Angelo A. Mustto - mustto@serv.com.ufrj.br
Gustavo C.R. Bodstein - gustavo@serv.com.ufrj.br
Universidade Federal do Rio de Janeiro, Departamento de Engenharia Mecânica
Cx. P. 68503 - 21945-970 - Rio de Janeiro, RJ, BRASIL

In this paper the Particle-Box Multipole Expansion Algorithm (PB) is used to calculate particle interactions and the results are compared to the direct Vortex-Vortex Biot-Savart Algorithm (VV). In a system with N particles, the PB scheme requires N 2 operations, whereas the PB scheme requires only NlogN. The PB scheme may replace the VV scheme in the Vortex Method Algorithm to compute the convection of discrete vortices that model the vorticity in 2-D, incompressible, unsteady flows around bodies. Two fluid domains are studied: in the first, the vortices are positioned randomly exterior to a circle and interior to a square region; in the second, the vortices lie in the wake region of a circular cylinder. The presence of the cylinder generates image vortices in its interior. The influence of the numerical parameters intrinsic to the PB scheme, such as the number of vortices, the number of boxes and the number of vortices per box, is studied for both domains; the numerical error and the speed-up obtained to calculate the vortex induced velocities are also determined.

Keywords: Vortices, Multipole expansion, Biot-Savart law, Circle theorem
 
 

Victor Koldaev - koldaev&iae.cta.br
Instituto de Aeronáutica e Espaço, Centro Técnico Aeroespacial, Divisão de Sistemas Espaciais, 12228-904, São José dos Campos, SP, Brasil

Maurício Guimarães da Silva - mgsilva@lcca7.feg.unesp.br
Universidade Estadual de São Paulo, Faculdade de Engenharia de Guaratinguetá, Av. Agenor Pires da Fonseca, 129, 12500000, Guaratinguetá, SP, Brasil.

The process of design of parachute recovery system for sub-orbital or orbital platforms includes a numerical technique application to predict the flow-fields around parachute and aeroelastic effects during space motion of system. This work presents mathematical description and numerical solution of parachute interaction with medium during its unsteady space motion along trajectory with the help of equations of ballistics, aerodynamics, material elasticity and canopy form-shaping. Presented solution takes into consideration the aerodynamic analysis with the 3D-flow simulation over the surface of canopy. The emphasis of the solver aerodynamic is on multiblock. Within each block, the governing equations were discretized using the Beam Warming implicit approximate factorization algorithm. The implicit Euler method is adopted for the time march and second order accurate central difference formulas are used to approximate the spatial derivatives that appear in the governing equations. The proposed method makes it possible to simulate the aeroelastic parachutes functioning. The example illustrates its wide latitude.

Keywords: System ballistics, Parachute elasticity, Aerodynamic analysis.
 
 

Daniel Strauss - daniels@aer.ita.cta.br
Centro Técnico Aeroespacial, Instituto Tecnológico de Aeronáutica
CTA/ITA/IEAA, 12228-900 - São José dos Campos - SP - Brasil

João Luiz F. Azevedo - azevedo@iae.cta.br
Centro Técnico Aeroespacial, Instituto de Aeronáutica e Espaço
CTA/IAE/ASE-N, 12228-904 - São José dos Campos - SP - Brasil

The paper describes the implementation details and validation results for an agglomeration multigrid procedure developed in the context of hybrid, unstructured grid solutions of aerodynamic flows. The procedure is applied to the solution of 2-D, laminar viscous flows of aerospace interest. The governing equations are discretized with an unstructured grid finite volume method which is capable of handling hybrid unstructured grids. Time march uses an explicit, 5-stage, Runge-Kutta time-stepping scheme. Convergence acceleration to steady state is achieved through the implementation of an agglomeration multigrid procedure which retains all the flexibility previously available in the unstructured grid code. The calculation capability created is applied to 2-D laminar viscous flows of aerospace interest. Studies of the various parameters affecting the multigrid acceleration performance are undertaken with the objective of determining optimal numerical parameter combinations.

Keywords: Agglomeration multigrid, Convergence acceleration, Unstructured grids, Finite volume Methods.
 
