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30720130101005

  1. 1. Journal of Mechanical Engineering and Technology (JMET), ISSN 2347-3924 (Print), ISSN 2347-3932 (Online), Volume 1, Issue 1, July -December (2013) 54 COMPUTATIONAL ANALYSIS OF F LOW BEHAVIOR OVER THE MULTISTAGE LAUNCH VEHICLE WITH STRAPONS SIVARAJ G1 , K.M. PARAMMASIVAM2 , M.GOKULRAJ 3 1 Department of Aeronautical Engineering, Bannari Amman Institute of Technology, Sathyamangalam-638402 2 Department of Aerospace Engineering, Madras Institute of Technology, Chennai-600036, India 3 Department of Aeronautical Engineering, Bannari Amman Institute of Technology, Sathyamangalam-638402, India, ABSTRACT Technology has the enemy of nature in one way. But sometimes technologies do come out as an exception to the above rule. In this paper conclude that multi-stage launch vehicle with strapons is a complex configuration to know the flow behaviour over it. Generally extensive wind tunnel testing is done to understand the flow characteristics of such a configuration with the Computation Fluid Dynamics (CFD) as a design tool, it is appropriate to make use of its technology to understand the complex flow behaviour over a multi-stage launch vehicle with strapons. In the present paper conclude the flow behaviour over a typical multi stage launch vehicle with strapons was known using commercial CFD software. This involves choice of flow model, discretization, grid generation, solution technique and analysis of results. Grid generation and body shape generation are done in Structured and an unstructured grid on 2D, and it is generated to know the effect of flow behaviour. Both Euler and Navier-Stokes solvers are attempted. Sensitivity of results on turbulence models is analyzed. Keywords: Computation Fluid Dynamics, flow behaviour, grid generation, strapons, 1. INTRODUCTION In spaceflight, a launch vehicle is a rocket used to carry payloads from the Earth's surface into outer space. A launch system includes the launch vehicle, launch pad and other infrastructure. Usually the payload is a satellite placed into orbit, but some spaceflights are sub-orbital while others enable spacecraft to escape Earth orbit entirely. A launch vehicle which carries its payload on a suborbital trajectory is often called a sounding rocket. JOURNAL OF MECHANICAL ENGINEERING AND TECHNOLOGY (JMET) ISSN 2347-3924 (Print) ISSN 2347-3932 (Online) Volume 1, Issue 1, July-December (2013), pp. 54-65 © IAEME: http://www.iaeme.com/JMET.asp JMET © I A E M E
  2. 2. Journal of Mechanical Engineering and Technology (JMET), ISSN 2347-3924 (Print), ISSN 2347-3932 (Online), Volume 1, Issue 1, July -December (2013) 55 A space launch vehicle, during its atmospheric flight, presents a variety of aerodynamic problems for which solutions are to be obtained through analytical and Referencetechniques. Generally, the problems are complex and three dimensional in nature and quite often involve multi body interactions etc. Currently, a large amount of research work is going on to develop a reusable launch system i.e. a vehicle which is capable of launching into space more than once. This contrasts with expendable launch systems, where each launch vehicle is launched once and then discarded. The orbiter, which includes the main engines, and the two solid rocket boosters, are reused after several months of refitting work for each launch. The external fuel drop tank is however discarded. The present work can be carried out by using wind tunnel for conducting experiments to know the flow past launch vehicles and to know the pressure distribution over the surfaces or by using the commercial software’s for CFD applications. One such software Fluent is used for this present problem. The benefits involved are the use of computer-based computational fluid dynamics methods which will accelerate the design process, reduce preliminary development testing, and help create reliable, high-performance designs of space launch vehicles and their components. In addition to design verification and optimization, CFD can be used to simulate anomalies that occur in actual space vehicle tests or flights to fully understand the anomalies and how to correct them. The result is a more reliable and trouble-free space vehicle. A booster rocket is either the first stage of a multi-stage launch vehicle, or else a strap-on rocket used to augment the core launch vehicle's takeoff thrust and payload capability. Boosters are generally necessary to launch spacecraft into Earth orbit or beyond. Strap-on boosters are sometimes used to augment the payload or range capability of jet aircraft. For thorough understanding of the complex flow field around typical Space launch vehicles at zero angle flight, axi-symmetric computational simulation of the flow field can be made useful along with the Referencetesting. This can be done either by developing a code or by available commercial CFD software. In the present study computation has been attempted for flow over a typical SLV model with the commercially available software FLUENT 6.3.26 2. COMPUTATIONAL SETUP 2.1 Grid Generation The grid for the typical SLV model was generated using the GAMBIT software. Structured and unstructured grid was used for the analysis of the flow field around the model. 2.2 Grid Generation for Space Launch Vehicle Model A structured grid was generated for 2D simulation of flow around the SLV model. The grid was made very fine at the geometry surfaces and coarsens away from the body. The overall domain was selected based on several iterations, boundary conditions and finally a domain extending 5 times the major length of the SLV model (length L) ahead of the nose center, 5 times the length to the center line of the geometry and five times length behind the geometry was chosen. The extents of the domain were evaluated from inviscid simulation of the problem. A total of 50,000 cells were used in the grid system. The grid system with boundary conditions, in the vicinity of the geometry and a close up surface grid near main body and strap-on nose regions respectively. In order to capture the shocks more accurately, finer mesh cells were created near the surface of the model using appropriate edge mesh distribution.
