Composites are hybrids of two or more materials such as reinforced plastics, metals or ceramics. The reinforcements may be in the form of fibers, particles, whiskers or lamellae and are incorporated in a suitable matrix, thereby providing a material that combines the most useful properties of the constituents. Generally, the properties of a composite are superior to those of its individual constituents.
MATRIX OF FIBER REINFORCED PLASTICS
The primary functions of the matrix are to transfer stresses between the reinforcing fibers (hold fibers together) and protect the fibers from mechanical and environmental damages.
Most matrices are made up of resins for their wide variation in properties and relatively low cost. Common resin materials include
Resin Matrix
Epoxy
Phenols
Polyester
Polyurethane
Vinyl Ester
Among these resin materials, polyesters are the most widely used. Epoxies, which have higher adhesion and less shrinkage than polyesters, come in second for their higher cost. In the aerospace industry, epoxy is used as a structural matrix material or as structural glue.
Although less common, non-resin matrices (mostly metals) can still be found in applications requiring higher performance at elevated temperatures, especially in defence industry.
(b) Metal Matrix
Aluminium
Copper
Lead
Magnesium
Nickel
Silver
Titanium
(c) Non-metal Matrix
Ceramics
REINFORCEMENTS
The reinforcements of composites may be in the form of particles, whiskers, fibers, lamella or a mesh as shown in Fig.1.1 They may increase the strength, stiffness or modify the failure mechanism advantageously. There can be special cases where the fibers may conduct or resist heat and electricity. Whiskers are single crystal fibers or platelets almost from defects. Filamentary whiskers may have diameters of 0.1 to 1 μm and a length of order of 100 μm and 1000 μm with strength levels approaching theoretical values in those cases which are finer and free from defects. Whiskers of metals, inter-metallic, oxides, carbides and nitrides are frequently used as reinforcements.
FIG.1.1. COMPOSITE MATERIAL WITH DIFFERENT TYPES OF REINFORCEMENTS AND FIBER ORIENTATIONS
A variety of continuous fibers of glass, carbon, Kevlar (aramid), silicon carbide, alumina, boron, tungsten etc. are used as reinforcements. These fibers exhibit different combinations of mechanical properties, physio-chemical properties, and electromagnetic properties. Some properties of various types of fibers are shown in Table 1.1. The cost and easy availability of these fibers is also important in the selection of the fibers for reinforcement. Composites can be classified roughly into three types according to the fiber types:
Particulate
Short Fiber
Long Fiber
Laminate
Particulate composite
It consists of the composite material in which fiber materials are roughly round. An example of this type of composite would be a reinforced concrete where the cement is the matrix and the sand serves as the fiber. Lead particle in copper matrix is another example where both the matrix and fiber are metals.
Short and long fiber composites
These composites are composites in which the fiber matrix has the strength to diameter ratio (L/D) greater than 10. Short fiber composites are generally taken to have L/D of approximately 100 and while long fiber would have L/D approximately = ∞. Fiber glass filler for both panels is an example of short fiber composite. Carbon Fibers Aramid fibers are examples of long fiber composites. The properties of various fibers are given in table 1.1.
Laminate Composites
It is the type of composite that uses the fiber material in form of a sheet instead of round particles or fiber. Formica countertop is a good example of this type of composite. The matrix material is usually phenolic type thermoset polymer. The fiber could be of any material from craft paper (Formica) to canvas (canvas phenolic) to glass.
The composites are non-homogeneous; the resulting properties will be the combination of the properties of the constituents materials, the different type of loading may fall on different component of the composite to take the load. This implies that material properties of composite materials may be different in tension and in compression as well as in bending.
Material Form Diameter Filament (µm) Density (gm/cm³) Young’s modulus (GPa) U.T.S. (GPa) Melting/ Softening/ Decomposition ºC
CARBON :
HIGH STRENGTH Tow 7-8 1.7-1.8 220-250 2.5-3.5 3650
HIGH MODULUS Tow 7-8 1.8-1.9 340-380 2.2-2.4 3650
ULTRAHIGH MODULUS Tow 8-9 1.9-2.1 520-550 1.8-1.9 3650
GLASS :
E-GLASS Tow 9 2.6 72 3.45 700
S-GLASS Tow 9 2.49 87 4.6 840
QUARTZ Tow 10 2.2 69 0.9 1650
KEVLAR-49 : Tow 16 1.44 124 4.1 250
TABLE1.1. TYPICAL PROPERTIES OF FIBERS
FIG.1.2. CLASSIFICATION OF COMPOSITES ON BASIS OF FIBERS
FRP REINFORCEMENT
FRP composite materials are comprised of high strength continuous fibers, such as glass, carbon, or steel wires, embedded in a polymer matrix. The fibers provide the main reinforcing elements while the polymer matrix (epoxy resins) acts as a binder, protects the fibers, and transfers loads to and between the fibers.
