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Thermal effects play significant role in the design of a structure
when it is supposed to work in hot environment. For this purpose a
thorough analysis of structure is necessary that could precisely
consider the effect of temperature change. At first an accurate
thermal analysis is required that may be further followed by stress
or dynamic analysis. In this work, shell finite element developed
by Rolfes et al to perform 3D thermal analysis is studied using
Ansys and results are found good when comparing with the 3D solid
element result. The interlaminar shear stress analysis of composite
plate considering thermal and mechanical loading is studied under
different set up of laminate scheme and the results have been found
good when compared with the available 3D elasticity results. Then,
the effect of thermal stresses on natural frequency of a structure
is analyzed considering as thermally prestresssed case under
different laminate schemes and boundary conditions. Finaly, A
stringer stiffened panel is analyzed under thermal and mechanical
load. Interlaminar stresses and natural frequency are evaluated by
considering the effect of thermal loading.
Skin-stiffener structures are extensively used in the aerospace
field due to their structural efficiency in terms of
stiffness/weight and strength/weight ratios. The application of
such panels is primarily within fuselages and wing boxes, where the
weight saving potential of composite materials compared with
aluminum alloys is well known. However, design of composite panels
involves the optimization of a large number of variables such as
ply thickness and plate widths. Further complication arises when
the expert knowledge required for laminate design is considered and
when the panel is constrained by buckling under axial compression.
In this study, the behavior of Grid Stiffened Composite Cylinder is
examined under the axial compressive load. Compressive load causes
buckling and develop stresses in the structure. For the buckling
analysis two approaches are used.(1)Analytical (2)FEM. An
analytical smeared stiffener theory is used to determine buckling
load and then FEM results were compared to gain the confidence on
the developed models. The validated FEM model and analytical
smeared stiffener theory is used to conduct parametric analysis.
Composite stiffened panels have been used in aerospace industries
for past decades. Due to light weight nature composite panels are
prone to buckling. In this study buckling load for Ortho-grid panel
and Iso-grid panel is determined. This was accomplished by
developing a composite laminate for both skin and stiffener using
MSC PATRAN/NASTRAN was used for linear buckling analysis of
composite panels. The buckling loads of two panels were determined
based on finite element analysis results, including geometric
dimension, thickness of the skin, number of laminae, ply stacking
sequence, thickness and height of stiffener. Parametric studies of
general grid stiffened panel was conducted using skin thickness,
stiffener thickness and stiffener height as design variables.
Conclusions drawn from these results are presented. All three
buckling modes appeared and maximum buckling load was observed in
skin buckling mode for Iso-grid panel.
Modal parameters of spacecraft structures are important indicators
of their dynamic characteristics which can be used for vibration
reduction and control. In the present research modal parameters of
a rocket are identified by using continuous wavelet transformation.
The advantages of the wavelet transformation for the modal
parameters identification over other methods is study. Through
finite element method transient dynamic response of a system is
determined then continuous wavelet transformation is applied to
find the modal parameters like modal frequency, mode shapes and
damping ratio.
Numerical simulations of sheet metal forming process based on
finite element method (FEM) is widely applied for its powerful
capability in forming process prediction. Since there are
parameters which could affect the result of forming process, it
becomes important to approach a set of parameters to improve the
formability. In the present work, first, a comprehensive literature
review was made for different optimization methodologies in sheet
metal forming process. Then we proposed an optimization methodology
using Response Surface methodology (RSM) and Genetic Algorithm (GA)
for the optimization of sheet metal forming process, and the theory
of RSM and GA are illustrated. The presented method was first
applied to an example from literature, the results verified the
feasibility of the proposed methodology. Then the methodology was
applied to variable binder force optimization of the NUMISHEET'93
2D draw bending problem. The work indicated that the proposed
optimization methodology is efficient and universal, which means it
can also be applied in other applications of aerospace industry.
The aim of this work is to know more about the underprediction of
springback in numerical simulation of sheet metal forming. There
are a lot of factors which influence the accuracy of springback
with numerical simulation.One factor responsible for the inaccurate
springback prediction is the assumption of constant Young's modulus
in FE analysis. Young's modulus, however, is found to decrease
during plastic deformation. Therefore, an accurate model is needed
to be implemented in FE analysis which can predict the decrease in
Young's modulus. With passage of time, Young's modulus is found to
restore to its initial value. The present study focuses on this
aspect of springback prediction. Bake Hardenable (BH) steel has
been selected to investigate the degradation and recovery in
Young's modulus. The decrease in Young's modulus and its recovery
has been measured experimentally by two methods i.e. the dynamic
method (Impulse Excitation Technique) and a static method (Tensile
Test).
Lighter and stiffer structural construction is required for space
rockets than those used in aviation. A reduction of only 1 Kg in
the weight of a rocket intended for a flight to Mars and back
results in saving 2 tons of propellant. At present, sandwich
technology is extensively used in Aeronautics and Aerospace fields.
Sandwich panels with thin composite face sheets are used due to
their ability to produce structures with high stiffness to weight
and strength to weight ratios. In this research, Finite Element
(FE) model is generated for a particular test configuration of
honeycomb sandwich structure with metallic (Aluminium) skins. In
this modelling effort, the specimen is modelled in such a way to
simulate and analyze in a true sense. Stability and non linear
analysis is carried out using FEA software SAMCEF. With geometric
and material nonlinearity, the results are then compared with the
experimental ones. Force displacement curves acquired from FEA and
3point bending test seem good enough for comparison.
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