 

Enda Dimitri Vieira Bigarelli { biga@h8.ita.br
Centro T_ecnico Aeroespacial, Instituto Tecnol_ ogico de Aeron_autica
CTA/ITA/IEAA { 12228-900 { S~ ao Jos_e dos Campos, SP, Brasil

Lucas Rubiano de Souza Cruz { rubiano@h8.ita.br
Centro T_ecnico Aeroespacial, Instituto Tecnol_ ogico de Aeron_autica
CTA/ITA/IEAA { 12228-900 { S~ ao Jos_e dos Campos, SP, Brasil

Jo~ ao Luiz F. Azevedo { azevedo@iae.cta.br
Centro T_ecnico Aeroespacial, Instituto de Aeron_autica e Espa¢co
CTA/IAE/ASE-N { 12228-904 { S~ ao Jos_e dos Campos, SP, Brasil

This work presents a study of VLS aerodynamics which uses the capability implemented at IAE to simulate 3-D ows over typical launch vehicle con_gurations at angle of attack. This capability is further used to determine normal loads over the VLS main body con_guration at an angle-of-attack ight. The numerical simulations performed use the compressible Euler formulation, discretized in a _nite di_erence context for general curvilinear coordinates. A 5-stage, explicit Runge-Kutta time-march procedure is used and the spatial discretization employs central di_erences. Numerical results are compared with available experimental data for the VLS and they are used to assess the aerodynamic characteristics of the VLS central body at angle of attack.

Keywords: Launch vehicle, CFD, Three-dimensional ow, Flow at angle of attack, Supersonic ow.
 
 

Horacio P. Burbridge - willybur@ssdnet.com.ar
Armando M. Awruch - amawruch@adufrgs.ufrgs.br
Programa de Pós-graduação em Engenharia Civil, Universidade Federal do Rio Grande do Sul, Av. Osvaldo Aranha 99, 3° Andar, Porto Alegre, RS, Brasil.

An algorithm to simulate 3-D high compressible flows of viscous and non-viscous fluids is presented in this work. The time integration procedure was obtained from an expansion in Taylor series of the governing equations, while spatial discretization was carried out using the Finite Element Method (FEM) based on the classical Bubnov-Galerkin technique. In order to obtain considerable improvements in CPU time and memory, and to take advantage from the fast vectorial processors existing in modern supercomputers, an analytical evaluation of element matrices was adopted, deriving the corresponding expressions from the eight-node isoparametric brick element. One practical example is also presented in this paper in order to show the excellent computational performance and the good agreements with results obtained previously by other authors.

Keywords: Computational fluid dynamics, Compressible flows, Finite element method,Taylor-Galerkin.
 
 

João Batista Aparecido - jbaparec@dem.feis.unesp.br
Edson Del Rio Vieira - delrio@dem.feis.unesp.br
João Batista Campos Silva ? jbcampos@dem.feis.unesp.br
Universidade Estadual Paulista - Unesp - Departamento de Engenharia Mecânica
Cx. P. 31 ? 15385-000 - Ilha Solteira, SP, Brasil

The detailed knowledge of a flow field inside an incompressible subsonic wind or water tunnel contraction is of interest in designing this kind of equipment due to the needing to guarantee some flow characteristics such as: low turbulence level, small thickness of boundary layer and uniform velocity profile at inlet of the test section. Several approximate methods have been proposed in the technical literature to obtain two-dimensional flow fields inside axis-symmetric wind tunnel contraction. The purpose of this work is to apply the generalized integral transform technique (GITT) to analyze a low speed wind tunnel contraction with rectangular cross section by solving the streamline equation for the flow in a known axially non-symmetric contraction geometry.

Keywords: Contraction, Generalized integral transform technique, Hydrodynamic tunnel.
 
 

João B. P. Falcão F. - joaobpff@uol.com.br
Centro Técnico Aeroespacial (CTA)
Instituto de Aeronáutica e Espaço (IAE), Divisão de Sistemas Aeronáuticos (ASA-L)
Pça. Marechal Eduardo Gomes, 50
CEP: 12228-904, São José dos Campos, São Paulo

Proceeding with solutions to make it possible to install a large wind tunnel facility in the CTA campus (Aeronautical Technical Center in São José dos Campos), environmental noise is analyzed. The proposed wind tunnel presents a 2,0 x 2,4 m 2 test section, with Mach number range from 0.2 to 1.3. The total installed power is 70 MW and the wind tunnel energy generation system comprises one main compressor, many auxiliary power systems, dryers, cooling tower, associated pumps, and an injection system. All these components are large noise sources. Since the wind tunnel installation is designed to be close to other researches sectors and not far from the residential area with a hospital in the Campus, the environmental noise level evaluation is very important. In this aim, it was developed a simplified mathematical model to determine local noise levels, considering all sources of noise in the site, together with factors of attenuation and amplification (distance from the sources, wave reflection, the presence of the atmosphere, used materials in the buildings, etc.). Some different scenarios are also analyzed in order to determine best solutions.

