Honeycomb and Laminate Repairs

External Internal Scarf Patch Adhesive Composite skin Core Core splice adhesive Repair core Repair plug heating device. Typical curing temperature for composite materials is 250 °F or 350 °F. The temperature of large parts that are cured in an oven or autoclave might be different from that of an oven or autoclave during the cure cycle, because they act like a heat sink. The part temperature is most important for a correct cure, so thermocouples are placed on the part to monitor and control part temperature. The oven or autoclave air temperature probe that measures oven or autoclave temperature is not always a reliable device to determine part curing temperature. The oven temperature and the part temperature can be substantially different if the part or tool acts as a heat sink.

The elevated cure cycle consists of at least three segments: • Ramp up: The heating device ramps up at a set temperature typically between 3 °F to 5 °F per minute. • Hold or soak: The heating device maintains the temperature for a predetermined period. • Cool down: The heating device cools down at a set temperature. Cool down temperatures are typically below 5 °F per minute. When the heating device is below 125 °F, the part can be removed. When an autoclave is used for curing parts, make sure that the pressure in the autoclave is relieved before the door is opened. [Figure 7-53] The curing process is accomplished by the application of heat and pressure to the laminate. The resin begins to soften and flow as the temperature is increased. At lower temperatures, very little reaction occurs. Any volatile contaminants, such as air and/or water, are drawn out of the laminate with vacuum during this time. The laminate is compacted by applying pressure, usually vacuum (atmospheric pressure); autoclaves apply additional pressure, typically 50–100 psi. As the temperature approaches the final cure temperature, the rate of reaction greatly increases, and the resin begins to gel and harden. The hold at the final cure lets the resin finish curing and attain the desired structural properties.

Composite Honeycomb Sandwich Repairs A large proportion of current aerospace composite components are light sandwich structures that are susceptible to damage and are easily damaged. Because sandwich structure is a bonded construction and the face sheets are thin, damage to sandwich structure is usually repaired by bonding. Repairs to sandwich honeycomb structure use similar techniques for the most common types of face sheet materials, such as fiberglass, carbon, and Kevlar®. Kevlar® is often repaired with fiberglass. [Figure 7-54] Damage Classification A temporary repair meets the strength requirements, but is limited by time or flight cycles. At the end of the repair’s life, the repair must be removed and replaced. An interim repair LIB ERT Y 2008 Tap test with tap hammerInstrumented tap testCoin tap test 7-33 8 64 2 0 PRESSURE I0 Breather cloth Heat blanket Breather cloth Thermocouple Repair area restores the required strength to the component. However, this repair does not restore the required durability to the component. Therefore, it has a different inspection interval and/or method. A permanent repair is a repair that restores 0.50 inch minimum Partial depth core replacement Full depth core replacement the required strength and durability to the component. The repair has the same inspection method and interval as the original component.

Sandwich Structures Minor Core Damage (Filler & Potting Repairs) A potted repair can be used to repair damage to a sandwich honeycomb structure that is smaller than 0.5 inch. The honeycomb material could be left in place or could be removed and is filled up with a potting compound to restore some strength. Potted repairs do not restore the full strength of the part. Potting compounds are most often epoxy resins filled with hollow glass, phenolic or plastic microballoons, cotton, flox, or other materials. The potting compound can also be used as filler for cosmetic repairs to edges and skin panels. Potting compounds are also used in sandwich honeycomb panels as hard points for bolts and screws. The potting compound is heavier than the original core and this could affect flight control balance. The weight of the repair must be calculated and compared with flight control weight and balance limits set out in the SRM.

7-34 Damage Requiring Core Replacement & Repair to One or Both Faceplates Note: the following steps are not a substitution for the aircraft specific Structural Repair Manual (SRM). Do not assume that the repair methods used by one manufacturer are applicable to another manufacturer. Step 1: Inspect the Damage Thin laminates can be visually inspected and tap tested to map out the damage. [Figure 7-55] Thicker laminates need more in-depth NDI methods, such as ultrasonic inspection. Check in the vicinity of the damage for entry of water, oil, fuel, dirt, or other foreign matter. Water can be detected with X-ray, back light, or a moisture detector.

