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Ames Research Center has the lead role among NASA centers to conduct research in computational fluid dynamics. The past, the present, and the future prospects in this field are reviewed. Past accomplishments include pioneering computer simulations of fluid dynamics problems that have made computers valuable in complementing wind tunnels for aerodynamic research. The present facilities include the most powerful computers built in the United States. Three examples of viscous flow simulations are presented: an afterbody with an exhaust plume, a blunt fin mounted on a flat plate, and the Space Shuttle. The future prospects include implementation of the Numerical Aerodynamic Simulation Processing System that will provide the capability for solving the viscous flow field around an aircraft in a matter of minutes.

Aeronautical applications of simulation technology at Ames Research Center are described. The largest wind tunnel in the world is used to determine the flow field and aerodynamic characteristics of various aircraft, helicopter, and missile configurations. Large computers are used to obtain similar results through numerical solutions of the governing equations. Capabilities are illustrated by computer simulations of turbulence, aileron buzz, and an exhaust jet. Flight simulators are used to assess the handling qualities of advanced aircraft, particularly during takeoff and landing.

Supersonic flow past a blunt body is considered, where the flow contains an embedded subsonic region which lies between the shock wave and the body surface and is bounded by sonic lines from the body to the shock. A numerical approach is taken, which uses a basic finite difference scheme that solves the unsteady fluid dynamic equations in integral form. The unsteady equations are everywhere hyperbolic in time so no distinction need be made between subsonic and supersonic regions. Solutions to the mixed elliptic and hyperbolic steady flow equations are approached asymptotically in time. The method is illustrated for two-dimensional flows.

Computer simulations of the flow field around the Space Shuttle Orbiter are described. Results of inviscid calculations are presented for the shock wave pattern and bottom centerline pressure distribution at 30 deg angle of attack. Results of viscous calculations are presented for wall pressure and heat transfer distributions for simple configurations representative of regions where shock wave-boundary layer interactions occur. The computer codes are verified by comparisons with wind-tunnel data and can be applied to flight conditions.

A simple eddy-viscosity model is shown to make it possible to calculate numerically the mean properties of a turbulent wake. Although the structure of the Reynolds stress terms is not resolved, the results obtained are adequate for predicting velocity profiles and displacement thicknesses.

Turbulent skin friction and heat transfer prediction on flat plates and wind tunnel walls at supersonic and hypersonic Mach numbers, using Van Driest theory

Shock shape and surface pressure measurements on hypersonic pointed cones in helium

Skin friction and heat transfer data for hypersonic turbulent layer on flat plates, wind tunnel wall, and cones

Integer relations method of determining inviscid shock layer flow and pressure distributions over blunt bodies

Numerical solutions for blunt axisymmetric bodies in supersonic spherical source flow

Numerical methods and computer programs for two-dimensional and axisymmetric supersonic flow over blunt-nosed and flared bodies

Flow field around spherical blunt bodies in gas equilibrium mixtures resembling various planetary atmospheres - inverse method of flow field calculation

An efficient numerical method for calculating plane, axisymmetric, and fully three-dimensional blunt-body flow is presented. It is a second-order-accurate, time-dependent finite-volume procedure that solves the Euler equations in integral conservation-law form. These equations are written with respect to a Cartesian coordinate system in which an embedded mesh adjusts in time to the motion of the bow shock that is automatically captured as part of the weak solution. With such an adjusting mesh, oscillations in flow properties near the shock are shown to be virtually eliminated. The scheme uses a time-splitting concept that accelerates the convergence appreciably. Comparisons are made between computed and experimental results.