JETIR Research Journal

© 2019 JETIR May 2019, Volume 6, Issue 5 www.jetir.org (ISSN-2349-5162) Design Optimization and Analysis of Rocket Structure for Aerospace Applications 1Chethan S, 2Niranjan Hiremath 1M.Tech Machine Design Student, 2Associate Professor 1,2 School of Mechanical Engineering, REVA University, Rukmini Knowledge Park, Kattigenahalli, Yelahanka, Bengaluru, India

Abstract: The aim of this study is to perform techniques such as pressure mapping of the missile structure, Inertia Relief Analysis for Rocket Structure in the effective way with the low cost. And the analysis plays important role in the structural design and safety. Due to the advances with numerical software’s, the simulation helps in estimating the safety of the structures without actual prototype built up and testing. Firstly we design the Rocket parts by the help of modelling tools using solid edge v19, Catia v6 and assembled it later. In order to find pressure distribution on the missile surface and to find the stress and deformation of missile structure by pressure mapping using CFD mesh in Acusolve on to the Rocket Structure, Pressure values in “X, Y and Z” co-ordinate with magnitudes are obtained in CSV file format. And later we study behavior of the missile parts like Nose, Nose cone, Body, Tail and Fins when the surface remains same from nose to tail and the fins with the varying thickness. And we use the Maraging Steel 250 grade as the material and its properties is assigned to the mesh. And the mesh thickness of the outer shell material is increased for different cases along with a speed of Mach number 0.75 and 5. And structural changes are seen when the loads are applied to the mesh. Characteristics individualities of displacement, stress distribution and high stress locations are determined and results are extracted in terms of plots using the tools Hypermesh for Pre-processing, Optistruct, Acuconsole for Solving and Hyper-view, Acu- Field View for Post Processing.

IndexTerms - Rocket Structure, Missile, CFD mesh, Pressure mapping of the Missile Structure, Inertia Relief Analysis.

I. Introduction

Missile, is a guided self-propelled system, as opposed to an unguided self-propelled munition, referred to as a rocket (although these too can also be guided) Missiles have four system components: targeting or missile guidance, flight system, engine, and warhead. Missiles come in types adapted for different purposes: surface-to-surface and air-to-surface missiles (ballistic, cruise, anti- ship, anti-tank, etc.), surface-to-air missiles (and anti-ballistic), air-to-air missiles, and anti-satellite weapons. A Rocket is a vehicle which acquires push by the response of the rocket to the discharge of plane of quick moving liquid fumes from rocket engine. Solid fuel rockets make their fumes by the ignition of strong charge grain. The subsequent gasses are extended through the spout whose capacity is to change over this inward weight into a supersonic fumes speed. Rocket engine fumes are shaped altogether from force conveyed inside of the rocket before use. Rocket motors work by activity and response. Rocket motors push rockets forward by ousting their fumes the other way at fast. Rockets depend on energy, airfoils, helper response motors, gimbaled push, force wheels, diversion of the fumes stream, charge stream, turn, and/or gravity to help control flight.

The first rockets were used as propulsion systems for arrows, and may have appeared as early as the 10th century in Song dynasty China. However more solid documentary evidence does not appear until the 13th century. The technology probably spread across Eurasia in the wake of the Mongol invasions of the mid-13th century. Usage of rockets as weapons before modern rocketry is attested in China, Korea, Indian subcontinent, and Europe. One of the first recorded rocket launchers is the "wasp nest" fire arrow launcher produced by the Ming dynasty in 1380. Iron-cased rockets, known as Mysorean rockets, were developed in Kingdom of Mysore by the mid-18th century in India. William Congreve, son of the Comptroller of the Royal Arsenal, Woolwich, London, became a major figure in the field. From 1801, Congreve researched on the original design of Mysore rockets and set on a vigorous development program at the Arsenal's laboratory.

At the beginning of the 20th century, there was a burst of scientific investigation into interplanetary travel, largely driven by the inspiration of fiction by writers such as Jules Verne and H. G. Wells as well as philosophical movements like Russian cosmism. Scientists seized on the rocket as a technology that was able to achieve this in real life, a possibility first recognized in 1861 by William Leitch. II. Literature Review

Geyergy kuilanoff and Richard m. Drake fluor Daniel [1] This paper presents criteria and procedures for the design of structures and components for wind generated missiles. Methods for determining missile-induced loading, calculated structural response, performance requirements, and design considerations are covered. The presented criteria is applicable to Safety-Related concrete buildings as a whole and to all their exposed external components including walls, roofs, and supporting structural systems and elements.

