Grid generation about complex three-dimensional aircraft configurations

The problem of obtaining three dimensional grids with sufficient resolution to resolve all the flow or other physical features of interest is addressed. The generation of a computational grid involves a series of compromises to resolve several conflicting requirements. On one hand, one would like the grid to be fine enough and not too skewed to reduce the numerical errors and to adequately resolve the pertinent physical features of the flow field about the aircraft. On the other hand, the capabilities of present or even future supercomputers are finite and the number of mesh points must be limited to a reasonable number: one which is usually much less than desired for numerical accuracy. One technique to overcome this limitation is the 'zonal' grid approach. In this method, the overall field is subdivided into smaller zones or blocks in each of which an independent grid is generated with enough grid density to resolve the flow features in that zone. The zonal boundaries or interfaces require special boundary conditions such that the conservation properties of the governing equations are observed. Much work was done in 3-D zonal approaches with nonconservative zonal interfaces. A 3-D zonal conservative interfacing method that is efficient and easy to implement was developed during the past year. During the course of the work, it became apparent that it would be much more feasible to do the conservative interfacing with cell-centered finite volume codes instead of the originally planned finite difference codes. Accordingly, the CNS code was converted to finite volume form. This new version of the code is named CNSFV. The original multi-zonal interfacing capability of the CNS code was enhanced by generalizing the procedure to allow for completely arbitrarily shaped zones with no mesh continuity between the zones. While this zoning capability works well for most flow situations, it is, however, still nonconservative. The conservative interface algorithm was also implemented but was not completely validated

Analysis of a finite difference grid

Some means of assessing the suitability of a mesh network for a finite difference calculation are investigated in this study. This has been done by a study of the nonlinear truncation errors of the scheme. It turns out that the mesh can not be properly assessed a priori. The effect of the mesh on the numerical solution depends on several factors including the mesh itself, the numerical algorithm, and the solution. Several recommendations are made with regard to generating the mesh and to assessing its suitability for a particular numerical calculation

CNSFV code development, virtual zone Navier-Stokes computations of oscillating control surfaces and computational support of the laminar flow supersonic wind tunnel

The work performed during the past year on this cooperative agreement covered two major areas and two lesser ones. The two major items included further development and validation of the Compressible Navier-Stokes Finite Volume (CNSFV) code and providing computational support for the Laminar Flow Supersonic Wind Tunnel (LFSWT). The two lesser items involve a Navier-Stokes simulation of an oscillating control surface at transonic speeds and improving the basic algorithm used in the CNSFV code for faster convergence rates and more robustness. The work done in all four areas is in support of the High Speed Research Program at NASA Ames Research Center

Hypersonic blunt body computations including real gas effects

The recently developed second-order explicit and implicit total variation diminishing (TVD) shock-capturing methods of the Harten and Yee, Yee, and van Leer types in conjunction with a generalized Roe's approximate Riemann solver of Vinokur and the generalized flux-vector splittings of Vinokur and Montagne for two-dimensional hypersonic real gas flows are studied. A previous study on one-dimensional unsteady problems indicated that these schemes produce good shock-capturing capability and that the state equation does not have a large effect on the general behavior of these methods for a wide range of flow conditions for equilibrium air. The objective of this paper is to investigate the applicability and shock resolution of these schemes for two-dimensional steady-state hypersonic blunt body flows. The main contribution of this paper is to identify some of the elements and parameters which can affect the convergence rate for high Mach numbers or real gases but have negligible effect for low Mach number cases for steady-state inviscid blunt body flows

Constant temperature hot wire anemometry data reduction procedure

The theory and data reduction procedure for constant temperature hot wire anemometry are presented. The procedure is valid for all Mach and Prandtl numbers, but limited to Reynolds numbers based on wire diameter between 0.1 and 300. The fluids are limited to gases which approximate ideal gas behavior. Losses due to radiation, free convection and conduction are included

Lagrangian computation of inviscid compressible flows

A Lagrangian method is developed to solve the Euler equations of gas dynamics. The solution of the equations is obtained by a numerical computation with the well-known Flux-Corrected-Transport (FCT) numerical method. This procedure is modified so that the boundary treatment is accurate and relatively simple. Shock waves and other flow discontinuities are captured monotonically without any type of fitting procedures. The Lagrangian method is employed so that the problem of mesh generation is completely avoided. The method is applicable to all Mach numbers except the low subsonic range where compressibility effects are small. The method is applied to a one-dimensional Riemann problem (shock tube) and to a two-dimensional supersonic channel flow with reflecting shock waves

Concepts for radically increasing the numerical convergence rate of the Euler equations

Integral equation and finite difference methods have been developed for solving transonic flow problems using linearized forms of the transonic small disturbance and Euler equations. A key element is the use of a strained coordinate system in which the shock remains fixed. Additional criteria are developed to determine the free parameters in the coordinate straining; these free parameters are functions of the shock location. An integral equation analysis showed that the shock is located by ensuring that no expansion shocks exist in the solution. The expansion shock appears as oscillations in the solution near the sonic line, and the correct shock location is determined by removing these oscillations. A second objective was to study the ability of the Euler equation to model separated flow

High- Resolution Shock-Capturing Schemes for Inviscid and Viscous Hypersonic Flows

A class of high-resolution implicit total variation diminishing (TVD) type algorithms suitable for transonic multidimensional Euler and Navier-Stokes equations has been extended to hypersonic computations. The improved conservative shock-capturing schemes are spatially second- and third-order and are fully implicit. They can be first- or second-order accurate in time and are suitable for either steady or unsteady calculations. Enhancement of stability and convergence rate for hypersonic flows is discussed. With the proper choice of the temporal discretization and implicit linearization, these schemes are fairly efficient and accurate for very complex two-dimensional hypersonic in viscid and viscous shock interactions. This study is complemented by a variety of steady and unsteady viscous and inviscid hypersonic blunt body flow computations. Due to the inherent stiffness of viscous flow problems, numerical experiments indicated that the convergence rate is in general slower for viscous flows than for inviscid steady flows

Numerical Investigation of the Flow Angularity Effects of the NASA Langley UPWT on the Ares I DAC1 0.01-Scale Model

Investigation of the non-uniform flow angularity effects on the Ares I DAC-1 in the Langley Unitary Plan Wind Tunnel are explored through simulations by OVERFLOW. Verification of the wind tunnel results are needed to ensure that the standard wind tunnel calibration procedures for large models are valid. The expectation is that the systematic error can be quantified, and thus be used to correct the wind tunnel data. The corrected wind tunnel data can then be used to quantify the CFD uncertainties