Abstract: The strength of a welded truss type fuselage of a light aircraft – named SAFAT 01 – is considered in this paper. The aircraft is a monoplane with high strut-braced wings configuration with flaps. The fuselage is of welded tubular steel fabric-covered construction. According to its production contract; the aircraft is fully produced and assembled in Sudan whereas the documentation is limited to technical side only with no information available about design procedures and calculations. This makes it difficult to further modify or upgrade the aircraft. The fuselage geometry has been modeled using a CAD program. The main dimensions have been obtained using 2-D drawings and the missing dimensions and data due to lack of documents were extracted experimentally from a built aircraft. Aerodynamic loads were determined using computational fluid dynamic program for the horizontal tail. Static structural analysis was conducted by finite element method (FEM) using fastened connection property between the tubes. The results were observed in three main parts; rear part which supports the vertical and horizontal stabilizers and the rear landing gear; the mid part which provides cantilever reaction for the wing and supports front landing gear, and the front part on which the engine is mounted. The Von Mises stresses, displacements and principal stresses were observed in the three parts and found acceptable except for a small region near the attachment between wing and fuselage. However, further experimental validation is needed. Presently, experimental and dynamic analyses are being conducted and the results will be published later.
Abstract: Signal processing is an important element used for identifying damage in any SHM-related application. The method here is used to extract features from the use of different types of sensors, of which there are many. The responses from the sensors are also interpreted to classify the location and severity of the damage. This paper describes the signal processing approaches used for detecting the impact locations and monitoring the responses of impact damage. Further explanations are also given on the most widely-used software tools for damage detection and identification implemented throughout this research work. A brief introduction to these signal processing tools, together with some previous work related to impact damage detection, are presented and discussed in this paper.
Abstract: Due to discontinuity of mechanical properties in composite laminates, failure occurs in different damage mechanisms. Delamination growth of adjacent layers is a major failure mechanism in laminates with various layup configurations. Pre existing delamination may initiate in composite laminate before use, due to impact in assembly and fabrication process. Cyclic compressive loading may cause delamination growth due to both post-bucking behavior and fatigue nature of loading. In this paper, a 3D mixed-mode interface element model has been developed to simulate the growth of multiple delaminations under compressive cyclic loading. For this purpose, the presented model should be able to handle the geometry nonlinearity of post-buckling and material nonlinearity of cohesive zone constitutive law under cyclic loading at interfaces. Because of mixed-mode condition of stress field at the delamination-front of post-buckled laminates, a mixed-mode bilinear constitutive law has been used as user material in this model. Paris Law has been used to relate the energy release rate to the fatigue crack growth in cohesive zone. A composite laminate with pre-existing delamination under buckling load, available from the literature has been reproduced with the present approach. Finally, laminates containing multiple delaminations in various interface layers have been analyzed under compressive fatigue loading. It is shown that the pre-existing delamination with more depth from the surface of laminate causes more initial static and fatigue delamination growth rate.
Abstract: This paper aims to demonstrate the structure analysis of strut-braced wing of a typical manufactured Light-Aircraft by using FEM software (MSC PATRAN/NASTRAN) and determine the safety margin in all of its components, which are useful to determine the structure strength requirements. The geometrical model of the wing was created in CATIA and then exported to PATRAN, which is the modeler to build the finite element model. PATRAN model geometry was modified and prepared to create the mesh. The structural components have various functions and shapes, thus different element mesh was created. After creating the finite element model for all parts, the elements and material properties were assigned and the model was fixed at the spar root edge and strut-braced end, and loaded by distributing the inertia load and aerodynamic load, calculated using (CFD), acting on the rib edge. Then the model was submitted to NASTRAN for linear static analysis. The obtained Stress Results and Safety Margins of each part were calculated and found to be acceptable.
Abstract: This article focuses on the application of the Fourier-expansion based differential quadrature method (FDQM) for the buckling analysis of ring-stiffened composite laminated cylindrical shells. Displacements and rotations are expressed in terms of Fourier series expansions in longitudinal direction and their first order derivatives are approximated with FDQM in circumferential direction. The 'smeared stiffener' approach is adopted for the stiffeners modeling. Two FORTRAN programs prepared for linear and nonlinear analysis and results were compared by ABAQUS finite element software. Buckling loads of stiffened and unstiffened shells considering the effects of changes in shell and stiffener geometric and material properties and also shell lay-ups are investigated.
