AERODYNAMIC SHAPING OF PAYLOAD FAIRING FOR A LAUNCH VEHICLE Irish Angelin S* 1, Senthilkumar S 2

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1 e-issn , p-issn IJESR/May 2014/ Vol-4/Issue-5/ Irish Angelin S et al./ International Journal of Engineering & Science Research AERODYNAMIC SHAPING OF PAYLOAD FAIRING FOR A LAUNCH VEHICLE Irish Angelin S* 1, Senthilkumar S 2 1 Post Graduate Scholar, Dept. of Aeronautical Engineering, Nehru Institute of Engineering and Technology, Coimbatore, Tamil Nadu, India. 2 Asst. Prof, Dept. of Aeronautical Engineering, Nehru Institute of Engineering and Technology, Coimbatore, Tamil ABSTRACT Nadu, India. Simulations have been carried out for the payload fairing of a launch vehicle. The payload fairing faces maximum unsteadiness during transonic region of the launch vehicle flight which induces high acoustic loading. Study has been done for a vehicle s payload fairing at various Mach numbers in transonic zone ranging from 0.7 to 1.2. Analysis is been done aiming to reduce the unsteady pressure levels and thereby reduce the aeroacoustic loading. Aeroacoustic will be higher in the separated region and in and around the shock location. The pressure increase across the shock is studied. The area of separation zone and the magnitude of Shock Strength should be reduced as much as possible. Pressure distribution, shock strength and separation zones are studied through flow simulation. The aerodynamic characteristics of PLF are improved by considering the fairing shapes by means of maintaining the length of the cylinder, Nose Cone angle and Boattail angle which are the main parameters that decide the shock formation on the payload fairing and the separation region. The Mach number which gives a weak shock, thereby with a minimum separation area has been proposed. Keywords: Shock strength, shock location, separation, pressure distribution. 1. INTRODUCTION There is a significant interest in the aerodynamic shaping of the PLF which has been an important area of research for many years because of its importance in the launch vehicle. A serious instability can happen at transonic zone because of the shock formation and its strength. It can lead to separation in the flow also. Therefore, it is necessary to decrease the shock strength and the separation region. R.C.Mehta [4], studied the shock induced separation flow, surface pressure distribution over the bulbous heatshield at transonic region. TomoyaOchinero, and Thomas Deiters [10], designed an asymmetric payload fairing with an aerodynamic characteristic and flow analysis is been done in transonic Mach numbers. James C. Newman III [3], studied a complex geometrical configuration and flow physics by considering Nonlinear state equations. The difficulty in the application of sensitivity analysis was the challenge studied. Thereby the aerodynamic performance was increased. EndaDimitri Vieira Bigarella [2]., obtained pressure distribution over a blunt body configuration with boattail in the transonic flow regime using inviscid and viscous analysis. In the present study the flow analysis is done in a transonic Mach numbers aiming to achieve a minimum separation region and reduced shock strength over the heatshield of the vehicle. CFD analysis had been carried out for the present study. Turbulence is modeled by k-ε turbulence model. Grid convergence study has been carried out for validation. Fine grid generation is employed over the surface of the heatshield to capture the flow field accurately.the discretization procedure for the domain has been performed using CFD code. Mesh size helps to obtain finer and coarser mesh near the body and away from the body respectively. NS Turbulent Solver is been used for the present simulation. 2. CONFIGURATION The geometry of the payload faring is of primary importance as it has to accommodate a large payload and be of aerodynamic shape so that the loads on the launch vehicle do not exceed than the design value [9]. In this present work only axisymmetric forebody is considered. Thus, the construction of three-dimensional heatshield geometry has been simplified. PLF consists of four main parts: spherical nose cap, nose cone, cylindrical portion and boat tail. The shaping of the PLF is done by choosing the nose cone angle, nose cap radius, boat tail angle and cylinder. This section sets up the geometrical features of the launch vehicle PLF considered for the CFD simulation. The two-dimensional view of the geometry is shown in the below Fig 1. *Corresponding Author 295

2 The geometrical configuration is tabulated in Table 1. Table 1: Geometrical details Fig 1: Two-dimensional view of the body Core cylinder diameter D Nose cap radius 0.33D Length of the cylinder portion 2.34D Payload shroud diameter 1.25D For the above geometry grid has been generated and grid independent study has been carried out foe validation. A fine mesh is generated and grid is refined at various iterations so as to capture the flow field properly. For a set of transonic Mach numbers; 0.7 to 1.2 a comparative study is carried and the results are discussed below. 3. RESULT AND ANALYSIS As stated in previous section, simulations have been carried out for various Mach numbers in transonic region ranging from 0.75 to 1.2. The results in terms of pressure distribution over the surface, shock location, separation length, pressure rise due to shock and shock strength have been studied and compared. 3.1 Pressure Distribution and Shock Location The pressure distribution over the surface is influenced by the curvature discontinuity of the vehicle exclusively at various junctions like nose cap-nose cone junction, nose cone-cylinder junction and cylinder boattail junction. Approach Mach number to nose cap is invariably subsonic. At the stagnation point, the flow speed is brought to zero. The flow accelerates along nose cap and nose cone and becomes supersonic as it approaches nose cone-cylinder junction. Expansion at the nose cone cylinder junction increases flow Mach number. As there is no surface or curvature gradient in cylindrical section, flow tries to regain free stream properties. In lower transonic Mach number, supersonic flow in the beginning of cylindrical portion is brought to freestream value through a transonic terminal shock. The flow behavior at the cylinder boattail junction depends entirely on the approach Mach number. The variation of pressure, P along the non-dimensional length, X/D is shown in Figure 2 to Figure 4 for Mach numbers range of 0.7 to 0.8, 0.85 to 1.0 and 1.05 to 1.2 respectively. Fig 2: Pressure Distribution over Heat Shield at M=0.7 to 0.8 Copyright 2013 Published by IJESR. All rights reserved 296

