ESDU AERO W.S.05.03.01
Pressure distribution on blunt noses of two-dimensional sections at zero incidence in supersonic flow.
Abstract:ESDU Aero W.S.05.03.01 gives empirically-derived curves of static pressure as a fraction of the stagnation pressure behind a normal shock plotted against the ordinate of the nose measured from the centreline for three shapes: flat, semi-circular, and semi-elliptical with nose length to thickness ratio of 0.25. An expression is given for the stagnation pressure behind a normal shock or ESDU Aero S.00.03.15 may be used to evaluate it graphically. For the semi-circular shape some variation with Mach number was found and guidance is given on the magnitude of that effect for Mach numbers up to 6. It is noted that due to this effect the drag of that nose shape should be obtained from ESDU Aero W.S.02.03.10 rather than by integration of the data given in these curves. For the other shapes data applied to Mach numbers only up to 2 but showed no Mach number effect and it is suggested they would apply to higher Mach numbers. For an elliptical-shaped nose section with length to thickness ratio greater than 0.25 there was found to be a greater effect of Mach number and a single curve could not be given; a simple empirical approach is suggested relating the pressure coefficient at any point to the slope and the stagnation pressure coefficient there. Values given by Newtonian impact theory are suggested to define maxima to which pressure coefficients tend as Mach number increases. The application of the data to three-dimensional wings is discussed.
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