5 Component Form and Manufacture 5.1 Introduction Because fiber reinforcement is essentially a one-dimensional strengthening process,a major function of the component-forming process is to orientate the fibers in the matrix in the appropriate directions and proportions to obtain the desired two-dimensional or 3-dimensional mechanical properties.The forming process must also produce the shape of the component and develop the required properties of the matrix and the fiber/matrix bond.The forming process must not damage the fibers and must ensure that they are reasonably evenly distributed in a matrix,free from significant voiding or from large areas devoid of fibers. The simplest method that satisfies these requirements is to infiltrate an appropriately aligned fiber bed with a liquid,which is then converted by chemical reaction (in the case of thermosets)or simply by cooling (in the case of thermoplastics)to form a continuous solid matrix with the desired properties. Techniques based on liquid resin are known as liquid molding,with several subcategories according to various modifications of the process. Altematively,sheets of aligned fibers may be pre-coated with matrix precursor and the continuous matrix formed by flowing the coatings together(and curing,if a thermoset matrix)under heat and pressure.In this widely used form,the material is known as pre-preg(pre-impregnated). There are several methods that can be used to arrange the fibers when forming the composite structure.The main method for the manufacture of aircraft components is laminating woven cloth,or aligned fiber sheets,with the fibers orientated in appropriate directions in each layer. There are also several methods based on continuous fiber tow or yarn;these include: (1)filament winding onto a rotating mandrel;(2)braiding onto a rotating mandrel (the process of braiding is covered in detail in Chapter 14);(3)tow placement;and (4)pultrusion. The main differences between the use of thermosets and thermoplastic matrices are the need for extended times to cure (cross-link)the thermosets and the relatively high viscosities of the thermoplastics melts and the consequential requirement for high processing temperatures and pressures.Table 5.1 lists generic aircraft components made using these manufacturing procedures. 113
5 Component Form and Manufacture 5.1 Introduction Because fiber reinforcement is essentially a one-dimensional strengthening process, a major function of the component-forming process is to orientate the fibers in the matrix in the appropriate directions and proportions to obtain the desired two-dimensional or 3-dimensional mechanical properties. The forming process must also produce the shape of the component and develop the required properties of the matrix and the fiber/matrix bond. The forming process must not damage the fibers and must ensure that they are reasonably evenly distributed in a matrix, free from significant voiding or from large areas devoid of fibers. The simplest method that satisfies these requirements is to infiltrate an appropriately aligned fiber bed with a liquid, which is then converted by chemical reaction (in the case of thermosets) or simply by cooling (in the case of thermoplastics) to form a continuous solid matrix with the desired properties. Techniques based on liquid resin are known as liquid molding, with several subcategories according to various modifications of the process. Alternatively, sheets of aligned fibers may be pre-coated with matrix precursor and the continuous matrix formed by flowing the coatings together (and curing, if a thermoset matrix) under heat and pressure. In this widely used form, the material is known as pre-preg (pre-impregnated). There are several methods that can be used to arrange the fibers when forming the composite structure. The main method for the manufacture of aircraft components is laminating woven cloth, or aligned fiber sheets, with the fibers orientated in appropriate directions in each layer. There are also several methods based on continuous fiber tow or yarn; these include: (1) filament winding onto a rotating mandrel; (2) braiding onto a rotating mandrel (the process of braiding is covered in detail in Chapter 14); (3) tow placement; and (4) pultrusion. The main differences between the use of thermosets and thermoplastic matrices are the need for extended times to cure (cross-link) the thermosets and the relatively high viscosities of the thermoplastics melts and the consequential requirement for high processing temperatures and pressures. Table 5.1 lists generic aircraft components made using these manufacturing procedures. 113
114 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES Table 5.1 Typical Aircraft Fiber Composite Forms Made by the Different Techniques,as Listed Type of Structure Typical Application Laminates Sheets,thick monolithic Wing skins Sheets,integrally stiffened Tail skins Sandwich panels Control surfaces,floor sections Shells Fuselage sections Beams Spars/ribs Complex forms Aerofoils Filament Wound Closed shells Pressure vessels Open shells Radomes Rocket motors Tubes Drive shafts Secondary formed tubes Helicopter blades Braided Tubes Drive shafts Complex tubes Curved pipes Truss joints Ducts Closed shells Pressure vessels Secondary formed Fuselage frames Aircraft propellers Helicopter blades Tow Placed See laminates See laminates Complex wraps Grips Shafts Ducts Pultrusion Beams Floor beams Stringers Spars Ribs Longerons Considerable structural and cost efficiency can be obtained by using the composite in the most highly stressed regions,for example,in the upper and lower surfaces of components subject to bending or buckling.This is achieved by using a sandwich construction,as also listed in Table 5.1,with the composite laminate forming the outer skins,which are bonded to a metallic or polymeric