FRP composites can be manufactured on site using the wet lay-up process in which a dry fabric, made of carbon or glass, is impregnated with epoxy and bonded to prepared concrete substrate. Once cured, the FRP becomes an integral part of the structural element, acting as an externally bonded reinforcing system. FRP composites can also be prefabricated in a manufacturing facility in which the material is pultruded to create different shapes that can be used for strengthening applications, such as rods, bars and plates.
The most common FRP systems for concrete strengthening applications are carbon fiber based (CFRP). Carbon has superior mechanical properties and higher tensile strength, stiffness, and durability compared with glass fiber based systems. The use of prefabricated CFRP bars and plates is typically limited to straight or slightly curved surfaces; for example, the top side or underside of slabs and beams. Prefabricated FRP elements are typically stiff and cannot be bent on site to wrap around columns or beams.
FRP fabric, on the other hand, is available in continuous unidirectional sheets supplied on rolls that can be easily tailored to fit any geometry and can be wrapped around almost any profile. FRP fabrics may be adhered to the tension side of structural members (e.g. slabs or beams) to provide additional tension reinforcement to increase flexural strength, wrapped around the webs of joists and beams to increase their shear strength, and wrapped around columns to increase their shear and axial strength and improve ductility and energy dissipation behavior.
The adhesive systems used to bond FRP to the concrete substrate may include a primer that is used penetrate the concrete substrate and improve bond of the system; epoxy putty to fill small surface voids in the substrate and provide a smooth surface to which the FRP system is bonded; saturating resin used to impregnate the fabric and bond it to the prepared substrate; and protective coating to safeguard the bonded FRP system from potentially damaging environmental and mechanical effects. Most epoxies for FRP strengthening systems are adversely affected by exposure to ultraviolet light, but can be protected using acrylic coatings, cementitious coatings, and other types of coatings.
The resins and fiber for a FRP system are usually developed as one system, based on materials and structural testing. Mixing or replacing a component of one FRP system with a component from another system is not acceptable and can adversely affect the properties of the cured system.
The bond between FRP system and the existing concrete is critical, and surface preparation is essential to most applications. Any existing deterioration or corrosion of internal reinforcement must be resolved prior to installation of the FRP system. Failure to do so can result in damage to the FRP system due to delamination of the concrete substrate.
FRP APPLICATION
FRP systems provide a very practical tool for strengthening and retrofit of concrete structures, and are appropriate for:
Flexural strengthening,
Shear strengthening, and
Column confinement and ductility improvement.
FRP systems have also been successfully used for seismic upgrading of concrete structures. These applications include mitigating brittle failure mechanisms such as shear failure of unconfined beam-column joints, shear failure of beams and/or columns, and lap splice failure. FRP systems have also been to confine columns to resist buckling of longitudinal steel bars. These FRP schemes increase the global displacement and energy dissipation capacities of the concrete structure, and improve its overall behavior.
Because of the resistance to corrosion, FRP composites can be utilized on interior and exterior structural members in all almost all types of environments.
ADVANTAGES OF THE COMPOSITE MATERIAL
Fiber reinforced plastic (FRP) composite possess high strength and stiffness.
Fatigue strength to weight ratio as well as fatigue damage tolerance of many composite laminates is excellent compared to metals.
Fiber reinforced possesses high internal damping. This leads to better vibration energy absorption within the material and result in reduce transmission of noise, vibration and hardness (NVH).
FRP construction is low maintenance cost.
In space vehicle , reduction in weight is linked to fuel savings
FIG.1.3. APPLICATION OF FRP MATERIALS IN A PASSENGER AIRCRAFT STRUCTURE
CARBON MATRIX COMPOSITES
Carbon is a remarkable material. It includes materials ranging from lubricants to diamonds to structural fibers. The forms of carbon matrices resulting from the various carbon–carbon manufacturing processes tend to be rather weak, brittle materials. Thermal conductivities range from very low to high, depending on precursor materials and processes. As for ceramics, in situ matrix properties are difficult to measure.
