FAA Joint Advanced Materials & Structures (JAMS) Center of Excellence 9 th Annual Technical Review Meeting April 9 th, 2013

Transcription

1 ABSTRACT FAA Joint Advanced Materials & Structures (JAMS) Center of Excellence CACRC DEPOT BONDED REPAIR INVESTIGATION ROUND ROBIN TESTING John Tomblin and Lamia Salah National Institute for Aviation Research, Wichita, KS Curtis Davies, Lynn Pham FAA William J. Hughes Technical Center, International Airport, NJ Major technological advances using fiber-reinforced composite materials in airframe components to improve performance and promote energy efficiency have been achieved in the last 50 years leading to the introduction of these materials for the first time, in load-bearing wing and fuselage structures. Durability, repairability, and maintainability are key elements in the continued airworthiness of composite structures. Challenges associated with composite repair and supportability of composite structures are of particular interest and must be addressed during the design phase of the component. Rigorous material, structural and repair process substantiation, validation and implementation are crucial to ensure the structural integrity of bonded composite structures and repairs. The fundamental objective of this research work is to evaluate the CACRC (commercial aircraft composite repair committee) standards developed for repair of composite structures, the potential variability in the repair performance associated with technician training, as well as that resulting from repair implementation at various operator depots. The static and residual strength after cyclic loading of OEM (original equipment manufacturer) and field bonded repairs as well as the fatigue performance of these repairs under severe temperature and moisture environments are investigated and key elements in the process and implementation of bonded repairs are reported. RESEARCH OBJECTIVES The use of fiber-reinforced composites in aircraft structural components has significantly increased in the last few decades due to their improved specific strength and stiffness, the ability to be tailored to design requirements in various loading directions, manufacturability, the possibility of building a composite part in a single integral shell, and superior corrosion resistance and fatigue endurance when compared to metallic components. These features offer a great potential for weight savings, reduction in maintenance and operating costs resulting in more cost efficient and profitable airframes[1]. To capitalize of the performance benefits these materials offer, numerous technological challenges must be overcome in various disciplines, one of the most critical being repair and maintenance. As most repair and maintenance depots and facilities prepare for this new generation of composite airframe components, it is essential that the infrastructure for supportability is in place to maintain these structures and ensure their structural integrity. The fundamental objective of this research effort is to evaluate the existing CACRC standards for repair of composite structures. This study will be used to assess the repair process variability among multiple operator depots, using SRM(structural repair manual)-like 1

2 procedures, referencing CACRC repair techniques provided to all the depots. The variability associated with technician training, identified as a key element in the repair performance in a previous study will also be investigated. The ultimate strength and durability of OEM repairs will be compared to that of the CACRC repairs and to a set of control and open-hole panels to establish the residual strength of these repaired components. This study will also be used to evaluate the environmental effects on the ultimate strength and residual strength after fatigue of bonded repairs. Furthermore, weak interfacial bonds between composite substrates resulting from a deficient process or a compromised interface are not detectable with the current inspection methods and are likely to degrade as the component is in service, subjected to loading and the environment. It is therefore essential to interrogate the performance of these weak joints and draw attention to the need for appropriate training in the composite repair community. In the second phase of this research work, the static and residual strength after fatigue of damaged and substandard repairs bonded using deficient processes is investigated. The results may be used to identify the degree of criticality of the different steps within a bonded repair and the consequences of poor repair processes on the performance of the repaired component. The ultimate goal is to identify key elements in the implementation of bonded repairs that ensure repeatability and structural integrity of these repairs and to provide recommendations pertaining to the validity of the existing CACRC standards, repair technician training and repair process control. RESEARCH METHODOLOGY Details of the research investigative plan, panel manufacture and specimen design and validation were previously discussed [2, 3]. The large beam repair element used for the round robin exercise is shown in figure 1. The element is 11.5"x 48" with 4-ply facesheets and reinforced 8 pcf, 2-inch thick core under the loading and support points to prevent core crushing. The parent facesheet material is T300/ 934 3KPW bonded to the core using FM 377S adhesive (OEM). CL Symmetric Core Fill (4 places) 1/8" cell, 3 pcf 3/16" cell, 7 pcf Ø7.50 Ø Figure 1. Large Beam Configuration 2

