Composite Fasteners, Machining, and Safety

Composite Patch Bonded to Aluminum Structure Composite materials can be used to structurally repair, restore, or enhance aluminum, steel, and titanium components. Bonded composite doublers have the ability to slow or stop fatigue crack growth, replace lost structural area due to corrosion grindouts, and structurally enhance areas with small and negative margins. Boron epoxy, GLARE®, and graphite epoxy materials have been used as composite patches to restore damaged metallic wing skins, fuselage sections, floor beams, and bulkheads. As a crack growth inhibitor, the stiff bonded composite materials constrain the cracked area, reduce the gross stress in the metal, and provide an alternate load path around the crack. As a structural enhancement or blendout filler, the high modulus fiber composites offer negligible aerodynamic resistance and tailorable properties.

Surface preparation is very important to achieve the adhesive strength. Grit blast silane and phosphoric acid anodizing are used to prepare aluminum skin. Film adhesives using a 250 °F (121 °C) cure are used routinely to bond the doublers to the metallic structure. Critical areas of the installation process include a good thermal cure control, having and maintaining water-free bond surfaces, and chemically and physically prepared bond surfaces. Secondarily bonded precured doublers and in-situ cured doublers have been used on a variety of structural geometries ranging from fuselage frames to door cutouts to blade stiffeners. Vacuum bags are used to apply the bonding and curing pressure between the doubler and metallic surface.

Fiberglass Molded Mat Repairs Fiberglass molded mats consists of short fibers, and the strength is much less than other composite products that use continuous fibers. Fiberglass molded mats are not used for structural repair applications, but could be used for non- structural applications. The fiberglass molded mat is typically used in combination with fiberglass fabric. The molded mats are impregnated with resin just like a wet layup for fiberglass fabric. The advantage of the molded mat is the lower cost and the ease of use. Radome Repairs Aircraft radomes, being an electronic window for the radar, are often made of nonconducting honeycomb sandwich structure with only three or four plies of fiberglass. The skins are thin so that they do not block the radar signals.

The thin structure, combined with the location in front of the aircraft, makes the radome vulnerable to hail damage, bird strikes, and lightning strikes. Low-impact damage could lead to disbonds and delamination. Often, water is found in the radome structure due to impact damage or erosion. The moisture collects in the core material and begins a freeze- thaw cycle each time the airplane is flown. This eventually breaks down the honeycomb material causing a soft spot on the radome itself. Damage to a radome needs to be repaired quickly to avoid further damage and radar signal obstructions. Trapped water or moisture can produce a shadow on the radar image and severely degrade the performance of the radar.

To detect water ingression in radomes, the available NDE techniques include x-ray radiography, infrared thermography, and a radome moisture meter that measures the RF power loss caused by the presence of water. The repairs to radomes are similar to repairs to other honeycomb structures, but the technician needs to realize that repairs could affect the radar performance. A special tool is necessary to repair severely damaged radomes. [Figure 7-68] Transmissivity testing after radome repair ensures that the radar signal is transmitted properly through the radome. Radomes have lightning protections strips bonded to the outside of the radome to dissipate the energy of a lightning 20 psi air Drill holes Inject resin Delamination Skin Injection gun 7-40 strike. It is important that these lightning protection strips are in good condition to avoid damage to the radome structure.

Typical failures of lightning protection strips that are found during inspection are high resistance caused by shorts in the strips or attaching hardware and disbonding of the strips from the radome surface. [Figures 7-69] External Bonded Patch Repairs Repairs to damaged composite structures can be made with an external patch. The external patch repair could be made with prepreg, a wet layup, or a precured patch. External patches are usually stepped to reduce the stress concentration at the edge of the patch. The disadvantages of the external patch are the eccentricity of the loading that causes peel stresses and the protrusion of the patch in the air stream. The advantage of the external patch is that it is easier to accomplish than a flush scarf-type repair.