PROPERTIES OF CARBON MATRIX COMPOSITES
Carbon matrix composites consist of carbon or ceramic fibers embedded in carbon matrices. The most important subset of carbon matrix composites is carbon–carbon composites, which have carbon fiber reinforcements. As for other composites, there are a wide range of materials that fall in this category. The variables affecting properties include type of fiber, reinforcement form and volume fraction, and matrix characteristics. Historically, carbon–carbon composites were first used because of their excellent resistance to high-temperature ablation (erosion at high temperature). Strengths and stiff nesses of early materials were low, but these properties have steadily increased over the years. One of the most significant limitations of carbon–carbon composites is oxidation, which begins at a temperature threshold of approximately 370∘C (700∘F) for unprotected materials. Addition of oxidation inhibitors in the matrix and ceramic coatings raises the threshold substantially. In inert atmospheres or vacuum, carbon–carbon composites retain their properties to temperatures as high as 1380∘C (2500∘F). In recent years, silicon carbide fiber-reinforced carbon has been used in production aircraft gas turbine engine flaps. A significant advantage of this material is that its CTE is closer to that of ceramic coatings like SiC, reducing coating cracking and spallation.
MECHANICAL PROPERTIES OF CARBON MATRIX COMPOSITES
Carbon matrices are weak, brittle, low-stiffness materials. As a result, transverse and through-thickness elastic moduli and strength properties of unidirectional carbon–carbon composites are low. Fabric-reinforced materials are weak in the through-thickness direction. Because of this, three-dimensional reinforcement forms are often used. In the direction of fibrous reinforcement, it is possible to obtain moduli as high as 340 GPa (50 Msi), tensile strengths as high as 700 MPa (100 ksi), and compressive strengths as high as 800 MPa (110 ksi). In directions perpendicular to fiber directions, elastic moduli are in the range of 10 MPa (1.5 ksi), tensile strengths 14 MPa (2 ksi), and compressive strengths 34 MPa (5 ksi). The material used on the Space Shuttle Orbiter, was developed decades ago. It is reinforced with carbon fibers made from a rayon precursor, which is relatively weak. This material is called reinforced carbon–carbon (abbreviated RCC), which is redundant. It is sometimes called reusable carbon–carbon. The second material, designated ACC-4, is reinforced with T300, a PAN-based standard modulus fiber. In-plane strength and moduli are much greater for the latter composite.
PHYSICAL PROPERTIES OF CARBON MATRIX COMPOSITES
There are a wide variety of carbon–carbon composites, which have a wide variety of physical properties. Properties depend on the type of fibers used and the precursor and processes employed to produce the matrix. Densities generally fall in the range of 1.4–1.8 g/cm3. In-plane thermal conductivities range from less than 10 to about 350 W/m⋅K. Through-thickness conductivities is generally very low. CTEs are also very low, generally ranging from -1 to +1 ppm/K
FACTORS AFFECTING FAILURE
High performance fiber reinforced plastic components can be considerably weakened by the introduction of holes and cut-outs, due partly to the large stress concentrations 24 that occur in the region around such discontinuities and partly to a lack of plasticity. For example, by virtue of the high degree of anisotropy of unidirectional carbon fiber reinforced plastics (CFRP), the tensile elastic stress concentration factor due to a circular hole in a large sheet can be as large as 8 in contrast to the much lower value of 3 normally 25 associated with isotropic materials . Furthermore, as most isotropic materials exhibit some plasticity, yielding can take place in highly stressed regions and the effect of stress concentrations on the final net failing stress is small; such is not the case for unidirectional CFRP which is generally elastic to failure 26 and the effect of stress concentration is to give rise to correspondingly low net failing stress. It comes as no surprise, therefore, that the efficiency of bolted joints in unidirectional CFRP is very low indeed.
However, if in the region of the bolt holes the degree of anisotropy could be reduced and some plastic or pseudo-plastic behavior introduced then efficiency might be expected to increase considerably. Fortunately, such joint ‘softening’ can be readily achieved by the incorporation of fiber oriented in different directions and, moreover, with the use of pre-impregnated unidirectional sheet material to form multi-directional multi-ply laminates, this can be done without a significant increase in fabrication complexity. (It is worth noting here that joint softening might also be achieved by other techniques such as by the incorporation of other types of fibrous material, for example glass fiber or Kevlar. However, this is not under consideration at present but may be the subject of a further investigation.