114 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES Table 5.1 Typical Aircraft Fiber Composite Forms Made by the Different Techniques, as Listed Type of Structure Typical Application Laminates Sheets, thick monolithic Sheets, integrally stiffened Sandwich panels Shells Beams Complex forms Filament Wound Closed shells Open shells Tubes Secondary formed tubes Braided Tubes Complex tubes Closed shells Secondary formed Tow Placed See laminates Complex wraps Pultrusion Beams Wing skins Tail skins Control surfaces, floor sections Fuselage sections Spars/ribs Aerofoils Pressure vessels Radomes Rocket motors Drive shafts Helicopter blades Drive shafts Curved pipes Truss joints Ducts Pressure vessels Fuselage frames Aircraft propellers Helicopter blades See laminates Grips Shafts Ducts Floor beams Stringers Spars Ribs Longerons Considerable structural and cost efficiency can be obtained by using the composite in the most highly stressed regions, for example, in the upper and lower surfaces of components subject to bending or buckling. This is achieved by using a sandwich construction, as also listed in Table 5.1, with the composite laminate forming the outer skins, which are bonded to a metallic or polymeric
COMPONENT FORM AND MANUFACTURE 115 composite honeycomb or polymeric foam core.The metallic honeycomb is generally an aluminum alloy such as 5052,often with a coating or anodized layer to resist corrosion.The composite honeycomb would generally be glass- reinforced epoxy or phenolic;however,the most usual honeycomb material is Nomex,which is the trade name for a composite based on random meta-aramid fibers in a phenolic matrix.The foam core used for aerospace applications is generally made of PVC,but this material is not generally used in applications exposed to high temperatures.Polyetherimide (PED)and polymethacrylimide (PMI)polyimide foams are alternative cores for higher-temperature applications. This chapter deals primarily with pre-preg laminating procedures in some detail because this is the prime method for manufacturing aircraft composite components.Methods based on liquid resin are then considered,followed by details of the various processes,resin transfer and infusion,and filament winding and pultrusion.Finally,the particular processes for manufacturing with thermo- plastic resins are covered. 5.2 Outline of General Laminating Procedures Most reinforced-plastic components based on long fibers are manufactured by some form of laminating procedure.In this process,sheets of reinforcement, pre-coated with resin(pre-preg)or with resin freshly applied,are forced against the surface of a mold under the required conditions of pressure,temperature,and time.Chapter 3 provides details of some of the cloth materials available,and details of the pre-pregging process are provided later in this chapter. 5.2.1 Open Die Molding Open die molding involves the use of only one mold surface,over which the layers of fiber are placed or "laid-up."If dry cloth is used,the resin may be applied by brushing or spraying.With care and suitable materials,this method (which is still widely used outside the aircraft industry)can produce good-quality parts.However,handling wet resins can be messy and can raise occupational health and safety (OH&S)concerns.In addition,a particular concern with the use of wet lay-up in aircraft-part production is the lack of repeatability of the process, especially the control of resin content and therefore the weight,thickness,and mechanical properties.Some smaller companies,notably in the German Glider Industry,have adopted wet pre-preg dispensing machines,which saturate reinforcement fabric on demand with a controlled amount of liquid resin, normally epoxy,and hardener.This solution is cheap and flexible,and it does not require cold storage. Various methods are engaged to apply pressure to consolidate the lay-up.In contact molding,which is generally used only for fairly low-stress applications of
COMPONENT FORM AND MANUFACTURE 115 composite honeycomb or polymeric foam core. The metallic honeycomb is generally an aluminum alloy such as 5052, often with a coating or anodized layer to resist corrosion. The composite honeycomb would generally be glassreinforced epoxy or phenolic; however, the most usual honeycomb material is Nomex, which is the trade name for a composite based on random meta-aramid fibers in a phenolic matrix. The foam core used for aerospace applications is generally made of PVC, but this material is not generally used in applications exposed to high temperatures. Polyetherimide (PEI) and polymethacrylimide (PMI) polyimide foams are alternative cores for higher-temperature applications. This chapter deals primarily with pre-preg laminating procedures in some detail because this is the prime method for manufacturing aircraft composite components. Methods based on liquid resin are then considered, followed by details of the various processes, resin transfer and infusion, and filament winding and pultrusion. Finally, the particular processes for manufacturing with thermoplastic resins are covered. 