External Bonded Repair With Prepreg Plies The repair methods for carbon, fiberglass, and Kevlar ® are similar. Fiberglass is sometimes used to repair Kevlar ® material. The main steps in repairing damage with an external patch are investigating and mapping the damage, removal of the damage, layup of the repair plies, vacuum bagging, curing, and finish coating. Step 1: Investigating & Mapping the Damage Use the tap test or ultrasonic test to map out the damage. Step 2: Damage Removal Trim out the damage to a smooth round or oval shape. Use scotch or sand paper to rough up the parent surface at least 1 inch larger than the patch size. Clean the surface with an approved solvent and cheese cloth.

Step 3: Layup of the Repair Plies Use the SRM to determine the number, size, and orientation of the repair plies. The repair ply material and orientation must be the same as the orientation of the parent structure. The repair can be stepped to reduce peel stresses at the edges. Step 4: Vacuum Bagging A film adhesive is placed over the damaged area and the repair layup is placed on top of the repair. The vacuum bagging materials are placed on top of the repair (see Prepreg Layup and Controlled Bleed Out) and a vacuum is applied. Step 5: Curing the Repair The prepreg patch can be cured with a heater blanket that is placed inside the vacuum bag, oven, or autoclave when the part can be removed from the aircraft. Most prepregs and film adhesives cure at either 250 °F or 350 °F. Consult the SRM for the correct cure cycle.

Step 6: Applying Top Coat Remove the vacuum bag from the repair after the cure and inspect the repair, remove the patch if the repair is not satisfactory. Lightly sand the repair and apply a protective topcoating. External Repair Using Wet Layup & Double Vacuum Debulk Method (DVD) Generally, the properties of a wet layup repair are not as good as a repair with prepreg material; but by using a DVD method, the properties of the wet layup process can be improved. The DVD process is a technique to remove entrapped air that causes porosity in wet layup laminates. The DVD process is often used to make patches for solid laminate structures for complex contoured surfaces. The wet layup patch is prepared in a DVD tool and then secondary bonded to the aircraft structure. [Figure 7-70] The laminating process is similar to a standard wet layup process. The difference is how the patch is cured.

Double Vacuum Debulk Principle The double vacuum bag process is used to fabricate wet layup or prepreg repair laminates. Place the impregnated fabric within the debulking assembly, shown in Figure 7-70. To begin the debulking process, evacuate the air within the inner flexible vacuum bag. Then, seal the rigid outer box onto the inner vacuum bag, and evacuate the volume of air between the rigid outer box and inner vacuum bag. Since the outer box is rigid, the second evacuation prevents atmospheric pressure from pressing down on the inner vacuum bag over the patch. This subsequently prevents air bubbles from being pinched off within the laminate and facilitates air removal by the inner vacuum. Next, heat the laminate to a predetermined debulking temperature in order to reduce the resin viscosity and further improve the removal of air and volatiles from the laminate. Apply the heat through a heat blanket that is controlled with thermocouples placed directly on the heat 7-41 Inner vacuum bag extends past rigid boxInner vacuum bag Rigid box with two layers of breather and vacuum bagRigid outer box: manufacture sides from 2" x 4", top from 1" plywood, drill 1/4" air holes on each side.

Box top Nails Side boards (wooden 2" x 4") Air holes (approxmately 0.25" diameter) 1" thick plywood blanket. Once the debulking cycle is complete, compact the laminate to consolidate the plies by venting the vacuum source attached to the outer rigid box, allowing atmospheric pressure to reenter the box and provide positive pressure against the inner vacuum bag. Upon completion of the compaction cycle, remove the laminate from the assembly and prepare for cure. DVD tools can be purchased commercially but can also be fabricated locally from wood two-by-fours and sheets of plywood, as illustrated in Figure 7-70.