3 The proposed test matrix is summarized in Table 1. Four repair systems are considered, an OEM repair system using the parent material and adhesive for repair (350 F cure repair, labeled as OEM-R1), a wet lay-up repair system using Tenax HTA tex f3000t0 fabric with EA9396 C2 laminating resin and EA9696 adhesive (labeled as OEM-R2) and two CACRC field repair systems using Hexcel M20/G904 prepreg and EA9695 NW adhesive (250 F cure repair, labeled as CACRC-R1) and using Tenax HTA tex f3000t0 fabric with Epocast 52A/B (200 F cure repair, wet lay-up) (labeled as CACRC-R2). These repairs will be compared to baseline pristine unrepaired coupons, unrepaired coupons with a 2.5 hole diameter and OEM repaired coupons. All materials have been supplied by the OEM and all panel fabrication has been conducted at NIAR/ NCAT facility using OEM approved processes to ensure that the resulting repair elements are representative of production materials and processes. Table 1. CACRC Round Robin Test Matrix Repair Station Coupon Configuration Repair Material Loading Mode Experience Static Static Fatigue Level RTA ETW ETW N/A Pristine/ Undamaged N/A Compression N/A 2.5" hole N/A- Open Hole Compression 3 3 OEM/ NIAR Repair/ 2.5" hole OEM-R1 Compression 3 3 OEM/ NIAR Repair/ 2.5" hole OEM-R2 Compression 3 3 OEM/ NIAR Repair/ 2.5" hole OEM-R2 Tension 3 3 OEM/ NIAR Repair/ 2.5" hole CACRC-R1 Compression 3 3 OEM/ NIAR Repair/ 2.5" hole CACRC-R1 Tension 3 3 OEM/ NIAR Repair/ 2.5" hole CACRC-R2 Compression 3 3 OEM/ NIAR Repair/ 2.5" hole CACRC-R2 Tension 3 3 Field Station 1 Repair/ 2.5" hole CACRC-R1 Compression M1 3 Field Station 1 Repair/ 2.5" hole CACRC-R2 Compression M1 3 Field Station 1 Repair/ 2.5" hole CACRC-R1 Compression M2 3 Field Station 1 Repair/ 2.5" hole CACRC-R2 Compression M2 3 Field Station 2 Repair/ 2.5" hole CACRC-R1 Compression M1 3 Field Station 2 Repair/ 2.5" hole CACRC-R2 Compression M1 3 Field Station 2 Repair/ 2.5" hole CACRC-R1 Compression M2 3 Field Station 2 Repair/ 2.5" hole CACRC-R2 Compression M2 3 Field Station 3 Repair/ 2.5" hole CACRC-R1 Compression M1 3 Field Station 3 Repair/ 2.5" hole CACRC-R2 Compression M1 3 Field Station 3 Repair/ 2.5" hole CACRC-R1 Compression M2 3 Field Station 3 Repair/ 2.5" hole CACRC-R2 Compression M2 3 Field Station 4 Repair/ 2.5" hole CACRC-R1 Compression M1 3 Field Station 4 Repair/ 2.5" hole CACRC-R2 Compression M1 3 Field Station 4 Repair/ 2.5" hole CACRC-R1 Compression M2 3 Field Station 4 Repair/ 2.5" hole CACRC-R2 Compression M2 3 A detailed repair procedure, referencing the relevant SAE CACRC standards was drafted and reviewed by the FAA technical monitors, industry points of contact and all participating OEM and airline depots before performing the repairs. The same procedures were used by all participating airline depots to conduct the repairs. Detailed repair checklists containing part specific information along with the CACRC standards were provided to the depot mechanics. Half of the repairs were conducted by mechanics with minimal experience and the other half by experienced mechanics. Repair kits, containing the CACRC repair materials were prepared and shipped to all participating depots as illustrated in figure 2. The CACRC materials used for repair are Hexcel M20/G904 prepreg with EA9695 NW 0.05 psf film adhesive approved per SAE AMS 3970 [4, 3

