, is made of three different layers, two thin face-sheets and a low density core. It has a relative high bending stiffness, lightweight, and high energy absorbing capacity, sandwich structures have been used widely in aircraft structures.

For this project, the sandwich panel will be tested using a ballistic impact test for the ballistic resistance and angle effect of penetration test. Ballistic impact is concerned with the effects of projectile on the target. The ballistic impact is a very importance study because it ensures the armor plate is made from a safe and optimum material for protection.

Problem Statement

Composite sandwich panels are of considerable interest in aircraft structures such as impact resistant fuselage panels and as low weight panels in wing leading edges and flaps. They are susceptible to impact damage due to their thin composite skins, but with suitable core materials may be designed to absorb impact energy.

Typical high velocity impact tests on composite sandwich panels show that impact damage and failure, particularly with hard projectiles, is very localized consists of penetration of the outer composite skin, damage to the core and, if impact energy is high enough, core penetration followed by inner skin damage or fracture.

Composite is used for energy absorbing application. Sandwich structure is used to enhance composite energy absorption capability. However, in order to simulate this local damage and failure detailed FE models are required to predict and measure energy absorption capability. The model was validated with experimental tests by comparing numerical and experimental result.

With reference to Wan Awis (2010) researches have done, there are four main problems relating to armor panels.

The first is cost. Excessively complex armor preparations, mostly those depending totally on synthetic fibers, can be responsible for a distinguished amount of the total vehicle cost, and make its produce non-profitable.

The second is weight. Protective panel for heavy but movable armed equipment, for example Armoured Personnel Carrier (APC) and tanks, is identified. Such armor generally includes a solid layer of alloy steel, which is planned to give protection adjacent to heavy and explosive projectiles. Armor for light vehicles is predictable to avoid penetration of bullets of any kind, yet when collision at a momentum in the span of 700 to 1000 meters per second. On the other hand, due to weight limitation it is hard to defend light vehicles from high caliber armor-piercing projectiles, such as of 12.7 and 14.5 mm, because the weight of standard armor to endure such projectile is such as to delay the mobility and operation of vehicles.

A third is relates to ceramic plates operate for individual and light vehicle armor, which plates have be present discover to be susceptible to spoil from mechanical impacts caused by rocks, falls, etc.

A fourth is density. A solid armor panel, with air spaces between it's a variety of layers, enlarge the aim profile of the vehicle. In case of civilian retrofitted armored vehicle which are equipped with internal armor, there is basically no room for a solid panel in nearly all the areas requiring armor.


To simulate the damage of composite sandwich structures subjected to high-velocity impact using finite element.

To determine the energy absorption capability of the components on the behavior of the sandwich panel under impact load using MSC-Dytran.

To develop a of energy-absorption model of the impact composite structure.

To validate a numerical model with actual experiment.

Scope of Research

To characterize a mechanical behavior of carbon fiber panel by using tensile, compression testing and determine the fiber volume force and density.

Design and validate the numerical model.

Conduct a ballistic impact test simulation.

Using the experiments data to calculate the energy absorption on the impact on the difference angle.



2.1 Introduction to Composite

Sandwich composite material has been widely used in the aircraft and aerospace industries because of its high strength to weight and stiffness to weight ratios. Typically sandwich composite is formed by bonding thin, strong facesheets to a thick, lightweight core. Each components of this composite is relatively weak and flexible but when working together they provide and extremely stiff, strong and lightweight structure. (Thomson et al, 2005)

A composite is a mixture of two phases called the matrix phase and are embedded in other materials, reinforcing phase, is in the form of fibers, sheets or particles. The most usual components of sandwiches are a honeycomb core and carbon-fibre skins due to their high specific stiffness and strength. Sandwich structures are very sensible to such loads. (Aktay et al, 2005)

Characteristically, the matrix is usually a ductile or strong material while reinforcing materials are strong with low densities. If the composite is fabricated and designed properly, it combines the toughness of the matrix with the strength of reinforcement to achieve a combination of desirable properties. The disadvantage is that such composite are regularly more expensive than conventional materials. (Schaffer et al, 1999)

2.2 Fiber-Reinforced Composite

Fiber-reinforced composite materials have been used in such diverse application as spacecraft, automobiles, sporting goods, aircraft, off-shore structures, civil infrastructure, electronics, and marine vehicles. (Agarwal and Broutman, 1990)

Fiber-reinforced composites often aim to improve the strength to weight and stiffness to weight ratios (i.e. desire light-weight structures that are strong and stiff). Glass or metal fibers are generally embedded in polymeric matrices. Fibers are available in 3 basic forms (Shackelford, 1996):

Continuous - Fibers are long, straight and generally layed-up parallel to each other.

