Introduction
1.1
Composite materials
Composite Materials, [1], (or simply Composites) are formed by combining two or more materials in order to obtain better properties than those of the individual com-ponents used alone. Unlike metallic alloys, here each component remains separate and distinct in the final product and retains its chemical, physical and mechanical properties.
The aeronautic industry, always looking for lighter materials, started being in-terested in composites during World War II with the introduction of fiberglass. In this material, the first fiber reinforced plastic (FRP), thin glass fibers with high tensile strength (reinforcement) are embedded in an homogeneous plastic polymer (matrix) that keeps them together, redistributes stresses and carries compression and shear loads. The resulting material presents extremely high strength-to-weight and stiffness-to-weight ratios, allowing the creation of very light structures if com-pared to classic construction metals, including aluminum. Later, in the 60’s and 70’s, the space age led to the development of new materials, in particular new rein-forcing fibers such as carbon, boron and aramid. Lighter and stronger than glass,
carbon fibers took the lead also over the others thanks to their superior processing capabilities and lower cost, quickly becoming the most used reinforcing fibers in in-dustry. After an initial lack of confidence, Carbon Fiber Reinforced Plastics (CFRP) are nowadays replacing metals even in primary structures of vehicles and aircraft. Examples are military vehicles like the Bell-Boeing V-22 Osprey (Figure 1.1), where 43% of the airframe structure is made of composite materials, [2], or civil aircraft like the Boeing 787 Dreamliner (Figure 1.2) in which the percentage of composites reaches 50% of the structure’s weight and 80% of its volume, [3].
Figure 1.1: A Bell-Boeing V-22 Osprey during landing
Many types of composite materials exist. Their definition is very wide and includes materials used by mankind since its very first age, like the combination of straw and mud used to form bricks for building construction. Also the combination of concrete and rocks or, in modern times, concrete and steel bars is a common example of composites, but the modern aerospace and automotive industry uses very particular and high performance materials that will be briefly described below. They are usually classified by the material used as reinforcement, the material used as matrix and the methodology of reinforcement.
Figure 1.2: Materials used in the structure of the Boeing 787 Dreamliner
Commonly used matrices are:
• Metals (MMC, Metallic Matrix Composites). Usually Aluminum or Tita-nium. Reinforcing metals is a good way to push further their limits in critical zones without adding resisting material and therefore weight. MMC however tend to be very expensive and difficult to fabricate: the reinforcement has to resist the high temperatures of molten metal, maintain its position during the infusion and be properly treated to avoid galvanic issues.
• Ceramic (CMC, Ceramic Matrix Composites). Ceramic materials are ex-tremely hard and resistant to chemical attack and high temperature, but also brittle. Reinforcing them can greatly increase their tenacity, but this class of materials is still used only for very specialized applications, usually related to high temperature conditions.
• Carbon (Carbon-Carbon). Carbon-Carbon is a very particular class of com-posites in which both reinforcement and matrix are made of carbon. It presents high thermal resistance and a low coefficient of thermal expansion, but lacks of impact resistance. It’s used for special applications such as high-performance brake discs or space application like the leading edge of the Space Shuttle orbiter’s wings (Figure 1.3).
Figure 1.3: Carbon-Carbon is used on the leading edge of the Space Shuttle orbiter to resist the intense heat encountered when re-entering the atmosphere
• Polymer (PMC, Polymer Matrix Composites). Polymeric materials present very poor properties, but are very light, cheap and tend to stick very well to an eventual fiber reinforcement, a very important property in composites. Despite their individual properties, a properly reinforced polymer is an in-credibly good material and in fact PMC represent the greatest percentage of composites used by the automotive and mechanical industry. Different types of polymers are used. The main distinction is between thermoplastic, that melt and can be formed at high temperature, and thermosetting that instead harden irreversibly when heated. The most commonly used one is Epoxy resin, a thermosetting polymer that presents very good mechanical properties, high
tenacity, good resistance to chemical attacks and very good adhesion to fibers. Its defects are the sensibility to humidity and ultra-violet light, but usually these problems can be avoided by simply painting and coating the part. A subsequent classification is based on the way in which the matrix is reinforced. The most common classes are:
• Particles. The easiest way to reinforce a material is to add randomly dis-tributed particles in it. In this way the isotropy is usually maintained and different properties can be improved depending on the size, the distribution and the nature of the reinforcing particles.
