Development of Carbon Fibre Metal Laminates (CFML): Design, Fabrication and Characterisation

Abstract Fibre Metal Laminates (FMLs) are hybrid materials consisting of metal layers bonded to fibre-reinforced polymer layers. CFML is an FML developed at NAL consisting of thin aluminum foil combined with carbon-epoxy and glass-epoxy prepreg materials. CFML is proposed as the candidate material for the leading edges of wing and empennage of an aircraft as it has superior characteristics in terms of shape retention (due to highly linearly elastic material like carbon/epoxy), energy absorption capability (due to layered structure and plastic deformation), lightning protection (due to the presence of aluminum layers), and also due to its cost effectiveness (lightweight construction and simple production techniques). An important design issue is the internal residual stresses built into the laminate during curing due to differential coefficients of thermal expansion of the different material systems. This report discusses the methods and issues involved in the fabrication of CFML. CFML laminates were fabricated and Tensile, Compression, ILSS and Flexure testing of standard specimens for different lay-ups were done. The failure modes exhibited during these tests indicate that these materials could be better in energy absorption. However, these conclusions need to be validated with the experiments to quantify their energy absorption capability

Development of CGLARE: Design, Fabrication and Characterisation

Fibre Metal Laminates (FMLs) are hybrid materials consisting of metal layers bonded to fibre-reinforced polymer layers. CGLARE is an FML developed at NAL consisting of thin aluminum foil combined with carbon-epoxy and glass-epoxy prepreg materials. CGLARE is proposed as the candidate material for the leading edges of wing and empennage of an aircraft as it has superior characteristics in terms of shape retention (due to highly linearly elastic material like carbon/epoxy), energy absorption capability (due to layered structure and plastic deformation), lightning protection (due to the presence of aluminum layers), and also due to its cost effectiveness (lightweight construction and simple production techniques). This paper describes the issues regarding the development of CGLARE such as surface preparation of aluminum foils and bonding of aluminum with glass. Tensile, Compression, ILSS and Flexure testing of ASTM standard CGLARE specimens for different layups have been done. An important design issue is the internal residual stresses built into the laminate during curing due to differential coefficients of thermal expansion of the different material systems. The paper presents these results that indicate some properties of these material systems that could be exploited for energy absorption in the leading edges of the aircraft

NALLA-2 : Strain measurements

Strains were monitored for NALLA2 Structural components at critical locations : 60 locations for wing, 50 for fuselage, 6 each for horizontal tail, vertical tail, landing gear and nose gear and 20 locations for wing fuselage joint. While the wing fuselage joint was tested upto the ultimate load, the rest of the components were loaded upto the limit load only. The measured strains have proved the linear behaviour upto the limit Load, with zero yield, validating the structural integrity, of the components

Experimental Characterisation of GLass Aluminum REinforced (GLARE™) laminates

Fibre metal laminates such as GLARE™ have found promising application in the aerospace industry. These laminates were developed at the structures and materials laboratory of Delft University of Technology, Netherlands. GLARE™ is a material belonging to the family of Fibre Metal Laminates consisting of thin aluminum layers bonded with unidirectional S2-Glass fibres with an adhesive. Aluminum and S2-Glass when combined as a hybrid material can provide best features of the both metals and composites. These materials have excellent fatigue, impact and damage tolerance characteristics and a lower density compared to aluminum. GLARE™ has found major application in front and aft upper fuselage, leading edges of empennages of advanced civil aircrafts like A380. This document looks into the evaluation of two configuration of GLARE™ for its mechanical and impact characteristics. The mechanical characterisation was carried out for tensile, compression, Flexure, ILSS, Open Hole Tension, Open Hole Compression and Shear (Iosipescu). The impact behaviour were characterised based on a low velocity drop weight impact carried on these laminates. The study shows that the basic properties evaluated were more dictated by the property of the S2-Glass used. The studies show that GLARE™ laminates posses’ high impact damage resistance compared to other composite material. All the test datas generated for this study will be brought out in this document

Mechanical properties of carbon-epoxy composites

This report presents the mechanical properties of carbon epoxy composites which were established for use in the design of 1:1.405 scale LCA high speed air intake model. The carbon fabrics and resin used in impregnation of specimens were G803W 5H satin weave and G807 8H satin weave supplied by M/s Brochier SA France, and Epoxy LY556 with Hardner HY951 supplied by CIBA (India). The mechanical properties evaluated included tensile strength, tensile modulus and horizontal shear strength (short beam method). The effect of post curing at 70°C was also studied

Non Destructive Evaluation of MiG-21 Compsoite Rudder Part II - APT Section Adhesive bond qulality Evaluation

This report contains the results of adhesive bond quality evaluation in the aft section of MiG-21 composite rudder using Fokker Bond Tester. Complete test recordings in respect of four different rudders are included in this report

Fractographic Analysis of Unidirectional Carbon Fiber Reinforced Plastic (CFRP) Composite Laminates Failed under Compression Loads

This paper describes fractographic features observed in unidirectional CFRP laminates failed under compression loads. The morphology of fracture surfaces was studied using scanning electron microscope. Typical fracture modes are presented. It has been demonstrated that macroscopy coupled with fractography yield useful data from which correlation of macro/microscopic details with loading conditions could be established. Detailed study was conducted to map the crack propagation direction

Non-destructive evaluation of co-cured wing for SARAS

After considerable experience with the development of pre-preg based (mostly Fiberdux T300-914C) autoclave13; moulded carbon epoxy composite airworthy components for both Tejas (LCA) and SARAS programmes, the Advanced Composites Division (ACD) has now come out with an ingenious Vacuum Enhanced Resin Infusion Technology (VERITy) route for the development of Co-cured Carbon Wing for the SARAS aircraft using Carbon Fabric and Epoxy resin. Non-Destructive Evaluation (NDE) has played a crucial role in the formulation of the VERITy process. Also, NDE will be used in the qualification of the wing components and an approach has been evolved

NON- Destructive Evaluation of Mig 21 composite Rudder

This report presents the preliminary studies carried out in connection with the setting up of a procedure, using Fokker Bond Tester, for the Non-Destructive Evaluation of adhesive bonding in MiG-21 Composite Rudder Aft Section

Damage Growth Studies on Composite Flap Structures Under Fatigue Loading

The rib-skin construction using carbon fiber composites has been an attractive design option for aircraft control surfaces like aileron, elevator, rudder, flap etc. The major concern in such structures is the debonding between skin and rib flange which can occur due to low velocity impacts like tool drops, run way debris etc. Such debonds which occur in service are barely visible and may not get detected till the next inspection schedule. The integrity of the structure in the intervening period is of great concern to the designers. In the present work, the structural integrity of a composite flap structure having multiple debonds at the rib skin interfaces under fatigue loading has been addressed. The flap is subjected to cyclic loading at design limit load for 110000 cycles using a whiffle tree mechanism. The strains have been monitored at different locations to understand the behaviour of structure during the test using strain gauges and Fiber Bragg Grating (FBG) sensors. Ultrasonic A-scan was used to monitor the defect growth after each block of 1000 cycles. The growth of debonds was not significant during the fatigue testing. The strain levels did not change appreciably throughout the test period indicating the damage tolerance capacity of the flap structure. The low growth of debonds was attributed to the low level of strains in the structure since the flap design is driven by stiffness considerations