Abstract
A routine numerical method of solving integral equations arising in steady supersonic flow past flexible delta wings, with supersonic and subsonic leading edges, is presented. The wing is subdivided into diamond shaped panels, by a network of equi‐spaced Mach lines, and the effect of placing unit downwash at each panel in turn is calculated. The resulting influence function, which has been tabulated, is then formed into the required matrix of influence coefficients. Within the limitations of linearized theory the results are valid for all Mach numbers and mode shapes, including chordwise bending. Although applied here to delta wings the method is equally suited to wings of arbitrary plan form. An aeroelastic efficiency calculation, for a sonic leading edge 45 deg. delta wing of representative torsional stiffness, is included for comparison with strip theory.
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