Tension
As with conventional materials, the tensile load required to fail an otherwise uniform, plain laminate, through a cross-section at which holes occur (net section) is less than at the section at which there are no geometric inclusions (gross section).
Bearing
A substantial degree of lateral constraint at the ends of a compressive specimen is necessary to prevent premature end failure due to a breakdown in the fiber resin interface and a consequent brush-like failure.
Bearing of a pin in a hole gives rise to similar compressive stresses in the material around the hole and it is therefore to be expected that lateral constraint will influence the magnitude of the composite ultimate bearing strength.
In the view of the above case, degree of lateral constraint is likely to be an important parameter in the determination of bearing strength.
Shear
The shear strength normally quoted for a composite is the interlaminar shear strength. Unfortunately, except for some unidirectional materials in which isotropy can be assumed on all planes normal to the fiber direction, it is of little use in the estimation of joint shear strengths where failure is due to inplane shear stresses. In-plane shear strength can be measured using the rail shear test but, due to shear stress concentrations around a loaded hole, it is unlikely to be a representative test. Shear strengths are therefore measured using pin shear pull-out.
STRESS CONCENTRATION
A stress concentration (often called stress raisers or stress risers) is a location in an object where stress is concentrated. An object is stronger when force is evenly distributed over its area, so a reduction in area, e.g., caused by a crack, results in a localized increase in stress. A material can fail, via a propagating crack, when a concentrated stress exceeds the material’s theoretical cohesive strength. The real fracture strength of a material is always lower than the theoretical value because most materials contain small cracks or contaminants (especially foreign particles) that concentrate stress. Fatigue cracks always start at stress raisers, so removing such defects increases the fatigue strength.
1.8.1. CAUSES
1) Geometric discontinuities cause an object to experience a local increase in the intensity of a stress field. Examples of shapes that cause these concentrations are cracks, sharp corners, holes, and changes in the cross-sectional area of the object. High local stresses can cause objects to fail more quickly, so engineers must design the geometry to minimize stress concentrations.
2) Due to discontinuities in applied loads.
3) Material discontinuities which may occur while manufacturing.
1.8.2. PREVENTION
A counter-intuitive method of reducing one of the worst types of stress concentrations, a crack, is to drill a large hole at the end of the crackThe drilled hole, with its relatively large diameter, causes a smaller stress concentration than the sharp end of a crack. This is however, a temporary solution that must be corrected at the first opportune time. It is important to systematically check for possible stress concentrations caused by cracks—there is a critical crack length of 2a for which, when this value is exceeded, the crack proceeds to definite catastrophic failure. This ultimate failure is definite since the crack will propagate on its own once the length is greater than 2a. (There is no additional energy required to increase the crack length so the crack will continue to enlarge until the material fails.) The origins of the value 2a can be understood through Griffith’s theory of brittle fracture.
Another method used to decrease the stress concentration is by creating the fillet at the sharp edges. It gives smooth flow of stress streamlines. In a threaded component force flow line is bent as it passes from shank portion to threaded portion as a result stress concentration takes place. To reduce this a small undercut is taken between shank and threaded portion.
OBJECTIVE AND SCOPE OF THE PROJECT
An empirical study was conducted to scrutinize the effect of multiple holes on the mechanical properties of CFRP (Carbon Fiber Reinforced Polymer) composites.
The hole stress concentration effects on the tensile strength and of Carbon Fiber Reinforced Polymers (CFRP) composites were reviewed with consideration being given to a multiple holes drilled through the specimen with discrete arrangements.
The main objective of this study was to investigate the effect of variously arranged hole arrays on the strength and failure modes of laminated composites. The understanding of the interaction effect that occurred provided a guideline in designing fastened repairs in laminated composites.