• Short fibers. Reinforce with short fibers is an excellent way to improve resistance, especially to fracture and fatigue. Short fibers can be parallel or randomly distributed in the matrix (whiskers) to maintain the isotropy. • Sandwich. Sandwich is a special class of composites specifically made to be
stiff and resistant to flexure. It is made by attacking two thin panels, which resist normal loads, on top and bottom of a soft and lightweight core that resists shear loads (Figure 1.4).
Figure 1.4: Structure scheme and picture of a sandwich laminate with honeycomb core
• Long fibers. Reinforcing with long fibers parallel to each other is the best way to optimize the properties in a particular direction. The resulting material is highly anisotropic, in particular orthotropic, and allow to orient the fibers in the direction of the stress paths, increasing efficiency and saving weight.
Long unidirectional fibers are the most widely used form of reinforcement and many materials have been developed in the last few years. In particular the aerospace industry is interested in the following one:
• Carbon. A single carbon fiber has a diameter of 5 − 10 µm and is composed mostly of carbon atoms (92% at least, [4]). Several thousand fibers are bundled together to form a tow, which may be used by itself or woven into a fabric. Such tow is relatively expensive to produce but presents a set of excellent properties (high stiffness, high tensile strength, low weight, high chemical resistance, high temperature tolerance and low thermal expansion) that make it the most commonly used reinforcing fiber in a wide set of high-performance applications such as military, aerospace, motorsports (Figure 1.5) etc.
(a) Rear spoiler of a supercar (b) Detail of a bycicle frame
Figure 1.5: Examples of parts made in carbon fibers. Nowadays the woven fabric texture is also used for aesthetic reasons.
• Glass. Glass fibers are not as strong and as stiff as carbon fibers, but are cheaper to produce and still have excellent mechanical properties. When glass is formed into long fibers it loses its brittleness, resulting in fact less brittle than carbon and, unlike this, it is a good thermal and electrical insulator. Another advantage over carbon is the absence of galvanic issues when coupled with aluminum: for this reason it is common practice to protect a carbon fiber layup with several glass fiber plies.
• Aramid. Aramid is a class of strong synthetic fibers. They present good mechanical properties, good heat resistance but, most important, extremely good resistance to impacts. This type of fibers is widely used in military applications as ballistic protection and can be used to reinforce a carbon fiber layup from impact damage.
• Boron. Boron fibers are obtained by deposition of boron particles over an extremely thin tungsten fiber. This process is very complex and expensive, but the resulting fiber presents extremely good mechanical properties (high modulus and high strength, also in compression) and heat resistance. Its cost however limits its use only to very limited high-performance applications that justify the expense.
This research is focused on the most common type of composites: long carbon fibers in epoxy resin matrix (CFRP, Carbon Fiber Reinforced Plastic). The main reason of its success is to be found in the excellent properties along the fiber direction, such as extremely high strength-to-weight and stiffness-to-weight ratios. In addic-tion, its orthotropic nature combined with the manufacturing procedure allows to easily tailor the material depending on the load path adding material only where is really needed, further increasing efficiency and reducing weight. The manufacturing procedure brings an additional advantage, the so called consolidation of parts: an
entire assembly of metal parts can be replaced by a single piece made of composite material, reducing complexity, costs and weight. Finally this material presents very high resistance to chemical attacks and is practically insensitive to fatigue damage, helping to save on maintenance costs.
The main disadvantage has to be found in its very poor resistance to impact damage. Impacts can create severe damage that can be barely visible or invisible to a visual inspection, but that may cause great reduction in load carrying capability, especially under compressive loads. Another disadvantage is the complexity in cre-ating analytic or numerical models able to properly describe and predict behavior, resistance and the effects of damage. This problem, together with the limited oper-ative experience, is what nowadays limits the extensive use of composite materials, forcing companies to use very high safety coefficients to obtain certification.