Chopped - Fibers are short and generally randomly distributed (fiberglass).

Woven - Fibers come in cloth form and provide multidirectional strength.

Figure 2.1 Schematic Illustration of Fiber Types: Unidirectional, Chopped & Woven. (Shackelford, 1996)

Table 2.1 Properties of Selected Fiber-Reinforcing Materials. (Shackelford, 1992)


Dispersed Phase


[Mpa (ksi)]

Compressive Strength

[Mpa (ksi)]

Precent Elongation at Failure


69x103 (10x103)




72.4x103 (10.5x103)




85.5x103 (12.4x103)



Ceramic fiber

C (graphite)

360-380 x103

(49-55 x103)




430x103 (62x103)



Ceramic whisker


430x103 (62x103)



Polymer fiber


131x103 (19x103)



Metal Filament


410x103 (60x103)



Concrete aggregate

Crushed stone and sand

34-69 x103 (5-10 x103)




Table 2.1 presents a listing of commonly used fiber materials; commonly utilized fibers include E-Glass (low cost), Kevlar (very low density) and Carbon (high strength and modulus). Whiskers are small, single crystal fibers that have a nearly perfect crystalline structure. (Shackelford, 1992)

Table 2.2 Properties of Composite Reinforcing Fibers (Gerstle, 1991)




































HS graphite







HM graphite







E is Young's Modulus, is breaking stress, is breaking strain and is density.

As seen in the Table 2.2, the fibers used in modern composites have strengths and stifnesses far above those of traditional bulk material. The high strengths of glass fibers due to processing that avoids the internal or surface flaws which normally weaken glass, and the strength and stiffness of the polymeric aramid fiber is a consequences of the nearly perfect alignment of the molecular chains with the fiber axis. (Gerstle, 1991)

Figure 2.2 Fiber Orientation in Fiber Reinforced Composites. (Elgun, 1999)

2.3 Carbon Fibers

Carbon fibers refer to fibers which are at least 92 % carbon in composition. They can be short or continuous; their structure can be crystalline, amorphous, or partly crystalline. (Fitzer, 1990)


Figure 2.3 Carbon Fiber (Wikipedia)

Table 2.3 Properties of Various Fibers and Whiskers. (Askeland, 1989)


Density (g/cm3)

Tensile Strength (GPa)

Modulus of Elasticity (GPa)

Ductility (%)

Melting Temp (°C)

Specific Modulus (106 m)

Specific Strength (104 m)









































Carbon (high-strength)








Carbon (high-modulus)
















































































Al2O3 whiskers








BeO whiskers








B4C whiskers








SiC whiskers








Si3N4 whiskers








Graphite whiskers








Cr whiskers








Table 2.3 compares the mechanical properties, melting temperature, and density of carbon fibers with other types of fibers. There are numerous grades of carbon fibers. Table 2.3 only shows the two high-performance grades which are labeled "high strength" and "high modulus". Among the fibers (not counting the whiskers), high strength carbon fibers exhibit the highest strength while high modulus carbon exhibits the highest modulus of elasticity. Moreover, the density of carbon fibers is quite low, making the specific modulus (modulus/density ratio) of high-modulus carbon fibers exceptionally high. The polymer fibers, such polyethylene and Kevlar fibers, have densities even lower than carbon fibers, but their melting temperatures are low. The ceramic fibers, such as glass, SiO2, Al2O3, and SiC fibers, have densities higher than carbon fibers. (Askeland, 1989)

2.3.1 Classification and Types of Carbon Fibers

Based on final heat treatment temperature, strength and modulus, carbon fibers can be classified into the following categories (Hegde et al, 2004):

Base on precursor fiber material:

Gas-phase-grown carbon fibers

PAN-based carbon fibers

Rayon-based carbon fibers

Mesophase pitch-based carbon fibers

Pitch-based carbon fibers

Isotropic pitch-based carbon fibers

Base on carbon fibers properties:

Super high-tensile (SHT) (tensile strength > 4.5GPa)

Low modulus and high-tensile (HT) (modulus < 100GPa, tensile strength > 3.0GPa)

Ultra-high-modulus (UHM) (modulus > 450GPa)

High-modulus (HM) (modulus between 350 - 450GPa)

Intermediate-modulus (IM) (modulus between 200 - 350GPa)

Base on final heat treatment temperature:

Type-I, high-heat-treatment carbon fibers (HTT), where final heat treatment temperature should be above 2000°C and can be associated with high-modulus type fiber

Type-II, intermediate-heat-treatment carbon fibers (IHT), where final heat treatment temperature should be around or above 1500°C and can be associated with high-strength type fiber

Type-III, low-heat-treatment carbon fibers, where final heat treatment temperature not greater than 1000°C. These are low modulus and low strength material.