DetlefK uhl [2] The acceleration of these gases through the engine exerts force “thrust" on the combustion chamber and nozzle, propelling the vehicle according to Newton's Third Law. This actually happens because the force pressure time’s area on the combustion chamber wall is unbalanced by the nozzle opening, this is not the case in any other direction. The shape of the nozzle also generates force by directing the exhaust gas along the axis of the rocket.

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© 2019 JETIR May 2019, Volume 6, Issue 5 www.jetir.org (ISSN-2349-5162) L.C. Scalabrin and J.L.F. Azevedo, P.R.F. Teixeira, A.M. Awruch [3] Aerodynamic flow simulations over the first Brazilian satellite launch vehicle, VLS, during its first-stage flight are presented. The three dimensional compressible flow is modeled by the Euler equations and a Taylor-Galerkin finite element method with artificial dissipation is used to obtain the numerical solution. Transonic and supersonic results for zero angle-of-attack are presented and compared to available experimental results. The influence of mesh refinement and artificial dissipation coefficient on the transonic flow results are discussed. The results obtained for the supersonic simulations present good agreement with experimental data. The transonic simulation results capture the correct trends but they also indicate that this flight condition requires more refined meshes.

III. Objectives

The specific objectives of the current work are as follows

i. To implement the techniques such as structural (static and dynamic) analysis to the sub sonic missile in the effective way with the low cost. To design the missile model by the help of modelling tools using solid edge v19, Catia v6.

ii. Is to achieve the Maneuvering in all 6 Dof , The missile flight range, Universal for multiple platforms, Fire and forget principle of operation, Shorter flight times leading to lower target dispersion and quicker engagement, Transport launch canister (TLC) for transportation, Storage and launch.

iii. To study behaviour of the missile when the thickness of the outer shell material is increased and structural changes when the loads are applied. To analyse by applying the loads on each parts of the rocket model by using AL-2024, AL-7075, Aluminim-copper alloy-AA2014, Titanium alloy(23Ti6Al4rELI alloy), Maraging Steel and the comparison is made by applying same value of loads, boundary conditions and constraints on the model for the materials considered separately and selected the best feasible material for the missile structural.

iv. To obtain the results when load is applied on the first stage of the rocket model comparing the displacement and contour results for the materials respectively using the Hypermesh For Pre-processing, Optistruct, Acuconsole For Solving and Hyper-view, Acu-Field View For Post Processing.

3.1 Methodology

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In this design and modelling of the rocket structure is done through Solid Edge and Catia V6 tools. Here in this proposed rocket structure we have five main parts that is nose, nose cone, body, tail and fin. These parts are individually design first with the respective dimensions. After the modelling of each part they are assembled and we get the final design of the rocket structure.

4.1 Design of Nose Part

Nose part is the foremost part in a rocket structure where the nozzle is attached. First we design the nose part of the rocket.

Fig 1: Design of Nose Part of Rocket

Fig 1.1: Diameter of Nose Part of Rocket

4.2 Design of Nose Cone

The second part after nose is nose cone design. Nozzles are carried in this nose cone part. The nose cone is of the most pivotal piece of a rocket. The nose cone of a rocket goes about as an approach to punch an opening in the environment.

Fig 2: Design of Nose Cone Part of Rocket

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Fig 2.1: Inner diameter of Nose Cone Part

Fig 2.1.1: Outer diameter of Nose Cone

4.3 Design of Main Body

The third part is body of a rocket is one of the more credible part. The reason for the body is to house the fuel. It is frequently as an empty chamber in light of the fact that it diminishes the sum surface space that is in contact with the air. This thusly minimizes drag.

Fig 3: Design of Main Body Part of Rocket

4.4 Design of Tail Part and Fin Part

The fourth part is tail part. The design of tail is part is crucial reason that the climate vane bolt focuses into the wind that is the tail of the bolt which has a much bigger surface region than the sharpened stone. The streaming air confers a more significant power to the tail than the head, and in this manner the tail is pushed away. There is a point on the bolt where the surface region is the same on one side as the other. This spot is known as the focal point of weight. The final part in the design is Fin or blade part of the rocket. The reason for putting blades on a rocket is to give strength during flight, that is to permit the rocket to keep up its introduction and strategic flight way.