Patch Installation on the Aircraft After the patch comes out of the DVD tool, it is still possible to form it to the contour of the aircraft, but the time is typically limited to 10 minutes. Place a film adhesive, or paste adhesive, on the aircraft skin and place the patch on the aircraft. Use a vacuum bag and heater blanket to cure the adhesive. [Figures 7-71 and 7-72] 1. Insulation 2. Heat blanket 3. Caul plate 4. Nonporous film 5. Porous film 6. Patch laminate 7. Bagging materials 8. Bagging film V1: Inner vacuum V2: Outer vacuum Bottom plate 1 2 3 4 5 6 5 7 8 V2 V1 Vacuum port Rigid outer box Vacuum port External Repair Using Precured Laminate Patches Precured patches are not very flexible and cannot be used on highly curved or compound curved surfaces. The repair steps are similar as in External Bonded Repair With Prepreg Plies, except step 3 and 4 that follow.

Step 3: A Precured Patch Consult the SRM for correct size, ply thickness, and orientation. You can laminate and cure the precured patch in the repair shop and secondary bond to the parent structure, or obtain standard precured patches. [Figure 7-73] Step 4: For a Precured Patch Apply film adhesive or paste adhesive to the damaged area and place the precured patch on top. Vacuum bag the repair and cure at the correct temperature for the film adhesive or paste adhesive. Most film adhesives cure at either 250 °F or 350 °F. Some paste adhesives cure at room temperature although an elevated temperature could be used to speed the curing process.

Bonded versus Bolted Repairs Bonded repair concepts have found applicability in both types of manufacturing assembly methods. They have the advantage of not introducing stress concentrations by drilling fastener holes for patch installation and can be stronger than original part material. The disadvantage of bonded repairs is that most repair materials require special storage, handling, and curing procedures. Bolted repairs are quicker and easier to fabricate than bonded repairs. They are normally used on composite skins thicker than 0.125-inch to ensure sufficient fastener bearing area is available for load transfer. They are prohibited in honeycomb sandwich assemblies due to the potential for moisture intrusion from the fastener holes and the resulting 7-42 Vent outer box after 60 minutes 140 120 100 80 60 40 20 Ramp rate 1 °F to 5 °F per minute Inner bag Full vacuum Outer box Full vacuum Hold at 125 °F for 90 minutes ± 5 minutes Inner bag Full vacuum Outer box No vacuum 0 10 20 30 40 50 60 70 80 90 100 110 120 SMP-0029M1-9 Double Vacuum Debulk Cycle—Laminate Thickness <16 plies 140 120 100 80 60 40 20 0 10 20 30 40 50 60 70 80 90 100 110 120 130 140 150 160 170 180 SMP-0029M1-10 Double Vacuum Debulk Cycle—Laminate Thickness <16 plies Vent outer box after 60 minutes Inner bag Full vacuum Outer box No vacuum Ramp rate 1 °F to 5 °F per minute Inner bag Full vacuum Outer box Full vacuum Hold at 125 °F for 150 minutes ± 5 minutes core degradation. Bolted repairs are heavier than comparable bonded repairs, limiting their use on weight-sensitive flight control surfaces.

Honeycomb sandwich parts often have thin face sheets and are most effectively repaired by using a bonded scarf type repair. A bonded external step patch can be used as an alternative. Bolted repairs are not effective for thin laminates because of the low bearing stress of the composite laminate. Thicker solid laminates used on larger aircraft can be up to an inch thick in highly loaded areas and these types of laminates cannot be effectively repaired using a bonded scarf type repair. [Figure 7-74] Bolted Repairs Aircraft designed in the 1970s used composite sandwich honeycomb structure for lightly loaded secondary structure, but new large aircraft use thick solid laminates for primary structure instead of sandwich honeycomb. These thick solid laminate structures are quite different from the traditional sandwich honeycomb structures used for flight controls, landing gear doors, flaps, and spoilers of today’s aircraft.