Keywords: wind tunnel, noise, environmental, mathematical model
 
 

Dong Ho Choi - dchoi@usp.br
Universidade de São Paulo, Depto. de Engenharia Mecatrônica e Sistemas Mecânicos
Av. Prof. Mello Moraes, 2231 - São Paulo, SP 05508-900 BRASIL

The free expansion of Argon injected into a vacuum chamber is simulated using the Direct Simulation Monte-Carlo method. Domains with dimensions up to ten times the hole diameter were simulated with uniform cells of 40 µm length. The flow is shown to start subsonic, reaching the sonic condition at the exit of the injection hole and continuing to accelerate into the supersonic range. No Mach disks were observed due to the reduced size of the simulation domain. The decay of density along the center line of the injection axis is compared with available experimental data to validate the code. Axisymmetric 2-D distributions of density, temperature, pressure, Mach number, Knudsen number are plotted for some typical conditions. The pressure and the maximum Mach number scale directly with the mass flow rate. The Knudsen number scales inversely with the mass flow rate. Depending on the mass flow rate, the flow regime can traverse the whole range from continuum to free molecular. A discussion is made of the applicability of the continuum and equilibrium hypotheses in the simulation of rarefied gas flows in vacuum chambers.

Keywords: Rarefied, Gasdynamics, Supersonic, Expansion, Simulation
 
 

Paulo G.P Toro - (toro@iae.cta.br)
Instituto de Aeronaútica e Espaço - IAE
Centro Técnico Aeroespacial - CTA - São José dos Campos - SP 12228-904 - BRAZIL

Marco A.S. Minucci - (sala@ieav.cta.br)
Instituto de Estudos Avançados - IEAv
Centro Técnico Aeroespacial - CTA - São José dos Campos - SP 12231-970 - BRAZIL

Leik N. Myrabo - (myrabl@rpi.edu)
Henry T. Nagamatsu - (nagamh@rpi.edu)
Department of Mechanical Engineering, Aeronautical Engineering, and Mechanics
Rensselaer Polytechnic Institute, Troy, NY 12180-3590 - USA

The feasibility of transatmospheric flight is limited by phenomena such as aerodynamic drag and heating. For take off, escape from and flight through the earth's atmosphere the drag on the body should be reduced. For re-entry vehicle into the earth's atmosphere at a hypersonic speeds, it is important to have a large nose radius and high aerodynamic drag. An efficient hypersonic vehicle design has to combine a low drag coefficient with low heat transfer. A 6-in. diameter aluminum blunt body model was fabricated and fitted with pressure transducers and heat flux gauges over its forebody surface. A 6-in. long spike was placed at the stagnation point. Spiked blunt body with and without cooling gas flowing out of the spike were tested in the RPI 24-in. Diameter Hypersonic Shock Tunnel. Freestream Mach 10 flow, with a stagnation temperature about 800 K, was selected to conduct the pressure and heat transfer measurements over the model. The spiked blunt body with and without cooling gas were very similar to each other. The measured pressure and heat transfer data indicate that the aerodynamic drag and heating of the spiked blunt body with cooling gas is lower than the aerodynamic drag and heating for the spiked blunt body without cooling gas. When the sonic cooling gas is injected through the physical spike; heat transfer over the model forebody surface decreases below that of the spiked blunt body with no cooling gas.

Keywords: spiked blunt body; hypersonic experimental investigation; pressure and heat transfer measurements.
 
 

Alexandre P. Antunes - alex@iae.cta.br
Centro Técnico Aeroespacial, Instituto Tecnológico de Aeronáutica
CTA/ITA/IEAA - 12228-900 -São José dos Campos, SP, Brazil.

Edson Basso - basso@iae.cta.br
Centro Técnico Aeroespacial, Instituto Tecnológico de Aeronáutica
CTA/ITA/IEAA - 12228-900 -São José dos Campos, SP, Brazil.

João Luiz F. Azevedo - azevedo@iae.cta.br
Centro Técnico Aeroespacial, Instituto de Aeronáutica e Espaço
CTA/IAE/ASE-N - 12228-904 -São José dos Campos, SP, Brazil.

