The experimental work described in this paper, which was divided into two sections, was done to determine the usefulness of photoelasticity as a method for analysing fuselage frame problems. In the first part, a series of tests was made by the photoelastic method on models without any supporting shell, which were in equilibrium under the applied i loads, to determine the bending moment distribution in various types of frames and rings.

In all cases the theoretical bending moments were calculated from the principles of minimum strain energy. The experimental results agreed with theory to within 5 per cent, except close to the null points and joints.

The second part was devoted to shell-supported models, and the shear loads were reacted out through the skin to firm supports at the end of adjacent bays. Again the bending moments were determined for the entire frame from the isochromatic fringe pattern.

The shell-supported tests proved to be of considerable interest when compared with the theoretical curves of Ref. 4 for the corresponding relative stiffness parameters. In some cases agreement was good but in others there were marked differences. A possible reason for this was that compression effects in the skin were altering the distribution and value of the bending moments in the frame.

It was felt that there was plenty of scope for more interesting work on fuselage frame analysis, from the standpoint of shear flow determination, intermediate frame reaction, and skin-to-frame relative stiffness effects. All of these might feasibly be examined photoelastically with a certain degree of advantage over lengthy and laborious theoretical investigations.

This is a discussion of the various types of flow which may occur on thin swept wings when the air stream is not attached to the wing surface everywhere. In particular, the behaviour of vortex sheets which arise from such mixed flows is examined.

The purpose of this paper is to illustrate the research and development work necessary to improve existing products and to anticipate demands for new products facilitating the advance in efficiency and performance of aircraft. The experience of Thornton Aero Engine Research Laboratory, in which the author has worked since it came into operation in 1941, has been drawn on almost entirely.

There has been much discussion of late on the nature of the noise from aircraft travelling with velocities greater than that of sound. It is possible that these phenomena can largely be explained in terms of the Doppler effect. For a source of sound travelling with a velocity greater than that of sound relative to the observer, the Doppler effect takes a very interesting form. It is discussed mathematically in Ref. 1.

The Melvill Jones equation for the profile drag of a single aerofoil is adapted to the case of an aerofoil in cascade, where the static pressure may be permanently raised (compressor), or lowered (turbine). A simplified procedure for measuring the drag is then described, assuming that the total pressure wake has the form of an error curve. A table of multiplying factors is given, for compressible flow up to an outlet Mach number of 0·9. Many published measurements of cascade drag have ignored this factor, with a consequent error of up to 25 per cent.

In his technical note on “Dimensions required for Inspection” (July Journal), Mr. G. A. G. Saywell stated that drawings calling up various machined fittings often lack any reference to “practical” measurement by which the operator, toolmaker and inspection department may “set out” and check the work concerned. This statement does not really do justice to the draughtsman who, in preparing the detail drawing for manufacture, has to give attention to all aspects of manufacturing.