Step 2: Remove Water From Damaged Area Water needs to be removed from the core before the part is repaired. [Figure 7-56] If the water is not removed, it boils during the elevated temperature cure cycle and the face sheets blow off the core, resulting in more damage. Water in the honeycomb core could also freeze at the low temperatures that exist at high altitudes, which could result in disbonding of the face sheets. Step 3: Remove the Damage Trim out the damage to the face sheet to a smooth shape with rounded corners, or a circular or oval shape. Do not damage the undamaged plies, core, or surrounding material. If the core is damaged as well, remove the core by trimming to the same outline as the skin. [Figure 7-57] Step 4: Prepare the Damaged Area Use a flexible disc sander or a rotating pad sander to taper sand a uniform taper around the cleaned up damage. Some manufacturers give a taper ratio, such as 1:40, and others prescribe a taper distance like a 1-inch overlap for each existing ply of the face sheet. Remove the exterior finish, including conductive coating for an area that is at least 1 inch larger than the border of the taper. Remove all sanding dust with dry compressed air and a vacuum cleaner. Use a clean cloth moistened with approved solvent to clean the damaged area. [Figure 7-58] Step 5: Installation of Honeycomb Core (Wet Layup) Use a knife to cut the replacement core. The core plug must be of the same type, class, and grade of the original core. The direction of the core cells should line up with the honey comb of the surrounding material. The plug must be trimmed to the right length and be solvent washed with an approved cleaner.

For a wet layup repair, cut two plies of woven fabric that fit on the inside surface of the undamaged skin. Impregnate the fabric plies with a resin and place in the hole. Use potting compound around the core and place it in the hole. For a prepreg repair, cut a piece of film adhesive that fits the hole and use a foaming adhesive around the plug. The plug should touch the sides of the hole. Line up the cells of the plug with Adhesive film*Fabric prepreg Replacement core plug Adhesive** Adhesive film* Replacement core plug Adhesive** Section Through Repair Area Partial Depth Core Replacement Section A-A Section Through Repair Area Full Depth Core Replacement Section B-B * BMS 5-154, Grade 5 or two plies of Grade 3 ** BMS 5-90, Type III, Class 1, Grade 50, or BMS 5-90, Type IV * BMS 5-154, Grade 5 ** BMS 5-90, Type III, Class 1, Grade 50, or BMS 5-90, Type IV 7-35 A B A B Nonstructural sanding ply (adhesive film or fiberglass prepreg) Determine number of plies, orientation, and material from skin identification Orient repair plies in same direction as original layers Foaming adhesive BMS 5-90, Type III, Class 1, Grade 50, or BMS 5-90, Type IV Extra ply Prepreg plies 0.50 overlap (typical) Adhesive film Aeraded area. Do not damage ÿbers.

Taper sanded area Masking tape (remove after sanding) Core replacement* *Butt splicing shown. Vacuum gauge Repair Solid parting film Breather material Vacuum bag material Bleeder material Heat blanket Perforated parting film Vacuum probe Vacuum bag sealing compound Caul plate 8 64 2 0 I0 PRESSURE the original material. Vacuum bag the repair area and use an oven, autoclave, or heat blanket to cure the core replacement. The wet layup repair can be cured at a room temperature up to 150 °F. The prepreg repair must be cured at 250 °F or 350 °F. Usually, the core replacement is cured with a separate curing cycle and not co-cured with the patch. The plug must be sanded flush with the surrounding area after the cure. [Figure 7-59] Step 6: Prepare & Install the Repair Plies Consult the repair manual for the correct repair material and the number of plies required for the repair. Typically, one more ply than the original number of plies is installed. Cut the plies to the correct size and ply orientation. The repair plies must be installed with the same orientation as that of the original plies being repaired. Impregnate the plies with resin for the wet layup repair, or remove the backing material from the prepreg material. The plies are usually placed using the smallest ply first taper layup sequence. [Figure 7-60] Step 7: Vacuum Bag the Repair Once the ply materials are in place, vacuum bagging is used to remove air and to pressurize the repair for curing. Refer to Figure 7-61 for bagging instructions.