The present work is part of the effort for the development of computational tools necessary to simulate aerodynamic flows over aerospace geometries, especially those related to the first Brazilian Satellite Launch Vehicle, VLS. Aerodynamic flow simulations over the VLS during its first-stage flight are presented together with a validation effort of these results through
the use of wind tunnel experimental data available for the vehicle. The calculations use the Chimera technique together with block structured grids to discretize the computational domain. The current approach is based on the solution of the 3-D Euler equations in curvilinear coordinates. A finite difference method is applied to these equations and a centered spatial discretization is used. Artificial dissipation terms, based on a scalar, non-isotropic model, are added. The time march process is accomplished with a 5-stage, 2nd-order accurate, Runge-Kutta scheme.

Keywords: Multiblock technique, Chimera, VLS, Finite differences.
 
 

Yang Xuefeng - xuefeng@mecanica.ufu.br
Aristeu da Silveira Neto - aristeus@mecanica.ufu.br
Américo Scotti - ascotti@mecanica.ufu.br
Universidade Federal de Uberlândia, Faculdade de Engenharia Mecânica
38400-902 - Santa Mônica - Uberlândia - MG - Brasil

A mathematical model which is capable to predict the heat transfer and fluid flow of plasma arc encountered in welding process TIG is presented in this paper. The simulation results are compared with the experiment results available in literature and the agreements are considered as good.

Keywords: welding, numerical simulation, modeling, plasma
 
 

Paulo G. de P. Toro - toro@iae.cta.br
Instituto de Aeronaútica e Espaço - IAE - Centro Técnico Aeroespacial - CTA
São José dos Campos - SP 12228-904 - BRAZIL

Marco A.S. Minucci - sala@ieav.cta.br
Instituto de Aeronaútica e Espaço - IAE - Centro Técnico Aeroespacial - CTA
São José dos Campos - SP 12231-970 - BRAZIL

Heidi Korzenowski - heidi@univap.br
Universidade do Vale do Paraíba - UNIVAP
São José dos Campos - SP 12244-000 - BRAZIL

Leik N. Myrabo - myrabl@rpi.edu
Henry T. Nagamatsu - nagamh@rpi.edu
Department of Mechanical Engineering, Aeronautical Engineering, and Mechanics
Rensselaer Polytechnic Institute, Troy, NY 12180-3590 - USA

The feasibility of hypersonic flight is limited by phenomena such as aerodynamic drag and heating. For take off, escape from and flight through the earth's atmosphere the drag on the body should be reduced. For re-entry vehicle into the earth's atmosphere at a hypersonic speeds, it is important to have a large nose radius and high aerodynamic drag. An efficient hypersonic vehicle design has to combine a low drag coefficient with low heat transfer. A small ballistic re-entry vehicle SARA configuration is numerically (low drag case) and experimentally (high drag case) investigated by using the Euler equations and the RPI 24-in. diameter Hypersonic Shock Tunnel (RPI-HST), respectively. The governing equations are discretized in a cell centered, finite volume procedure for unstructured triangular grids. Spatial discretization uses an upwind scheme. Time march uses, an explicit 2nd-order accurate, 5-stage Runge-Kutta time stepping scheme. A 6-in. diameter aluminum "double Apollo disc" model was fabricated and fitted with piezoelectric pressure transducers and thin-film platinum heat gauges over its forebody surface. Numerical (inviscid simulations) and experimental results are presented for the hypersonic flow. Freestream Mach 10 flow was selected to conduct numerical and experimental investigations.

Keywords: Re-entry vehicle, numerical investigation, experimental investigation, hypersonic flow
 
 

Paulo S. Zidanski - zidanski@aer.ita.cta.br
M. A. Ortega - ortega@aer.ita.cta.br
Nide G. C. R. Fico Jr. - nide@aer.ita.cta.br
Instituto Tecnológico de Aeronáutica, Divisão de Engenharia Aeronáutica,
São José dos Campos, 12228-900, SP, Brasil.

This paper reports on a series of numerical simulations of laminar flows over two-dimensional cavities. Five different aspect ratios, ranging from [1:9.6] to [1:28], were considered. The calculations pointed out that for the aspect ratios studied the flow topology was sensitive to this parameter. The aspect ratio plays a direct influence upon the number and the position of the flow vortices inside the laminar cavity. The numerical results were very sensitive to the inlet velocity profile, which is closely related to be boundary layer thickness upstream of the cavity. The SIMPLER numerical algorithm developed by Patankar was used to solve the discrete equations on a staggered grid. The interpolation functions are based upon the power law scheme.