  3. 3. Journal of Mechanical Engineering and Technology (JMET) ISSN 2347-3932 (Online), Volume 1, Issue 1, July Fig. 1 Fig. 2 Grid cell distributions at the nose of the main body Fig. 3 Journal of Mechanical Engineering and Technology (JMET), ISSN 2347 , Volume 1, Issue 1, July -December (2013) 56 Fig. 1 2D unstructured grid for complete body Grid cell distributions at the nose of the main body Fig. 3 2D structured grid for complete body ISSN 2347-3924 (Print),
  4. 4. Journal of Mechanical Engineering and Technology (JMET), ISSN 2347-3924 (Print), ISSN 2347-3932 (Online), Volume 1, Issue 1, July -December (2013) 57 Fig. 4 Close view of 2D structured grid for complete body Firstly 2D unstructured grid was generated, total of 50,000 cells were used in the grid system. Figures 1, 2, 3 & 4 show the grid system with boundary conditions, in the vicinity of the geometry and a close up surface grid near main body and strap-on nose regions respectively. In order to capture the shocks more accurately, finer mesh cells were created near the surface of the model using appropriate edge mesh distribution. But when it is iterated for solution in fluent, reverse flow was encountered. This resulted in change of grid. Structured grid was generated with approximate No. of cells are 64500. This grid was shown in Figure 4.The smallest size of grid cell chosen is 2.730508e-005. For this grid the extent of outer domain for the front side is 0.25 times the length of the body and on the rear side it is 5 times the length of the body, on the top side also it is five times length of the body. When this grid was initialized for the simulation, this showed a positive trend and all the residuals converged. 2.3 Solution Methodology Using Fluent The solution method in FLUENT can be broadly divided into three parts namely: Pre – processing, Solver and Post processing. Pre – processing of the problem was done in GAMBIT as discussed in detail in the preceding sections. Once the problem is meshed and the boundary conditions are specified the meshed geometry is then imported as a ‘mesh file’ into FLUENT. FLUENT uses a control-volume-based technique to convert the following governing equations as conservation of mass, conservation of momentum and conservation of energy to algebraic equations that can be solved numerically. This control volume technique consists of integrating the governing equations about each control volume, yielding discrete equations that conserve each quantity on a control volume basis. FLUENT has two solvers: Segregated solver and Coupled solver. Using either method, FLUENT will solve the governing integral equations for the conservation of mass and momentum, and (when appropriate) for energy and other scalars such as turbulence and chemical species. In both cases a control-volume-based technique is used that consists of: Division of the domain into discrete control volumes using a computational grid, Integration of the governing equations on the individual control volumes to construct algebraic equations for the discrete dependent variables such as velocities, pressure, temperature, and conserved scalars and Linearization of the discretized equations and solution of the resultant linear equation system to yield updated values of the dependent variables.
