This paper determines the sheet thickness variation near a circular hole to yield a constant hoop stress in a sheet subjected to uniform radial stress at infinity. A similar analysis is presented for a variable thickness sheet plus an annular reinforcement. The stresses due to the remote application of shear and uniaxial tension are also determined.

The results of extensive measurements which are presented here show the presence of secondary flow in the laminar corner boundary layer which is opposite in rotation sense to that found in the turbulent corner flow. This secondary flow is associated with progressive changes of the form of the velocity profiles. These changes become marked prior to transition and may be linked with the transition process. The laminar corner flow separates readily under the influence of an adverse streamwise pressure gradient.

Relationships for the sectional drag of wings are derived by using the linearised wing theory. It is suggested that some of these results may prove useful for checking the accuracy of numerical lifting-surface theories and for formulating approximate theories. Two main problems are considered, namely the lifting (no thickness) case and the thickness (zero lift) problem. In both cases, general planform shapes are examined and it is shown that particularly simple results are achieved for planar wings with a spanwise axis of symmetry.

The linearised theory of supersonic flow is applied to a plane wing with a pointed apex whose leading edge is subsonic near the apex but supersonic further outboard. The change in character of the leading edge may occur either where the edge is kinked or at a point where its slope is continuous. In each case, the theoretical pressure field due to incidence is found to be singular along the downstream, outward-going Mach line through the point at which the edge changes character. The form of the singularity is different for the two cases, and somewhat unexpected. Measurements of the pressure on a plane wing with a curved, mixed leading edge showed no abrupt variation near the Mach line on which the singular behaviour was predicted.

It is shown that a pair of supersonic inviscid conical flow fields can exist on either side of a conical deflagration wave. The configuration is relevant to the base flow question and indicates how base drag may be alleviated by burning. Results of exact computations are presented as well as those derived from a slender-body theory.

Hypersonic conical flows over delta wings are treated in the thin-shock-layer approximation due to Messiter. The equations are hyperbolic throughout, even in regions where the full equations are elliptic, and have not hitherto been solved for flows with attached shock waves. The concept of the simple wave has been used to construct a class of solutions for such flows; they contain discontinuities in flow variables and shock slope but, for the case of flow over a delta wing with lateral symmetry, agreement with results of numerical solutions of the full equations is good. The method is applied to plane delta wings at yaw, and to wings with anhedral and dihedral. For the flow at the tip of a rectangular wing, it is shown that two distinct solutions may be constructed.

Fatigue tests have been made in constant amplitude loading and in narrow band random loading on FV520B steel lugs of width 27 mm loaded by means of a pin 12·7 mm in diameter. The test frequency was 220 Hz, a mean tensile stress of 309 N/mm2 was used and tests extended for endurances of up to 109 cycles.

For endurances up to 106 cycles there was good agreement between constant amplitude and random loading results with both stresses expressed in root mean square values. Beyond this endurance, the random loading results lay below those obtained in constant amplitude loading. Predictions made using the Palmgren-Miner cumulative damage hypothesis in conjunction with the endurance curve in constant amplitude loading gave good agreement with the results in random loading over the whole range of endurances tested.