COMPRESSIBLE boundary layers have been extensively stud-ied for a variety of applications from flat plates to swept wings and for subsonic Mach numbers to hypersonic Mach numbers. For low subsonic Mach numbers, it is often possible to assume incompressible flow. Hypersonic boundary layers at high Mach numbers generally involve strong interaction and strong viscous dissipation in a high-temperature, chemically reacting flow and require a more detailed set of equations. For transonic and supersonic flows, however, the boundary-layer equations are reasonably accurate, though changes in viscosity and thermal conductivity within the boundary layer must be taken into account.

In previous computations by other authors, the compressible boundary-layer equations have been solved with a variety of approximations. The viscosity–density parameter ρµ/ρeµe appearing in the transformed boundary-layer equations is computed using Sutherland’s law, the power law, or a linear law. In some studies, the coefficient of proportionality for the linear law or the exponent in power law is such that the parameter ρµ/ρeµe is approximately constant, where e is the edge conditions in the boundary layer. 1− 4 In addition, using a range of values of the exponent ω in the power law5 or Sutherland’s law (Van Driest6) allows this parameter to vary within the boundary layer (see also Refs. 7–11). More recently, Moraes et al. 12 (see also Ref. 13) gave a complete solution of the boundary-layer equations, while incorporating change of viscosity using Sutherland’s law. In the current study, as in Refs. 6 and 12, Sutherland’s law for viscosity is used along with the ideal gas assumption. In this study, the boundary-layer equations are solved, fully accounting for the change of viscosity and thermal conductivity within the boundary layer. The exact solution obtained from this computation is used for a comparison with simpler subsonic low Mach number solutions.