separation vectors between points 1 and 2 along the axis direction x x x 2 u2212 x 1 and axis direction y y y 2 u2212 y 1 of the simply supported panel coordinate system the equation represents the boundary layer pressure fluctuations on the outside of the aircraft referring to fig5 the joint acceptance function for the fundamental panel mode m 1 n 1 of a simply supported plate was calculated at angles u03b1 0 u00b0 5 u00b0 10 u00b0 15 u00b0 30 u00b0 and 45 u00b0 for the case described below fig6 shows the u03b4db reduction of the panel response for the various flow angles the reduction in panel response associated with a given mode is simply the ratio of the joint acceptance with flow angle 0 u00b0 to the joint acceptance with flow angle u03b1 u03b4 u2062 u2062 db 10 u2062 u2062 log u2062 j 11 2 u2061 u03c9 o j 11 2 u2061 u03c9 u03b1 although the example is for a particular panel size and aircraft flight condition the invention is applicable to any structure having a boundary layer flow the reduction in panel response or lack thereof is a function of the natural frequency of the panel for a panel with a very low natural frequency there would be a reduction in panel response of up to 5 db for a flow angle of 30 u00b0 for a panel mode natural frequency greater than 800 hz there would actually be a slight increase in the panel response due to flow angle internal pressurization of an aircraft fuselage at altitude is expected to produce a panel mode natural frequency of approximately 400 hz for the skin panels described at 400 hz there would.