They present a challenge to repair and are difficult to repair with a bonded repair method. Bolted repair methods have been developed to repair thicker solid laminates. 7-43 Table 2 X Lightly loaded structures – laminate thickness less than 0.1" Highly loaded structures – laminate thickness between 0.125" – 0.5" Highly loaded structures – laminate thickness larger than 0.5" High peeling stresses Honeycomb structure Dry surfaces Wet and/or contaminated surfaces Disassembly required Restore unnotched strength Bonded versus bolted repair BondedBolted Bolted repairs are not desirable for honeycomb sandwich structure due to the limited bearing strength of the thin face sheets and weakened honeycomb structure from drilling holes. The advantage of a bolted repair is that you need to select only patch material and fasteners, and the repair method is similar to a sheet metal repair. There is no need for curing the repair and storing the prepreg repair material and film adhesives in a freezer. Patches may be made from aluminum, titanium, steel, or precured composite material.

Composite patches are often made from carbon fiber with an epoxy resin or fiberglass with an epoxy resin. You can repair a carbon fiber structure with an aluminum patch, but you must place a layer of fiberglass cloth between the carbon part and the aluminum patch to prevent galvanic corrosion. Titanium and precured composite patches are preferred for repair of highly loaded components. Precured carbon/epoxy patches have the same strength and stiffness as the parent material as they are usually cured similarly. Titanium or stainless steel fasteners are used for bolted repairs of a carbon fiber structure. Aluminum fasteners corrode if used with carbon fiber. Rivets cannot be used because the installation of rivets using a rivet gun introduce damage to the hole and surrounding structure and rivets expand during installation, which is undesirable for composite structures because it could cause delimination of the composite material.

Repair Procedures Step 1: Inspection of the Damage The tap test is not effective to detect delamination in thick laminates unless the damage is close to the surface. An ultrasonic inspection is necessary to determine the damage area. Consult the SRM to find an applicable NDI procedure. Step 2: Removal of the Damage The damaged area needs to be trimmed to a round or rectangular hole with large smooth radii to prevent stress concentrations. Remove the damage with a sander, router, or similar tool. Step 3: Patch Preparation Determine the size of the patch based on repair information found in the SRM. Cut, form, and shape the patch before attaching the patch to the damaged structure. It is easier to make the patch a little bigger than calculated and trim to size after drilling all fastener holes. In some cases, the repair patches are stocked preshaped and predrilled. If cutting is to be performed, standard shop procedures should be used that are suitable for the patch material. Titanium is hard to work and requires a large powerful slip roller to curve the material.

Metal patches require filing to prevent crack initiation around the cut edges. When drilling pilot holes in the composite, the holes for repair fasteners must be a minimum of four diameters from existing fasteners and have a minimum edge distance of three fastener diameters. This is different from the standard practice for aluminum of allowing a two diameter distance. Specific pilot hole sizes and drill types to be used should follow specific SRM instructions. [Figure 7-75] Step 4: Hole Pattern Lay Out To locate the patch on the damaged area, draw two perpendicular centerlines on the parent structure and on the patch material that define the principal load or geometric directions. Then, lay out hole pattern on the patch and drill pilot holes in the patch material. Align the two perpendicular centerlines of the patch with the lines on the parent structure and transfer the pilot holes to the parent material. Use clecos to keep the patch in place. Mark the edges of the patch so that it can be returned to the same location easily.

Step 5: Drilling & Reaming Holes in Patch & Parent Structure Composite skins should be backed up to prevent splitting. Enlarge the pilot holes in the patch and parent materials with a drill 1⁄64 undersize and then ream all holes to the correct size. A tolerance of +0.0025/–0.000-inch is usually recommended for aircraft parts. For composites, this means interference fasteners are not used. 7-44 Three rows of fasteners are required Edge distance is three times the diameter of the fastener Spacing four to six times the diameter of the fastener Hi-lok or lock bolt Damage is cut out to a smooth rectangular shape Radius of repair plate corner is 0.5" Step 6: Fastener Installation Once fastener holes are drilled full size and reamed, permanent fasteners are installed. Before installation, measure the fastener grip length for each fastener using a grip length gauge. As different fasteners are required for different repairs, consult the SRM for permissible fastener type and installation procedure. However, install all fasteners wet with sealant and with proper torque for screws and bolts.