5,2 Outline of General Laminating Procedures Most reinforced-plastic components based on long fibers are manufactured by some form of laminating procedure. 1 In this process, sheets of reinforcement, pre-coated with resin (pre-preg) or with resin freshly applied, are forced against the surface of a mold under the required conditions of pressure, temperature, and time. Chapter 3 provides details of some of the cloth materials available, and details of the pre-pregging process are provided later in this chapter. 5.2.1 Open Die Molding Open die molding involves the use of only one mold surface, over which the layers of fiber are placed or "laid-up." If dry cloth is used, the resin may be applied by brushing or spraying. With care and suitable materials, this method (which is still widely used outside the aircraft industry) can produce good-quality parts. However, handling wet resins can be messy and can raise occupational health and safety (OH&S) concerns. In addition, a particular concern with the use of wet lay-up in aircraft-part production is the lack of repeatability of the process, especially the control of resin content and therefore the weight, thickness, and mechanical properties. Some smaller companies, notably in the German Glider Industry, have adopted wet pre-preg dispensing machines, which saturate reinforcement fabric on demand with a controlled amount of liquid resin, normally epoxy, and hardener. This solution is cheap and flexible, and it does not require cold storage. Various methods are engaged to apply pressure to consolidate the lay-up. In contact molding, which is generally used only for fairly low-stress applications of
116 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES glass/polyester composites,the pressure is developed by hand-rolling over a sheet of plastic film placed over the surface of the lay-up. The bag procedure involves the use of a flexible plastic membrane that is formed over the surface of the lay-up to form a vacuum-tight bag.In vacuum bagging,the bag is evacuated and atmospheric pressure used to consolidate the lay-up against the surface of the mold.The vacuum initially removes most of the air and volatile materials.Vacuum bagging is an inexpensive and versatile procedure;however,it can provide only limited consolidation pressure and may produce voided laminates due to the enlargement of the bubbles(formed by any residual gases or volatile material)trapped in the resin in regions where the bag is unable to apply pressure,for example,because of local bridging.To minimize this problem,autoclave procedures,described later,are used to manufacture most of the high-quality laminates used in the aircraft industry. Alternatively,pressure may be applied to the surface of an open mold by means of a flexible plunger mounted in a press,by gas-bags,or by thermal expansion of an entrapped rubber or metallic insert. Temperature,generally required to cure the resin,can be applied to the open mold in various ways,including exteral methods such as hot-air blowers and ovens or interally by electric elements or steam or oil pipes buried in the mold. Temperatures up to 180C may be required in aerospace-grade epoxy resin systems. 5.2.2 Compression Molding Compression or matched-die molding involves the use of matching male and female dies that close to form a cavity of the shape of the component (Fig.5.1). The dies,generally made of tool steel,can be internally heated,if required,by electric elements or steam,or hot oil pipes.The fiber layers are placed over the lower mold section,and the two halves of the mold are brought together in a press.Lands built into the mold usually control the thickness of the part. Advantages of matched-die molding include excellent dimensional control;high- quality surface finish,produced on both surfaces;high production rates;and good consolidation and high fiber content. However,the cost of the matching dies(with hardened faces)is very high,and the size of the available hydraulic presses used to apply the closing pressure limits the size of parts that can be produced. Wet laminating procedures may be used,in which case the dry fiber is laid in the mold and the resin added.High-quality fiber composite components are generally based on the use of pre-pregs or by the use of a solid,but uncured,resin film that is laid on the mold surface,followed by dry fiber layers or a fiber preform. Alternatively,a liquid resin can be injected into the sealed and evacuated mold cavity,as discussed later
116 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES glass/polyester composites, the pressure is developed by hand-rolling over a sheet of plastic film placed over the surface of the lay-up. The bag procedure involves the use of a flexible plastic membrane that is formed over the surface of the lay-up to form a vacuum-tight bag. In vacuum bagging, the bag is evacuated and atmospheric pressure used to consolidate the lay-up against the surface of the mold. The vacuum initially removes most of the air and volatile materials. Vacuum bagging is an inexpensive and versatile procedure; however, it can provide only limited consolidation pressure and may produce voided laminates due to the enlargement of the bubbles (formed by any residual gases or volatile material) trapped in the resin in regions where the bag is unable to apply pressure, for example, because of local bridging. To minimize this problem, autoclave procedures, described later, are used to manufacture most of the high-quality laminates used in the aircraft industry. Alternatively, pressure may be applied to the surface of an open mold by means of a flexible plunger mounted in a press, by gas-bags, or by thermal expansion of an entrapped rubber or metallic insert. Temperature, generally required to cure the resin, can be applied to the open mold in various ways, including external methods such as hot-air blowers and ovens or internally by electric elements or steam or oil pipes buried in the mold. Temperatures up to 180 °C may be required in aerospace-grade epoxy resin systems. 