1.2
Manufacturing of composites
Modern composites are typically sold in the form of thin layers (called laminae or plies) of fibers parallel to each other (unidirectional) or woven into a fabric. Fibers can be dry or already embedded in uncured viscous matrix (pre-preg).
There are numerous methods for fabricating composite components, developed to meet specific design or manufacturing challenges. Selection of a method for a particular part, therefore, will depend on the materials, the part design and end-use or application. Composite fabrication processes involve some form of molding, to shape the resin and reinforcement. A mold tool is required to give the unformed resin/fiber combination its shape prior to and during cure. The most basic fabri-cation method for thermoset composites is hand layup, which typically consists of laying one by one dry or pre-preg plies by hand onto a mold tool until the desired thickness, strength and stiffness are reached, forming a laminate stack (Figure 1.6). Particular attention is required during manufacturing: air bubbles and inclusions,
even dust particles, may remain trapped between plies, forming micro-voids in the final product that greatly reduce the overall properties. For this reason the process is usually done in a controlled environment like a clean-room. In the case of dry plies, resin is applied after layup is complete (resin infusion), otherwise in the case of pre-preg the layup can pass directly to the curing phase.
Figure 1.6: Deposition of a composite ply on a mold tool
Several curing methods are available. The most basic is simply to allow cure to occur at room temperature. Cure can be accelerated, however, by applying heat, typically with an oven, and pressure, by means of a vacuum. For the latter, a vacuum bag (Figure 1.7), with breather assemblies, is placed over the layup and attached to the tool, then evacuated using a vacuum pump before cure.
The vacuum bagging process consolidates the plies of material and significantly reduces voids due to the off-gassing that occurs as the matrix progresses through its chemical curing stages. Many high-performance thermoset parts require heat and high consolidation pressure to cure conditions that require the use of an autoclave. Autoclaves, generally, are expensive to buy and operate. Manufacturers that are
Figure 1.7: A technician applies a vacumm bag in a Lamborghini™ facility
equipped with autoclaves usually cure a number of parts simultaneously. Computer systems monitor and control autoclave temperature, pressure, vacuum and inert atmosphere, which allows unattended and/or remote supervision of the cure process and maximizes efficient use of the technique. When heat is required for cure, the part temperature is ramped up in small increments, maintained at cure level for a specified period of time defined by the resin system, then ramped down to room temperature, to avoid part distortion or warp caused by uneven expansion and contraction. When this curing cycle is complete and after parts are demolded, some parts go through a secondary cure, during which they are subjected for a specific period of time to a temperature higher than that of the initial cure to enhance properties.
The stacking sequence greatly influences the overall properties of the part. Uni-directional plies are stacked one above the other orienting each ply in a specific direction to tailor the properties where is really necessary. For example if the stress path has a main orientation, most plies will be oriented with the fibers parallel to it: this can be taken as reference orientation, defined as 0◦. A laminate however must
have plies with different orientations, usually at 90◦ or ±45◦ with respect to the ref-erence, to resist secondary or unexpected loads and eventual misalignments. ±45◦ oriented plies are also used to increase impact resistance. Particular attention is necessary when choosing the stacking sequence: the Stiffness Matrix of a composite laminate normally presents coupling terms (for example coupling between traction and bending) that can be eliminated if the layup is both balanced (for every +x◦ ply a −x◦ ply is also present) and symmetric (symmetry with respect to the middle plane).
Figure 1.8: Example of multiple plies stacked one above the other with different orientations to form a balanced and symmetric layup
1.3
Damage in composites
CFRP presents particular types of damage related to their multi-phase nature. The most important will be briefly described.
