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The flow about multi-element airfoil configurations is investigated using the unsteady Reynolds averaged Navier-Stokes equations. An explicit scheme is used to advance the solution in time while a finite difference scheme is applied to discretize the flux terms. An algebraic and two one-equation turbulence models are used to model turbulence. The domain about each multi-element airfoil is discretized with structured Chimera grids. The multi-element configurations presented in this paper include two airfoils with slotted flaps and an airfoil with a 50% chord vented aileron deflected at 90 degrees. Subsonic flow computations are performed for attached and separated flow conditions. The computational results obtained with the CRTVD code developed at Wichita State University are in good correlation with wind tunnel data and with computational results obtained with the INS2D computer code developed at NASA Ames research center.

Flow field computations and, in particular, that of pressure, skin friction, and heat transfer (for high speed flights) are the primary parameters in the design of aerospace vehicles. Most computational schemes based on either the inviscid Euler equations or various forms of the Navier-Stokes equations are remarkably accurate in the predictions of pressure distributions. However, computations of skin friction and heat transfer particularly at high speeds have been a source of considerable difficulty. Problems arise not only due to the grid resolution but also due to the particular numerical scheme employed. To address the difficulty associated with accurate computations of the velocity and temperature gradients, a comparative investigation of several Navier- Stokes codes is undertaken. Previous studies with regard to the effect of grid resolution are incorporated into the current investigation.

Inviscid flow fields have been computed for blunt axisymmetric shapes exposed to a hypersonic environment. A fully implicit scheme using a flux vector splitting technique was used to obtain a finite difference formulation. To increase computational efficiency, an approximate factorization scheme was used. Flow-field solutions were obtained for chemically reacting air using a decoupled approach. Flow fields were computed for blunted, flared shapes over a range of Mach numbers from 2 to 20.