  5. 5. Journal of Mechanical Engineering and Technology (JMET), ISSN 2347-3924 (Print), ISSN 2347-3932 (Online), Volume 1, Issue 1, July -December (2013) 58 The segregated solver is the solution algorithm in which, the governing equations are solved sequentially (i.e., segregated from one another). The coupled solver solves the governing equations of continuity, momentum, and (where appropriate) energy and species transport simultaneously (i.e., coupled together). Governing equations for additional scalars will be solved sequentially (i.e., segregated from one another and from the coupled set) using the procedure described for the segregated solver. Because the governing equations are non-linear (and coupled), several iterations of the solution loop must be performed before a converged solution is obtained. Each iteration consists of the steps outlined below: Fluid properties are updated based on the current solution (If the calculation has just begun the fluid properties will be updated based on the initialized solution). The continuity, momentum, and energy and species equations are solved simultaneously. Where appropriate, equations for scalars such as turbulence and radiation are solved using the previously updated values of the other variables. A check for convergence of the equation set is made. These steps are continued until convergence criteria are met. In both the segregated and coupled solution methods the discrete, non-linear governing equations are linearized to produce a system of equations for the dependent variables in every computational cell. The resultant linear system is then solved to yield an updated flow-field solution. The manner in which the governing equations are linearized may take an ‘Implicit’ or ‘Explicit’ form with respect to the dependent variable (or set of variables) of interest. 2.4 Solution Methodologies For The Typical SLV Model The steps of setting up a problem in FLUENT are discussed briefly: defining geometry, importing and checking the grid, selection of solver formulation and equations to be solved (laminar/turbulent/inviscid etc.), material properties, specification of operating and boundary conditions, specification of numerical properties (under-relaxation factors, CFL etc.) and initialization of variables. The criterion for convergence was in the order of 10-3 for continuity, x, y and z velocities and 10-5 for energy calculation and turbulence quantities. The residuals were monitored in the graphics window of FLUENT. In addition the net mass flow rate was monitored for convergence. For all the cases iterations continued till the convergence or near convergence were reached. 2.5 Solver Settings Computations were carried out with double precision 2D models with steady, coupled, explicit solver scheme. The viscous model chosen for the problem was the standard Spalart –Allmaras model with turbulent intensity and viscosity ratio as inputs. Standard wall functions were used for the near wall treatment of the flow. 2.6 Materials Selection For the present simulations, Ideal gas condition was used. The ideal gas law for compressible flows was used for air. Since the flow was dependent on temperature, the Sutherland viscosity model with three coefficients was used. Sutherland’s viscosity law resulted from a kinetic theory by Sutherland (1893) using an idealized intermolecular-force potential.
  6. 6. Journal of Mechanical Engineering and Technology (JMET), ISSN 2347-3924 (Print), ISSN 2347-3932 (Online), Volume 1, Issue 1, July -December (2013) 59 2.7 Operating Conditions The input of the operating pressure is of great importance when density with the ideal gas law is being computed. The criteria for choosing a suitable operating pressure are based on the Mach-number regime of the flow and the relationship used to determine density. For the present case where ideal gas law is used and the flow Mach number is greater than 0.1 the operating pressure is set to 101325. 2.8 Boundary Conditions Pressure far field boundary condition was set for the inflow (surface facing the flow), where it is need to specify the free stream static pressure and Mach number. Pressure outlet boundary condition was set to the out flow (surface from which the flow leaves), where the variable will be extrapolated from the interior cells. Wall boundary condition is assigned for the model surfaces and domain boundary extents other than inflow and outflow. 2.9 Post processing The simulation setup can be stored as case and data file. The auto save option can be used to save the results of the iterations from step to step. The case file includes the information for the grid, the boundary conditions and the solver settings. The data file stores information about the data in each node of the cells. The contour plots, vector plots and the surface data plots etc. of pressure, velocity and density etc. can be checked during the solution process and at convergence. These plots can be saved as image files and the data from the surface plots can be written on to a file and plotted. Points, lines, rakes and planes can be created in the flow domain to analyze the properties at the desired locations. FLUENT offers a very good range of post processing options which can be used to analyze the computational data, compare computational results with the Referenceresults as desired. For the present analysis points, lines and planes were used to analyze the properties at the desired locations in the flow domain. Multigrid option was also used to reduce the computational time. Computations were performed to understand the flow field around a scaled down model of a typical space launch vehicle with strapons. Computations using the commercially available software FLUENT 6.3.26 were carried out for two dimensional and three dimensional fully developed flows. A validation test was performed by referring to the same type of model for the same Mach number. 