5.2.2 Compression Molding Compression or matched-die molding involves the use of matching male and female dies that close to form a cavity of the shape of the component (Fig. 5.1). The dies, generally made of tool steel, can be internally heated, if required, by electric elements or steam, or hot oil pipes. The fiber layers are placed over the lower mold section, and the two halves of the mold are brought together in a press. Lands built into the mold usually control the thickness of the part. Advantages of matched-die molding include excellent dimensional control; highquality surface finish, produced on both surfaces; high production rates; and good consolidation and high fiber content. However, the cost of the matching dies (with hardened faces) is very high, and the size of the available hydraulic presses used to apply the closing pressure limits the size of parts that can be produced. Wet laminating procedures may be used, in which case the dry fiber is laid in the mold and the resin added. High-quality fiber composite components are generally based on the use of pre-pregs or by the use of a solid, but uncured, resin film that is laid on the mold surface, followed by dry fiber layers or a fiber preform. Alternatively, a liquid resin can be injected into the sealed and evacuated mold cavity, as discussed later
COMPONENT FORM AND MANUFACTURE 117 Fig.5.1 Matched-die mold and resulting top-hat stiffened component. 5.2.3 Wrapping Wrapping is an alternative procedure to filament winding,described later,for producing tubular components.A pre-preg sheet,either wrap sheet or cloth,is wrapped onto a removable metal mandrel and cured under pressure.Special machines are available to perform the wrapping operations.The pressure during an elevated temperature cure may be applied by the use of shrink film(applied by a tape-winding machine),vacuum bag,or autoclave.Alternatively,a silicon- rubber bladder may be placed over the mandrel before the wrapping of the laminate.Pressure is applied to the laminate through-inflation of the bladder that forces the laminate against an outer mold surface.This technique is often used to make fishing rods,golf clubs,and tennis rackets. 5.3 Laminating Procedures For Aircraft-Grade Composite Components Major aircraft manufacturers and their subcontractors,especially in the United States,use B-staged epoxy pre-preg as their preferred material form.In this material,the reinforcement is pre-impregnated by a supplier with a resin already
COMPONENT FORM AND MANUFACTURE 117 Fig. 5.1 Matched-die mold and resulting top-hat stiffened component. 5.2.3 Wrapping Wrapping is an alternative procedure to filament winding, described later, for producing tubular components. A pre-preg sheet, either wrap sheet or cloth, is wrapped onto a removable metal mandrel and cured under pressure. Special machines are available to perform the wrapping operations. The pressure during an elevated temperature cure may be applied by the use of shrink film (applied by a tape-winding machine), vacuum bag, or autoclave. Alternatively, a siliconrubber bladder may be placed over the mandrel before the wrapping of the laminate. Pressure is applied to the laminate through-inflation of the bladder that forces the laminate against an outer mold surface. This technique is often used to make fishing rods, golf clubs, and tennis rackets. 5.3 Laminating Procedures For Aircraft-Grade Composite Components Major aircraft manufacturers and their subcontractors, especially in the United States, use B-staged epoxy pre-preg as their preferred material form. In this material, the reinforcement is pre-impregnated by a supplier with a resin already
118 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES containing hardener.This has been partially cured(B-staged)such that the resin does not flow at room temperature,but at the same time it remains tacky(sticky to the touch).B-staged epoxy pre-pregs are normally staged(partially cured)to about 15%of full cure for hand lay-up,and up to 25%for automated lay-up. To protect this material and keep it from sticking to itself,a backing or release film is added to at least one side of the pre-preg before it is rolled up for storage or transport. 