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Sairajan K.K.,StructuResearch Group | Aglietti G.S.,University of Surrey | Mani K.M.,StructuResearch Group
Acta Astronautica | Year: 2016

The emerging field of multifunctional structure (MFS) technologies enables the design of systems with reduced mass and volume, thereby improving their overall efficiency. It requires developments in different engineering disciplines and their integration into a single system without degrading their individual performances. MFS is particularly suitable for aerospace applications where mass and volume are critical to the cost of the mission. This article reviews the current state of the art of multifunctional structure technologies relevant to aerospace applications. © 2015 IAA. Published by Elsevier Ltd. All rights reserved.


Gupta A.K.,StructuResearch Group | Patel B.P.,Indian Institute of Technology Delhi | Nath Y.,University of Toronto
European Journal of Mechanics, A/Solids | Year: 2015

Abstract The objective of this paper is to investigate the progressive failure behaviour of laminated cylindrical/conical panels under meridional compression considering geometric nonlinearity and evolving material damage. The evolving microscopic damage such as fiber breakage, matrix cracking, fiber matrix debonding etc. is modeled through a generalized macroscopic continuum theory within the framework of irreversible thermodynamics. The analysis is carried out using field consistent finite element approach based on first-order shear deformation theory. The nonlinear governing equations are solved using the Newton-Raphson iterative technique coupled with the adaptive displacement control method to trace the equilibrium path. The damage evolution equations are solved at every Gauss point using Newton-Raphson iterative technique within each iteration of a loading/displacement increment. To accurately model the transverse shear strain energy, shear correction factors are calculated using layers' properties and lamination scheme. The detailed study is carried out to highlight the influences of evolving damage, span-to-thickness ratio, lamination scheme, radius-to-span ratio, boundary conditions and semi-cone angle on the postbuckling response and failure load of laminated panels. © 2015 Elsevier Masson SAS.


Li G.,StructuResearch Group
Journal of Mechanics of Materials and Structures | Year: 2012

In this paper, closed-form solutions for the adhesive stresses in bonded composite single-strap butt joints have been obtained. Two strategies were used for deriving the adhesive peel stress. The solutions are applicable to a butt joint made from different adherend and doubler laminates, as well as the unbalanced single-lap joints. In addition, three-dimensional finite element models of the unit-width composite joints were created for analyzing the adhesive stresses under a plane strain condition. A total of six joint conditions, three joint configurations and each with two layup sequences, were studied. Consistency in the peel stress predictions was obtained from the two theoretical strategies. Good agreement has been achieved between the theoretical and finite element results. The effects of the doubler thickness and laminate layup sequence on the adhesive stress variation can be displayed. The theoretical solution would provide a solid foundation for supporting the practical composite joint assessment. © 2012 by Mathematical Sciences Publishers.


Sairajan K.K.,StructuResearch Group | Nair P.S.,Satellite Center
Composites Part B: Engineering | Year: 2011

Reduction of spacecraft structure mass is an important design goal as it helps to increase the payload fraction, improve agility and also reduce launch cost. The spacecraft is subjected to mechanical loads during launch and thermal loads during on orbit operations. A typical spacecraft contains a main load bearing base structure, which supports all primary payloads and connects the satellite with the launch vehicle. The dimensional stability and tolerance requirements of the payload interfaces are very stringent and there is a need to improve these under the specified thermal environment. This paper explains a novel design consisting of a bonded assembly of metal and laminated composites to achieve a dimensionally stable, low mass base structure without altering the interfaces and overall dynamic behavior of the spacecraft. The design was checked for its performance under all critical loading conditions and was found to meet all requirements including stiffness and stability. Thermal distortion analysis showed that radial distortion is an order of magnitude lower than that of the existing design. The structure was fabricated and it showed compliance to dimensional accuracy requirements. The new base structure weighed 7.54 kg achieving a mass saving of 35% on the existing structure. © 2010 Elsevier Ltd. All rights reserved.


Anandatheertha S.,Indian Institute of Science | Naik G.N.,Indian Institute of Science | Gopalakrishnan S.,Indian Institute of Science | Rao P.S.,StructuResearch Group
Physica E: Low-Dimensional Systems and Nanostructures | Year: 2010

Many previous studies regarding the estimation of mechanical properties of single walled carbon nanotubes (SWCNTs) report that, the modulus of SWCNTs is chirality, length and diameter dependent. Here, this dependence is quantitatively described in terms of high accuracy curve fit equations. These equations allow us to estimate the modulus of long SWCNTs (lengths of about 100120 nm) if the value at the prescribed low lengths (lengths of about 510 nm) is known. This is supposed to save huge computational time and expense. Also, based on the observed length dependent behavior of SWCNT initial modulus, we predict that, SWCNT mechanical properties such as Young's modulus, secant modulus, maximum tensile strength, failure strength, maximum tensile strain and failure strain might also exhibit the length dependent behavior along with chirality and length dependence.