3 Fig 3: Pressure Distribution over Heat Shield at M=0.85 to 1.0 Fig 4: Pressure Distribution over Heat Shield at M=1.05 to 1.2 The initial peak value of P is due to stagnation point at the centre of spherical cap. The flow then expands over the spherical nose cap to a lower pressure value. At the nose cap cone junction, the flow is compressed due to the change in geometrical discontinuity with positive gradient and hence the pressure increases. In the cone region, the flow remains almost constant or slightly decreases. At the cone-cylinder junction, the pressure reduces rapidly due to the expansion in the flow and the flow later tries to recover. The sudden pressure jump in the cylinder forebody indicates the presence of transonic shock for the Mach numbers 0.7 to The location of shock move backward towards the boat tail region with increase in Mach number as shown above. It can be seen that the supersonic region increases with increase in free stream Mach number and the shock moves downstream in the Mach number range of 0.95 to 1.2. This can be seen in the Mach distribution shown in Fig 5. The flow expansion on the cone-cylinder junction is small for M=0.7 and expansion increases with Mach number up to 1.2. Fig 5: Mach number distributions over Heatshield Copyright 2013 Published by IJESR. All rights reserved 297

4 Flow separation zones are obtained for all Mach numbers. It depicts that for low Mach number of transonic regime there is no flow separation i.e. for M=0.7, 0.75 and 0.8. For Mach numbers 0.85 and 0.9 the flow separation is confined to a short distance and reattaches immediately on the cylinder portion and for Mach number 0.95, flow separation is found near the boat tail region. There is no flow separation found for Mach numbers 1.0 to 1.2. The shock is formed at the cylindrical portion for Mach numbers 0.7 to 0.9 and for Mach numbers 0.95 to 1.2 the shock is moved towards the boattail region. The flow separation distance for the respective Mach numbers is tabulated in Table 2. Table 2: Separation Length Mach no On Cylinder On Boattail D D D Furthermore, strength of shock is computed, and these are studies for shock wave analysis. For obtaining shock strength, pressure rise across the shock is derived from CFD results i.e. the pressure downstream of the shock and upstream of the shock is noted. Thus the shock strength is found by using the relation, where, P 1 = Upstream pressure of shock P 2 = Downstream pressure of shock The shock strength is tabulated in Table 3. The strength of shock is lower for Mach number 0.7 when compared with Mach number It can be seen from the table that, the shock strength decreases from Mach numbers 0.75 to Table 3: Shock strength 4. CONCLUSION Mach numbers Shock strength For Payload Fairing configuration with 20 nose cone and 20 boattail, simulations are carried out in transonic Mach numbers ranging from 0.7 to 1.2 in steps of The following observations have been made. The shock formation moves towards the aftbody of the cylinder as the free stream Mach number increases up to 1.2. The separation is found for Mach numbers of 0.85, 0.9 and 0.95, which is a lower transonic Mach number. For higher transonic Mach numbers there is only the formation of shock but there is no shock induced separation found over the body. The surface pressure distribution shows that there is a formation of expansion waves near the cone-cylinder region. As the Mach number increases the expansion wave also increases. The strength of shock wave varies with the Mach numbers. Copyright 2013 Published by IJESR. All rights reserved 298

5 REFERENCES [1] Jameson A. Efficient Aerodynamic Shape Optimization. 10 th AIAA/ISSMO Multidisciplinary Analysis and Optimization Conference, Albany, New York, [2] Dimitri E, Bigarella V. Centered and UpwingMultigrid turbulent Flow Simulations of Launch Vehicle Configurations. Journal of Spacecraft and Rockets 2007; 44(1). [3] Newman JC. Overview of Sensitivity Analysis and Shape Optimization for Complex Aerodynamic Configurations. Journal of Aircraft 1999; 36(1). [4] Mehta RC. Transonic Flow Simulation for a Bulbous Heat Shield. Journal of Spacecraft and Rockets 1997; 34(4): [5] Menter FR. Performance of Popular Turbulence Models for Attached and Separated Adverse Pressure Gradient Flows. AIAA Journal 1992; 30(8). [6] Newman JC, Taylor AC, Barnwell RW, Newman PA, Hou GJW. Overview of Sensitivity Analysis and Shape Optimization for Complex Aerodynamic Configurations. Journal Aircraft 1999; 36(1): [7] Reisenthel PH, Childa RE, Higgins JE. Surrogate-Based Design Optimization of a Large Asymmetric Launch Vehicle Payload Fairing. 45 th AIAA Aerospace Science Meeting and Exhibit 8-11 Jan, 2007, Reno, Nevada. [8] Spalart PR, Allmaras SR. A One-Equation Turbulence Model for Aerodynamic Flows. 30 th Aerospace Sciences Meeting and Exhibit, Jan 6-9, 1992, AIAA [9] Krivanek TM, Yount BC. Composite Payload Fairing Structural Architecture Assessment and Selection. [10] Ochinero T, Deiters T, John Higgins PE, Arritt B, Blades E, Newman J. Design and Testing of a Large Composite Asymmetric Payload Fairing, 50 th AIAA Structural Dynamics and materials conference 17 th, 4 7, Copyright 2013 Published by IJESR. All rights reserved 299