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Antoine Gueydan - One of the best experts on this subject based on the ideXlab platform.

  • secondary creep stage behavior of copper Clad Aluminum thin wires submitted to a moderate temperature level
    Materials Science and Engineering A-structural Materials Properties Microstructure and Processing, 2018
    Co-Authors: Antoine Gueydan, Eric Hug
    Abstract:

    Abstract This work focuses on the role of the microstructure on the creep behavior of thin copper-Clad Aluminum (CCA) wires. Creep tests were performed at 150 °C on hard drawn and annealed CCA and on pure metals. It is shown that the Cu-Al interface of CCA ensures a mechanical resistance leading to lower creep velocities than for metals. Creep mechanisms are driven by Aluminum at lower stresses and copper for higher stresses, independently of the physical nature of the interface.

  • Secondary creep stage behavior of copper-Clad Aluminum thin wires submitted to a moderate temperature level
    Materials Science and Engineering: A, 2018
    Co-Authors: Antoine Gueydan, Eric Hug
    Abstract:

    This work focuses on the role of the microstructure on the creep behavior of thin copper-Clad Aluminum (CCA) wires. Creep tests were performed at 150 degrees C on hard drawn and annealed CCA and on pure metals. It is shown that the Cu-Al interface of CCA ensures a mechanical resistance leading to lower creep velocities than for metals. Creep mechanisms are driven by Aluminum at lower stresses and copper for higher stresses, independently of the physical nature of the interface.

  • study of the intermetallic growth in copper Clad Aluminum wires after thermal aging
    Intermetallics, 2014
    Co-Authors: Antoine Gueydan, B Domenges
    Abstract:

    Abstract Study of the solid-state diffusion between copper and Aluminum was carried out in the temperature range [573–673] K in order to better understand the aging mechanisms which occur in copper-Clad Aluminum thin wires. A complete microscopic analysis was performed to evaluate the interface composition and corresponding microstructure. The intermetallic phases developed during annealing identified by TEM and X-Ray diffraction analysis are respectively Al 2 Cu, AlCu, and Al 4 Cu 9 . A fine layer containing nanometric copper grains was also depicted and identified as a diffusion-induced recrystallization region. These results agree with EDXS analysis and nanoindentation measurements. The effective heat of formation model was used to evaluate the first phase(s) which happens in the interface and the sequence formation of intermetallic compounds during annealing. This model finely describes the metallurgical aging of copper-Clad Aluminum wires and explains the presence of only three intermetallic compounds in the interface between copper and Aluminum.

Eric Hug - One of the best experts on this subject based on the ideXlab platform.

U.e. Enemuoh - One of the best experts on this subject based on the ideXlab platform.

  • Design, analysis and performance of adhesively bonded composite patch repair of cracked Aluminum aircraft panels
    Composite Structures, 2005
    Co-Authors: A. Chukwujekwu Okafor, Navdeep Singh, U.e. Enemuoh
    Abstract:

    Abstract During its service life, an aircraft is subjected to sever structural and aerodynamic loads. These loads can cause damage or weakening of the structure especially for aging military and civilian aircraft thereby affecting its load carrying capabilities. Hence, a repair or reinforcement of the damaged or weakened part of the structure to restore the structural efficiency and thus assure the continued airworthiness of the aircraft has become an important issue in recent years to military and civilian aircraft operators. The US Air Force in recent years has shown considerable interest in the use of advanced composites to repair cracked metallic aircraft structures to enhance their life. One issue preventing using bonded composite patches, as a standard means of repairing damaged metallic aircraft structures is the fact that the integrity of the repairs is unknown. In this paper the design, analysis and durability of adhesively bonded composite patch repairs of cracked aircraft Aluminum panels is reported. Pre-cracked 2024-T3 Clad Aluminum panels of 381 × 89 × 1.6 mm (15 × 3.5 × 0.063 in.) repaired with octagonal single sided boron/epoxy composite patch were used as test specimen. Two different composite ply configurations, 5- and 6-ply were investigated. Linear and non- linear finite element analyses were performed on the test specimen using 8-noded 24 degree of freedom (DOF) hexagonal elements for the Aluminum panel, boron/epoxy patch and adhesive material subjected to uni-axial tensile loading. The stress distributions obtained were used to predict the increase in strength and durability of the repaired structure. A comparison of the stress values at critical points was made. The analysis also was used to validate various assumptions made in the design of the composite patch. Experimental investigations were conducted on the cracked Aluminum panel repaired with a 5-ply composite patch as well as on two baseline-unpatched panels (one with a crack and one with no crack) by uni-axial tensile testing to validate the analytical results. The experiment was conducted on the Instron tension-testing machine. It was found that the maximum skin stress decreases significantly after the application of the patch and the region of maximum skin stress shifts from the crack front for an unpatched panel to the patch edges for a patched one.

A. Chukwujekwu Okafor - One of the best experts on this subject based on the ideXlab platform.

  • Evaluation of the Effects of Corrosion on Fatigue Life of Clad Aluminum Alloy 2024-T3-Riveted Lap Joints with Acoustic Emission Monitoring
    Journal of Failure Analysis and Prevention, 2012
    Co-Authors: A. Chukwujekwu Okafor, Christopher Nnadili
    Abstract:

    Corrosion affects the fatigue life of Clad Aluminum alloy-riveted lap joints, such as those found on an aircraft fuselage structure. Single-, double-, and triple-column-riveted lap joint specimens were fabricated and corroded in a Q-Fog accelerated corrosion chamber for five months using an ASTM G85-A5 prohesion test. Specimens were taken out of the chamber every 4 weeks, and the corrosion products which had been deposited on them were removed by immersion in concentrated nitric acid. For each corroded specimen, the mass loss with corresponding corrosion rate was determined. The specimens were fatigue loaded to failure on an MTS Universal Testing Machine with acoustic emission monitoring. Results indicate that exposure of lap joint specimens to this corrosive environment increased corrosion (mass loss), corrosion rate, and significantly reduced fatigue life. For a prolonged exposure in the corrosive environment, the fatigue life was reduced to zero, which has significant implication for aging aircraft. Acoustic emission monitoring successfully detected fatigue failure. Two failure modes, multisite crack damage and shear of the rivets, were observed.