Table 2.4 Characteristic and Applications of Carbon Fibers (

Physical strength, specific toughness, light weight

Aerospace, road and marine transport, sporting goods

High dimensional stability, low coefficient of thermal expansion, and low abrasion

Missiles, aircraft brakes, aerospace antenna and support structure, large telescopes, optical benches, waveguides for stable high-frequency (GHz) precision measurement frames

Good vibration damping, strength and toughness

Audio equipment, loudspeaker for Hi-fi equipment, pickup arms, robot arms

Electrical conductivity

Automobile hoods, novel tooling, casings and bases for electronic equipments, EMI and RF shielding, brushes

Biological inertness and x-ray permeability

Medical applications in prostheses, surgery and x-ray equipment, implants, tendon/ligament repair

Fatigue resistance, self-lubrication, high damping

Textile machinery, genera engineering

Chemical inertness, high corrosion resistance

Chemical industry, nuclear field, valves, seals, and pump components in process plants

Electromagnetic properties

Large generator retaining rings, radiological equipment

2.4 Sandwich Panel

Sandwich panels can be simply defined as a three-layer structure that consists of two thin, outer skins of high-strength material separated by a low-density and low-weight core material. The core material separates the face sheets that provide most of the strength to the structure. (Hoffart et al, 2008)

Sandwich panel cores are low in density and lightweight; but when combined with reinforcing fibers and resin, they become incredibly stiff, light and strong structures. The core helps absorb impact and distribute impact. In a honeycomb sandwich panel has much lower impact resistance as the skin is not in full contact with the honeycomb core although often stronger and stiffer. In a sandwich panel, the compression strength of the core helps prevent the sandwich from buckling, delaminating or wrinkling. (, 2010)

Sandwich panel normally consists of a low-density core material sandwiched between two high modulus face skins to produce a lightweight panel with exceptional stiffness as shown in Figure 2.4. The face skins act like the flanges of an I-beam to provide resistance to the separating face skins and carrying the shear forces. The faces are typically bonded to the core to achieve the composite action and to transfer the forces between the components. (Akour et al, 2010)

Figure 2.4 Illustration Sandwich Plate Geometry (Akour et al, 2010)

Figure 2.5 Diagram of an Assembled Composite Sandwich (A), and its Constituent Face Sheets or Skins (B) and Honeycomb Core (C) (Alternately: Foam Core) (Wikipedia)

2.4.1 Face sheets

The face sheets provide the flexural rigidity of the sandwich panel. It should also possess tensile and compressive strength. Since the carbon-epoxy composite has lower density than aluminum, significant weight savings can be realized by replacing them. The analysis of composite plates by Harris et al (1996) indicates that the sandwich plates with carbon epoxy face sheets have the lowest weight for different loading cases and that they are dimensionally more stable for a wide range of temperatures.

2.4.2 Cores

The purpose of the core is to increase the flexural stiffness of the panel. The core in general has low density in order to add as little as possible to the total weight of the sandwich construction. The core must be stiff enough in shear and perpendicular to the faces to ensure that face sheets are distant apart. In addition the core must withstand compressive loads without failure. The cores can be almost any material, but in general fall into the following four types. They are foam or solid core, honeycomb core, Web core and Corrugated or truss core. In Web core and truss core construction, a portion of the in-pane and bending loads are also carried by the core elements. (Sudharsan, 2003)

2.5 Honeycomb Sandwich Panel

Table 2.5 Standard Compositions of Honeycomb Sandwich Panels (Hexcel, 2001)

Panel composition

Key characteristics

Aluminium honeycomb core, aluminium skins.

Medium weight and stiffness at low cost.

Aluminium honeycomb core, woven glass fibre skins.

Lighter and less stiff than aluminium-skinned panels, at lower cost.

Non-metallic Nomex honeycomb core, unidirectional or woven glass fibre skins.

More resilient and higher cost than panels with aluminium honeycomb core. Unidirectional fibres give greater stiffness, at higher cost than woven fibres.

Non-metallic Nomex honeycomb core, unidirectional or woven carbon fibre skins.

