Hypersonic airbreathing propulsion has been under research since the mid 1960s and will
continue to be of interest as the imperative for cheaper, safer and more reliable access to
space grows. As such, the compression of high Mach number boundary layers in scramjet
intakes is an ongoing field of interest.
Even at high Mach numbers, laminar boundary layers under constant conditions are well
understood and well represented theoretically. Abrupt changes in pressure gradient, whether
from geometry or incident shock waves, cause significant deviation from these theories. This
thesis sets out to explore this modification.
The chosen, computational investigation used the change between zero and adverse linear
constant pressure gradient flows with the computational geometry consisting of a flat plate
section and a compression curve. The computations covered a range of adverse pressure
gradients [Mathematical equation appears here. To view, please open pdf attachment], Reynolds numbers (Re[Cartesian streamwise distance (m)]0 = 87500 to 90000000) and Mach
numbers (4 to 8). The effect of pressure gradient, Reynolds number and Mach number on
the equilibrium or non-equilibrium change between pressure gradients was investigated.
It was observed that flows of this type can be divided into three sections: a fully developed zero pressure gradient flow, a fully developed adverse pressure gradient flow and an
interaction region between the two.
The stronger the pressure gradient, the more non-equilibrium the response of the boundary layer to the change in pressure became. Inertial forces came to dominate in the interaction
region and viscous effects only reasserted themselves further downstream. This was illustrated
by the significant transverse variations in pressure occurring in this interaction region.
The investigation is extended to turbulent boundary layers modelled using the Baldwin-
Lomax turbulence model.