4 5,6,7,8] and Hexcel G904 D1070 TCT PW dry fabric, 193 g/m 2 using Tenax Fibers and Huntsmann Epocast 52A/B resin approved per AMS 2980 [13, 14, 15, 16, 17, 18]. Peel ply and perforated film for wet lay-up bagging were also provided to the depots to control the material process variability. Representative repair instructions are summarized in the following sections. Figure 2. CACRC prepreg and wet lay-up repair materials CACRC-R1 PREPREG REPAIR PROCEDURE 1 Thaw the repair kit from 0 F (-17.8 C) to room temperature for at least 10 hours prior to bonding the repair [4]. The kit supplied is contained in a sealed moisture proof bag and should be stored at or below 0 F (-17.8 C). The storage life for this repair system (prepreg and adhesive) is 12 months from the date of shipment and the out time is 10 days at room temperature ambient [8]. Please refer to the actual kit out time and shelf life as specified in each package (out time for material handling and kit preparation already accounted for). 2 Inspect the materials contained in the repair kit to ensure that all materials required for repair have been provided and cut the required adhesive and prepreg plies Materials Hexcel M20/G904 prepreg material (cut the required repair plies: 3.5, 4.5, 5.5 and 6.5 diameter) EA9695 NW 0.05 psf film adhesive material (cut two layers of adhesive 7.0 in diameter) Seal and store at room temperature until needed for the repair/ bonding operations. If the repair is planned for a later date, store in freezer at or below 0 F (-17.8 C) until needed. Allow 8-10 hours to thaw kit prior to bonding the repair. 3 Prepare the panel for repair by identifying the panel 0 degree direction (0 degree direction along the panel length). Mark panel (PNL) orientation as 0PNL, 90PNL, 45PNL as shown in figure 3. All repairs MUST be conducted on facesheet 2 (FS2) of the panels. The panel 0 direction is aligned with the panel 48" length. 4

5 4 Mask a 7" diameter area at the center of the panel in preparation for scarf sanding. Use the guidelines of ARP 4916 masking methods 4 or 5 using masking tape [15]. Using a mylar template (or equivalent), mark the inner 2.5" diameter scarf circle/ boundary and the outer 6.5" diameter scarf periphery. Please take a photograph of the prepared marked panel ready for sanding including a 12-inch scale for reference. 5 Scarf sand from an inner diameter of 2.5" to an outer diameter of 6.5" per AIR 5367 [16] sections 5.0 and 5.1 using a 0.5" " inch per ply overlap. (Use an air powered die grinder, or equivalent, with aluminum oxide sanding discs for scarf sanding) Start sanding with 80 grit abrasive paper and finish with 150 grit abrasive paper. Remove the remaining 2.5 diameter center disk after scarf sanding is complete. (The scarf boundary extends from the 2.5" inner diameter to the 6.5" outer diameter. Using the same procedure, abrade the area extending from the 6.5" inner diameter to the 7.0" outer diameter) WARNING: WHILE SANDING, IT IS MANDATORY TO WEAR PROTECTIVE GLOVES, SAFETY GLASSES, DISPOSABLE LAB COATS AND DUST MASKS FOR PROPER PROTECTION FROM SKIN CONTACT WITH DUST AND/OR DUST INHALATION CL Symmetric Core Fill (4 places) 3/16" cell, 8 pcf PNL 45 PNL 0 PNL Ø6.50 Ø Figure 3. Scarf Sanded Panel Ready for Repair 6 Remove all sanding debris using oil free compressed filtered air and a vacuum cleaner. Clean per ARP 4916 [15] Method 5 two cloth method, using approved solvents and wiping media per AMS 3819C [17] that will not leave a residue. Please take a photograph of the scarfed cleaned panel. 7 Use a ruler/ straight edge to ensure adequate core height with respect to the inner scarf boundary. Use a feeler gage (or equivalent) to measure the depth of the core with respect to the inner scarf boundary. Use filler plies if the depth of the core is lower than that of the inner scarf boundary. Do not use more than 3 filler plies. (The facesheet cured ply thickness is 0.008") 5