5.3.1 Pre-Preg Production A pre-preg can be made incorporating a variety of reinforcement fabrics and fiber types.Although it can be produced by the component fabricator,it is normally purchased from a materials-supply company.The following material forms are available as carbon/epoxy pre-pregs. Woven bi-directional cloth pre-preg is most commonly made from plain weave or satin weave fabrics,0.2-0.4 mm thick and up to 1200 mm wide.One common method of pre-impregnation is to infuse the cloth with matrix resin diluted with solvent to lower its viscosity.The pre-preg then passes through a heating tower to remove the solvent and stage the resin.The newer hot-melt method (See Fig.5.2)involves first continuously casting a B-staged resin film on a non-stick backing film of coated paper or polymer.A doctor blade is used to control the thickness of the resin film applied(the same method used to make adhesive film).The reinforcement is then sandwiched between two of these films as it passes through a pair of heated rollers.This process has an advantage over the solvent process in that it produces lower volatile emissions. Unidirectional pre-preg (warp sheet)is made by spreading and collimating many fiber tows(typically around 10 fibers in each tow)into a uniform sheet of Top Paper Reel Comb Impregnation Roll Pull Rolls Doctor Heat Cool Prepreg 88 Toke-up Reel Chill mpregnatio Filming Plate Plate Bottom Paper Roll Fig.5.2 Schematic illustration of hot-melt film pre-pregging process.Adapted from Ref.2
118 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES containing hardener. 2 This has been partially cured (B-staged) such that the resin does not flow at room temperature, but at the same time it remains tacky (sticky to the touch). B-staged epoxy pre-pregs are normally staged (partially cured) to about 15% of full cure for hand lay-up, and up to 25% for automated lay-up. To protect this material and keep it from sticking to itself, a backing or release film is added to at least one side of the pre-preg before it is rolled up for storage or transport. 5.3.1 Pre-Preg Production A pre-preg can be made incorporating a variety of reinforcement fabrics and fiber types. Although it can be produced by the component fabricator, it is normally purchased from a materials-supply company. The following material forms are available as carbon/epoxy pre-pregs. Woven hi-directional cloth pre-preg is most commonly made from plain weave or satin weave fabrics, 0.2-0.4 mm thick and up to 1200 mm wide. One common method of pre-impregnation is to infuse the cloth with matrix resin diluted with solvent to lower its viscosity. The pre-preg then passes through a heating tower to remove the solvent and stage the resin. The newer hot-melt method (See Fig. 5.2) involves first continuously casting a B-staged resin film on a non-stick backing film of coated paper or polymer. A doctor blade is used to control the thickness of the resin film applied (the same method used to make adhesive film). The reinforcement is then sandwiched between two of these films as it passes through a pair of heated rollers. This process has an advantage over the solvent process in that it produces lower volatile emissions. Unidirectional pre-preg (warp sheet) is made by spreading and collimating many fiber tows (typically around 104 fibers in each tow) into a uniform sheet of ~ Top Poper Reel Rber Rolls Doctor -~ Pr~reg Blade -~ Heat ~ Cool Take-up Reel ~f..~ ~- Filming Plate Plate Fig. 5.2 Schematic illustration of hot-melt film pre-pregging process. Adapted from Ref. 2
COMPONENT FORM AND MANUFACTURE 119 parallel fibers typically 0.125-0.25 mm thick and 300 or 600 mm wide.This is immediately pre-impregnated.Unidirectional pre-preg is the cheapest to make, and it provides laminates with the best mechanical properties.However,it may be difficult to lay into double-curved shapes.Other types of reinforcement architec- ture,such as multi-axial warp knit(also known as non-crimp,knitted,or stitched) fabrics can also be pre-impregnated,but the process becomes increasingly difficult as the fabric becomes thicker. The pre-preg with its non-stick backing films is then inspected for resin content,which is typically between 34%and 42%by weight for carbon pre- pregs,wound onto a roll,and sealed to prevent the absorption of water vapor. Some pre-pregs have up to 15%more resin than is required to form a laminate with the desired fiber/volume fraction.With these pre-pregs,the resin is required to bleed out of the laminate during curing.Low-bleed or non-bleed pre-pregs with a more viscous resin are now more popular. The standard pre-preg thickness for unidirectional materials is of the order of 0.125 mm.More recently,to cut costs,much larger tows are being used,resulting in much thicker pre-pregs;however,because it is more difficult to maintain fiber alignment in thick tows,there is some reduction in mechanical properties of the finished composite. 5.3.2 Pre-Preg Transport and Storage The major disadvantage of pre-preg(apart from the extra cost of creating it from the fiber and resin)is that once the hardener has been added,the resin begins to react.Therefore the material normally only has a limited"shelf'(storage)life and"shop"(usage)life before the resin has reacted sufficiently for the pre-preg to become stiff and intractable for lay-up,or for the quality of the resulting composite to suffer.Most pre-pregs need to be stored in a freezer,typically at around -20C,which halts or at least greatly slows down the curing reaction in the resin.Pre-pregs generally used in aerospace are cured at elevated temperatures,typically 120C or 180C for epoxy resins.Because the resin is designed to react at elevated temperature,the supplier can normally guarantee a shelf(freezer)life of 6 months to a year,and a shop life ("out"life at room temperature)of at least 2 weeks. If the distance from the supplier to the user is long,the pre-preg will need to be shipped in refrigerated shipping containers;or for smaller lots,in insulated packages containing dry ice (frozen carbon dioxide). 5.3.3 Cutting and Kitting When pre-preg is required for use,it is thawed to room temperature before being removed from its bag to avoid picking up condensation.The pre-preg is then moved into the cutting room,which like the lay-up room is maintained as a "clean room,"free of dust and with controlled temperature (around 20C)and
COMPONENT FORM AND MANUFACTURE 119 parallel fibers typically 0.125-0.25 mm thick and 300 or 600 mm wide. This is immediately pre-impregnated. Unidirectional pre-preg is the cheapest to make, and it provides laminates with the best mechanical properties. However, it may be difficult to lay into double-curved shapes. Other types of reinforcement architecture, such as multi-axial warp knit (also known as non-crimp, knitted, or stitched) fabrics can also be pre-impregnated, but the process becomes increasingly difficult as the fabric becomes thicker. The pre-preg with its non-stick backing films is then inspected for resin content, which is typically between 34% and 42% by weight for carbon prepregs, wound onto a roll, and sealed to prevent the absorption of water vapor. Some pre-pregs have up to 15% more resin than is required to form a laminate with the desired fiber/volume fraction. With these pre-pregs, the resin is required to bleed out of the laminate during curing. Low-bleed or non-bleed pre-pregs with a more viscous resin are now more popular. The standard pre-preg thickness for unidirectional materials is of the order of 0.125 mm. More recently, to cut costs, much larger tows are being used, resulting in much thicker pre-pregs; however, because it is more difficult to maintain fiber alignment in thick tows, there is some reduction in mechanical properties of the finished composite. 5.3.2 Pre-Preg Transport and Storage The major disadvantage of pre-preg (apart from the extra cost of creating it from the fiber and resin) is that once the hardener has been added, the resin begins to react. Therefore the material normally only has a limited "shelf' (storage) life and "shop" (usage) life before the resin has reacted sufficiently for the pre-preg to become stiff and intractable for lay-up, or for the quality of the resulting composite to suffer. Most pre-pregs need to be stored in a freezer, typically at around -20°C, which halts or at least greatly slows down the curing reaction in the resin. Pre-pregs generally used in aerospace are cured at elevated temperatures, typically 120°C or 180°C for epoxy resins. Because the resin is designed to react at elevated temperature, the supplier can normally guarantee a shelf (freezer) life of 6 months to a year, and a shop life ("out" life at room temperature) of at least 2 weeks. If the distance from the supplier to the user is long, the pre-preg will need to be shipped in refrigerated shipping containers; or for smaller lots, in insulated packages containing dry ice (frozen carbon dioxide). 5.3.3 Cutting and Kitting When pre-preg is required for use, it is thawed to room temperature before being removed from its bag to avoid picking up condensation. The pre-preg is then moved into the cutting room, which like the lay-up room is maintained as a "clean room," free of dust and with controlled temperature (around 20°C) and
120 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES humidity (e.g.,between 50-70%RH).The pre-preg is then unrolled onto the cutting table,with its backing paper still in place.Plies of the required size,shape, and fiber orientation are then cut from the roll;as an example,Figure 5.3 shows a ply stack for a wing rib.This can be done by hand-using a template,or with a die in a roller press;in all but the smallest operations,this is usually done by a numerically controlled flat-bed cutter similar to those used in the textile industry. Cutting is usually achieved using an oscillating blade,but sharp"draw knife" blades as well as lasers or water jets are also used.Some cutters can cut multiple layers of fabric.Some flat-bed cutters can also label the plies automatically.The various ply shapes are then labelled,if necessary,and assembled as part of a kit containing all the plies for a component,which may be delivered directly to the lay-up room or sealed and stored in a plastic bag in the freezer for later use. Abrasive water jet cutting uses a high-pressure water stream,perhaps up to 400 MPa,which is forced through a small sapphire orifice to produce a supersonic jet travelling at speeds up to 900 m s,carrying abrasive particles to form a powerful cutting jet.Most materials can be machined with the water jet's ability to revolve with the robotic end effector.The critical process parameters are speed;stand-off distance;impact angle;water-jet pressure;water flow rate; orifice diameter;abrasive particle shape,hardness and size;and nozzle mixing tube geometry and material.Generally,the impact angle can be optimized to produce the maximum removal rate.The work-piece material should be softer than the abrasive compound.Oscillation of the cutting head can also influence the quality of the cut. Laser cutting can be considered a thermal process as a portion of the beam energy is absorbed by the surface material,and this energy raises the temperature ● Stack..... Center Fig.5.3 Schematic diagram of a typical ply stack for a wing rib
120 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES humidity (e.g., between 50-70% RH). The pre-preg is then unrolled onto the cutting table, with its backing paper still in place. Plies of the required size, shape, and fiber orientation are then cut from the roll; as an example, Figure 5.3 shows a ply stack for a wing rib. This can be done by hand-using a template, or with a die in a roller press; in all but the smallest operations, this is usually done by a numerically controlled flat-bed cutter similar to those used in the textile industry. Cutting is usually achieved using an oscillating blade, but sharp "draw knife" blades as well as lasers or water jets are also used. Some cutters can cut multiple layers of fabric. Some flat-bed cutters can also label the plies automatically. The various ply shapes are then labelled, if necessary, and assembled as part of a kit containing all the plies for a component, which may be delivered directly to the lay-up room or sealed and stored in a plastic bag in the freezer for later use. Abrasive water jet cutting uses a high-pressure water stream, perhaps up to 400 MPa, which is forced through a small sapphire orifice to produce a supersonic jet travelling at speeds up to 900 m s-1, carrying abrasive particles to form a powerful cutting jet. Most materials can be machined with the water jet's ability to revolve with the robotic end effector. The critical process parameters are speed; stand-off distance; impact angle; water-jet pressure; water flow rate; orifice diameter; abrasive particle shape, hardness and size; and nozzle mixing tube geometry and material. Generally, the impact angle can be optimized to produce the maximum removal rate. The work-piece material should be softer than the abrasive compound. Oscillation of the cutting head can also influence the quality of the cut. Laser cutting can be considered a thermal process as a portion of the beam energy is absorbed by the surface material, and this energy raises the temperature oooo o°° Sta~l~" °" Center Fig. 5.3 Schematic diagram of a typical ply stack for a wing rib
COMPONENT FORM AND MANUFACTURE 121 of the material.A sufficient amount of such energy will cause local decomposition of the material.Some compromise is required when focussing the laser beam as minimum spot size(a result of using short focal length lenses)is achieved at the expense of depth of field.The creation of thermal energy during cutting can produce problems in the course of dealing with standard epoxy pre- preg systems producing local cure and toxic vapors. All methods of cutting for complex geometry flat shapes must be capable of operation with either a standard robot or gantry-type equipment. 5.3.4 Lay-Up Most aerospace components are still laid-up by skilled labor,although considerable efforts are being made to automate or mechanize the process,as described in the subsequent sections.Hand lay-up is very versatile because human hands make excellent grippers,eyes marvellous sensors,and the brain a powerful process control and quality control unit!Any residual dust or resin from previous use is cleaned off before a thin layer of release agent is applied to the surface,where necessary.The mold will then be moved into the lay-up clean room. The pre-preg plies are then applied to the mold in the correct position, orientation,and sequence according to a set of instructions sometimes called a ply book;these instructions may be viewed on a computer screen.The ply is located on the surface by reference to markings on the mold or with the aid of a rigid or flexible template.Many companies now have lay-up stations where an overhead projector rapidly scans a low-power laser beam to"draw"the outline of each ply on the mold surface.These machines can also project instructions for ply lay-up onto the mold. Typically,the lower backing paper is removed by the operator before lay-up, and the upper one after positioning and consolidating using rollers or other simple tools.For larger plies,two or more operators may be required to handle and position the tacky pre-preg.Where the mold surface is doubly-curved,the pre- preg needs to be further distorted,enabling it to fit the surface. Different types of material may be combined in the same lay-up as long as the materials are compatible.For instance,in sandwich structures,aluminium or Nomex honeycomb and adhesive films will normally be combined with carbon- epoxy pre-preg to form the structure.Different fibers such as glass and carbon may be combined to form hybrid lay-ups,and different reinforcement arrangements such as unidirectional tape and woven fabric may be combined. 5.3.5 Automated Forming of Pre-Preg Stacks To reduce lay-up times and consequently labor costs,automated or semi- automated methods have recently been introduced to aircraft component production lines