The lightest and stiffest panels, which is reflected in their cost.

Honeycomb is usually used as a core in sandwiched structures to meet design requirements for highly stressed structural components. When sandwiched between layers of carbon fiber, honeycomb shows extreme resistance to shear stresses. As a structural core material, it is used in all types of aerospace vehicles and supporting equipment where sandwich structure offers aerodynamics smooth surfaces, rigid panels of minimum weight, and high fatigue resistance. The same structural properties are also used for commercial applications for example snow and water skies, tools, floors and bulkheads. Honeycomb is also used where designs need a means of energy absorption. It also has a high stiffness and specific strength. Below are the features of honeycomb (Raymond et al, 2008):

Compatible with most adhesives used in sandwich composites

Good thermal stability

Over expanded cell configuration suitable for forming simple curves

Excellent dielectric properties

High strength to weight ratio

Corrosion resistant

High toughness

Thermally insulating

Excellent creep and fatigue performance

Densities as low as 32 kg/m3 (2.0 lb/ft3)

They are available in variety of materials for sandwich structures. They range from low strength and stiffness applications to high strength and lightweight applications such as aircraft industries. They can be formed to any shape or curve without excessive heating or mechanical force. Honeycombs have very high stiffness perpendicular to the faces and the highest shear stiffness and strength to weight ratios of the available core materials. The most commonly used honeycombs are made of aluminum or impregnated glass or aramid fiber mats such as Nomex and thermoplastic honeycombs. The main drawback is high cost and difficulty in handling. (Sudharsan, 2003)

2.5.1 Honeycomb Sandwich Panel Construction

A structural sandwich is a layered construction formed by bonding two thin facings to a thick one. The principal design concept is to space strong thin facings far enough separately with a thick core to ensure the combination will be stiff, to give a core that is strong and stiff enough to hold the facings plane with an epoxy resin ply, and to give a core material of enough shearing resistance. (Bob Burdon, 2009)

2.5.2 Aluminum Honeycomb

Aluminum honeycomb sandwich panels (AHSP) have been widely used in the fields of aerospace, defense, railway, marine, automotive, sport industry, and communication for the merits of light weight, high stiffness and high strength. But AHSP shows it's vulnerable before impact force. McGowan et al (1998) represented that manufacturing defect would happen to AHSP caused by careless impact. This damage almost didn't mark on the face sheet externally but it would cause strength reduction internally. It was also stated that the conditions and degrees of impact damage were variable according to the test methods such as drop weight impact test and air-gun test etc.

Santosa et al (1998) compared the impact strength between the honeycomb cell and foam. They found that the aluminum honeycomb could own better impact behavior than aluminum foam under unidirectional loading and a combination load of compressive and bending.

Reddy et al (1998) analyzed the impact conformation with the experiential equation derived. Papka et al (1997) evaluated the damage mechanism after impacting by optical analysis and ultrasonic c-scan. Kim et all (1995) found that test frequency is lower than resonance frequency of honeycomb structure by using mechanical impedance method in spite of various ways of finding the debonding honeycomb structure.

Kim et al (2000) calculated the elastic modulus, shear modulus, Poisson's ratio, compressive bending strength, shear bending strength and flexibility of the sort of cell. In the research triangular and star cell was weak compressive bending strength and shear bending strength. But it had high flexibility.

Features of aluminum honeycomb (Plascore, 2009):

High thermal conductivity

Use temperatures up to 350 ͦ F

Fungi resistant

Excellent moisture and corrosion resistance

Low weight/high strength

Flame resistant

2.6 Ballistic Impact

According to Abrate (1998), several definitions of ballistic impacts are used in the literature. Impacts resulting in complete penetration of the laminate are often called ballistic impacts, whereas non-penetrating impact are called low-velocity impacts. Although non-penetrating impacts were studied extensively, impact penetration in composite materials has received considerably less attention. A different classification consists of calling low-velocity impacts those for which stress wave, a shear wave, and Rayleigh waves propagate outward from the impact point. Compressive and shear waves reach the back face and reflect back. After many reflections through the thickness of the laminate, the plate motion is establish are called low-velocity impacts.

