30720130101005

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. PARAMMASIVAM

2, M.GOKULRAJ

3

1 Department of Aeronautical Engineering,

Bannari Amman Institute of Technology, Sathyamangalam-638402 2Department of Aerospace Engineering,

Madras Institute of Technology, Chennai-600036, India 3Department 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



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.

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),



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.



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.



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



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.



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



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



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



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



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.

30720130101005

Technology

multistage launch vehicle

typical space launch

reusable launch system

expendable launch systems

troublefree space vehicle

complex flow behaviour

core launch vehicles

technology jmet issn