This data led to better understanding of the effects of various damage scenarios and the corresponding repair design needed to restore a laminate’s structural integrity. The aircraft undergoes various inspection and maintenance procedures and also in order to design an aircraft the design specifications are restricted to a particular or rather a specific value in order for the aircraft to be efficient and majority of the aircraft failures which are aimed to be considered to be under Fail Safe condition after the occurrence of the failures. Holes are drilled for various design purposes as well as are considered under structural failures. Inspections such as NDT (i.e. Fiberscope) are carried out on internal complex structures of an aircraft for detection of cracks and flaws. Such inspections are required to have holes on the structure for non-destructive testing procedures. The hole pattern plays an essential part while drilling onto the structural components of the aircraft. An aircraft undergoes various loads after it is airborne. The hole on the structural component results in structural failure after it exceeds the assigned loading parameters. In order to delay the failure process, the holes are drilled in discrete patterns such that the distribution of the stress concentration factor around the circumference of the hole avoids the rapid failure of the component. The scope of the project is illustrated in Figure 1.4. below.
EXPERIMENTAL METHODOLOGY
FABRICATION OF THE LAMINATE
PREPARATION OF THE CARBON FIBRE :
Carbon Fibre mat of dimension 500mmX270mm were cut from the roll.
Seven such Carbon Fibre mats were initially required in order to prepare a carbon fibre reinforced polymer laminate.
The weight of all seven Carbon Fibre mats was measured using an electronic weighing machine.
PROCEDURE FOR FABRICATION OF CFRP LAMINATE :
The carbon fibre with the following properties:
Thickness 0.2mm
200gsm BIDIRECTIONAL
was employed in order to fabricate the required composite panel of dimensions 500mmX270mmX2mm.
The fibres were first put in place in the mould. The fibers were in the form of woven. Then the resin was impregnated. The impregnation of resin was done using rollers, brushes. The impregnation helped in forcing the resin inside the fabric. The laminates fabricated by this process were then cured under standard atmospheric conditions.
ADVANTAGES OF HAND LAY-UP PROCESS
The process resulted in low cost tooling with the use of room-temperature cure resins.
The process was simple to use.
Higher fibre contents and longer fibres was the output when compared to other processes.
DISADVANTAGES OF HAND LAY-UP PROCESS
Since the process was worked by hands/manually, there were safety and hazard considerations.
The resin needed to be less viscous so that it could have been easily worked by hands.
Uniform distribution of resin inside the fabric was not possible. It led to voids in the laminates.
The resin used was epoxy LY 556 and the hardener HY951 was added to provide good bond strength. The resin and hardener were mixed in 10:1 ratio and the layers of this mixture were applied between the layers of fibers to provide good bond. The curing was done at room temperature for 24 hours. A total of 7 layers were laid during the fabrication of this panel and all the layers were laid in the similar manner. The properties of the carbon fibers are given in Table 3.1. The fabricated laminate is shown in Figure 3.1.
The detailed procedure of the hand lay-up procedure of Carbon Fiber /Epoxy Matrix is given below:
The table is cleansed with a thinner.
Next, the first layer of Carbon Fiber is placed and then resin is applied over the first layer of Carbon Fiber.
Rollers are then used in order to squeeze out or eliminate the excess resin.
The second layer is thus placed post the previous step and resin is then applied over the second layer.
Rollers are used after the application of resin over every layer in order to eliminate the excess resin.
The Steps 1-4 are repeated for every alternating layer of Carbon Fiber mats and resin mixture until 7 layers of Carbon Fiber are completed.
A plastic film cover or Mylar sheet with wax applied on both the sides is then placed over the wet laid-up laminate and is sealed.
Heavy weights are then placed over the wet laid-up laminate in order to exert high pressure and the laminate is cured at room temperature for 24hours.
The above procedures resulted in a laminate thickness of 1.9- 2.4mm.
PROPERTIES OF CARBON FIBERS
Sr. No. Property Magnitude
1. Tensile strength (KN/mm2) 3.53
2. Tensile Modulus (KN/mm2) 230.284
3. Elongation % 1.52
4. Density (gm/cm3) 1.76
TABLE 3.1
FIG.3.1. FABRICATED LAMINATE
PREPARATION OF THE SPECIMENS
The total no. of 5 specimens were cut from the prepared composite laminate into the required dimensions by following the procedures mentioned in ASTM D3039 for tensile testing of FRP composite laminate. The dimensions of the specimen are shown in Figure 3.2.