A simple method can be used to evaluate a transition velocity beyond which stress wave effects dominate. A cylindrical zone immediately under the impactor undergoes a uniform strain across each cross section as the wave progresses from the front to the back face. With this simplifying assumption, the problem is reduced to that of the impact of a rigid mass on a cylindrical rod. Then, the initial compressive strain on the impacted surface is given by

where V is the impact velocity and c is the speed of sound in the transverse direction. Typically, critical strains between 0.5 and 1.0% are used to calculate the transition velocity. For common epoxy matrix composites, the transition to a stress wave-dominated impact occurs at impact velocities between 10 and 20m/s. Drop weight testers generally induce low-velocity impacts since a drop height of 5m will produce an impact velocity of 9.9m/s and most testers have a shorter drop height. (Caldwell et al, 1990)

In studying ballistic impacts, it is important to measure the residual velocity of the projectile accurately. This is a difficult task because many small particles, fibers, and shear plugs are pushed out by the projectile during penetration. This material can trigger the speed-sensing device being used and yield erroneous values. (Arndt and Coltman, 1990)

With high velocity impact, it is very important to measure the incident and residual velocities of the projectile in order to calculate the energy absorbed during the penetration process. Measuring the residual velocity of the projectile using optical sensor is difficult, since spalled material, shear plugs, and small particles can move ahead of the projectile as it exits the other side of the laminate, Zee et al (1991) developed a microvelocity sensor to measure the velocity of a projectile during the penetration process.

2.7 Energy Absorption

Cantwell and Morton (1985) calculated the ballistic perforation energy as the sum of the energy the target absorbs by flexure Ef, contact deformation Ec, delamination Ed and shear-out Es. For thicker targets, two distinct failure processes are observed for the upper and lower portions of the specimen. (Bless and Hartman, 1989, Bless et al. 1990, Lin et al.1990, Cantwell and Morton 1990).

A lower-bound estimate for the penetration energy for a thin laminate (4 to 16 plies) was obtained by Cantwell and Morton (1989). The hole produced by shearing of the fibers during perforation was shaped as a truncated cone starting with the diameter of the projectile and with a 45 ͦ half angle (Figure 2.6). The energy required to produce the hole was estimated by multiplying the fracture energy per unit area by the area of the frustum

where h is the laminate thickness and r is the projectile radius. For simple supported beams subjected to a central force P, the strain energy is given by

where the P is central force, L is length of the beam, E is Young's Modulus and I is moment of inertia. (Cantwell and Morton, 1989)

Figure 2.6 Shear Failure Mode (Abrate, 1998)

Kinetic energy absorbed by the target plate was determined by using the striking (Vs) and residual velocities (Vr) of the projectile. Energy absorption was calculated from the difference in kinetic energies of the projectile before impact and after perforation as given by the following formula (Wan Awis, 2009):

2.8 Finite Element Analysis

Matthew (2003) explained in his book Finite Element Modeling of Composite Materials and Structures that the finite element (FE) analysis is merely an alternative approach to solving the governing equations of a structural problem. hence, FE and classical methods will produce identical results for the same problem, provided the former method is correctly applied.

Prevorsek et al (1993) used a finite element model for simulating the deformation of a composite plate during ballistic impact, and a finite difference model for determining the temperature rise during the event. The analysis showed that a significant temperature rise occurs at the projectile-composite interface, but because of the short duration of the impact and the low thermal conductivity of the composite, this temperature rise is confined of a very small region around the interface. The volume of material affected is too small to have any effect on performance.

Buitrago et al (2009) was analyzed the perforation of composite sandwich structures subjected to high-velocity impact using finite element model in ABAQUS/Explicit. The experimental tests provided information only about the velocity of the projectile before the impact over the front skin and after the perforation of the back skin. However, the finite element model showed the evolution of the projectile while it was crossing through the sandwich plate.

Talebi et al (2009) investigated the effects of projectile nose angle on fabric impact; describe fabric deformation and failure under such type of loading. They obtained energy absorption trends in accordance with projectile nose angle and found the maximum projectile efficiency when penetrating into fabric armor.




Advanced composites are a comparatively new engineering material. As a result, reliable database of material properties are quite rare. Generating property databases, therefore, is often an important part of composite engineering projects. To do rightly in this research, the methodology for this project will guide this research from the beginning until the final result.

First of all, the acid digestion method must be done to determine the ply and weight of epoxy composite skin. The acid digestion method is conducted according to standard ASTM D3171. Then for the behavior of composite sandwich structure, the tensile test will be performing under the standard ISO 1924-2:1994 en / ASTM D 3518. The flow of this research is clearly shown in Figure 3.1.

