The present work is inserted into an effort to develop a Chimera flow simulation code capable of handling general launch vehicle configurations. The paper is primarily concerned with presenting results of laminar and turbulent viscous simulations of flows over the first Brazilian satellite launch vehicle, the VLS, during its first-stage flight. The finite difference method is applied to the governing equations written in conservation-law form for general body conforming curvilinear coordinates. The spatial discretization is accomplished with a central difference scheme in which artificial dissipation terms, based on a scalar, non-isotropic model, are added to the numerical scheme to maintain stability. The time march process is accomplished with a 5-stage, 2nd-order accurate, Runge-Kutta scheme. The results here included are indicative of the current status of the Chimera flow simulation capability under development by the authors. The results also highlight interesting features of the flow over the complete VLS and point out the importance of the inclusion of viscous effects for flow simulation over such complex vehicles.

Transonic and supersonic flow simulations over typical launch vehicle configurations are presented. A 3-D finite difference numerical code, written for general, curvilinear, body-conforming coordinate systems, is used. The code solves the thin-layer approximation for the laminar Navier-Stokes equations. Simulations are performed for a launcher and a sounding rocket configurations, currently under development at Instituto de Aeronáutica e Espaço. Calculations consider cases at angle of attack and at various freestream Mach numbers. Normal force coefficients are obtained such that the loads required for the design phase can be determined. Computational results are compared to available experimental data. In general, good results within engineering error margins are obtained.

Aerodynamic flow simulations over the first Brazilian satellite launch vehicle, VLS, during its first-stage flight are presented. The three dimensional compressible flow is modeled by the Euler equations and a Taylor-Galerkin finite element method with artificial dissipation is used to obtain the numerical solution. Transonic and supersonic results for zero angle-of-attack are presented and compared to available experimental results. The influence of mesh refinement and artificial dissipation coeffcient on the transonic flow results are discussed. The results obtained for the supersonic simulations present good agreement with experimental data. The transonic simulation results capture the correct trends but they also indicate that this flight condition requires more refined meshes.

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 governing equations are discretized using an unstructured grid finite volume method, which is capable of handling hybrid unstructured grids. A centered scheme as well as a second order version of Liou’s AUSM+ upwind scheme are used for the spatial discretization. The 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 validated considering 2-D laminar and turbulent viscous flows over a flat plate. Studies of the various parameters affecting the multigrid acceleration performance are undertaken with the objective of determining optimal numerical parameter combinations.

The results of a numerical study of premixed Hydrogen-air flows ignition by an oblique shock wave (OSW) stabilized by a wedge are presented, in situations when initial and boundary conditions are such that transition between the initial OSW and an oblique detonation wave (ODW) is observed. More precisely, the objectives of the paper are: (i) to identify the different possible structures of the transition region that exist between the initial OSW and the resulting ODW and (ii) to evidence the effect on the ODW of an abrupt decrease of the wedge angle in such a way that the final part of the wedge surface becomes parallel to the initial flow. For such a geometrical configuration and for the initial and boundary conditions considered, the overdriven detonation supported by the initial wedge angle is found to relax towards a Chapman-Jouguet detonation in the region where the wedge surface is parallel to the initial flow. Computations are performed using an adaptive, unstructured grid, finite volume computer code previously developed for the sake of the computations of high speed, compressible flows of reactive gas mixtures. Physico-chemical properties are functions of the local mixture composition, temperature and pressure, and they are computed using the CHEMKIN-II subroutines.