Details of Grant 

EPSRC Reference:	EP/H020853/1
Title:	Shock-induced separation of hypersonic transitional boundary layers
Principal Investigator:	Hillier, Professor R
Other Investigators:	Ganapathisubramani, Professor B
Researcher Co-Investigators:
Project Partners:
Department:	Aeronautics
Organisation:	Imperial College London
Scheme:	Standard Research
Starts:	01 July 2010	Ends:	30 September 2013	Value (£):	564,840
EPSRC Research Topic Classifications:	Aerodynamics
EPSRC Industrial Sector Classifications:	Aerospace, Defence and Marine
Related Grants:
Panel History:	Panel Date Panel Name Outcome 16 Dec 2009 Material, Mechanical & Medical Engineering Panel Announced
Summary on Grant Application Form
At high supersonic and hypersonic speeds, surface friction and surface heat transfer (kinetic heating) are dominant effects. These are consequences of the boundary layer, the thin layer of retarded flow near the body surface that arises through the action of viscosity. The limit states for a boundary layer are either the smooth-flowing laminar state or the random fluctuations of a turbulent boundary layer. Laminar boundary layers are characterised, amongst other properties, by low heat transfer and low skin friction. In a turbulent boundary layer typical values may be of order five times larger, so that it is critical to understand the likely boundary layer state. The phenomenon of transition is the process by which an initial laminar boundary layer changes to a turbulent one. Transition itself is a pacing item in aerodynamics and many aspects of transition are still not properly understood. A special feature of hypersonic flows is that the transition region may be very substantial in extent; it may cover a significant extent of the body surface so that critical flow interactions then occur in this transitional state. The effect of shock waves provides one such critical interaction. Shock waves, themselves, are an almost inevitable feature of supersonic flows. They arise when the flow is turned in such a manner as to raise its pressure significantly and are features of many practical configurations such as: intakes to air-breathing engines; deflection of control surfaces; the flow approaching leading edges of aerodynamic surfaces. If the pressure rise generated by a shock wave is sufficiently strong it can then generate a so-called separated flow. In this case the boundary layer no longer follows the surface; instead it rises above the surface with a 'trapped' recirculating flow beneath it. This is a complex flow; it may be unsteady and it is characterised, amongst other factors, by surface heat transfer rates that rise significantly above those of the attached boundary layer. Laminar boundary layers are much more readily separated than turbulent boundary layers and produce much larger separation zones. Transitional boundary layers therefore, potentially, occupy a midway state. Given that a transitional boundary layer may, in effect, be switching between laminar and turbulent states, the separated flow is likely to respond strongly to the instantaneous state of the approaching flow.This proposed study is therefore focussed on a benchmark experimental investigation of the separation of a transitional boundary layer by a shock wave. This is an interaction that can commonly occur but which, because of the challenging nature of transitional flows, has so far received only the most modest of attention, either experimentally or computationally. The objective here is to establish both the transitional boundary layer and the shock wave system with as high a precision as possible and in a manner that aids experimental repeatability; this requires the most careful design and management by the experimentalist. The ultimate objective is to generate high quality benchmark data that explains the basic flow physics and produces a well-defined test case for computational simulation and contributes to the engineering design database.
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