Finite Element Analysis of Interface ring of a
Rocket Launcher

Devika Venu; Dr. Alice Mathai; Prof. Jayasree Ramanujan

Finite Element Analysis of Interface ring of a Rocket Launcher

Devika Venu¹, Dr. Alice Mathai², Prof. Jayasree Ramanujan³
M.Tech. student, Dept. of Civil Engineering, National Institute of Technology, Trichy, Tamil Nadu, India¹
Associate Professor, Dept. of Civil Engineering, Mar Athanasius College of Engineering, Kerala, India²
Professor, Dept. of Civil Engineering, Mar Athanasius College of Engineering, Kothamangalam, Kerala,India ³

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Abstract

This paper aims at performing linear static analysis of laminated composite interface ring of rocket launcher and to investigate the effect of ply orientation of laminates on deformation and stress pattern. The interface ring is a part of conical shell, having a top radius of 367mm and bottom radius of 472mm. The shell is made of laminated carbon fibre plastic composites with distribution of total of 20 layers forming a total thickness of 2mm. The load of 200kN from the components above the ring acts as distributed compressive load on the top of the ring, the bottom of which is fixed. The analysis carried out on structure lead to best suited laminate and optimum ply orientation. The minimum deformation for laminated conical shells having angle plies is obtained with a ply orientation of 10 degree

Keywords

Interface ring; CFRP; ply orientation; deformation contour; stress contour

INTRODUCTION

Design of structures for space missions is a formidable challenge to design community due to contradictory requirements of low margin, low mass and cheaper but high reliability. These requirements are presently met by design and construction concepts of externally stiffened construction and integrally stiffened construction, such as, closely stiffened structures, ring stiffened structures, waffle, isogrid structures, sandwich, composite structures etc. The innovations in design and structures have parallely led to advancements in materials and manufacturing technologies and have substantially contributed to satisfy the specified requirements of the present space mission applications. The use of advanced composite materials has become essential in spacecraft structures for achieving low mass while satisfying structural and functional requirements related to stiffness, strength, dimensional stability and agility. Reduction of structural mass is important in many ways – it helps to increase payload fraction in a spacecraft, improves agility and also reduces the launch cost. These properties of composites would result in large weight savings thus replacing the conventional metallic designs.

A satellite launch vehicle is a complex transportation system which is tailored and engineered to function as a flying machine. This vehicle is made capable of transcending the harsh atmospheric regime and beyond near earth space and deliver the encapsulated satellite in the orbit; thus fulfilling the mission. Launch vehicles are built as an assemblage of several stages which may be solid, liquid, cryo/semi cryo, etc depending on the propulsion systems employed. In multi– stage launch vehicles, the inert stages, which have performed their function, are jettisoned sequentially as flight progresses. Fig.1 shows the different stages & structures in a launch vehicle. This is enabled by staging/separation systems which are essentially pyro/explosive based systems. Launch vehicle structures which have the primary role to provide/maintain necessary external aerodynamic shapes while withstanding the flight loads and environments may be broadly classified under two headings such as primary structures and secondary structures. Use of advanced composite materials has been the most contributing factor. Till nowadays the base structure, especially the interface ring remained metallic due to a number of practical considerations. Such metallic base structures are normally integrally machined parts and contribute to a significant part of the spacecraft structure mass. The upper interface of such structures made of aluminum alloy often poses serious problems of dimensional accuracy and stability when such requirements are mandatory. These base structures are used mainly for compatibility between the launch vehicle payload adaptor, clamp band release mechanism and other requirements. Recently there have been attempts on the use of light weight materials like composites for the construction of spacecraft structural systems. These are specially designed for meeting the stringent requirements of stiffness, dimensional stability, impact resistance and fatigue. The structure under study is a composite base structure of 600 kg class spacecraft. It is a conical shell type structure. Fig.1 shows exploded view of spacecraft structure showing base / interface ring. The spacecraft bus elements are attached to it through horizontal decks and radial panels. In addition, a heavy alignment sensitive payload is also attached at one end. All spacecraft loads are transferred to the launch vehicle adaptor during launch through this part. This interface structure is made of carbon fiber reinforced plastic (CFRP) [2].

II. METHODOLOGY

Interface ring of the rocket launcher was modelled and analysed using ANSYS software as CFRP material. ANSYS is a finite element analysis (FEA) software package. It uses a preprocessor software engine to create geometry. Then it uses a solution routine to apply loads to the meshed geometry. Finally it outputs desired results in post-processing. ANSYS includes a large number of different structural elements which are developed for specific applications. ANSYS is an FEA package which supports layered elements like SHELL99 for the analysis of shell structures and BEAM189 for the analysis of beam elements. The selection of an element for an application should be based on its capabilities, its cost (stiffness matrix generation, decomposition, stress calculation etc.) and the desired accuracy in the results.

