Development of Damage Tolerant Adhesive Bonded Repair

Transcription

1 Development of Damage Tolerant Adhesive Bonded Repair More Info at Open Access Database V. Srinivasa, Ramesh Kumar. M, Suresh Chand Jangir, Saji. D, Pitchai. P, Sudheendra. H.N Advanced Composites Division, CSIR-National Aerospace Laboratories, Bangalore , India Key word: Damage tolerance, Composite adhesive bonded repair, Ultrasonic inspection, joint Abstract: A very high percentage of composites being used in military and civil aircraft structures keep increasing due to the properties like high strength and low weight of the material. During the service life of the aircraft the composite parts are prone to damages while taxing, take off & landing and in the process of flight. The damages occur due to the above problem needs repair to continue the operation without major downtime. Damage tolerant repairs are adopted for composite parts manufactured at the works of the aircraft structures. The same approach of repair can be followed at the hanger to avoid the downtime of the aircraft. A well designed adhesive bonded joint repair will have almost same strength of the adjoining structure. Damage tolerance of adhesive bonded repair is a concern to be addressed in the case of in-situ repair scheme at the hanger. In this paper an attempt is made to develop damage tolerant bonded repair and evaluate its strength through destructive and non- destructive testing. Typical damage tolerant bonded repair are step lap repair and repairs followed by two different fabrication approach using carbon fibre reinforced plastics (CFRP) adherent used with epoxy based adhesive. The bonded repairs on laminates are evaluated by Ultrasonic Non-Destructive Evaluation (NDE) technique. Repair joint patch laminate strength are determined by static tensile testing and compared with the damaged laminate. To improve the joint more damage tolerant external cover plies were laid over the joint and the recovery strength was evaluated by applying load on the repair laminates. With the addition of two external cover plies the strength of the repair was increased by 12% compared with the artificially made damage laminate. Introduction: Composite are being used as primary, secondary or tertiary composite parts for aircraft structures in large numbers for both civil and military program. Autoclave moulding fabrication technology is widely adopted in aircraft industries for composite structures by co-curing method. Advanced Composites Division (ACD) is vigorously pursuing co-curing technology activity for a number of composite components/assemblies for both military and civilian aircraft development projects. Cocuring methodology has a number of advantages, including consolidation of parts to form an assembly in a single cure cycle. This minimizes the use of fasteners, eliminates stress concentration due to drilled holes, reduces assembly time and costs, and enhances the production rate. Fabrication activity has been vigorously pursuing co-curing technology, and a number of composite components/assemblies have been developed for both military and civilian aircraft. At the same time co-curing method of fabrication possesses challenges to Non Destructive Testing and Evaluation (NDT&E) of components and approve for assembly. During the service life of the aircraft these co-cured parts may go damages due to the runway debris, damages due to impact from the neighboring aircraft, bird impact and due to human negligent (tool drop) during servicing of an aircraft [1,2]. These damages must be repaired immediately in the field to avoid downtime period of an aircraft during the operation of the flight in the indent usage. In view of this the field repair is considered as most important and stand the necessity of the repair instantly

2 which meet the strength requirements, easy of manufacturing, inspection and at low cost [ 3]. At a later stage the permanent repair can be made at the OEM works, integral to the parent material with suitable design, testing and validation [4,5]. Experiment: Manufacturing of Repair laminates: The higher demand for the field repair of composite parts in both military and commercial aircraft makes it compulsory to use only proven composite field repair techniques to restore the structural integrity with additional weight to minimum. The damages are also sometimes critical that it needs attention immediately, in such case of field repair of these advanced composite materials requires an in-depth knowledge of composite structures and availability of repair materials, tooling, adhesives, etc for the repair. The restoration of structural integrity during the field repairs can be attained only from a properly designed repair [5]. In order to validate a field repair scheme, design database is required. The design database is evolved based on number of tests at the coupon, feature or component level [7,8]. The fabricated laminates for test purpose are damaged with a 50 mm diameter hole at the center. The damaged laminates are then repaired with the repair in two different methods one is using wet layup of carbon fabric and other using carbon Prepreg plies. Make a taper of ratio 1:20 or 3 taper around the cutout by machining and then prepare the surface. On the surface prepared, lay up the repair patches using the layup fixture and allow the repair to cure. After curing the laminates were inspected using non destructive evaluation[6]. Fig. 1: Schematic of repair laminates Fig. 2: Scarf region with 3 Taper machined in-situ on composite panel by special tool - ready to bond the patch laminate The following most commonly used field repair techniques on composites structures like Scarf, External bonded repairs by using various materials and matrix has been studied.