Step 7: Sealing of Fasteners & Patch Sealants are applied to bolted repairs for prevention of water/ moisture intrusion, chemical damage, galvanic corrosion, and fuel leaks. They also provide contour smoothness. The sealant must be applied to a clean surface. Masking tape is usually placed around the periphery of the patch, parallel with the patch edges and leaving a small gap between the edge of the patch and the masking tape. Sealing compound is applied into this gap. Step 8: Application of Finish Coat & Lightning Protection Mesh The repair needs to be sanded, primed, and painted with an approved paint system. A lightning protection mesh needs to be applied if composite patches are used in an area that is prone to lightning strikes.

Fasteners Used with Composite Laminates Many companies make specialty fasteners for composite structures and several types of fasteners are commonly used: threaded fasteners, lock bolts, blind bolts, blind rivets, and specialty fasteners for soft structures, such as honeycomb panels. The main differences between fasteners for metal and composite structures are the materials and the footprint diameter of nuts and collars. Corrosion Precautions Neither fiberglass nor Kevlar® fiber-reinforced composites cause corrosion problems when used with most fastener materials. Composites reinforced with carbon fibers, however, are quite cathodic when used with materials, such as aluminum or cadmium, the latter of which is a common plating used on fasteners for corrosion protection.

Fastener Materials Titanium alloy Ti-6Al-4V is the most common alloy for fasteners used with carbon fiber reinforced composite structures. Austenitic stainless steels, superalloys (e.g., A286), multiphase alloys (e.g., MP35N or MP159), and nickel alloys (e.g., alloy 718) also appear to be very compatible with carbon fiber composites. Fastener System for Sandwich Honeycomb Structures (SPS Technologies Comp Tite) The adjustable sustain preload (ASP) fastening system provides a simplified method of fastening composite, soft core, metallic or other materials, which are sensitive to fastener clamp-up or installation force conditions.

Clamping force can be infinitely adjustable within maximum recommended torque limits and no further load is applied during installation of the lock collar. The fastener is available in two types. The ASP® has full shank and the 2ASP® has a pilot type shank. [Figures 7-76 and 7-77] Hi-Lok® & Huck-Spin® Lockbolt Fasteners Most composite primary structures for the aircraft industry are fastened with Hi-Loks ® (Hi-Shear Corp.) or Huck- Spin® lockbolts for permanent installations. The Hi-Lok ® is a threaded fastener that incorporates a hex key in the 7-45 threaded end to react to the torque applied to the collar during installation. The collar includes a frangible portion that separates at a predetermined torque value. [Figure 7-78] The lockbolt incorporates a collar that is swaged into annular grooves. It comes in two types: pull and stump. The pull-type is the most common, where a frangible pintail is used to 1 Pin component installed clearance fit 2 Sleeve component threaded on pin 3 Torque controlled tool tightens sleeve react the axial load during the swaging of the collar. When the swaging load reaches a predetermined limit, the pintail breaks away at the breakneck groove. The installation of the Hi-Lok® and the pull-type Huck-Spin ® lockbolt can be performed by one technician from one side of the structure.

The stump-type lockbolt, on the other hand, requires support on the head side of the fastener to react the swage operation. This method is usually reserved for automated assembly of detail structure in which access is not a problem. The specific differences in these fasteners for composite structure in contrast to metal structure are small. For the Hi-Lok®, material compatibility is the only issue; aluminum collars are not recommended. Standard collars of A286, 303 stainless steel, and titanium alloy are normally used. The Huck-Spin® lockbolt requires a hat-shaped collar that incorporates a flange to spread the high bearing loads during installation. The lockbolt pin designed for use in composite structure has six annular grooves as opposed to five for metal structure. [Figures 7-79 and 7-80] Eddie-Bolt® Fasteners Eddie-Bolt® fasteners (Alcoa) are similar in design to Hi- Loks® and are a natural choice for carbon fiber composite structures. The Eddie-Bolt ® pin is designed with flutes in the threaded portion, which allow a positive lock to be made during installation using a specially designed mating nut or 4 Lock collar placed on pin 5 Lock collar swaged on pin splines 6 Pintail breaks off 7-46 HLH 103, HLH 104, HLH 110, HLH 111 or HLH 500 installation tool Large 130 flush shear head Lock grooves CP titanium flanged collar Pull grooves 1 Tool engages lockbolt pintail 2 Gap closes, collar swage begins 3 collar. The mating nut has three lobes that serve as driving ribs. During installation, at a predetermined preload, the lobes compress the nut material into the flutes of the pin and form the locking feature. The advantage for composite structure is that titanium alloy nuts can be used for compatibility and weight saving without the fear of galling. The nuts spin on freely, and the locking feature is established at the end of the installation cycle. [Figure 7-81] Cherry’s E-Z Buck® (CSR90433) Hollow Rivet The Cherry Hollow End E-Z Buck ® rivet is made from titanium/columbium alloy and has a shear strength of 40 KSI.

