The Experts below are selected from a list of 1539 Experts worldwide ranked by ideXlab platform
U.e. Enemuoh - One of the best experts on this subject based on the ideXlab platform.
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Design, analysis and performance of adhesively bonded composite Patch repair of cracked aluminum aircraft panels
Composite Structures, 2005Co-Authors: A. Chukwujekwu Okafor, Navdeep Singh, U.e. EnemuohAbstract: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.
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Design, analysis and performance of adhesively bonded composite Patch repair of cracked aluminum aircraft panels
Composite Structures, 2005Co-Authors: A. Chukwujekwu Okafor, Navdeep Singh, U.e. EnemuohAbstract: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.
Hari Bhogapurapu - One of the best experts on this subject based on the ideXlab platform.
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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, 2006Co-Authors: Chukwujekwu A Okafor, Hari BhogapurapuAbstract: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.
Navdeep Singh - One of the best experts on this subject based on the ideXlab platform.
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Design, analysis and performance of adhesively bonded composite Patch repair of cracked aluminum aircraft panels
Composite Structures, 2005Co-Authors: A. Chukwujekwu Okafor, Navdeep Singh, U.e. EnemuohAbstract: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.
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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, 2006Co-Authors: Chukwujekwu A Okafor, Hari BhogapurapuAbstract: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.