6 8 Mask the core per ARP 4916 and perform the water break test after cleaning using ARP 4916 section Envelope bag the panel and dry per ARP 4977 [18] section 4.2 under vacuum at 180 F for 2 hours minimum. 10 Final Clean per ARP 4916 [15] Method 5 using approved solvent and wiping media per AMS 3819C Class 1 only [17]. Record the time the final cleaning was completed. Repair application MUST be conducted minutes after final cleaning. WARNING: A CONTAMINATED/ INADEQUATELY PREPARED SCARFED SURFACE WILL YIELD A DEFECTIVE REPAIR. BE ABSOLUTELY CERTAIN TO REMOVE PREPREG AND ADHESIVE BACKING AND SEPARATOR FILMS. FAILURE TO REMOVE BACKING FILM WILL RESULT IN A DEFECTIVE REPAIR. 11 Layup the repair patch as follows: 2 layers of EA9695 NW 0.05 psf adhesive (7.0 diameter) Repair Ply 1: [0/90], 3.5 diameter Repair Ply 2: [±45], 4.5 diameter Repair Ply 3: [±45], 5.5 diameter Repair Ply 4: [0/90], 6.5 diameter Please take photographs of the adhesive and ply lay-up. 12 Following the guidelines of ARP 4916 methods 4 or 5, use masking tape [15] to cover 2 periphery around the repair. 13 Install eight thermocouples per repair per ARP 5144 [19] and Envelop bag using the no bleed method per ARP 5143 [20]. For proper heat application, the heat blanket must be larger than the repair area by a minimum of 2 on all repair sides. Use a second blanket to heat up the bottom facesheet of the panel. Perform a leak test to check the integrity of the bag, the leak rate should not exceed 5in Hg in 5min. Record the vacuum leak rate. Apply heat and cure per ARP Use two heat blankets for all repairs and ensure that the areas around and above the heat blanket are properly insulated to reduce heat losses. Ensure that the hot bond vacuum assisted machine and the thermocouples are calibrated according to the requirements of ARP The recommended cure cycle is shown in Table 2. Discontinue the cure if the vacuum level falls below 16 in Hg during the cure. Table 2. Cure Cycle for CACRC Prepreg Repair Material Cure Parameters Units Requirements Units Requirements Heating Rate C/min 1-3 F/min 2-5 Cure Temperature C F Cure Time min min Cure Pressure bar >0.75 in-hg >22 Cooling Rate C/min 5 max F/min 9 6

7 14 Inspect using tap testing, MIA, PE/TTU Ultrasonics per ARP 5089 [21]. CACRC-R2 WET LAY-UP REPAIR PROCEDURE 1 Inspect the materials contained in the wet lay-up repair kit to ensure that all materials required for repair have been provided. Materials Dry Carbon Fabric, supplied in a sealed moisture proof bag. The fabric can be stored, in a sealed dry plastic bag, for a minimum of two years from the manufacture date. Epocast 52A/B laminating resin, 1 quart kit supplied; The resin can be stored for six months at room temperature. Please refer to the actual kit out time and shelf life as specified in each package (out time for material handling and kit preparation already accounted for). 2 Prepare the panel (PNL) for repair by identifying the panel 0 degree direction (0 degree direction along the panel length). Mark panel orientation as 0PNL, 90PNL, 45PNL as shown in figure 4. All repairs MUST be conducted on facesheet 2 (FS2) of the panels. The panel 0 direction is aligned with the panel 48" length. 3 Mask an 8" diameter area at the center of the panel in preparation for scarf sanding. Use the guidelines of ARP 4916 methods 4 or 5 using masking tape [15]. Using a mylar template (or equivalent), mark the inner 2.5" diameter scarf circle/ boundary and the outer 6.5" diameter scarf periphery. Please take a photograph of the prepared marked panel ready for sanding including a 12-inch scale for reference. 4 Scarf sand from an inner diameter of 2.5" to an outer diameter of 6.5" per AIR 5367 [16] sections 5.0 and 5.1 using a 0.5" inch per ply overlap. (Use an air powered die grinder, or equivalent, with aluminum oxide sanding discs for scarf sanding) Start sanding with 80 grit abrasive paper and finish with 150 grit abrasive paper. Remove the remaining 2.5 diameter center disk after scarf sanding is complete. (The scarf boundary extends from the 2.5" inner diameter to the 6.5" outer diameter. Using the same procedure, abrade the area extending from the 6.5" inner diameter to the 7.0" outer diameter) WARNING: WHILE SANDING, IT IS MANDATORY TO WEAR PROTECTIVE GLOVES, SAFETY GLASSES, DISPOSABLE LAB COATS AND DUST MASKS FOR PROPER PROTECTION FROM SKIN CONTACT WITH DUST AND/OR DUST INHALATION 5 Remove all sanding debris using oil free compressed filtered air and a vacuum cleaner. Clean per ARP 4916 [15] Method 5 two cloth method, using approved solvents and wiping media per AMS 3819C [17] that will not leave a residue. Please take a photograph of the scarfed cleaned panel. 6 Use a ruler/ straight edge to ensure adequate core height with respect to the inner scarf boundary. Use a feeler gage to measure the depth of the core with respect to the inner scarf boundary. Use filler plies if the depth of the core is lower than that of the inner scarf boundary. Do not use more than 3 filler plies. 7