Keywords: viscous flow, incompressible flow, cavity, numerical method
 
 

João B. P. Falcão Filho - joaobpff@uol.com.br
Centro Técnico Aeroespacial, Instituto de Aeronáutica e Espaço
São José dos Campos -SP - 12228-904

M. A. Ortega - ortega@aer.ita.br
Nide G. C. R. Fico Jr. - nide@aer.ita.br
Instituto Tecnológico de Aeronáutica, Divisão de Engenharia Aeronáutica
São José dos Campos -SP- 12228-900

The ultimate aim of this research effort is the numerical simulation of the turbulent mixing of two parallel jets, one supersonic and the other subsonic, in the presence of a solid wall. The geometry of the flow is two-dimensional. This physical situation is the one to be used as a boostering device in a transonic wind tunnel and has the objective of extending the tunnel's envelope without an excessive increase of the main compressor power. The basics of a finite-difference code, using the implicit Beam and Warming algorithm, is already running including laminar viscous terms. This paper has the objective of reporting the actual status of the study including the validation of the code. The reference problem is a transonic nozzle flow. Special care is taken on the implementation of boundary conditions, especially for the exit section of the nozzle when both supersonic and subsonic regions are already established. Results for the implementation of boundary conditions using zeroth-order extrapolation or the characteristic relations are also reported.

Keywords: Viscous Compressible Flow, Finite-Difference Algorithm, Transonic Nozzle, Mixed Boundary Conditions
 
 

Paulo Afonso de Oliveira Soviero - soviero@aer.ita.br
Instituto Tecnológico de Aeronáutica
12228-900 - São José dos Campos, SP, Brazil
Fábio Henrique Lameiras Pinto - lameiras@netvale.com.br
Instituto de Aeronáutica e Espaço, Divisão de Ensaios em Vôo
12228-904 - São José dos Campos, SP, Brazil

Abstract.   The mathematical model for two-dimensional unsteady sonic flow, based on the classical diffusion equation with imaginary coefficient, is presented and discussed. The main purpose is to develop a rigorous formulation in order to bring into light the correspondence between the sonic, supersonic and subsonic panel method theory. Source and doublet integrals are obtained and Laplace transformation demonstrates that, in fact, the source integral is the solution of the doublet integral equation. It is shown that the doublet-only formulation reduces to a Volterra integral equation of the first kind and a numerical method is proposed in order to solve it. To the authors' knowledge this is the first reported solution to the unsteady sonic thin airfoil problem through the use of doublet singularities. Comparisons with the source-only formulation are shown for the problem of a flat plate in combined harmonic heaving and pitching motion.

Keywords: Unsteady aerodynamics, Sonic flow, Panel method, Linearized theory
 
 

Pedro Paglione - e-mail: paglione@era.ita.cta.br
Roberto M. Girardi - e-mail: girardi@aer.ita.cta.br
Instituto Tecnológico de Aeronáutica (ITA)
Praça Mal. Eduardo Gomes, 50 - 12.228-900 São José dos Campos, SP.

Modern wing plan forms are characterized by a kink in some place along the wing span, used to solve practical problems associated with the positioning of landing gear and aeronautical systems (such as hydraulic, fuel and anti-ice systems) inside the wing. A method for designing such wings, based on optimization techniques is presented in the present paper. An optimization code and an aerodynamic code, based on the well known vortex lattice method, are combined. The objective of the optimization procedure is the minimization of the induced drag and some constraints are imposed to be satisfied during this procedure, in order to avoid non practical solutions. The Bandeirante aircraft wing was used as the initial condition of the optimization procedure and two wing designs, subjected to two different sets of constraints, were obtained. A kink station was observed in both solutions and induced drag values were lower than the value calculated for the Bandeirante wing. Moreover, the designed wings have lower plan form area, which is advantageous because the profile drag is proportional to such area.