3. RESULT AND DISCUSSION Two dimensional computational simulations were carried out for studying the flow field around typical space launch vehicle geometry at supersonic Mach number. The results for the computations performed are discussed in detail in the following sections. The computations were performed on a work station Core 2 Duo processor with a bus speed of 2.0GHz and a RAM of 2GB. Typical times taken for inviscid problem were 250 iterations per hour and for viscous three dimensional problems were 150 iterations per hour. For all the cases, an average 3000 iterations were performed until the desired convergences are obtained. The inviscid analysis was performed for comparing the results obtained for viscous flows. 3.1 Validation of Computational Procedure Verification and validation are the primary means to assess accuracy and reliability in computational simulations. Several computational tests were performed on a typical space
  7. 7. Journal of Mechanical Engineering and Technology (JMET), ISSN 2347-3924 (Print), ISSN 2347-3932 (Online), Volume 1, Issue 1, July -December (2013) 60 launch vehicle at a Mach number of 2 for verifying and validating the computational grid and turbulence model that are going to be adopted for the present work. Inviscid, laminar, S-A, standard k-ε and standard k-ω models were tested for verifying the most suitable turbulence model for present case of a typical launch vehicle configuration. The solver settings operational conditions, material properties, and boundary conditions were set according to the present typical space launch vehicle problem. The problem was iterated till the residuals of continuity, momentum, and energy converged to a value of 10-3 and the scalars nut (SA) residuals converged to a value of 10-5 . After importing this grid to Fluent, all the residuals show a converging trend and matches with the Referenceresults. Hence this grid was tried for the selection of suitable turbulent model. Trials were made for Inviscid, Laminar, K Epsilon, K Omega and SA models. For K Epsilon case Cp Vs X/L plot, as shown in Figure 6a, Cp distribution varies from the Referenceresults of Reference and does not follow the trend. Density contours captured are plotted in Figure 6b. The main body region density distribution is higher near the nose and the shock formation near the nose of the booster is invisible. For K Omega case Cp Vs X/L plot, as shown in Figure 7a, Cp trend near the main body region is good when compared to the Referenceresults. But the Cp distribution near the location of the booster is much higher when compared with Referencevalue as well as numerical value. When density contours are observed, as given in Figure 7b, it can be clearly visualized that near the booster shock wave in front of the booster is standing at a distance. For SA method, Cp Vs X/L plot, as shown in Figure 8a, Cp trend near the main body region is good when compared to the Referenceresults. Near the booster location, Cp distribution is lesser than the Referenceresults but the nature of the curve is good. The variation of the Cp distribution near the booster location peak with the Density contours are given in Figure 8b, this contours gives the appropriate reasons for choosing it as the turbulent model. One can clearly observe the shock wave near the nose cone of the main body, high density region and near the booster also a shock wave can be visualized. After the solution was obtained from the different models, the pressure coefficients on the main body of the space launch vehicle were plotted and compared with the Referenceresults from reference Scalabrin et al. It is observed that standard SA model with appropriate turbulence specification method, turbulent intensity and turbulent viscosity ratio is most suitable model producing approximate results as of the Referenceresults. Figure 9 shows the comparison for Cp distributions obtained from the different models. From this plot we can see SA model results match well with Referenceresults. Spalart Allarmas method is chosen for the detail study, because of its simplicity. . High density is attained at the nose of the main body, shock wave is observed near the nose. Low density region near the boat tail. Circulation is observed near the booster nose. Shock wave formation at the nose of the booster can be seen. The entire flow phenomenon is visualized in these contours. Though flow behavior is captured through 2D simulations, Cp distribution is not following the Referenceresults trend.
  8. 8. Journal of Mechanical Engineering and Technology (JMET), ISSN 2347-3924 (Print), ISSN 2347-3932 (Online), Volume 1, Issue 1, July -December (2013) 61 Fig. 5 Pressure coefficient distributions on a vehicle core in a plane between two boosters for Mach 2 and zero angle of attack Ref. Scalabrin et al Fig. 6a Cp Vs X/L for K Epsilon case Fig. 6b Density contours for K Epsilon case -1 -0.5 0 0.5 1 1.5 0 0.5 1 1.5 Cp X/L Cp Vs X/L K Epsilon Cp
  9. 9. Journal of Mechanical Engineering and Technology (JMET), ISSN 2347-3924 (Print), ISSN 2347-3932 (Online), Volume 1, Issue 1, July -December (2013) 62 Fig. 7a Cp Vs X/L for K Omega case Fig. 7b Density contours for K Omega case Fig. 8a Cp Vs X/L for SA method -1 -0.5 0 0.5 1 1.5 2 0 0.5 1 1.5 Cp X/L Cp Vs X/L k Omega Cp -0.2 0 0.2 0.4 0.6 0.8 1 0 0.5 1 1.5 Cp X/L Cp Vs X/L SA Cp
  10. 10. Journal of Mechanical Engineering and Technology (JMET), ISSN 2347-3924 (Print), ISSN 2347-3932 (Online), Volume 1, Issue 1, July -December (2013) 63 Fig. 8b Density contours for SA method Fig. 9 Cp comparisons with the reference plot 4. CONCLUSION Computational studies were carried out to get an understanding of the flow field around typical space launch vehicle with strapons at Mach 2. Two dimensional simulations of the flow field using FLUENT 6.3.26 were performed. SA turbulent model was adopted to capture the flow field. Computations were validated through a simulation of flow field around the similar geometry at a Mach number 2 by earlier investigators. After a good agreement with reported results, simulation of the present case was carried out and compared with available experiments. The following important observations were made from the results obtained through computations and experiments: 1. The basic flow structure around a typical launch vehicle with strapons was captured through 2D computations. -1 -0.5 0 0.5 1 1.5 2 0 0.5 1 1.5 Cp X/L Cp vs X/L Ref plot K Epsilon K Omega SA Laminar Inviscid
  11. 11. Journal of Mechanical Engineering and Technology (JMET), ISSN 2347-3924 (Print), ISSN 2347-3932 (Online), Volume 1, Issue 1, July -December (2013) 64 2. Mach and density contours showed the bow shock wave structure, near the nose cone of the main body and near the booster nose. Shock and boundary layer interferences were observed near the booster nose. 3. Comparison of SA, k-ε and k-ω turbulence modeled viscous simulations around the typical space launch vehicle showed that the SA model predicts well for the flow field features of wall bounded shear flows. 4. A good comparison of computations and available Referenceresults were achieved. ACKNOWLEDGMENT We first thank our ‘GOD’, the supreme power for giving us a good knowledge and our parents for making us study in a renowned college we owe a great many thanks to my colleagues and friends for their help and encouragement. REFERENCES [1] Seiji Tsutsumi, Taro Shimizu, Ryoji Takaki, Eiji Shima, and Kozo Fujii., “ Numerical Study of Pressure Waves Generated byH-IIA Launch Vehicle at Lift Off”, Japan Aerospace Exploration Agencies , March 2008 [2] Enda Dimitri V. Bigarella, Jo ao Luiz F. Azevedo., “Numerical Study of Turbulent Flows over Launch Vehicle Configurations” Journal of Spacecraft and Rockets, Vol. 42, No. 2, March–April 2005 [3] E. Basso, J. L. F. Azevedo., “Three-Dimensional Viscous Flow Simulations over the VLS Using Overset Grids”, Journal of the Brazilian Society of Mechanical Sciences & Engineering, Vol. XXVI, No. 4, October-December 2004 [4] P. Moraes Jr and Pereira A. L., “Verification of the Pressure Equalization Inside the Satellite Compartment of the Brazilian Satellite