8 CL Symmetric Core Fill (4 places) 3/16" cell, 8 pcf PNL 45 PNL 0 PNL Ø7.50 Ø Figure 4. Scarf Sanded Panel Ready for Repair (last 7.5 diameter ring is abraded for the extra ply and not actually scarfed) 7 Perform the water break test after cleaning using ARP 4916 section Envelope bag the panel and dry per ARP 4977 [18] section 4.2 at 180 F under vacuum for at least 2 hours. Please report the method used for drying. 9 Final Clean per ARP 4916 [15] Method 5 using approved solvent and wiping media per AMS 3819C Class 1 only [17]. Record the time the final cleaning was completed. Repair application MUST be conducted minutes after final cleaning. WARNING: A CONTAMINATED/ INADEQUATELY PREPARED SCARFED SURFACE WILL YIELD A DEFECTIVE REPAIR. 10 To determine the resin weight required, determine the amount of fabric needed based on the number/size of repair plies required. Please account for any additional filler plies if necessary. Weigh the fabric and determine the amount of resin required for the repair according to ARP 5256 [22]. Mix the Resin per ARP 5256 as illustrated in figure 5. The mixing ratio for the Epocast 52 A/B is as follows: to 100 parts by weight of Epocast 52-A, add 41 parts by weight of Epocast 52-B. Record the time resin mixing is started. Mix both components for five minutes until a homogeneous mixture is obtained. Figure 5. Epocast 52A/B Resin Mixing 8

9 WARNING: THE REPAIR MUST COMPLETED WITHIN 80% OF RESIN POT LIFE TO ENSURE ADEQUATE TIME FOR LAMINATE CONSOLIDATION. 11 Impregnate the repair Plies and apply per ARP 5319 [23] section 3.2 using Epocast 52 A/B resin per AMS 2980/4A [13] and Tenax HTA tex f3000t0 fiber per AMS 2980/3A [12] as illustrated in figure 6. Cut the plies and prepare for repair application. Figure 6. Dry Fiber Impregnation with Epocast 52A/B Resin Record the repair start time. Match the parent laminate ply orientations and add one extra ply in the dominant direction. Repair Ply 1: [0/90], 3.5 diameter Repair Ply 2: [±45], 4.5 diameter Repair Ply 3: [±45], 5.5 diameter Repair Ply 4: [0/90], 6.5 diameter Repair Ply 5: [0/90], 7.5 diameter, this is the extra ply Please take photographs of the ply lay-up. 12 Following the guidelines of ARP 4916 methods 4 or 5, use masking tape [15] to cover 2 periphery around the repair. 13 Install eight thermocouples per repair per ARP 5144 [19] and Envelop bag using vertical bleed method ARP 5143 [20]. Use a second blanket to heat up the bottom facesheet of the panel. Bagging should occur at maximum 80% of the resin pot life. Perform a leak test to check the integrity of the bag, the leak rate should not exceed 5in Hg in 5min. Record the vacuum leak rate. Please take photographs of the complete bag and thermocouple assembly. Peel ply: Airtech Nylon 62g/m 2, release ply B Bleeder: One Ply of style 120 glass fabric 9

10 Perforated Film: airtech perforated release film P (FEP, hole diameter 1.143mm, 1.27% open area). 14 Apply heat and cure per ARP 5144 [19]. Use two heat blankets for all repairs and ensure that the areas around and above the heat blanket are properly insulated to reduce heat losses. Ensure that the hot bond vacuum assisted machine and the thermocouples are calibrated according to the requirements of ARP The recommended cure cycle is shown in table 3. Discontinue the cure if the vacuum level falls below 16 in Hg during the cure. Table 3. Cure Cycle for CACRC Wet Lay-Up Repair Material Cure Parameters Units Requirements Units Requirements Heating Rate C/min 1-3 F/min 2-5 Cure Temperature C 93 C -103 C F 200 F-217 F Cure Time min min min min Cure Pressure bar >0.75 in-hg >22 Cooling Rate C/min 5 max F/min 9 max 15 Inspect using tap testing, MIA, PE/TTU Ultrasonics per ARP 5089 [21]. Representative prepreg and wet lay-up repairs are shown in figures 7 and 8. Figure 7. CACRC Repaired Element (CACRC Prepreg Repair) Figure 8. CACRC Repaired Element (CACRC Wet Lay-up Repair) 10