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Fig 4: Design of Tail and Fin Parts of Rocket

Fig 4.1: Diameter of Tail Part of Rocket

Fig 5: Full Assembled Missile or Rocket Model

V. Pressure Mapping Of The Missile Structure

5.1 Input Details

Material used = Maraging Steel 250grade Young's Modulus = 210 GPa Poisson's ratio = 0.3 Density = 8000 kg/m3 Yield Strength = 1680 MPa

5.2 Analysis Procedure

A) To find pressure distribution on the missile surface 1. Design the CAD model adhering to the existing standards (Akash missile) 2. Mesh the model along with the fluid volume 3. Perform the CFD air flow analysis in Acusolve to obtain the pressure distribution over missile surface 4. Export Pressure values from AcuField-View in CSV Format

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© 2019 JETIR May 2019, Volume 6, Issue 5 www.jetir.org (ISSN-2349-5162) B) To find the stress and deformation of missile structure by pressure mapping 1. Missile Body is modeled with 2nd order Shell Elements (tria6) 2. Set up the boundary conditions and map the pressure values obtained from Acu-Field in the model in Hypermesh 3. Perform Inertia Relief Analysis to find out the maximum stresses and displacements 4. Perform Modal Analysis to find out the shapes and frequencies at which the structure will amplify the effect of a load

5.3 CFD Mesh Used For Finding Pressure Distribution on Missile Surface

In this process we import the missile.iges file to the Ansys and construct the virtual wind tunnel and five boundary layers are made in order to map the pressure around the missile structure. And later it is Export to obtain the Pressure values in Acusolve.

Fig 6: Fluid volume Mesh

Fig 6.1: Isometric sectional view

Fig 6.2: Boundary layer mesh in the missile wall

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Fig 6.3: CFD mesh 5.4 Solver Settings Used In Acusolve

5.4.1 Fluid Properties

Air at,

Density = 1.225 kg/m3 Specific Heat = 1005 J/kg-k Conductivity = 0.02521 W/m-k Viscosity = 1.781e-5 kg/m-s

5.4.2 Inlet Boundary Condition

Air velocity = 250 m/s and 1700m/s Mach number = 0.75 and 5 Air temperature = 273 k Pressure = 101325 N/m2

5.4.3 Outlet Boundary Condition

Open to atmospheric pressure

Fig 6.4: Acusolve Solver Settings View

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Fig 6.5: Velocity Streamline Flow

5.5 Acu-Field View Result

. Here the Pressure Distribution on Missile Surface can see Acu field view near the Nose, Nose Cone, Body, Tail and Fins sections.

Fig 6.6: Pressure Distribution on Nose and Nose Cone

Fig 6.7: Pressure Distribution on Body

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Fig 6.8: Pressure Distribution on Tail and Fins

5.6 Pressure Values Obtained From Acu-Field View

The Pressure values showing “X, Y and Z” co-ordinate with magnitudes in CSV file format is extracted in order use this in the loading conditions for different mesh thickness.

Fig 6.9: Co-ordinate with Magnitudes

VI. Meshing Each Parts Using Hypermesh

Before analysing the structure meshing should be done. Meshing of the parts is done by hypermesh tool. The model is meshed with tria6 elements using hypermesh with varying mesh size, the minimum sizes of elements are 3mm, 5mm, 7mm and maximum elements size is 10mm. And it as element count of 1, 20,268.

Fig 7: Fully Meshed Missile Model

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The Pressure Values are taken from the above “X, Y and Z” co-ordinate with magnitudes in CSV file format and applied to the mesh thickness for the above mentioned cases.

Fig 7.1: Image showing linear interpolation of pressure values input using CSV file

VII. Inertia Relief Analysis

The technique of inertia relief has been a well-known approach for the analysis of unsupported systems such as air vehicles in flight, automotive in motion, or satellites in space. The sum of forces and moments are calculated and applied to achieve an equilibrium state in inertia relief analysis. Inertia relief was applied to calculate load distribution in the missile support structure due to flight load imbalances. Although inertia relief approach has been widely employed in the simulation of unconstrained aircrafts and space vehicles, the published work has rarely been found. There is still lack of research on inertia relief analysis of diverse types of basic structures and critical structural considerations associated with inertia relief calculation. In this paper, inertia relief method is applied to analyse a variety of structures including spring-mass structures, truss structures, plate structures, and etc. The work is aimed to study various key issues associated with inertia relief analysis, such as conventional inertia relief and automatic inertia relief, the effect of constraints and mass distribution on inertia relief calculation, the accuracy of inertia relief, and critical considerations. Commercial finite element program MSC/NASTRAN is applied to generate numerical results.

7.1 Results of Inertia Relief Analysis

CASE 1: Mach number= 0.75 Thickness= 0.01m