3 Table 1 Fabrication details of the repair laminates Base laminate (3mm thick) Carbon UD Prepreg Hexply 914/34%/UD160/ AS4/12K-300mm Carbon UD Prepreg Hexply 914/34%/UD160/ AS4/12K-300mm Repair joint Repair patch Carbon UD Prepreg Hexply 914/34%/UD160/AS4/ 12K-300mm (Lay-up in situ) Carbon Fabric TWILL 2X2, mm t (Lay-up in situ) Carbon UD Prepreg Hexply 914/34%/UD160/AS4/ 12K-300mm Carbon UD Prepreg Hexply 914/34%/UD160/AS4/ 12K-300mm Bonding glue/resin Redux- 319A Ly5052+C H 5052 Wet lay-up Redux- 319A AV138M HV998 Curing 175 c vacuum 80 c Vacuum 175 c vacuum 60 c Vacuum Non Destructive Evaluation: The Non Destructive Testing and Evaluation (NDT&E) of the repaired panels have been carried out using water squirter Ultrasonic technique before repair and after repair along with the adhesive glue used in bonding the patch laminate. Before bonding the patch laminate was also inspected and evaluated. Since the patch laminate was manufactured by two methods the bond quality was varied among the manufactured repair laminates. The Ultrasonic inspection parameters adopted was Ultrasonic water squirter C-scan technique using 5.0 & 2.5MHz transducers. The laminates were mounted in perpendicular to the water squirter and the data was acquired and results were analyzed and evaluated. Results and Discussions: Case-1: Many repair patch laminates were fabricated the results of few repair patch laminates were discussed as examples. Panel-01 was tested using both the inspection frequencies of 2.5 & 5.0MHz with Ultrasonic water squirter C-scan system and the results are shown in Fig. 3. In the below figure the first plot shows the result of 2.5MHz frequency on the repair region (between inner diameter and outer diameter) was having good bonding for a length of 10mm from outer diameter of the repair patch. The second plot in Fig.3 shows the c-scan result of 5.0MHz for the same repair laminate, the variation observed in the bonded region is due to the higher frequency and the wavelength effect in the region i.e the sensitivity is high in this frequency for the thin laminate

4 configuration. Variation noticed in the bonded regions are due to repair patch laminate was fabricated by joint using wet / hand layup scheme has very low consolidation compared to the parent laminate. Drilled region of 50mm diameter Repair patch bonded region Fig. 3: C-scanning of Panel-01 at 2.5 & 5.0MHz Case-2: In the case of patch laminate was manufactured using pre-preg and cured in oven were bonded to the parent laminate joint. The adhesive used for bonding the patch with the parent laminate are provided in the above table 1 and the repair procedure are similar in all the cases. The bonding was made leaving the centre region of 50mm diameter (ID) till 170mm diameter (OD) of the repair laminate. In this case the bonding was good when compare with the case-1 in both the frequency of inspection. The results are shown in Fig. 4. Fig. 4: C-scan plots of Panel-02 at 2.5 & 5.0MHz Case-3: Few repair patch laminates were fabricated and the results of repair patch laminates were discussed. Panel-03 was tested using both the inspection frequencies of 2.5 & 5.0MHz with Ultrasonic water squirter C-scan system and the results are shown in Fig. 5. The plot shows the result of 2.5 and 5.0 MHz frequency on the repair region (between inner diameter and outer diameter) were having good bonding on the repair patch. Another c-scan image on the same figure shows the c-scan result of 5.0MHz of the same repair laminate. Variation noticed in the bonded regions are due to repair patch laminate was fabricated by external patch using wet / hand layup method has higher attenuation levels compared to the parent laminate.

5 Fig. 5: 2.5 & 5.0MHz C-scan image of Panel-03 Case-4: In the case of patch laminate was manufactured using pre-preg and cured in oven were bonded to the parent laminate joint. The adhesive used for bonding the patch with the parent laminate is same in all the cases and the repair procedure was also the same. Except the ID of 50mm diameter in the bonded region the bonding was good. The results are shown in Fig. 6. Fig. 6: Panel-03 C-scan image of 2.5 & 5.0MHz Conclusion : Adhesive bonded joints are widely used in aircraft structures, as they are not prone to stress concentration issues. Further they are effective in transferring loads through shear. Typical damage tolerant adhesive bonded joints are and step lap joint. Scarf joint was fabricated and tested for its static compressive strength and found that upto 80% of strength can be recovered through this repair procedure [9]. The effectiveness of damage tolerant repair was demonstrated through the introduction of cover plies. Scarf joint was evaluated by ultrasonic NDE and found that ultrasonic was an effective method for qualification of adhesive bonded repair. The virgin laminates were tested for strength and did not fail till 380KN load. The similar laminates having 50mm hole (considered as damage laminate) in the centre was tested and found that the laminate gave up at 320KN load. Similar strength test was conducted on the repaired patch laminate on the same machine and it with stand more than the damaged laminate (drilled hole laminate) i.e upto 330KN. This proves that the repair is stronger than the damaged laminate and it can withstand more than the damaged laminate. The field in-situ repair is having similar strength as that of the parent structures of the aircraft. References : 1. B.C Hoskin, A.A Baker, Composite for Aircraft Structures ; AIAA Education Series, Air Force Institute of Technology, Wright-Patterson Air Force Base, Ohio, 1986

6 2. Composite repair of military aircraft structures ; AGARD Conference proceedings, AGARD-CP-550, NATO, France, Keith B. Armstrong and Richard T. Barrett, Care and Repair of Advanced Composites ; Society of Automotive Engineers Inc., Warrendale, Pa, USA, Alan Baker, Francis Rose, Rhys Jones, Advances in the Bonded Composite Repair of Metallic Aircraft Structures ; Vol. 1 & 2, Elsevier, Recent Advances in Structural Joints and Repairs for Composite s ; Liyong Tong and Costas Soutis, Kluwer Academic Publisher, London, Teufel, P., M. Maxwell and R. Gardiner, Low Cost Composite for a General Aviation Aircraft Wing, AIAA/ICAS International Air and Space Symposium and Exposition, July 2003, American Institute of Aeronautics and Astronautics, Dayton, Ohio 7. R.D Adams, Adhesive bonding science, technology and applications ; Wood Head Publishing Limited, England, Introduction to Composite s Design ; Ever J. Barbero, CRC Press, Annual Book of ASTM Standards, Section 15, Vol & Section-3, Vol