Figure 3.1 Mechanical and Modeling Testing Flow on Carbon Fiber and Sandwich Composites

Acid Digestion Method

Chemical digestion is described in ASTM D3171. A chemical must be selected that does not damage the fibers. Typical selections consist of sulfuric acid and hydrogen peroxide for peek and polyimide, nitric acid for epoxy and others as described in ASTM D3171. The fiber volume fraction is calculated as

where Vf = volume fraction of fibers, Wf = weight of fibers, Wm = weight of matrix, = density of fibers and = density of matrix

Nitric Acid Digestion Method Procedure

The carbon fiber specimens were then placed in individual 100-ml glass beakers marked No. 1, 2, and 3.

These beakers were filled with 60 ml of 70% nitric acid, covered with a watch glass, and placed on a hot plate heated to approximately 120°C (250°F).

The specimens remained in the beakers on the hot plate for roughly 1 hour after the nitric acid began to boil, or until, based on a visual inspection, no epoxy matrix material remained, bonding the individual fibers together.

When this point had been reached, the beakers were removed from the hot plate and allowed to cool.

The nitric acid was carefully drained off so that all fibers remained in the beaker.

The nitric acid was poured into a waste container for disposal.

The beakers containing the carbon fibers were then refilled with 100 ml of distilled water.

The fibers were gently swirled in the beakers using a glass stirring rod to clean the acid and epoxy matrix residue from the fibers.

Next, the distilled water was carefully drained off and disposed of.

This process was repeated two more times, followed by a final rinse using 95% ethyl alcohol.

The beakers were then placed in a drying oven at 49°C (120°F) for a minimum of 8 hours to thoroughly dry the fibers.

The beakers were then placed in a sealed desiccator and allowed to cool to room temperature.

The beakers were then removed from the desiccator and the fibers in the individual beakers were weighed using the Mettler balance. This weight was recorded for each specimen.

The dry weight of each fiber and void volume specimen, the submerged weight of each specimen, the weight of the fibers alone, and the densities of the fiber and matrix materials were used to calculate the fiber and void volume percentages of each specimen as described in ASTM D 3171-76 (1992) and ASTM D 792-66 (1992).

Mechanical Testing

The mechanical testing of composite structures to obtain parameters such as strength and stiffness is a time consuming and often difficult process. It is, however, an essential process, and can be somewhat simplified by the testing of simple structures, such as composite sandwich structure. The data obtained from these tests can be directly related with varying degrees of simplicity and accuracy to any structural shape. Some, such as the tensile test, are widely recognized as standards, whereas there are dozens of different tests for the measurements of shear properties.

Tensile Testing

Tensile testing is runs according to ASTM standard D 3518 / ISO 14129:2002 en. Tensile testing utilizes the test geometry as shown in figure 3.2 and consists of two regions: a central region called the gauge length, within which failure is expected to occur, and the two end regions which are clamped into a grip mechanism connected to a test machine.

Figure 3.2 Typical Tensile Composite Test Specimen (Hodgkinson, 2000)

These ends are usually tabbed with a material such as aluminum, to protect the specimen from being crushed by the grips. This test specimen can be used for longitudinal, transverse, cross-ply and angle-ply testing.

Figure 3.3 Tensile Testing (

Figure 3.4: Screw Action Grips With a 50 mm Gauge Length Clip-on Extensometer on Specimen. (Hodgkinson, 2000)

Finite Element Modeling

The orthotropic nature of each layer of sandwich laminate is represented in MSC Dytran so that stacking sequence and material properties of the composite sandwich structure can be properly incorporated into the analysis.

For the methodology of finite element modeling, firstly, the composite must be created by MSC Patran. The ply of the composite epoxy skin must be determined. Secondly, the core will be model by using solid geometry. For the recently progress, the modeling are shown in figure below.

Making a composite model.

Figure 3.5: Create Geometry

Figure 3.6: Create Mesh Seeds

Figure 3.7: Create Surface Mesh

Figure 3.8: Apply Boundary Conditions to The Model

Figure 3.9: Apply Load to the Model

Figure 3.10: Check Element Normals

Figure 3.11: Define a Material Coordinate System

Figure 3.12: Apply the Material Coordinate System to the Elements

Figure 3.13: Analysis and Attach XBD Result File

The result for this modeling cannot obtain because during the analysis, the error is occurred.

Create Rigid penetration ball

Figure 3.14: Create Geometry (Point)

Figure 3.15: Create Geometry (Surface)

Figure 3.16: Create Mesh

Figure 3.17: Mirror/Transform Mesh

Figure 3.18: Equivalence All Nodes

Figure 3.19: Define Ball Material Properties

Figure 3.20:

Figure 3.21:

Figure 3.22: