《纺织复合材料》课程参考文献（Composite Materials Handbook，Volume 3）Chapter 12 Lessons Learned

MIL-HDBK-17-3F Volume 3.Chapter 12-Lessons Learned Other environmental effects worth noting include the effect of long term exposure to radiation.Ultra- violet rays from the sun can degrade epoxy resins.This is easily protected by a surface finish such as a coat of paint.Another factor is erosion or pitting caused by high speed impact with rain or dust particles. This is likely to occur on unprotected leading edges.There are surface finishes such as rain erosion coats and paints for preventing surface wear.Lightning strike is also a concern to composites.A direct strike can cause considerable damage to a laminate.Lightning strike protection in the form of conductive surfaces is applied in susceptible areas.In cases where substructure is also composite,the inside end of attachment bolts may need to be connected with each other and to ground by a conducting wire. 12.2.6 Joints 12.2.6.1 Mechanically-fastened joints Successful joint design relies on knowledge of potential failure modes.Failure modes depend on joint geometry and laminate lay-up for one given material.The type of fastener used can also influence the occurrence of a particular failure mode.Different materials will give different failure modes. Net-section tensile/compressive failures occur when the bolt diameter is a sufficiently large fraction of the strip width.For most successful designs,this fraction (D/W)is about one-quarter or more for near-isotropic lay-ups in carbon/epoxy systems that have a D/E of one-third or less. Shear-out and shear-out delamination failures occur because the bolt is too close to the edge of the laminate.Such a failure can be triggered when there is only a partial net-section tension or bearing fail- ure.D/t ratios should be 0.75 to 1.25. In some instances the bolt head may be pulled through the laminate after the bolt is bent and de- formed.This mode is frequently seen with countersunk fasteners and is highly dependent on the particu- lar fastener used. Bearing strength is a function of joint geometry,fastener and member stiffnesses.For a 0/45/90 family of laminates with 20-40%of 0 plies and 40-60%of t45plies,plus a minimum(10%)of 90 plies, the bearing strength is relatively constant.Fastener characteristics such as clamp-up force and head configuration have a significant effect.However,for a specific laminate family,a specific fastener,and equal thickness laminate joining members,the parameter with the greatest influence is D/t. Composite joints require smaller D/W and D/E ratios than do metals to get bearing failures. Composite joint strength characteristics differ from metals because the strength is influenced by the bypass load going around the joint.This occurs when two or more fasteners are arranged in a line to transfer the load through a joint.Since not all of the load is reacted by one fastener,some of the load by-passes it.The by-pass effects become prominent once the ratio of by-pass to fastener bearing load exceeds 20%. Titanium fasteners are the most common means of mechanical attachment in composites.This is because titanium is non-corrosive in the galvanic atmosphere created by the dissimilar materials.Tita- nium is closer to carbon on the cathodic scale. 12.2.6.2 Problems associated with adhesive bonding to peel-ply composite surfaces There are two schools of thought in regard to the adhesive bonding of fibrous composite laminates. One demands light but thorough mechanical abrasion,such as by low-pressure grit blasting,because the only such bonds never to fail prematurely were made to abraded surfaces on completely dry laminates. The other permits bonding directly to surfaces created by stripping off peel plies,with or without a drying requirement,using the justification that there is "adequate"initial strength,even though some of these joints have failed prematurely in service.It is also significant that no ultrasonic inspection technique has 12-6

MIL-HDBK-17-3F Volume 3, Chapter 12 - Lessons Learned 12-6 Other environmental effects worth noting include the effect of long term exposure to radiation. Ultra￾violet rays from the sun can degrade epoxy resins. This is easily protected by a surface finish such as a coat of paint. Another factor is erosion or pitting caused by high speed impact with rain or dust particles. This is likely to occur on unprotected leading edges. There are surface finishes such as rain erosion coats and paints for preventing surface wear. Lightning strike is also a concern to composites. A direct strike can cause considerable damage to a laminate. Lightning strike protection in the form of conductive surfaces is applied in susceptible areas. In cases where substructure is also composite, the inside end of attachment bolts may need to be connected with each other and to ground by a conducting wire. 12.2.6 Joints 12.2.6.1 Mechanically-fastened joints Successful joint design relies on knowledge of potential failure modes. Failure modes depend on joint geometry and laminate lay-up for one given material. The type of fastener used can also influence the occurrence of a particular failure mode. Different materials will give different failure modes. Net-section tensile/compressive failures occur when the bolt diameter is a sufficiently large fraction of the strip width. For most successful designs, this fraction (D/W) is about one-quarter or more for near-isotropic lay-ups in carbon/epoxy systems that have a D/E of one-third or less. Shear-out and shear-out delamination failures occur because the bolt is too close to the edge of the laminate. Such a failure can be triggered when there is only a partial net-section tension or bearing fail￾ure. D/t ratios should be 0.75 to 1.25. In some instances the bolt head may be pulled through the laminate after the bolt is bent and de￾formed. This mode is frequently seen with countersunk fasteners and is highly dependent on the particu￾lar fastener used. Bearing strength is a function of joint geometry, fastener and member stiffnesses. For a 0/±45/90 family of laminates with 20-40% of 0° plies and 40-60% of ±45° plies, plus a minimum (10%) of 90° plies, the bearing strength is relatively constant. Fastener characteristics such as clamp-up force and head configuration have a significant effect. However, for a specific laminate family, a specific fastener, and equal thickness laminate joining members, the parameter with the greatest influence is D/t. Composite joints require smaller D/W and D/E ratios than do metals to get bearing failures. Composite joint strength characteristics differ from metals because the strength is influenced by the bypass load going around the joint. This occurs when two or more fasteners are arranged in a line to transfer the load through a joint. Since not all of the load is reacted by one fastener, some of the load by-passes it. The by-pass effects become prominent once the ratio of by-pass to fastener bearing load exceeds 20%. Titanium fasteners are the most common means of mechanical attachment in composites. This is because titanium is non-corrosive in the galvanic atmosphere created by the dissimilar materials. Tita￾nium is closer to carbon on the cathodic scale. 12.2.6.2 Problems associated with adhesive bonding to peel-ply composite surfaces There are two schools of thought in regard to the adhesive bonding of fibrous composite laminates. One demands light but thorough mechanical abrasion, such as by low-pressure grit blasting, because the only such bonds never to fail prematurely were made to abraded surfaces on completely dry laminates. The other permits bonding directly to surfaces created by stripping off peel plies, with or without a drying requirement, using the justification that there is “adequate” initial strength, even though some of these joints have failed prematurely in service. It is also significant that no ultrasonic inspection technique has

MIL-HDBK-17-3F Volume 3,Chapter 12-Lessons Learned been able to distinguish between bonded joints which will fail in service and those which will not.In addi- tion,most traveler specimens do not represent the same cure conditions as experienced by adjacent large parts and,therefore,mechanical testing also often fails to identify defective bonding.One must de- pend on process control of techniques which can be relied upon 100 percent of the time and on thorough validation of the processes before committing them to production. Consider a surface,created by stripping off a peel ply,which is then bonded as part of an adhesive joint.The resulting adhesive bond "sticks"well enough to pass all inspections;however,may fail prema- turely at the interface between the laminate and the adhesive.All premature bond failures,other than those caused by incomplete cure,occur at the interface between the adhesive and the resin in the lami- nate.Structurally sound bonds either fail outside the joint area,cohesively within the layer of adhesive,or interlaminarly in the resin matrix between the surface fibers and the adhesive layer.These premature failures can occur either when uncured adhesive is bonded to precured laminates,or when uncured pre- preg is cured against cured adhesive films used to stabilize honeycomb cores and the like. There are several ways in which peel plies can create surfaces on which reliable durable bonds are not possible. The peel ply can be coated with a release agent,which transfers to the cured laminate when the peel ply is stripped off. The surface of the peel ply's fibers must be sufficiently inert that the ply can be removed without damaging the laminate.The grooves left in the laminate (or glue layer)by stripping off the peel ply may retain enough inert surface that the resin which is subsequently cured onto it may simply fail to adhere.Adhesion requires more than cleanliness:surface tension is also critical.In the absence of cohesion at the interface,a bonded joint relies only on mechanical interlocking,which is far weaker in peel than it is in shear. The peel-ply surface in the laminate consists of innumerable short grooves separated by sharp edges where the resin between the filaments in the peel ply fractured as the peel ply was stripped off.Moisture on (or in)the adhesive or the laminate can be trapped in these grooves.If this moisture cannot escape during the curing of the adhesive (or of a co-cured face sheet),the trapped moisture will result in a slick bond when examined microscopically after failure. It should be noted that the latter two mechanisms function without any contamination. One aircraft company's process specification has,for decades,required that any peel-ply surface to be bonded must first be thoroughly abraded to remove all traces of the texture of the peel ply.In the ab- sence of the ridges between the grooves,it is presumed that moisture could escape,as it turned to steam during cure,unless the part was too large and too poorly ventilated.Using these requirements,this air- craft company has had no disbonds in those secondarily bonded composite structures which were grit blasted before bonding.The same cannot be claimed for bonds made to unabraded (or only scuff- sanded)peel-ply surfaces.In two instances,on different aircraft types,disbonds were traced to transfer of release agents from silicone-coated peel plies,the use of which is now banned throughout all docu- ments,not only the approved materials lists. On another aircraft type,interfacial failures on peel-ply surfaces appear to be the result of prebond moisture,the exact origin of which has yet to be established.An accident with one test panel during pro- cess qualification by a supplier revealed the consequences of condensate on adhesive film (the roll of adhesive had not been properly sealed when returned to the freezer after the previous use).There was absolutely no adhesion between the resin and the adhesive,even though the lap-shear numbers seemed to be acceptable.Microscopic examination of the surfaces clearly showed perfect imprints of the peel ply texture on both surfaces,with the surface in all grooves as smooth as glass and all of the resin on one surface and all of the adhesive on the other.However,with thicker-than-normal(0.123 inch(3.2 mm)) adherends of the same unidirectional carbon/epoxy,bonds made with the same nonreleased peel ply and the same kind of adhesive achieved cohesive failure of the bond at the same strength level attained by 12-7

MIL-HDBK-17-3F Volume 3, Chapter 12 - Lessons Learned 12-7 been able to distinguish between bonded joints which will fail in service and those which will not. In addi￾tion, most traveler specimens do not represent the same cure conditions as experienced by adjacent large parts and, therefore, mechanical testing also often fails to identify defective bonding. One must de￾pend on process control of techniques which can be relied upon 100 percent of the time and on thorough validation of the processes before committing them to production. Consider a surface, created by stripping off a peel ply, which is then bonded as part of an adhesive joint. The resulting adhesive bond “sticks” well enough to pass all inspections; however, may fail prema￾turely at the interface between the laminate and the adhesive. All premature bond failures, other than those caused by incomplete cure, occur at the interface between the adhesive and the resin in the lami￾nate. Structurally sound bonds either fail outside the joint area, cohesively within the layer of adhesive, or interlaminarly in the resin matrix between the surface fibers and the adhesive layer. These premature failures can occur either when uncured adhesive is bonded to precured laminates, or when uncured pre￾preg is cured against cured adhesive films used to stabilize honeycomb cores and the like. There are several ways in which peel plies can create surfaces on which reliable durable bonds are not possible. • The peel ply can be coated with a release agent, which transfers to the cured laminate when the peel ply is stripped off. • The surface of the peel ply’s fibers must be sufficiently inert that the ply can be removed without damaging the laminate. The grooves left in the laminate (or glue layer) by stripping off the peel ply may retain enough inert surface that the resin which is subsequently cured onto it may simply fail to adhere. Adhesion requires more than cleanliness; surface tension is also critical. In the absence of cohesion at the interface, a bonded joint relies only on mechanical interlocking, which is far weaker in peel than it is in shear. • The peel-ply surface in the laminate consists of innumerable short grooves separated by sharp edges where the resin between the filaments in the peel ply fractured as the peel ply was stripped off. Moisture on (or in) the adhesive or the laminate can be trapped in these grooves. If this moisture cannot escape during the curing of the adhesive (or of a co-cured face sheet), the trapped moisture will result in a slick bond when examined microscopically after failure. It should be noted that the latter two mechanisms function without any contamination. One aircraft company’s process specification has, for decades, required that any peel-ply surface to be bonded must first be thoroughly abraded to remove all traces of the texture of the peel ply. In the ab￾sence of the ridges between the grooves, it is presumed that moisture could escape, as it turned to steam during cure, unless the part was too large and too poorly ventilated. Using these requirements, this air￾craft company has had no disbonds in those secondarily bonded composite structures which were grit blasted before bonding. The same cannot be claimed for bonds made to unabraded (or only scuff￾sanded) peel-ply surfaces. In two instances, on different aircraft types, disbonds were traced to transfer of release agents from silicone-coated peel plies, the use of which is now banned throughout all docu￾ments, not only the approved materials lists. On another aircraft type, interfacial failures on peel-ply surfaces appear to be the result of prebond moisture, the exact origin of which has yet to be established. An accident with one test panel during pro￾cess qualification by a supplier revealed the consequences of condensate on adhesive film (the roll of adhesive had not been properly sealed when returned to the freezer after the previous use). There was absolutely no adhesion between the resin and the adhesive, even though the lap-shear numbers seemed to be acceptable. Microscopic examination of the surfaces clearly showed perfect imprints of the peel ply texture on both surfaces, with the surface in all grooves as smooth as glass and all of the resin on one surface and all of the adhesive on the other. However, with thicker-than-normal (0.123 inch (3.2 mm)) adherends of the same unidirectional carbon/epoxy, bonds made with the same nonreleased peel ply and the same kind of adhesive achieved cohesive failure of the bond at the same strength level attained by

MIL-HDBK-17-3F Volume 3.Chapter 12-Lessons Learned metal-to-metal bonding (6,000 psi(40 MPa)or so).With normal thickness composite adherends,only half this strength was reached,because the resin between the surface fibers and adhesive layer then failed in peel,leaving resin clearly covering both surfaces.This problem can be minimized by maintaining very tight time limits between making parts and bonding them together,with a requirement to thoroughly dry everything before bonding if the time constraints are exceeded.Careful scheduling can avoid this added drying step.The same high-strength cohesive bond failures had previously been achieved by an- other supplier of composite structures using grit-blasted surfaces and 0.080 inch(2.0 mm)thick unidirec- tional laminates. In considering adhesive bond strength,it is vital to note that the specimen testing validates the proc- ess,NOT the part.There is no requirement for the specimen to look like the actual part.Indeed.in a properly designed bonded joint,the bond will not fail first.Consequently,the use of specimens which are "similar"to the part and which are evaluated in terms of the "adequacy"of the load carried in relation to the stresses in the part,is not sufficient to ensure the integrity of the bonded composite structure.This issue is complicated because,only with unidirectional tape laminates is it possible to develop sufficient load to fail a high-strength adhesive bond cohesively.Therefore,only such specimens can provide any assurance that the part they are intended to substantiate has been bonded properly.However,in real parts made from woven-fabric laminates,failures within bundles of fibers at 90 to the applied load will trigger interlaminate failures before such bond strengths can be attained. In all cases,the one condition which can be detected visually on test specimens and failed parts alike which is a guaranteed indicator of a defective bond is an interfacial failure with all of the resin on one side and all of the adhesive on the other,with a clear imprint of the peel ply texture on both surfaces. 12.2.7 Design The design of composite structure is complicated by the fact that every ply must be defined.Draw- ings or design packages must describe the ply orientation,its position within the stack,and its boundaries. This is straightforward for a simple,constant thickness laminate.For complex parts with tapered thick- nesses and ply build-ups around joints and cutouts,this can become extremely complex.The need to maintain relative balance and symmetry throughout the structure increases the difficulty. Composites can not be designed without concurrence.Design details depend on tooling and proc- essing as does assembly and inspection.Parts and processes are so interdependent it could be disas- trous to attempt sequential design and manufacturing phasing. Another factor approached differently in composite design is the accommodation of thickness toler- ances at interfaces.If a composite part must fit into a space between two other parts or between a sub- structure and an outer mold line,the thickness requires special tolerances.The composite part thickness is controlled by the number of plies and the per-ply-thickness.Each ply has a range of possible thick- nesses.When these are layed up to form the laminate they may not match the space available for as- sembly within other constraints.This discrepancy can be handled by using shims or by adding "sacrifi- cial"plies to the laminate(for subsequent machining to a closer tolerance than is possible with nominal per-ply-thickness variations).The use of shims has design implications regarding load eccentricities. Another approach is to use closed die molding at the fit-up edges to mold to exact thickness needed. The anisotropy of special laminates,while more complicated,enables a designer to tailor a structure for desired deflection characteristics.This has been applied to some extent for aeroelastic tailoring of wing skins. Composites are most efficient when used in large,relatively uninterrupted structures.The cost is also related to the number of detail parts and the number of fasteners required.These two factors drive de- signs towards integration of features into large cocured structures.The nature of composites enables this possibility.Well designed,high quality tooling will reduce manufacturing and inspection cost and rejection rate and result in high quality parts. 12-8

MIL-HDBK-17-3F Volume 3, Chapter 12 - Lessons Learned 12-8 metal-to-metal bonding (6,000 psi (40 MPa) or so). With normal thickness composite adherends, only half this strength was reached, because the resin between the surface fibers and adhesive layer then failed in peel, leaving resin clearly covering both surfaces. This problem can be minimized by maintaining very tight time limits between making parts and bonding them together, with a requirement to thoroughly dry everything before bonding if the time constraints are exceeded. Careful scheduling can avoid this added drying step. The same high-strength cohesive bond failures had previously been achieved by an￾other supplier of composite structures using grit-blasted surfaces and 0.080 inch (2.0 mm) thick unidirec￾tional laminates. In considering adhesive bond strength, it is vital to note that the specimen testing validates the proc￾ess, NOT the part. There is no requirement for the specimen to look like the actual part. Indeed, in a properly designed bonded joint, the bond will not fail first. Consequently, the use of specimens which are “similar” to the part and which are evaluated in terms of the “adequacy” of the load carried in relation to the stresses in the part, is not sufficient to ensure the integrity of the bonded composite structure. This issue is complicated because, only with unidirectional tape laminates is it possible to develop sufficient load to fail a high-strength adhesive bond cohesively. Therefore, only such specimens can provide any assurance that the part they are intended to substantiate has been bonded properly. However, in real parts made from woven-fabric laminates, failures within bundles of fibers at 90° to the applied load will trigger interlaminate failures before such bond strengths can be attained. In all cases, the one condition which can be detected visually on test specimens and failed parts alike which is a guaranteed indicator of a defective bond is an interfacial failure with all of the resin on one side and all of the adhesive on the other, with a clear imprint of the peel ply texture on both surfaces. 12.2.7 Design The design of composite structure is complicated by the fact that every ply must be defined. Draw￾ings or design packages must describe the ply orientation, its position within the stack, and its boundaries. This is straightforward for a simple, constant thickness laminate. For complex parts with tapered thick￾nesses and ply build-ups around joints and cutouts, this can become extremely complex. The need to maintain relative balance and symmetry throughout the structure increases the difficulty. Composites can not be designed without concurrence. Design details depend on tooling and proc￾essing as does assembly and inspection. Parts and processes are so interdependent it could be disas￾trous to attempt sequential design and manufacturing phasing. Another factor approached differently in composite design is the accommodation of thickness toler￾ances at interfaces. If a composite part must fit into a space between two other parts or between a sub￾structure and an outer mold line, the thickness requires special tolerances. The composite part thickness is controlled by the number of plies and the per-ply-thickness. Each ply has a range of possible thick￾nesses. When these are layed up to form the laminate they may not match the space available for as￾sembly within other constraints. This discrepancy can be handled by using shims or by adding "sacrifi￾cial" plies to the laminate (for subsequent machining to a closer tolerance than is possible with nominal per-ply-thickness variations). The use of shims has design implications regarding load eccentricities. Another approach is to use closed die molding at the fit-up edges to mold to exact thickness needed. The anisotropy of special laminates, while more complicated, enables a designer to tailor a structure for desired deflection characteristics. This has been applied to some extent for aeroelastic tailoring of wing skins. Composites are most efficient when used in large, relatively uninterrupted structures. The cost is also related to the number of detail parts and the number of fasteners required. These two factors drive de￾signs towards integration of features into large cocured structures. The nature of composites enables this possibility. Well designed, high quality tooling will reduce manufacturing and inspection cost and rejection rate and result in high quality parts

MIL-HDBK-17-3F Volume 3.Chapter 12-Lessons Learned 12.2.8 Handling and storage Epoxy resins are the most common form of matrix material used in composites.Epoxies are perish- able.They must be stored below freezing temperature and even then have limited shelf life.Once the material is brought out of storage there is limited time it can be used to make parts(30 days is common). For very complex parts with many plies,the material's permissible out-time can be a controlling factor.If the material is not completely used,it may be returned to storage.An out-time record should be kept.In addition,freezer storage of these materials is usually limited by the vendor to 6 to 12 months.Overage material will produce laminates with a high level of porosity. The perishability of the material also requires that it be shipped refrigerated from the supplier.Upon arrival at the contractor's facility,there must be provisions to prevent it being left on-dock for long periods of time Tack is another composite material characteristic that is unique.Tack is"stickiness"of the prepreg.It is both an aid and a hindrance.Tack is helpful to maintain location of a ply once it is placed in position.It also makes it difficult to adjust the location once the ply has been placed. 12.2.9 Processing and fabrication Composite parts are fabricated by successive placement of plies one after the other.Parts are built-up rather than machined down.Many metal fabrication steps require successive removal of material starting from large ingots,plates,or forgings.Prepreg "tape"material typically comes in rolls of relatively thin strips(0.005-0.015 inches or 0.13-0.38 mm).These strips are a variety of widths:3",6",and 36". Prepreg "fabric"is usually thicker than tape(0.007-0.020 inches or 0.18-0.51 mm)and usually comes in 36-inch(0.9 m)wide rolls. Fabrication of a detail part requires the material to be taken out of the freezer in a sealed bag and allowed to come to room temperature prior to any operations.Placement of the prepreg on the tool(if not automated)requires care.The plies must be aligned properly to the desired angle and stacked in the prescribed sequence.Prepreg plies come with a backing material to keep them from sticking together on the rolls.This backing material must be removed to prevent contamination of the laminate.Care must be exercised when handling the material to prevent splinters from piercing the hands. Part lay-up(particularly when done by hand)can lead to air entrapment between plies.This creates difficulty when the part is cured because the air may not escape,causing porosity.Thus.thick parts are normally pre-compacted using a vacuum periodically during the lay-up. Some prepreg materials contain an excess of resin.This excess is expected to be"bled"away during cure.Bleeder plies are placed under the vacuum bag to soak up the excess resin.However,most cur- rent prepreg materials are "net resin"so no bleeding is required. Composite processing requires careful attention to tool design.The tools must sustain high pres- sures under elevated temperature conditions.The composite material has different expansion character- istics than most tooling materials,thus thermal stresses are created in the part and in the tool.Tool sur- faces are treated with a release agent to facilitate removal of the part after cure.Tools must also be pres- sure tight because autoclave processing requires application of a vacuum on the laminate as well as posi- tive autoclave pressure.Lastly,tool design must account for the rate of manufacture and the number of parts to be processed. Prepreg material is not fully cured.Curing requires application of heat and pressure that is usually performed in the autoclave.Autoclaves typically apply 85 psi(590 kPa)pressure up to 350F(180C). They can go beyond these values if required for other materials(such as polyimides),but they must be qualified for higher extremes.Autoclave size may limit the size of a part to be designed and manufac- tured.Very large autoclaves are available,but they are expensive and costly to run.Common problems 12-9

MIL-HDBK-17-3F Volume 3, Chapter 12 - Lessons Learned 12-9 12.2.8 Handling and storage Epoxy resins are the most common form of matrix material used in composites. Epoxies are perish￾able. They must be stored below freezing temperature and even then have limited shelf life. Once the material is brought out of storage there is limited time it can be used to make parts (30 days is common). For very complex parts with many plies, the material's permissible out-time can be a controlling factor. If the material is not completely used, it may be returned to storage. An out-time record should be kept. In addition, freezer storage of these materials is usually limited by the vendor to 6 to 12 months. Overage material will produce laminates with a high level of porosity. The perishability of the material also requires that it be shipped refrigerated from the supplier. Upon arrival at the contractor's facility, there must be provisions to prevent it being left on-dock for long periods of time. Tack is another composite material characteristic that is unique. Tack is "stickiness" of the prepreg. It is both an aid and a hindrance. Tack is helpful to maintain location of a ply once it is placed in position. It also makes it difficult to adjust the location once the ply has been placed. 12.2.9 Processing and fabrication Composite parts are fabricated by successive placement of plies one after the other. Parts are built-up rather than machined down. Many metal fabrication steps require successive removal of material starting from large ingots, plates, or forgings. Prepreg "tape" material typically comes in rolls of relatively thin strips (0.005-0.015 inches or 0.13 - 0.38 mm). These strips are a variety of widths: 3", 6", and 36". Prepreg "fabric" is usually thicker than tape (0.007-0.020 inches or 0.18 - 0.51 mm) and usually comes in 36-inch (0.9 m) wide rolls. Fabrication of a detail part requires the material to be taken out of the freezer in a sealed bag and allowed to come to room temperature prior to any operations. Placement of the prepreg on the tool (if not automated) requires care. The plies must be aligned properly to the desired angle and stacked in the prescribed sequence. Prepreg plies come with a backing material to keep them from sticking together on the rolls. This backing material must be removed to prevent contamination of the laminate. Care must be exercised when handling the material to prevent splinters from piercing the hands. Part lay-up (particularly when done by hand) can lead to air entrapment between plies. This creates difficulty when the part is cured because the air may not escape, causing porosity. Thus, thick parts are normally pre-compacted using a vacuum periodically during the lay-up. Some prepreg materials contain an excess of resin. This excess is expected to be "bled" away during cure. Bleeder plies are placed under the vacuum bag to soak up the excess resin. However, most cur￾rent prepreg materials are "net resin" so no bleeding is required. Composite processing requires careful attention to tool design. The tools must sustain high pres￾sures under elevated temperature conditions. The composite material has different expansion character￾istics than most tooling materials, thus thermal stresses are created in the part and in the tool. Tool sur￾faces are treated with a release agent to facilitate removal of the part after cure. Tools must also be pres￾sure tight because autoclave processing requires application of a vacuum on the laminate as well as posi￾tive autoclave pressure. Lastly, tool design must account for the rate of manufacture and the number of parts to be processed. Prepreg material is not fully cured. Curing requires application of heat and pressure that is usually performed in the autoclave. Autoclaves typically apply 85 psi (590 kPa) pressure up to 350°F (180°C). They can go beyond these values if required for other materials (such as polyimides), but they must be qualified for higher extremes. Autoclave size may limit the size of a part to be designed and manufac￾tured. Very large autoclaves are available, but they are expensive and costly to run. Common problems

MIL-HDBK-17-3F Volume 3,Chapter 12-Lessons Learned that occur in autoclave operations include blown vacuum bags,improper heat-up rates,and loss of pres- sure. Once the part is cured it may still require drilling,trimming and machining.Drilling of composites re- quires very sharp bits,careful feed and speed,and support of the back face to prevent splintering.Wa- ter-jet cutters are very useful for trimming.Machining produces a fine dust that requires protection for the operator's safety. 12.2.9.1 Quality control The quality control function for composite materials starts at a much earlier phase than for metals. There is much coordination and interaction occurring between the material supplier and the user before the material is ever shipped.These controls are defined by the material and process specifications and in some cases design allowables requirements.The supplier is often required to perform chemical and me- chanical tests on the material prior to shipment.These involve the individual material constituents,the prepreg,and cured laminates. Material processing and handling must be monitored throughout the various manufacturing phases. Receiving inspections are performed on the prepreg and cured laminates when the material first comes in.From this time on the material is tracked to account for its shelf life and out-time. Quality control activities include verification of the ply lay-up angle,its position in the stack,the num- ber of plies,and the proper trim.During lay-up it is necessary to ensure all potential contaminates and foreign materials are not allowed to invade the material. The curing process is monitored to ensure proper conformance to time-temperature-pressure profiles. These records are maintained for complete traceability of the parts. After the part is cured,there are a number of methods to verify its adequacy.One of the most com- mon is Through-Transmission-Ultrasonics (TTU).Parts with high porosity or delaminations can not transmit sound as well as unflawed parts.Thus ultrasound transmission is attenuated in a flawed part. Other techniques used to verify part quality include traveler specimens,specimens cut from excess mate- rial on the part,tracer yarns within the laminate,and in some cases proof loading.Visual inspections, thickness measurements,and tap testing also serve to interrogate composite parts. One of the most crucial aspects of quality control is information on the effect of defects.It is not enough to discover a flaw or suspected non-conformity.There must also be sufficient information to evaluate the impact of that rejection.The quality control function in its entirety includes the dispositioning of exposed non-conformances.Dispositioning includes acceptance as-is,repair or rework,and scrap- page.If proper dispositioning is not possible because of a lack of knowledge about the effect of defects, an inordinate expense will be incurred scrapping or reworking affected parts. 12-10

MIL-HDBK-17-3F Volume 3, Chapter 12 - Lessons Learned 12-10 that occur in autoclave operations include blown vacuum bags, improper heat-up rates, and loss of pres￾sure. Once the part is cured it may still require drilling, trimming and machining. Drilling of composites re￾quires very sharp bits, careful feed and speed, and support of the back face to prevent splintering. Wa￾ter-jet cutters are very useful for trimming. Machining produces a fine dust that requires protection for the operator's safety. 12.2.9.1 Quality control The quality control function for composite materials starts at a much earlier phase than for metals. There is much coordination and interaction occurring between the material supplier and the user before the material is ever shipped. These controls are defined by the material and process specifications and in some cases design allowables requirements. The supplier is often required to perform chemical and me￾chanical tests on the material prior to shipment. These involve the individual material constituents, the prepreg, and cured laminates. Material processing and handling must be monitored throughout the various manufacturing phases. Receiving inspections are performed on the prepreg and cured laminates when the material first comes in. From this time on the material is tracked to account for its shelf life and out-time. Quality control activities include verification of the ply lay-up angle, its position in the stack, the num￾ber of plies, and the proper trim. During lay-up it is necessary to ensure all potential contaminates and foreign materials are not allowed to invade the material. The curing process is monitored to ensure proper conformance to time-temperature-pressure profiles. These records are maintained for complete traceability of the parts. After the part is cured, there are a number of methods to verify its adequacy. One of the most com￾mon is Through-Transmission-Ultrasonics (TTU). Parts with high porosity or delaminations can not transmit sound as well as unflawed parts. Thus ultrasound transmission is attenuated in a flawed part. Other techniques used to verify part quality include traveler specimens, specimens cut from excess mate￾rial on the part, tracer yarns within the laminate, and in some cases proof loading. Visual inspections, thickness measurements, and tap testing also serve to interrogate composite parts. One of the most crucial aspects of quality control is information on the effect of defects. It is not enough to discover a flaw or suspected non-conformity. There must also be sufficient information to evaluate the impact of that rejection. The quality control function in its entirety includes the dispositioning of exposed non-conformances. Dispositioning includes acceptance as-is, repair or rework, and scrap￾page. If proper dispositioning is not possible because of a lack of knowledge about the effect of defects, an inordinate expense will be incurred scrapping or reworking affected parts