FIG.3.2. SPECIMEN DIMENSIONS
The water-jet cutting technique was used to cut all the specimens to avoid generation of micro-cracks while cutting and to attain good surface finish. After the successful cutting of specimens, holes of 5 mm diameter was introduced into the specimens by using vertical drilling machine. To eject an extremely high velocity jet through a nozzle orifice, highly pressurized water jet was passed through it at a high pressure setting of 37,000PSI. Due to heat dissipation laser cutting was not preferred as it causes delamination and various other damages to the material. The Aluminum tabs of dimensions 50mmX25mmX2 mm were attached at both edges of specimens to provide proper grip between the specimens and testing machine jaws. The specimens before test are shown in Figure 3.3. A Standard Epoxy Araldite as an adhesive in order to clamp the specimens and provide the necessary bonding between the Aluminium tabs and the composite laminate.
FIG.3.3. LARGE SAMPLE FOR CONSISTENCY
Before drilling holes the pitch, row distances were referred. And the values taken for pitch and row distances are 2.5D each. Where D is the diameter of the hole. The diameter of was found to be in the range of 4-6 mm. So as to accommodate the holes on the specimen holes were drilled with diameter of 5mm. For the above chosen diameter, an appropriate W/D ratio was taken into consideration. A standard vertical drilling machine to drill circular holes onto the specimen, illutsrated in the Figure 3.4.. A system of high rpm and low feed was used to have the least impact the specimen. One, two, four and five central hole configurations were drilled on the specimen. Five Configurations of hole orientation patterns is illustrated in Figure 3.5, in a carbon/epoxy composite laminate with a layup of [0/±90] were studied. Configuration A has one hole and is designated as SH. Configuration B has two holes vertically placed and is designated as DH. Configuration C has three holes placed on the three edges of a triangle and is designated as TH. Configuration D and E have four holes placed as a square and a diamond array respectively and are designated as FH and FDH respectively
FIG.3.4. DRILLING OF HOLES
FIG.3.5. SPECIMEN CONFIGURATION
TENSILE TEST AND TEST SETUP
A Universal testing machine was used to test the tensile properties of all the specimens. The ASTM D3039 procedure was followed to perform the tensile test of all the specimens. The tensile test was performed at room temperature by placing the specimen in the jaws of the machine and it was pulled until failure.
Specimens were fixed properly before starting the machine to avoid any kind of slippage during testing. The cross-head travel was maintained as 0.1inch/min throughout the test. The test setups of notched specimens are shown in Figure 3.6.
FIG.3.6. TEST SETUP
PROPERTIES OF ARALDITE LY556/HARDENER HY951
Anhydride-cured, low viscosity standard matrix system with extremely long post life.
The reactivity of the system is adjustable by variation of the accelerator content.
The system is easy to process, has good fibre impregnation properties and exhibits excellent mechanical, dynamic and thermal properties.
It has an excellent chemical resistance especially to acids at temperatures up to 80°C.
Viscosity at 25°C. MPa.s 10,000-12,000
Density at 25°C. g/cc 1.15-1.2
Vapour Pressure at 25°C. Pa <0.01
TABLE 3.2 ARALDITE LY556 PROPERTIES
Viscosity at 25°C. MPa.s 10-20
Density at 25°C. g/cc 0.97-0.99
Vapour Pressure at 25°C. Pa -390
TABLE 3.3 HARDENER HY951 PROPERTIES
Araldite LY556 100 parts by weight
Hardener HY951 10-12 parts by weight
TABLE 3.4.MIXING RATIO
Property Specific Gravity Tensile Strength Tensile Modulus
Units – MPa GPa
Value 1.28 82.74 3.792
TABLE 3.5.PROPERTIES OF EPOXY MATRIX
RESULTS AND DISCUSSIONS
INTRODUCTION
This section contains complete and detailed results, interpretations and graphs for achieving the scope of this project, study on failure of CFRP composites with multiple holes. The outputs were analyzed and interpreted. Parameters such as elongation, strain, load, and ultimate or peak tensile strength were tabulated.
TABULATIONS OF RESULTS
Tables have been prepared according to the output data from the Universal Testing Machine. Parameters such as elongation, load, Young’s modulus, strain, ultimate strength were tabulated.