COMPONENT FORM AND MANUFACTURE 121 of the material. A sufficient amount of such energy will cause local decomposition of the material. Some compromise is required when focussing the laser beam as minimum spot size (a result of using short focal length lenses) is achieved at the expense of depth of field. The creation of thermal energy during cutting can produce problems in the course of dealing with standard epoxy prepreg systems producing local cure and toxic vapors. All methods of cutting for complex geometry flat shapes must be capable of operation with either a standard robot or gantry-type equipment. 5.3.4 Lay-Up Most aerospace components are still laid-up by skilled labor, although considerable efforts are being made to automate or mechanize the process, as described in the subsequent sections. Hand lay-up is very versatile because human hands make excellent grippers, eyes marvellous sensors, and the brain a powerful process control and quality control unit! Any residual dust or resin from previous use is cleaned off before a thin layer of release agent is applied to the surface, where necessary. The mold will then be moved into the lay-up clean room. The pre-preg plies are then applied to the mold in the correct position, orientation, and sequence according to a set of instructions sometimes called a ply book; these instructions may be viewed on a computer screen. The ply is located on the surface by reference to markings on the mold or with the aid of a rigid or flexible template. Many companies now have lay-up stations where an overhead projector rapidly scans a low-power laser beam to "draw" the outline of each ply on the mold surface. These machines can also project instructions for ply lay-up onto the mold. Typically, the lower backing paper is removed by the operator before lay-up, and the upper one after positioning and consolidating using rollers or other simple tools. For larger plies, two or more operators may be required to handle and position the tacky pre-preg. Where the mold surface is doubly-curved, the prepreg needs to be further distorted, enabling it to fit the surface. Different types of material may be combined in the same lay-up as long as the materials are compatible. For instance, in sandwich structures, aluminium or Nomex honeycomb and adhesive films will normally be combined with carbonepoxy pre-preg to form the structure. Different fibers such as glass and carbon may be combined to form hybrid lay-ups, and different reinforcement arrangements such as unidirectional tape and woven fabric may be combined. 5.3.5 Automated Forming of Pre-Preg Stacks To reduce lay-up times and consequently labor costs, automated or semiautomated methods have recently been introduced to aircraft component production lines
122 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES Instead of shaping and consolidating(laying up)each ply separately by hand,a flat stack can be assembled by manual or mechanical means.This flat stack can then be formed into the required shape using various methods;pressing,stamping, or diaphragm-forming.One version of the diaphragm-forming process is illustrated in Figure 5.4.A flat pre-preg stack is laid up and placed over a male- forming die.A diaphragm is fitted and sealed to the forming box.A vacuum is then applied to the box cavity.Because they are not extensible in the fiber direction,the plies must deform by shear to conform to the shape of the tool.It may be necessary to heat the flat pre-preg stack to a temperature above room temperature to assist forming.An infrared heating source is often used for this purpose. This process is most attractive for deep draws,and consequently the shear deformation required can be considerable.There are three main modes of deformation:intraply shear (a trellising action in which the fiber tows pivot at the crossover points),slippage between plies,and ply out-of-plane bending.The main problem is to avoid wrinkling of the plies caused by the development of compressive residual stresses.Computer simulation to assist in predicting the optimum conditions for forming is a recent development discussed later in this chapter. 5.3.6 Automated Lay-Up Lay-up of large components such as wing skins requires automation because. owing to the time required for hand lay-up,materials may be close to their out-life when the task is nearing completion. Fig.5.4 Schematic diagram of the diaphragm-forming process;below,carbon fiber-epoxy rib made using this process
122 COMPOSITE MATERIALS FOR AIRCRAFT STRUCTURES Instead of shaping and consolidating (laying up) each ply separately by hand, a flat stack can be assembled by manual or mechanical means. This flat stack can then be formed into the required shape using various methods; pressing, stamping, or diaphragm-forming. One version of the diaphragm-forming process is illustrated in Figure 5.4. A flat pre-preg stack is laid up and placed over a maleforming die. A diaphragm is fitted and sealed to the forming box. A vacuum is then applied to the box cavity. Because they are not extensible in the fiber direction, the plies must deform by shear to conform to the shape of the tool. It may be necessary to heat the fiat pre-preg stack to a temperature above room temperature to assist forming. An infrared heating source is often used for this purpose. This process is most attractive for deep draws, and consequently the shear deformation required can be considerable. There are three main modes of deformation: intraply shear (a trellising action in which the fiber tows pivot at the crossover points), slippage between plies, and ply out-of-plane bending. The main problem is to avoid wrinkling of the plies caused by the development of compressive residual stresses. Computer simulation to assist in predicting the optimum conditions for forming is a recent development discussed later in this chapter. 5.3.6 Automated Lay-Up Lay-up of large components such as wing skins requires automation because, owing to the time required for hand lay-up, materials may be close to their out-life when the task is nearing completion. I Fig. 5.4 Schematic diagram of the diaphragm-forming process; below, carbon fiber-epoxy rib made using this process