The E-Z Buck® rivet is designed to be used in a double flush application for fuel tanks. The main advantage of this type of rivet is that it takes less than half the force of a solid rivet of the same material. The rivets are installed with automated riveting equipment or a rivet squeezer. Special optional dies ensure that the squeezer is always centered during installation, avoiding damage to the structure. [Figure 7-82] Blind Fasteners Composite structures do not require as many fasteners as metal aircraft because stiffeners and doublers are co-cured with the skins, eliminating many fasteners. The size of panels on aircraft has increased in composite structures, which causes backside inaccessibility. Therefore, blind fasteners or screws and nutplates must be used in these areas. Many manufacturers make blind fasteners for composite structures; a few are discussed below.

4 Pintail fractures at the break notch 5 Tool anvil reverses off swaged collar 6 Swage process complete Installation complete 7-47 Blind Bolts The Cherry Maxibolt ® is available in titanium for compatibility with composite structures. The shear strength of the Maxibolt® is 95 KSI. It can be installed from one side with a G-83 or equivalent pneumatic-hydraulic installation tool, and is available in 100° flush head, 130° flush head and protruding head styles. [Figure 7-83] The Alcoa UAB™ blind bolt system is designed for composite structures and is available in titanium and stainless steel. The UAB™ blind bolt system is available in 100° flush head, 130° flush head, and protruding head styles.

The Accu-Lok ™ Blind Fastening System is designed specifically for use in composite structures in which access is limited to one side of the structure. It combines high joint preload with a large diameter footprint on the blind side. C2C2 C1 D Grip range 100° ± 1° 100° ± 2° 0.028 0.037 0.046 B REF 0.028 0.037 0.046 1/8 (–4) 5/32 (–5) 3/16 (–6) 7/32 (–7) C1 DIA 0.195 0.189 0.247 0.242 0.302 0.297 0.328 0.323 C2 DIA 0.195 0.189 0.247 0.242 0.302 0.297 0.328 0.323 D DIA 0.132 0.129 0.162 0.159 0.195 0.191 0.227 0.224 Rivet Diameter A REF CSR 90433 A (ref) for manufactured head B (ref) for shop-formed head 2,500 2,700 3,000 3,750 1/8" (–4) 5/32" (–5) 3/16" (–6) 7/32" (–7) Hollow End E-Z Buck® Nominal Diameter Upset Load (Lb) + 200 Lb 839B10-4 839B10-5 839B10-6 839B10-7 1/8" 5/32" 3/16" 7/32" 839B1-4 839B1-5 839B1-6 839B1-7 Rivet Diameter 1/4" Diameter Mount 3/16" Diameter Mount Cherry Flaring Snap Die Part Numbers Flushness +0.005 −0.000 Flushness +0.015 −0.000 Head dimple Squeezer yoke or riveting machine Cherry flaring snap die Hollow End E-Z Buck® Composite material Cherry snap die (optional) (839B3 = 3/16" shank size) (839B13 = 1/4" shank size) Note: 1 die fits all fastener diameters.

Manufactured head Shop formed head 7-48