`A. Composite Modeling in ANSYS

Composite materials can be modeled in ANSYS using specialized elements called layered elements. After building the model using these elements, any structural analysis like static, nonlinear static, Eigen buckling, nonlinear buckling, modal analysis, dynamic analysis etc. can be performed. SHELL 99 is a linear Layered Structural Shell Element is an 8-node, 3-D shell element with six degrees of freedom at each node. It is designed to model thin to moderately thick plate and shell structures with a side-to-thickness ratio of roughly 10 or greater. The element has six degrees of freedom at each node: translations in the nodal x, y, and z directions and rotations about the nodal x, y, and z-axes.

Fig.2 shows the geometry of the layered shell element where, LN = Layer number and NL = Total Number of Layers. The element is defined by eight nodes, average or corner layer thicknesses, layer material direction angles, and orthotropic material properties. Mid side nodes may not be removed from this element. A triangular-shaped element may be formed by defining the same node number for nodes K, L and O. BEAM189 is an element suitable for analyzing slender to moderately stubby/thick beam structures. This element is based on Timoshenko beam theory. Shear deformation effects are included. BEAM189 is a quadratic (3-node) beam element in 3-D. BEAM189 has six or seven degrees of freedom at each node, with the number of degrees of freedom depending on the value of KEYOPT(1). When KEYOPT(1) = 0 (the default), six degrees of freedom occur at each node. These include translations in the x, y, and z directions and rotations about the x, y, and z directions. When KEYOPT(1) = 1, a seventh degree of freedom (warping magnitude) is also considered. This element is well-suited for linear, large rotation, and/or large strain nonlinear applications.

III. LINEAR STATIC ANALYSIS OF INTERFACE RING

As per the Research and Development Programs conducted at ISRO recently, the problem may be identified as an interface ring possessing the following conditions regarding dimensions of interface ring modeled as conical shell with top radius = 367 mm, bottom radius = 472mm, height = 290mm, thickness = 2mm, no. of layers = 20, thickness of each laminae = 0.0001mm [2].

Elements used are SHELL 99 for conical shell and BEAM 189 for stiffeners. Boundary condition of the ring is fixed at bottom using fixtures. Loading along longitudinal direction of intensity 200 kN as distributed compressive load on the top of the ring. Fig.4 shows the loading diagram of conical shell in ANSYS window. The properties of the CFRPcomposite used for the analysis are given in Table 1 below. It is evident that CFRP possess improved properties in comparison with other conventionally used materials.

IV. RESULTS OF LINEAR STATIC ANALYSIS

A. Mesh Convergence

The accuracy of the finite element solution depends on the discretisation which is characterized by the finite element mesh and the choice of elements. Various mesh sizes are selected arbitrarily so that the entire area of the shell is divided proportionally. The one that gave minimum value of deformation for the given conditions was selected as the optimum mesh size and the same was taken for analysis of optimum ply orientation and study of laminate behaviour [7]. Various mesh sizes were randomly selected and design loads were tried upon the model and deformation values for respective mesh sizes were tabulated as shown in Table 2.

On comparing the deformations shown by the interface ring with progressively increasing mesh sizes, the mesh convergence was obtained at a mesh size 96 x10. Hence for further analysis optimum mesh 96 x10 was selected.

B. Types of Laminates and their Behaviour

The behaviour of different laminates under the design loads was studied with respect to deformation observed on the shell. The comparison on deformation shown by cross ply and angle ply laminates were done for both symmetric and anti symmetric arrangement. Fig .5 shows the layer stacking sequence of cross ply and angle ply laminates. For cross ply laminates, 00 ply and 900 ply are placed alternately. For angle ply laminates, plies are oriented in a random fashion.

The deformation observed on different configuration of laminates for symmetric and anti symmetric arrangements were tabulated as in table 3 and explained in Fig. 6

On comparing cross ply and angle ply laminates, cross ply anti symmetric laminates has found to give better result. Using the criteria for selection and taking into account, the advantages of cross ply laminate over other configuration, an improved configuration that employs the best suited orientation is adopted. Herein, for further analysis, the improved configuration of laminates is designated as [ 02/902/±?/902/02]s .

C. Determination of optimum Ply Orientation

Anti-symmetric cross ply laminates were found to have minimum deformation on comparison with other types. Hence optimum ply orientation was tried upon varying the values of θ in the orientation configuration stated as [ 02/902/±?/902/02]s. The axial deformation was obtained for different ply orientations on analysis under design loads and was tabulated as in Table 3. A graph was plotted between deformation and ply orientation and a smooth continuous curve was obtained. In Fig. 6, the curve initially sags down up to 100 and then rises upwards.

The minimum deformation under axial loading was obtained for ? = 10 .? The corresponding laminate configuration is designated as [02/902/±10/902/02]s .

V. CONCLUSION

Stiffened composite laminated conical shells are integral part of spacecrafts and launch vehicle structures. The best suited laminate for interface ring was obtained as cross plies comparing with angle plies. The optimum ply orientation for laminated conical shells having angle plies was obtained with a ply orientation of 10 degree.

Tables at a glance


Table 1	Table 2	Table 3

Figures at a glance


Figure 1	Figure 2	Figure 3	Figure 4


Figure 5	Figure 6	Figure 7

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