Keywords: Wing design, Optimization technique, Induced drag
 
 

Fernandes, M. S. - mfernand@mecanica.ufu.br
Universidade Federal de Uberlândia, MG, LEDIF
Faculdade de Engenharia Mecânica - João Naves de Ávila, SN

Biage, M. - mbiage@mecanica.ufu.br
Universidade Federal de Uberlândia, MG, LEDIF
Faculdade de Engenharia Mecânica - João Naves de Ávila, SN

O objetivo do presente trabalho foi o de desenvolver um código computacional, utilizando a técnica espectral da colocação, para simular camadas de mistura compressível e turbulenta, em desenvolvimento temporal e espacial, descritas pelas equações governantes completas, na forma conservativa. Para isto, implementou-se o método espectral da colocação de Chebyshev, aliado ao esquema de integração temporal de Runge-Kutta, utilizando uma formulação explícita. Para a discretização das derivadas parciais envolvidas no equacionamento, utilizou-se a Transformada Rápida de Chebyshev (TRC) que permitiu obter um significativo ganho na velocidade de processamento. O código computacional foi desenvolvido para simular escoamentos compressíveis, em altos e baixos números de Reynolds, com e sem efeitos de estratificação. Os resultados revelaram vários aspectos das estruturas dos escoamentos que, na sua grande maioria, concordam com as conclusões apresentadas por várias pesquisas sobre camadas de mistura. Em particular, são ilustrados os detalhes do comportamento do campo de densidade nas camadas de mistura espaciais e temporais, evidenciando os efeitos da taxa de densidade na camada turbulenta.

Palavras-chave: Camada de Mistura, Chebyshev, Colocação Espectral, Escoamento Compressível e Turbulento
 
 

Olympio Achilles de Faria Mello - oamello@iae.cta.br
João Luiz Filgueiras de Azevedo - azevedo@iae.cta.br
Instituto de Aeronáutica e Espaço, Centro Técnico Aeroespacial
12228-904 - São José dos Campos - SP, Brazil

A simulation of high angle-of-attack flow over a hemisphere-cylinder body is presented. The Reynolds-averaged Navier-Stokes equations are solved using a diagonal form of an alternating-direction implicit (ADI) approximate factorization procedure. A modified Baldwin-Lomax turbulence model is used. Computed pressure coefficient distributions at angles of attack up to 19 degrees are compared with experimental data. Qualitative evaluation of main flow features, based on local Mach number and skin friction distributions are also analyzed with respect to experimental visualization results. The present work is used to validate the method for simulation of high angle-of-attack flow around sounding rockets and launch vehicles.

Keywords: Separated flow, Transonic flow, Computational aerodynamics.
 
 

Danton J.F.Villas Boas - danton@iae.cta.br
Khoze Kessaev
Mario Niwa - niwa@iconet.com.br
Instituto de Aeronáutica e Espaço - IAE Centro Técnico Aeroespacial
12228-904 - São José dos Campos - SP - Brasil

It is known that an underexpanded gas jet entering into a thin-wall cylindrical cavity (resonance tube) can cause fast and strong heating of the walls, up to a temperature which can exceed several times the gas jet stagnation temperature (Sprenger, 1954). Recently, an application of this phenomenon used for ignition of rocket-engine gas and liquid propellants was proposed by Niwa et al., 2000, since the temperature inside the resonance tube can achieve values large enough to ignite the propellant mixture. In the present work, a summary of the resonance tube theoretical model (Kessaev, 1990) is presented and not obvious peculiarities of tube wall heating are revealed. Results from a theoretical model are compared with those from experiments. Characteristics such as tube length effect, temperature distributions along the tube wall and location of maximum temperature region are studied in detail.

Keywords: resonance tube, propellant ignition, shock waves, gas-dynamical heating.
 
 

Leonor Camila Q. Yagua - camila@aer.ita.cta.br
Centro Técnico Aeroespacial, Instituto Tecnológico de Aeronáutica
CTA/ITA/IEA - 12228-900 - São José dos Campos, SP, Brasil
João Luiz F. Azevedo - azevedo@iae.cta.br
Centro Técnico Aeroespacial, Instituto de Aeronáutica e Espaço
CTA/IAE/ASE-N - 12228-904 - São José dos Campos, SP, Brasil

This work presents numerical simulations of the flow over the first Brazilian satellite launch vehicle, VLS. The algorithm solves the thin-layer Navier-Stokes equations for compressible flows using the Chimera technique. The computational code considers a finite difference formulation and the time discretization uses an explicit method. The spatial discretization uses a centered scheme in which the artificial dissipation terms are explicitly added. Pressure coefficient results obtained in the present simulations are compared to experimental data.

Keywords: CFD, VLS, Multiblock Grids, Numerical Simulation, Chimera Technique.