Property Configuration- A Configuration- B Configuration- C Configuration- D Configuration- E
Max. Elongation (mm) 9.9
10.6
6.8
8.5
6.6
Max. Load (KN) 16.58
17.16
9.88
13.36
8.12
Strain at Ultimate strength (no units) 0.066
0.070667
0.045333
0.056667
0.044
Ultimate Strength (N/〖mm〗^2) 331.6
343.2
197.6
267.2
162.4
Rate of feed (mm/min) 5 5 5 5 5
TABLE4.1. PARAMETERS OBTAINED
The test results of each configuration are given in Figure 4.1. The test data scatter of two specimens is also shown. The data indicates that configuration B has the highest strength, whereas configuration E, having the four holes in a diamond array, has the lowest strength.
All of the test specimens exhibited an edge delamination before the presence of the visible damage around the hole. The final failure occurs in the neighbourhood of the hole for all of the specimens except for configuration A, the specimen with a single hole. For the single hole specimen, multiple failures are observed not only in the neighbourhood of the hole but also at locations away from the hole. For this specimen, failure away from the hole can be attributed to the effect of edge delamination.
FIG.4.1. TEST RESULTS
A comparison of configurations A and B is shown in Figure 4.3. For configuration B, significant failure was observed along the longitudinal direction. This signifies that the composite failed due to ply splitting. Configuration B has more strength because of an additional hole in-line with the load, resulting in the reduction of peak stress at the hole.
Single and Double Hole
FIG.4.2.FAILURE COMPARISONS OF LAMINATES WITH VARIOUS CONFIGURATIONS
The graphs corresponding to configurations A and B are given in the Figure 4.2.
CONFIGURATION A
CONFIGURATION B
FIG 4.3.STRESS V/S STRAIN GRAPH COMPARISION
Figure 4.3(a) shows the failure comparison of configurations C and D specimens. Significant fibre breakage is observed between the two side-by-side holes in configuration D. Configuration D exhibits similar failure pattern as configuration B, more inclined towards longitudinal direction. Although configuration D laminate has four hole the laminate strength is higher than configuration C because of the fibre breakage between the holes. Because of fibre breakage between holes, the strength of configuration C is lower than configuration. D. For the four-hole specimens, Figure 4.3(b), extensive fibre breakages are shown in configuration E resulting in the lowest strength among all of the configurations.
(b)
FIG4.4.FAILURE COMPARISONS OF LAMINATES WITH VARIOUS CONFIGURATIONS
The graphs corresponding to the above configurations C, D, and E are given in the Fig 4.4.
CONFIGURATION C
CONFIGURATION D
CONFIGURATION E
FIG.4.5.STRESS VS STRAIN GRAPHS OF VARIOUS CONFIGURATIONS
In the below Figure 4.5 the Load Vs Deflection of various configurations is given.
CONFIGURATION A
CONFIGURATION B
CONFIGURATION C
CONFIGURATION D
CONFIGURATION E
FIG.4.6.LOAD VS DEFLECTION GRAPHS OF VARIOUS CONFIGURATIONS
CONCLUSION:
A study is conducted to investigate the effects of multiple holes on composite laminate strength experimentally and using finite element method. The multiples holes are in analogy with the circular cut-outs created on the damaged laminate during composite structural repair. The finite element study is extended to study the effects of different shape configurations on the laminates with multiple holes. The stress concentrations occur at the edge of the hole periphery where the fibres are tangent to the hole. The magnitude of stress concentrations in the holes depend on the arrangement of the holes in the laminate. The peak stresses at the hole with multiple hole arrangements is studied.
All the test specimens exhibited edge delamination before the presence of visible damage around the hole. The final failure occurs near the hole for all specimens. The final failure of SH specimen reveal a significant delamination at the neighbourhood of the hole and free edge. This may result higher strength of the specimen.
So, to summarize:
The peak stress occurs at the periphery of the hole.
The arrangement of holes in the laminate plays a significant role is determining the peak stresses.
Obtained in the following order Configuration B (Top and bottom hole) > A (single hole) >D (squared array) >C (Triangular pattern)> E (Diamond array). This sequence is in agreement with the test indication where configuration B has the highest strength and Configuration E has the least strength.
FUTURE ENHANCEMENTS
This project can be used as a base for future projects for development and comparison. It can be done by the following ways,
The tests can be carried out with different displacement rates and the results can be compared.
The progressive failure of the composites can be studied and analysed in detail using FEM (Finite Element Method).
The test can be carried out under uniaxial compressive load.
Different composite material can be used to find the progressive failure.