11 CACRC DEPOT REPAIRS/ DEPOT MECHANIC SURVEYS A total of 13 depot mechanics participated in the study either performing the repairs or participating in the CACRC surveys. Preliminary data summary gathered is summarized in table 4 and provides information on the mechanics training, years of experience, number of repairs performed during that time and the estimated repair rework rate. 70% of the mechanics have an airframe (or an Airframe & powerplant license) license and all mechanics received on the job training. The CACRC standards were provided to the repair mechanics along with repair checklists for each one of the repaired beams. The checklists provided contained part specific information such as stacking sequence, material details, etc. The checklists referenced the CACRC repair techniques for taper sanding, part masking, solvent cleaning, bagging, drying, etc... Some of the CACRC standards were difficult to interpret, had missing or incomplete information as well as outdated nomenclature ("mushroom sanding disk holder"). From the surveys conducted and the discussions with depot personnel, there was a general consensus on the need for teamwork and repair material and procedure standardization. More accessibility to engineering documentation and data is needed as well as training with OEM documents and SRMs, training to a particular repair manual due to the differences between the various OEM repair manuals but also the differences between the repair manuals used for different parts. Depot personnel may spend years to become familiar with a particular SRM. There was an overall consensus for the need for material standardization for improved repair process efficiency. Most of the time spent preparing for a repair is used to interpret the SRM, search for materials, prepare the tooling required whereas performing the repair itself only takes a fraction of the total time. Furthermore, it should be noted that many repairs are performed on similar parts at an OEM, whereas at an airline depot, a mechanic may only repair a given part occasionally (practice/training needed on the same part). In addition, in a depot environment, the repairs must be performed within a given timeframe in the case of AOGs (Aircraft on Ground). Some recommended topics in training include a brief overview of the history of composites in airframe structures, working/ training on example parts, composite part identification (know what to look for, material type, style ), computer training for depot personnel to access SRMs and find required documentation, understanding the differences between wet lay-up and prepreg repairs, cure temperature and the results if the part is either overcured or undercured, examples of bad processes and the consequences on the repair performance, repair pass-fail criteria, etc. Training should stress the importance of all repair steps and the potential implications on safety if inadequate processes are used. 11

12 Table 4. CACRC Depot Repair Mechanic Experience Mechanics Company Certification/ Qualification Program Years of Experience Number of Repairs Performed Rate of Rework Mechanic 1 OJT, OEM fiberglass class Worked on metals initially 23 years working on AOG 5000 repairs, 60% wet lay-up, 40% prepreg repairs Less than 10% Mechanic 2 OJT, Operator basic course Minimal Undergoing Training -- Mechanic 3 OJT, Operator basic course 16 years of experience with composites Mechanic 4 OJT, Operator Composite Classes 15 years of experience in composites Mechanic 5 Mechanic 6 Mechanic 7 Mechanic 8 Mechanic 9 OJT, 2 classes 1 week each Basic Composites I/II 3 years in composites OJT, Operator basic Composite Course (40 hours)/ Advanced Course (40 hours), OEM composite class (120 hours) OJT, operator general composites course (3 days) and advanced composites course (5 days) OJT, operator basic course 5 days, advanced course 5 days, Advanced Composites hands on course 1 week OJT 20 years of experience in composites 24 years of experience in composites 13 years of experience in composites 10 years in aircraft industry, 3.5 years of experience in composites early in career Mechanic 10 OJT 2 years of experience in composites Mechanic 11 OJT 3 years of experience in composites 780 repairs Mechanic 12 Mechanic 13 OJT OJT 20 years of experience in aviation, 10 years of experience in composites 24 years of experience in aviation, 15 years of experience in composites 700 repairs, 40% wet lay-up, 60% prepreg repairs 1700 repairs, 50% wet lay-up, 50% prepreg repairs 500 repairs, 60% wet lay-up, 40% prepreg repairs 4000 repairs, 67% wet lay-up, 33% prepreg repairs 2500 repairs, 10% wet lay-up, 90% prepreg repairs 3500 repairs, 50% wet lay-up, 50% prepreg repairs 72 repairs, Over 95% wet lay-up repairs 310 repairs, Over 95% wet lay-up repairs 2000 repairs 1800 repairs, 45% wet lay-up, 55% prepreg repairs -- Less than 1% rework -- Less than 1% rework Less than 5% rework Less than 5% rework -- Minimal Less than 10% rework Less than 5% rework Less or equal 2% 12