  • Design, analysis and performance of adhesively bonded composite patch repair of cracked Aluminum aircraft panels
    Composite Structures, 2005
    Co-Authors: A. Chukwujekwu Okafor, Navdeep Singh, U.e. Enemuoh
    Abstract:

    Abstract During its service life, an aircraft is subjected to sever structural and aerodynamic loads. These loads can cause damage or weakening of the structure especially for aging military and civilian aircraft thereby affecting its load carrying capabilities. Hence, a repair or reinforcement of the damaged or weakened part of the structure to restore the structural efficiency and thus assure the continued airworthiness of the aircraft has become an important issue in recent years to military and civilian aircraft operators. The US Air Force in recent years has shown considerable interest in the use of advanced composites to repair cracked metallic aircraft structures to enhance their life. One issue preventing using bonded composite patches, as a standard means of repairing damaged metallic aircraft structures is the fact that the integrity of the repairs is unknown. In this paper the design, analysis and durability of adhesively bonded composite patch repairs of cracked aircraft Aluminum panels is reported. Pre-cracked 2024-T3 Clad Aluminum panels of 381 × 89 × 1.6 mm (15 × 3.5 × 0.063 in.) repaired with octagonal single sided boron/epoxy composite patch were used as test specimen. Two different composite ply configurations, 5- and 6-ply were investigated. Linear and non- linear finite element analyses were performed on the test specimen using 8-noded 24 degree of freedom (DOF) hexagonal elements for the Aluminum panel, boron/epoxy patch and adhesive material subjected to uni-axial tensile loading. The stress distributions obtained were used to predict the increase in strength and durability of the repaired structure. A comparison of the stress values at critical points was made. The analysis also was used to validate various assumptions made in the design of the composite patch. Experimental investigations were conducted on the cracked Aluminum panel repaired with a 5-ply composite patch as well as on two baseline-unpatched panels (one with a crack and one with no crack) by uni-axial tensile testing to validate the analytical results. The experiment was conducted on the Instron tension-testing machine. It was found that the maximum skin stress decreases significantly after the application of the patch and the region of maximum skin stress shifts from the crack front for an unpatched panel to the patch edges for a patched one.

Chukwujekwu A Okafor - One of the best experts on this subject based on the ideXlab platform.

  • acoustic emission monitoring of tensile testing of corroded and un corroded Clad Aluminum 2024 t3 and characterization of effects of corrosion on ae source events and material tensile properties
    40TH ANNUAL REVIEW OF PROGRESS IN QUANTITATIVE NONDESTRUCTIVE EVALUATION: Incorporating the 10th International Conference on Barkhausen Noise and Micr, 2014
    Co-Authors: Chukwujekwu A Okafor, Shridhar Natarajan
    Abstract:

    Corrosion damage affects structural integrity and deteriorates material properties of Aluminum alloys in aircraft structures. Acoustic Emission (AE) is an effective nondestructive evaluation (NDE) technique for monitoring such damages and predicting failure in large structures of an aircraft. For successful interpretation of data from AE monitoring, sources of AE and factors affecting it need to be identified. This paper presents results of AE monitoring of tensile testing of corroded and un-corroded Clad Aluminum 2024-T3 test specimens, and characterization of the effects of strain-rate and corrosion damage on material tensile properties and AE source events. Effect of corrosion was studied by inducing corrosion in the test specimens by accelerated corrosion testing in a Q-Fog accelerated corrosion chamber for 12 weeks. Eight (8) masked dog-bone shaped specimens were placed in the accelerated corrosion chamber at the beginning of the test. Two (2) dog-bone shaped specimens were removed from the corrosion...

  • design and analysis of adhesively bonded thick composite patch repair of corrosion grind out and cracks on 2024 t3 Clad Aluminum aging aircraft structures
    Composite Structures, 2006
    Co-Authors: Chukwujekwu A Okafor, Hari Bhogapurapu
    Abstract:

    Abstract Many military and commercial aging aircrafts flying beyond their design life may experience severe crack and corrosion damage, and thus lead to catastrophic failures. In this paper, the design, fabrication and analysis of adhesively bonded thick composite patch repair of circular corrosion grind-out and a crack propagating on the periphery of the corrosion grind-out on thick 2024 T3 Clad Aluminum aircraft panel is presented. Thick orthogonal composite patch configurations of 7–25 plies were designed separately for crack and corrosion grind-out using CRAS. Using the principles of superimposition a single patch was designed to repair both the crack and corrosion grind-out. Finite element analysis (FEA) was performed on the test specimen subjected to uniaxial tensile loading. Stress distribution and displacements were obtained and analyzed. Dog-bone shaped tensile test panels were fabricated with damage and repaired with boron/epoxy patch of 11 plies. The patched and unpatched panels were subjected to tensile tests. The experimental and the FEA results show that the maximum skin stress decreases significantly and shifted away from damaged area after the application of composite patch. The load carrying capacity of patched specimen significantly increased over that for unpatched specimen.