Experimental Results from Three Cone-Flow Waveriders.
by J. Pike.
Agard Conference Proceedings 30, Hypersonic Boundary Layers and Flow Fields, Royal Aeronautical Society, London, Ref. 12, p. 20, 1-3 May 1968.

The 3ft x 4ft high supersonic speed tunnel at R.A.E. Bedford has been used to obtain experimental results from two waverider models with sharp leading edges and a third with a rounded leading edge. At a particular incidence and Mach number (M=4) the flow supported by the compression surface of the model can be predicted theoretically. This predicted flow is shown to agree closely with experimental measurements of lower surface pressures, the shock wave shape and the surface streamline pattern. At other incidences or Mach numbers the flow cannot be prediced theoretically. The experimental results show however, that for high incidence at M=4 the pressures are remarkably uniform, and at constant incidence with Mach number decreasing, the shape of the shock wave changes smoothly, including its detachment from the leading edge. Leading edge rounding is shown to affect pressures only close to the leading edge.

A wing surface which is designed to be the same shape as a stream surface in the flow downstream of a conical or other shock wave, must be expected to support the theoretical flowfield at the design Mach number and incidence. However it is less clear that the flow will change smoothly when small changes are made to surface shape or design conditions. Here it is shown that a non-conical stream surface shape from a Mach 4 cone flow generates the correct conical shockwave at the design conditions. The tests also show that typical cone-flow waveriders have flows which change smoothly from the design flow under the influence of boundary layer development, leading edge blunting, Mach number changes and model incidence changes.

An explanation of the models and tunnel tests is included in this Imperial War Museum film

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Last amended: Dec 2013.