13 CONCLUSIONS The fundamental objective of this research work is to evaluate the adequacy of the existing CACRC standards for composite repair and to identify areas of improvement, if any, to ultimately yield robust repair procedures and standardized techniques that can be used across OEMs, airlines and repair stations. From the surveys conducted to date, the feedback received from depot personnel addresses a growing need for composite repair technician training and certification as well as periodic certification validation to ensure the structural integrity of bonded repairs due to the complexity and the skill set required for proper implementation of bonded repairs to composite substrates. 13

14 REFERENCES 1 Tenney, D. R. Davis, J. G., Jr., Pipes, R. B., and Johnston, N., NASA Composite Materials Development: Lessons Learned and Future Challenges, NASA Report LF , Tomblin, J. et. al.,"cacrc Depot Bonded Repair Investigation," FAA Joint Advanced Materials and Structures Center of Excellence 8th annual review meeting, Tomblin, J. et. al.,"cacrc Depot Bonded Repair Investigation," FAA Joint Advanced Materials and Structures Center of Excellence 7th annual review meeting, SAE AMS 3970/1a "Carbon Fiber Fabric Repair Prepreg, 120 C (250 F) Vacuum Curing - Part 1 - General Requirements SAE AMS 3970/2a "Carbon Fiber Fabric Repair Prepreg, 120 C (250 F) Vacuum Curing - Part 2 - Qualification Program for Fiber, Fabric, Carbon Prepreg, Film adhesive and Non- Structural Glass prepreg SAE AMS 3970/3a "Carbon Fiber Fabric Repair Prepreg, 120 C (250 F) Vacuum Curing - Part 3 - Purchasing Specification for Carbon Prepreg SAE AMS 3970/4a "Carbon Fiber Fabric Repair Prepreg, 120 C (250 F) Vacuum Curing - Part 4 - Purchasing Specification for Film Adhesive SAE AMS 3970/6 "Carbon Fiber Fabric Repair Prepreg, 120 C (250 F) Vacuum Curing - Part 6 - Carbon Fiber Fabric Reinforced Epoxy Prepreg for Repair, Plain Weave, 193 gm/m2, Adhesive film for repair, non structural glass fiber fabric reinforced epoxy prepreg, SAE AMS 2980 Technical Specification: Carbon Fiber Fabric and Epoxy Resin Wet Lay-Up Repair Material Part 0 - Introduction 10 SAE AMS 2980/1A Technical Specification: Carbon Fiber Fabric and Epoxy Resin Wet Lay-Up Repair Material Part 1 - General Requirements 11 SAE AMS 2980/2A Technical Specification: Carbon Fiber Fabric and Epoxy Resin Wet Lay-Up Repair Material Part 2 Qualification Program 12 SAE AMS 2980/3A Technical Specification: Carbon Fiber Fabric and Epoxy Resin Wet Lay-Up Repair Material Purchasing Specification - Fabric 13 SAE AMS 2980/4A Technical Specification: Carbon Fiber Fabric and Epoxy Resin Wet Lay-Up Repair Material Purchasing Specification - Resin 14 SAE AMS 2980/5 Carbon Fiber Fabric and Epoxy Resin Wet Lay-Up Repair Material Part 5 Material Specification: Carbon Fiber Fabrics, Plain Weave, 193g/m 2 and epoxy 15 SAE ARP 4916 Masking and Cleaning of Epoxy and Polyester Matrix Thermosetting Composite Materials 16 AIR 5367 Machining of Epoxy and Polyester Matrix Thermosetting Composite Structures 17 SAE AMS 3918C Cloths, Cleaning For Aircraft Primary and Secondary Structural Surfaces 18 SAE ARP 4977 Drying of thermosetting composites 19 SAE ARP 5144 Heat Application for Thermosetting Resin Curing 20 SAE ARP 5143 Vacuum bagging of thermosetting composite repairs 21 SAE ARP 5089 Composite Repair NDI/ NDT Handbook 22 SAE ARP 5256 Mixing Resins, adhesives and potting compounds 23 SAE ARP 5319 Impregnation of Dry Fabric and Ply Lay-Up 14