Journal of Aerospace Science and Technology
Volume:11 Issue: 2, Summer and Autumn 2017

The purpose of this paper is to estimate the fatigue life of an airplane wing with laminated composite skin, subjected to variable mechanical and thermal loads. To achieve this aim,at first, the three-dimensional model of airplane wing was drawn in CATIA software. Then, by transferring the model to the ABAQUS software, the finite element model of the wing wascreated. Here, the spars and ribs weremade ofaluminum T7075 and skin waslaminated composite with uni-directional and woven roving carbon epoxy layers. The convergence behavior of the structure wasexamined for selecting an appreciate element numbers. Finally, transient dynamic analysis followed by fatigue analysis of the wing structure was carried out. By performing fatigue simulation, the number of loading cycles resulted in failure of the structural components and wing skin panel was predicted. By comparing the simulation results with experimental research carried out by other researchers, the validity of the presented simulation method was demonstrated. The results of this research indicated that the replacement of traditional metallic wing skin with laminated composite skin causes a significant increase in fatigue life of the wing,besides a considerable weight reduction.

The aim of this study is to investigatethe effective parameters on vibrations of circular cylindrical shells with fixed rotary speed andresting elastic foundation by means of analytical and finite element numerical simulation. First, the governing equations are derived using the theory of Donnell, considering the centrifugal forces,Coriolis acceleration, and the initial annular tension. Then, the analytical solution for cylindrical shells isintroduced under simply supported conditions. Further, the effect of parameters such as therotational speed of the shell, its lay-up, fiber angle, and the stiffness of the elastic foundation on the values of natural frequency and the critical velocity of the shells are studied. The analytical solution results are in good compatibility with the results achieved from the finite element method.

Conventional quaternion based methods have been extensively employed for spacecraft attitude control where the aerodynamic forces can be neglected. In the presence of aerodynamic forces, the flight attitude control is more complicated due to aerodynamic moments and inertia uncertainties. In this paper, a robust nero-adaptive quaternion controller based on back-stepping technique for vehicle with aerodynamic actuators is proposed. The presented control lawconsists of a neural network based adaptive part and an additional term which ensures the robustness of the system. Actually, the first term is designed to approximate and cancel out the matched uncertainties and the second term is used toensure the robustness of system against approximation error of the neural network.The Lyapunov direct method is applied to derive the learning laws for the neural network weights and adaptive gain. Also,theultimately boundedness of the error signals is guaranteed based on theLyapunov’s stability criterion. The benefit of the presented method is evaluated through simulation of an aerodynamic control vehicle.

The present work derives the exact analytical solutions for buckling and post-buckling analysis of nano-composite beams reinforced by single-walled carbon nanotubes (SWCNTs) based on the Euler-Bernoulli beam theory and principle of virtual work. The reinforcements are considered to be aligned in the polymeric matrix either uniformly distributed (UD) or functionally graded (FG) distributed through the thickness direction of the beam. In FG beams, material properties vary gradually along the thickness direction. The effective material properties of the nano-composite beam are predicted based on the extended rule of mixture. Also, by applying von Kármán assumptions, the geometric nonlinearities are taken into consideration. The developed governing equations are solved by utilizing analytical methods and exact closed form solutions for buckling and post-buckling loads of functionally graded carbon nanotube-reinforced composite (FG-CNTRC) beam with different boundary conditions are obtained. By comparing the present post-buckling load results with the ones reported in the literature, the accuracy and reliability of the current method are demonstrated. Eventually, the numerical results are provided and the effects of CNTs distribution, CNTs volume fraction, slenderness ratio, maximum deflection of the beam and boundary conditions on the post-buckling characteristics of the CNTRC beam are discussed.

Composite stiffened cylindrical shells are widely used as primary elements in aerospace structures. In the recent years, there has been a growing research interest in optimum design of composite stiffened cylindrical shell structures for stability under buckling load. This paper focuses upon the development of an efficient optimization of ring-stringer stiffened cylindrical shell. The optimization problem used in this study involves weight minimization of ring-stringer stiffened composite cylindrical shell with buckling load and stress, which are considered as design constraints. The proposed methodology is based on Particle Swarm Optimization (PSO) algorithm. The material of shell is composite, but the material of stiffeners is considered to be isotropic. The approach adopted in modeling utilizes the Rayleigh-Ritz energy method and the stiffeners are treated as discrete members. In addition, a 3-D Finite Element (FEM) model of the ring-stringer stiffened cylindrical shell is developed that takes into consideration the exact geometric configuration. The results obtained using the Rayleigh-Ritz energy method are compared with those using 3-D FE model. The proposed methodology is implemented on the ring-stringer stiffened cylindrical shell using the PSO algorithm. The obtained results show a 13% reduction in the weight of the ring-stringer stiffened cylindrical shell whilst all the design constraints are satisfied. In addition, the results show that the proposed methodology provides an effective way of solving composite stiffened cylindrical shell design problems.

In this paper, optimization of the sound transmission loss of finite rectangular anisotropic laminated composite plate with simply supported boundary conditions has been developed to maximize transmission loss. Appropriate constraints were imposed to prevent the occurrence of softening effect due to optimization. For this purpose, optimization process was incorporated into comprehensive finite element software. The transmission loss (TL) obtained from the numerical solution was compared with those of other authors indicated good agreement. The discrete frequencies have been chosen based upon the sound transmission class with A-weighting constant. Several traditional composite materials have been studied and the results have shown that in the mass control region, the optimization of stacking sequence and optimal thickness has not been an effective contribution to improve the transmission loss. The results also show that, the lamina thickness optimization has an important effect on improving the transmission loss, but the advantage of low weight composite material is compromised by optimization.