Murthy M.V.V.S.,StructuResearch Group | Renji K.,StructuResearch Group | Gopalakrishnan S.,Indian Institute of Science
Composite Structures | Year: 2015

Spectral elements are found to be extremely resourceful to study the wave propagation characteristics of structures at high frequencies. Most of the aerospace structures use honeycomb sandwich constructions. The existing spectral elements use single layer theories for a sandwich construction wherein the two face sheets vibrate together and this model is sufficient for low frequency excitations. At high frequencies, the two face sheets vibrate independently. The Extended Higher order SAndwich Plate theory (EHSaPT) is suitable for representing the independent motion of the face sheets. A 1D spectral element based on EHSaPT is developed in this work. The wave number and the wave speed characteristics are obtained using the developed spectral element. It is shown that the developed spectral element is capable of representing independent wave motions of the face sheets. The propagation speeds of a high frequency modulated pulse in the face sheets and the core of a honeycomb sandwich are demonstrated. Responses of a typical honeycomb sandwich beam to high frequency shock loads are obtained using the developed spectral element and the response match very well with the finite element results. It is shown that the developed spectral element is able to represent the flexibility of the core resulting into independent wave motions in the face sheets, for which a finite element method needs huge degrees of freedom. © 2015 Published by Elsevier Ltd.


Deshpande S.,StructuResearch Group | Patnaik M.N.M.,StructuResearch Group | Shankar Narayan S.,StructuResearch Group | Mittal V.,StructuResearch Group
Journal of Spacecraft Technology | Year: 2014

Conventionally during spacecraft vibration test the input at its global modes is notched to dynamic CLA base forces and moments. Presently at ISAC, the base force during spacecraft vibration test is estimated by indirect methods like from armature current, interface strains etc. How ever these conventional methods have limitations and fail to accuratelyestimate the base force / moment in certain cases. This necessitates the requirement for direct base force measurement. Force measurement devices (FMDs) are designed, developed and implemented for spacecraft vibration tests for mini and micro satellites. Basically it uses tri-axial force gauges capable of measurement in all three directions. Detailed design and configuration of the FMD and its application are discussed in this paper. It also gives the details of the implementation of force measurement of for spacecraft vibration tests. Spacecraft CG response and effective masses for major modes estimated from the base force measurement are brought out and compared with theoretical predictions.


Divya K.,StructuResearch Group | Renji K.,StructuResearch Group
Journal of Spacecraft Technology | Year: 2014

Spacecraft experience many types of dynamic loads during launch. The low frequency transients are normally expressed in terms of sine loads. They are simulated using shaker systems through a sine sweep excitation. Though sine sweep excitation is applied during the sine vibration testing of the spacecraft, the acceleration responses are theoretically estimated using steady state estimation techniques. The steady state responses of a single degree-of-freedom system are very much comparable with the responses due to sine sweep excitation and the relations between them are well established. In this work, investigations are carried out on the above behaviour in a multi-degree-of- freedom system like spacecraft. It is shown that when the modes are well separated the response characteristics are similar to that of the single degree-of-freedom system. When the modes are very close, separated by 0.2 Hz, the response of the second mode is very much influenced by the transient response component of the first mode. A suitable time step for estimating the response to sine sweep excitation is also discussed.


Peereswara Rao M.V.,StructuResearch Group | Harursampath D.,Indian Institute of Science | Renji K.,StructuResearch Group
Composite Structures | Year: 2012

This work focuses on the formulation of an asymptotically correct theory for symmetric composite honeycomb sandwich plate structures. In these panels, transverse stresses tremendously influence design. The conventional 2-D finite elements cannot predict the thickness-wise distributions of transverse shear or normal stresses and 3-D displacements. Unfortunately, the use of the more accurate three-dimensional finite elements is computationally prohibitive. The development of the present theory is based on the Variational Asymptotic Method (VAM). Its unique features are the identification and utilization of additional small parameters associated with the anisotropy and non-homogeneity of composite sandwich plate structures. These parameters are ratios of smallness of the thickness of both facial layers to that of the core and smallness of 3-D stiffness coefficients of the core to that of the face sheets. Finally, anisotropy in the core and face sheets is addressed by the small parameters within the 3-D stiffness matrices. Numerical results are illustrated for several sample problems. The 3-D responses recovered using VAM-based model are obtained in a much more computationally efficient manner than, and are in agreement with, those of available 3-D elasticity solutions and 3-D FE solutions of MSC NASTRAN. © 2012 Elsevier Ltd.


Kelvina Florence S.J.,StructuResearch Group | Renji K.,StructuResearch Group
Journal of Sound and Vibration | Year: 2016

Modal density is an important parameter in Statistical Energy Analysis (SEA) based response estimation. Many space structures use composite cylinders. Modal densities of such structural elements are not reported. In this work an expression for modal density of composite cylindrical shells is derived. Its characteristics and sensitivity to various parameters are discussed. The frequency at which the modal density has a maximum is derived. Modal densities of typical composite cylinders are obtained. It is shown that computing modal density considering an equivalent isotropic cylinder can lead to significant errors. © 2015 Elsevier Ltd. All rights reserved.

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