Step 8: Curing the Repair The repair is cured at the required cure cycle. Wet layup repairs can be cured at room temperature. An elevated temperature up to 150 °F can be used to speed up the cure. 7-36 Time Temperature (°F) 250 175 100 70 Temperature (°C) 121 80 38 21 Note: Keep a minimum vacuum of 22 inches of mercury during the cure cycle. NO SCALE Ramp up Ramp down Soak Increase the temperature 2 °F to 5 °F (0.5 °C to 3 °C) per minute Decrease the temperature 5 °F/minute (3 °C/minute) maximum Below 125 °F (52 °C) release the pressure and remove the layup and vacuum bag materials Hold for 90 to 150 minutes at 260 °F + 6 °F (126 °C + 6 °C) The prepreg repair needs to be cured at an elevated cure cycle.

[Figure 7-62] Parts that can be removed from the aircraft could be cured in a hot room, oven, or autoclave. A heating blanket is used for on-aircraft repairs. Remove the bagging materials after curing and inspect the repair. The repair should be free from pits, blisters, resin- rich and resin-starved areas. Lightly sand the repair patch to produce a smooth finish without damaging the fibers. Apply top finish and conductive coating (lightning protection). Step 9: Post Repair Inspection Use visual, tap, and/or ultrasonic inspection to inspect the repair. Remove the repair patch if defects are found.

[Figure 7-63] Perform a balance check if a repair to a flight control surface was made, and ensure that the repaired flight control is within limits of the SRM. Failure to do so could result in flight control flutter, and safety of flight could be affected. RepairHeat affected area Heat blanket area 7-37 Solid Laminates Bonded Flush Patch Repairs New generation aircraft have fuselage and wing structures made from solid laminates that are externally stiffened with co-cured or co-bonded stringers. These solid laminates have many more plies than the face sheets of honeycomb sandwich structures. The flush repair techniques for solid laminate structures are similar for fiberglass, Kevlar ®, and graphite with minor differences.

A flush repair can be stepped or, more commonly, scarved (tapered). The scarf angles are usually small to ease the load into the joint and to prevent the adhesive from escaping. This translates into thickness-to-length ratios of 1:10 to 1:70. Because inspection of bonded repairs is difficult, bonded repairs, as contrasted with bolted repairs, require a higher commitment to quality control, better trained personnel, and cleanliness. The scarf joint is more efficient from the viewpoint of load transfer as it reduces load eccentricity by closely aligning the neutral axis of the parent and the patch. However, this configuration has many drawbacks in making the repair.

First, to maintain a small taper angle, a large quantity of sound material must be removed. Second, the replacement plies must be very accurately laid up and placed in the repair joint. Third, curing of replacement plies can result in significantly reduced strength if not cured in the autoclave. Fourth, the adhesive can run to the bottom of the joint, creating a nonuniform bond line. This can be alleviated by approximating the scarf with a series of small steps. For these reasons, unless the part is lightly loaded, this type of repair is usually performed at a repair facility where the part can be inserted into the autoclave, which can result in part strength as strong as the original part.

There are several different repair methods for solid laminates. The patch can be precured and then secondarily bonded to the parent material. This procedure most closely approximates the bolted repair. [Figure 7-64] The patch can be made from prepreg and then co-cured at the same time as the adhesive. The patch can also be made using a wet layup repair. The curing cycle can also vary in length of time, cure temperature, and cure pressure, increasing the number of possible repair combinations. Scarf repairs of composite laminates are performed in the sequence of steps described below. Step 1: Inspection & Mapping of Damage The size and depth of damage to be repaired must be accurately surveyed using appropriate nondestructive evaluation (NDE) techniques. A variety of NDE techniques can be used to Repair plies Adhesive Laminate parent material.

inspect for damage in composite structures. The simplest technique is visual inspection, in which whitening due to delamination and/or resin cracking can be used to indicate the damage area in semitransparent composites, such as glass-polyester and glass-vinyl ester laminates. Visual inspection is not an accurate technique because not all damage is detectable to the eye, particularly damage hidden by paint, damage located deep below the surface, and damage in nontransparent composites, such as carbon and aramid laminates. A popular technique is tap testing, in which a lightweight object, such as a coin or hammer, is used to locate damage. The main benefits of tap testing are that it is simple and it can be used to rapidly inspect large areas. Tap testing can usually be used to detect delamination damage close to the surface, but becomes increasingly less reliable the deeper the delamination is located below the surface. Tap testing is not useful for detecting other types of damage, such as resin cracks and broken fibers.