Launch Vehicle”, Journal of the Brazilian Society of Mechanical Sciences & Engineering, Vol. XXVII, No. 4, October-December 2005 [5] Scalabrin L. C., Azevedo J. L. F.,. Teixeira P. R. F and Awruch A. M.,” Three Dimensional Flow Simulations with the Finite Element Technique over a Multi- Stage Rocket”, Journal of the Brazilian Society of Mechanical Sciences & Engineering, Vol. XXVI, No. 2, April-June 2004 [6] Enda Dimitri V. Bigarella, Jo ao Luiz F. Azevedo., “ A Numerical Study of Turbulent Flows over Launch Vehicle Configurations”, 41st Aerospace Sciences Meeting and Exhibit, 6-9 January 2003 [7] Alexandre P. Antunes, Edson Basso and Joao Luiz F. Azevedo, “chimera simulations of viscous flows over a complex Satellite launcher configuration”, 19th Applied Aerodynamics Conference, 11 -14 June 2001 [8] Rogerio Pirk , Wim Desmet , Bert Pluymers, Paul Sas and Luis C. S. Goes., “Vibro- acoustic Analysis of the Brazilian Vehicle Satellite Launcher (VLS) fairing”, proceedings of ISMA, volume V, 2000 [9] T. Shivananda, S. McKeel, M. Salita, and E. Zabrensky, “Space Launch Vehicle Aerodynamics Comparison of Engineering and CFD Predictions with Wind Tunnel Data”, 39th Aerospace Sciences Meeting & Exhibit8-11 January 2001
  12. 12. Journal of Mechanical Engineering and Technology (JMET), ISSN 2347-3924 (Print), ISSN 2347-3932 (Online), Volume 1, Issue 1, July -December (2013) 65 [10] K. P. Singh, T. S. Prahlad, “Numerical Simulation of Inviscid Supersonic Flow Over a Launch Vehicle With Strap-On Boosters”, AIAA 25th Aerospace Sciences Meeting, January 12-15, 1987 [11] T. P. Shivananda, E. F. Zabrensky, S. A. McKeel, M. L. Papay, “comparison of engineering and CFD predictions and wind tunnel data for a launch vehicle Configuration”, AIAA- 97-2251, 1997 [12] Jerry E. Deese, D. L. Pavish, Jerry G. Johnson, and Bharat K. Soni, “Flow field Predictions for Multiple Body Launch Vehicles”, AIAA 10th Applied Aerodynamics Conference, June 22-24, 1992 [13 ] S. K. Chakrabartty*, K. Dhanalakshmi and J. S. Mathur, “Computation of flow past aerospace vehicles”, Computational and Theoretical Fluid Dynamics Division, National Aerospace Laboratories [14] Robert L. Stallings Jr., “Reynolds number effects on aerodynamic characteristics at large angles of attack”, Journal of Spacecraft, Vol. 17, No 2, Mar-Apr 1980 [15] Jerry M.Allen, “Vortex Development on slende missiles at supersonic speeds”, Journal of Spacecraft, Vol.17, No4, july-Au 1980 [16] Yanta W.J. and Wardlaw A.B., “Flow Field about and Forces on Slender Bodies at High Angles of Attack”, AIAA Journal, Vol. 19, No 3, March 1981, pp. 296- 302. [17] “Computational fluid dynamics (CFD) in launch vehicle applications”, PRACTICE NO. PD-AP-1311, NASA. [18] Mehta, R.C., “Computational Investigation of Flow Oscillations Over Reentry Capsules”., Computational Fluid Dynamics Journal 15(2):34, pp. 247-260, Jun-2006. [19] Degani D. and Zilliac G.G. , “ ReferenceStudy of Non Steady Asymmetric Flow Around on Ogive Cylinder at Incidence”, AIAA Journal, Vol. 28, No 4, 1990, pp 642- 649. [20] Smith E.H., Hebbar S.K. and Platzer M.F., “Aerodynamic Characteristics of a Canard Controlled Missile at High Angles of Attack”, Journal of Spacecraft and Rockets, Vol. 31, No.5, Sep-Oct 1994, pp 766-772. [21] Vashchenkov, P., Kashkovsky, A., Ivanov, M and Krylov, A., “Numerical Analysis of High Altitude Aerodynamics of Reentry Vehicles”, AIAA 2005-3409, 2005. [22] Hamdi T. Hemdan, “Similarity solutions for hypersonic flow past slender bodies of revolution at small incidence”, Acta Astronautical Vol. 35, No. 8, PP. 501-508, 1995.

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