Abstract: The high velocity impact response of composite laminated plates has been experimentally investigated using a nitrogen gas gun. Tests were undertaken on fibre-metal laminate (FML) structures based on Kevlar-29 fiber/epoxy-Alumina resin with different stacking sequences of 6061-T6 Al plates. Impact testing was conducted using a cylindrical shape of 7.62 mm diameter steel projectile at 400m/s velocity, which was investigated to achieve complete perforation of the target. The numerical parametric study of ballistic impacts caused by similar conditions in experimental work is undertaken to predict the ballistic limit velocity, energy absorbed by the target, and comparisons between simulations by using ANSYS AUTODYN 3D v.12.1 software and experimental work to study the effects of the shape of the projectile with different (4, 8, 12, 16 and 20mm) thicknesses on the ballistic limit velocity. While only one thickness was used with 24mm of back stacking sequence, it was not penetrated. The sequence of the Al plate position (front, middle and back) inside laminate plates of the composite specimen was also studied. The Al back stacking sequence plate for the overall results obtained was the optimum structure to resist the impact loading. The simulation results obtained of the residual velocity hereby are in good agreement with the experimental results with an average error of 1.8%. The energy absorption was obtained with 7.3% and 2.7% of the back to front and back to middle of the Al stacking sequence respectively. Hence, the back Al stacking sequence is considered the optimum position for resisting the impact loading. The data showed that these novel sandwich structures exhibit excellent energy-absorbing characteristics under high-velocity impact loading conditions. Hence, it is considered suitable for aerospace applications.
Abstract: In this study, finite element method is used to investigate the fracture analyses, crack growth trajectory and fatigue life of curved stiffened panels repaired with composite patches subjected to combined tension and shear cyclic loadings. For this purpose, 3-D finite element modeling are performed for consideration of real 3-D crack-front in general mixed-mode conditions. Contact elements are used between the crack surfaces on two crack sides to prevent interferences of crack surfaces and a complementary program was developed to handle the automatic fatigue crack growth modeling. The effects of various patch layups and shear-tension loading ratios on fracture parameters of the aluminum panel are investigated. It is shown that in low shear to tension ratios like 0.4, the patch layup of 4 (perpendicular to the initial crack) is more efficient than the patch with layups angle along the tension loading. As the shear to tension ratio increases, effect of patch layups with orientations of almost perpendicular to the crack trajectory on fatigue crack growth life is increased comparing with the patch layups parallel to the tension orientation like 4.
Abstract: The fatigue damage of titanium has been studied on thin plate specimens subjected to alternating plane bending in argon gas. Fatigue strength in argon gas at Nf = 108 cycles was obtained to be 102 MPa. Fatigue behavior of titanium in argon gas has been attributed to the degradation of grain boundary cohesion with argon gas atoms/molecules. Fatigue cracks were propagated partly in intergranular mode. It has been plausible that argon gas atoms/molecules could penetrate into the distorted regions close to grain boundary through lattice defects and degrade grain boundary cohesion. Grain boundaries have been preferentially damaged in argon gas. The results in argon gas have been compared with those obtained in vacuum and in air.
Abstract: In this paper, detailed analysis of the tip flow of an axial compressor rotor blade has been carried out using the commercial CFD package ANSYS CFX. The rotor blade was designed such that it is reminiscent of the rear stages of a multi-stage axial compressor. The effects of varying tip gaps are studied using CFD simulations for overall pressure rise and flow physics of the tip flow at the design point and near the peak pressure point. Rig tests of a low speed research compressor rotor with 3% tip clearance provided characteristics plots for validation of the CFD results. With increase in clearance from 1% to 4%, the rotor pressure rise at the design point was observed to decrease linearly. Increase in the clearance increases the cross flow across the tip; however, the magnitude of the average jet velocity crossing the tip decreases. The tip leakage vortex was observed to stay close to the suction surface with increase in clearance.
Abstract: The paper investigates the effects of various gas turbine operating and health conditions on its hot section component’s creep life via a simple relative creep life parameter known as Creep Factor. Using the Creep Factor, the correlation between individual gas turbine operating and health parameter and component’s creep life was established and the weight of the impact was measured. Analytical-parametric-based creep life estimation model combined with the Creep Factor approach was developed and integrated with an existing engine performance model to allow the estimation of various hot section component creep lives and the computation of the Creep Factors. The impact analysis was carried out on the high pressure turbine blade of a model turbo-shaft helicopter engine. The results indicate that for a clean engine, the change in engine rotational speed was seen to provide the highest impact on changing the blade’s creep life consumption while for a degraded engine, the presence of compressor fouling has the highest threat in changing the blade’s creep life. The analysis also shows that the Creep Factor is a good indicator of creep life consumption and provides a good technique to rank the influencing factor according to the threat they imposed.