More advanced NDE techniques for inspecting composites are impedance testing, x-ray radiography, thermography, and ultrasonics. Of these techniques, ultrasonics is arguably the most accurate and practical and is often used for surveying damage. Ultrasonics can be used to detect small delaminations located deep below the surface, unlike visual inspection and tap testing. Step 2: Removal of Damaged Material Once the scope of the damaged area to be repaired has been determined, the damaged laminate must be removed. The edges of the sound laminate are then tapered back to a shallow angle. The taper slope ratio, also known as the scarf angle, should be less than 12 to 1 (< 5°) to minimize the shear strains along the bond line after the repair patch is applied. The shallow angle also compensates for some errors in workmanship and other shop variables that might diminish patch adhesion. [Figure 7-65] 7-38 Initially, machine scarf to a knife’s edge steeper than required.

Continue working scarf back to scarf outline dimension. Finished scarf slope Sanding disk Scarf outline periphery Sanding disk holder Nonstructural sanding ply (adhesive film or fiberglass prepreg) Extra repair ply Third repair ply Second repair ply First repair ply Adhesive film Masking tape (3.0 to 4.0 wide) Taper sand Step 3: Surface Preparation The laminate close to the scarf zone should be lightly abraded with sandpaper, followed by the removal of dust and contaminates. It is recommended that, if the scarf zone has been exposed to the environment for any considerable period of time, it should be cleaned with a solvent to remove contamination.

Step 4: Molding A rigid backing plate having the original profile of the composite structure is needed to ensure the repair has the same geometry as the surrounding structure. Step 5: Laminating Laminated repairs are usually done using the smallest ply-first taper sequence. While this repair is acceptable, it produces relatively weak, resin-rich areas at each ply edge at the repair interface. The largest ply first laminate sequence, where the first layer of reinforcing fabric completely covers the work area, produces a stronger interface joint. Follow the manufacturer’s SRM instructions. Selection of the reinforcing material is critical to ensuring the repair has acceptable mechanical performance. The reinforcing fabric or tape should be identical to the reinforcement material used in the original composite. Also, the fiber orientation of the reinforcing layers within the repair laminate should match those of the original part laminate, so that the mechanical properties of the repair are as close to original as possible.

Step 6: Finishing After the patch has cured, a conducting mesh and finish coat should be applied if needed. Trailing Edge & Transition Area Patch Repairs Trailing edges of control panels are highly vulnerable to damage. The aft 4 inches are especially subject to ground collision and handling, as well as to lightning strike. Repairs in this region can be difficult because both the skins and the trailing edge reinforcement may be involved. The repairs to a honeycomb core on a damaged edge or panel are similar to the repair of a sandwich honeycomb structure discussed in the Damage Requiring Core Replacement and Repair to One or Both Faceplate Repair sections. Investigate the damage, remove damaged plies and core, dry the part, install new core, layup the repair plies, curing and post inspection. A typical trail edge repair is shown in Figure 7-66.

Resin Injection Repairs Resin injection repairs are used on lightly loaded structures for small damages to a solid laminate due to delamination. Two holes are drilled on the outside of the delamination area and a low-viscosity resin is injected in one hole until it flows out the other hole. Resin injection repairs are sometimes used on sandwich honeycomb structure to repair a facesheet disbond. Disadvantages of the resin injection method are that the fibers are cut as a result of drilling holes, it is difficult to remove moisture from the damaged area, and it is difficult to achieve complete infusion of resin. [Figure 7-67] 7-39