Cracks can start in the matrix both under tensile and compressive loads while fibers can break in tension or buckle (fiber micro-buckling or kinking) if compressed. Increasing the load, cracks can keep growing parallel or perpendicular to fibers in a composite lay-up, causing first of all an overall reduction of stiffness. If a crack is contained in a single ply it can cause the separation between fiber and matrix, a phenomenon called fiber pull-out, while those cracks that turn and grow at the in-terface between two different plies can cause their separation, a phenomenon called delamination (Figure 1.9).
(a) Representation of cracks in plies or-thogonal to the loading direction (90◦)
(b) Real microscopy of a composite spec-imen where cracks have grown causing delamination, [5]
Figure 1.9: Different stages of damage in composites
To better understand the evolution and the dynamic of damage in composites it is useful to observe the case of uniaxial tension loading on a [0◦/90◦]xS laminate
shown in Figure 1.10. In this case the 90◦ plies start accumulating cracks up to a level indicated as CDS, Crack Density Saturation. When this condition is reached the 90◦ plies present so many cracks that the totality of the load is carried by the
0◦ plies. Increasing the load, cracks start coupling and growing, causing debonding at the interface between matrix and fibers and delamination. Eventually the load will reach the ultimate strength of fibers leading to the final fracture of the part.
Figure 1.10: Evolution of damage in a [0◦/90◦]xS composite layup, [7]
Delamination is also a typical consequence of impacts and is one of the most dangerous form of damage to deal with, especially in the presence of compressive loads: it is hard to be detected and causes a thick laminate to behave like multiple thinner panels that can easily buckle, causing catastrophic failure. Delamination can also be caused by the characteristic phenomenon called free edge effect, [6]. This phenomenon is a consequence of the orthotropy of the material, in particular the difference in Poisson’s ratio in the two orthogonal reference directions, and is present when two plies with different fiber’s orientations are stacked one above the
other. When such laminate is subjected to in-plane loading very high (a singularity for the equilibrium equations) out-of-plane tensile stress σ3 originates at the free
edge, followed by inter-laminar shear τ13 and τ23 to maintain balance (Figure 1.11).
These stress components are particularly dangerous because tend to separate the plies, initiating delamination and reducing the overall strength of the part. This phenomenon can be controlled through the choice of materials, ply orientations, stacking sequence, and ply thickness. (Figure 1.12).
y/b
Sigma Z
0 0.2 0.4 0.6 0.8 1
(a) σ3 stress at the interface between two plies with different orientations. The stress singularity is visible at the free edge.
0 0.2 0.4 0.6 0.8 1 y/b
Tau ZY
(b) τ23 stress at the interface between two plies with different orientation. Shear must be zero at the free edge for equilibrium.
Figure 1.12: Free edge effect visual explanation
1.4
Research goals
This research is inspired by a practical problem: during several end-life inspections performed on aircraft wings, delamination damage has been located around bolted joints on composite structural panels, in particular at the connection between wing and fuselage. To safely extend the operative life and increase the safety of those aircraft a research program has been commissioned to evaluate the effects of this type of damage and the conditions in which it may bring to dangerous failure.
In order to do that, single and double-lap bolted joint bearing tests will be performed using modified composite specimens in which a known and consistent delamination has been introduced around the bolt’s hole. In particular the hole of interest is countersunk. The desired delamination is circular, coaxial with the hole, 1 inch in diameter and approximately in the middle plane of the specimen.
The problem is that at present time there is no ASTM standard procedure to create such damage in a specimen. The traditional method of introducing delamina-tion during manufacturing, interposing Teflon inserts between plies has limitadelamina-tions for this particular application. Firstly, using small circular Teflon inserts embedded into the laminate requires great control in positioning to properly cut and drill the specimens. Uncertainty in this phase may alter the results in unpredictable ways.
Most important, the presence of Teflon in the cracked region will change the friction between the two fracture surfaces, greatly influencing the crack growth.
For these reasons several alternative methods to introduce a controlled delam-ination in a bearing test specimen with a countersunk hole have been proposed. The research is focused in finding a method that guarantees damage to be identi-cal in every specimen, even if parameters like material or stacking sequence were changed. The method should also ensure the absence of damage besides the desired delamination, such as cracks in the matrix.
