In a decade when aircraft design procedures involve analyses which display ever increasing accuracy to satisfy growing operational demands, it is not surprising that the boundaries of compromise rapidly converge upon the envelope of acceptability. A finished aircraft becomes the unique solution to a well-defined design problem–or what is even more frightening, it becomes a solution to a unique problem. If the problem changes after the solution has been produced… c'est la vie.

In Part I the analysis was given of a circular hole (radius a) reinforced by a compact ring symmetrically placed either side of the shell skin which was assumed to be attached to the ring at its shear centre and centroid. The shell radius/thickness ratio (R/t) was assumed not to be large enough for non-linear effects to be significant.

A “light wire” is one for which the integrated wire weight and tangential fluid friction forces are much smaller than the weight of the load which it supports, so that only the aerodynamic force acting normal to the wire need be considered. The equations for this simplified model were first formulated by Glauert, and have been further studied by Bryant et al and Neumark, among others. Ref. 4 contains a comprehensive list of references.

In this note Glauert's equations are used to obtain numerical results for the wire shape, a wire shape integral which is of interest for trailing wire antennas, and the steady state aerodynamic forces which act upon the aircraft. Computer solutions taking cable weight into account have indicated that the results for a weightless cable are realistic for many cases.

The problem of flexural vibration of a rotating cantilever blade is non-linear when the vibration takes place in a plane other than perpendicular to the plane of rotation. The non-linearity arises from the effects of Coriolis accelerations due to the rotation of the blade mounted on the periphery of a disc. Lo and Renbarger neglected the non-linear terms in their investigation and arrived at a simple relation for the frequency of vibration in terms of the stationary frequency of the beam, the rotating speed and the orientation of the plane of vibration. Lo in his analysis simplified the non-linear problem by assuming the blade to be rigid everywhere excepting at the root and presented the solution in a phase plane. Isakson and Eisley have considered a similar model to determine the natural frequencies in coupled bending and torsion of twisted rotating cantilever blades.

In the theory of the incompressible laminar boundary layer, Stratford's method is widely regarded as one of the best practical methods for the prediction of separation. The method may also be used in conjunction with the Illingworth-Stewartson transformation (described below) for the prediction of the separation positions of certain compressible boundary-layer flows. For one such case, for which an “exact” solution has been computed, it has been suggested, in ref. 4, that the method, in its simplified form at any rate, is not quite as accurate as had been anticipated. It is the point of this note to show that the full Stratford method does in fact give an excellent prediction of the separation position for this flow.

In spite of the large number of published works on panel flutter there appears to be a wide gap in the literature concerning the flutter of skew panels. For instance, recently the flutter behaviour of skew panels in supersonic flow has been presented for a simply-supported boundary condition using double Fourier sine series to represent the deflection surface. Apart from this publication there is practically no literature concerning flutter of skew panels except refs. 3-4 which consider the flutter of skew panels clamped on all the edges. The method used in ref. 3 is the common 4-mode analysis by using the Iguchi functions for representing the deflections and in ref. 4 the same problem is solved by the use of beam characteristic functions. One inherent difficulty in these conventional methods, was that no single function could be chosen to represent the deformation which satisfied various boundary conditions, with the result that the entire analysis may have to be repeated with different assumed functions for accommodating different boundary conditions. Hence, a general method was proposed in ref. 5 for the study of panel flutter problems of arbitrary geometry by the Matrix Displacement Method, which permits application to problems with practically any geometrical boundary conditions on any or all the sides.

Many of the flow patterns associated with experiments in fluid mechanics include vortices, or vortex sheets, and frequently knowledge as to the position and strength of these vortices is desirable. Wool tufts and spinners are useful in determining vortex position, but can give no indication of the vortex strength or circulation. Yawmeters cannot readily give vortex position, although (with a fair amount of labour) the circulation can be determined.

It seems to me that the report contains 3 salient points in reference to domestic air services. These are:

1. 1. The unprofitability of domestic services generally.

2. 2. Competition on the trunk routes may have been beneficial to the customer in the short term but it may also have led to higher fares. (This need not however be true of the long term.)

3. 3. The absence of any coherent proposals for exploiting VTOL/STOL development in expanding domestic air services.

It is undoubtedly true to say that with the death of Archibald Reith Low on the 21st January 1969, there departed from the scene the last of the early pioneers of British aviation.

Born in Aberdeen on 31st December 1878, he was brought up with four brothers and three .sisters at the Manse at Fountainbridge, Edinburgh, where his father was Minister of the Church of Scotland. His mother was Jane Stuart Reith, aunt of the present Lord Reith, who was thereby his first cousin.

In most dynamic analyses, it is customary to use a coarser mass idealisation than structural idealisation especially if stresses are to be involved. Compatibility between the stiffness matrix and the smaller mass matrix M is achieved by reduction of the stiffness matrix. Guyan has shown that the reduction is equivalent to a transformation of the original matrix K :

where KR is the reduced stiffness matrix and T is the transformation matrix which contains stiffness terms.

A convenient aid to the investigation of the aeroelastic properties of aircraft is provided by wind tunnel tests on scale representative aeroelastic models. A wide variety of tests is commonly made ranging from tests on models of component parts of an aircraft (wing, tailplane, fin, etc.) to tests on models of the whole aircraft. In many of these tests it is adequate to mount the model structure rigidly to the tunnel wall or sting, but in other cases the model must be provided with rigid body freedoms for representative results to be obtained. This is generally necessary for aeroelastic models of complete aircraft, and particularly for models of slender configuration.

In an investigation, on the characteristics of the wake from slender cone-cylinders at large angles of attack, the strength and position of the vortices in the wake were determined for a large range of flow conditions. Two types of experiments were conducted, namely schlieren studies and wake traverses.

Schlieren photographs (such as Fig. 1) clearly reveal straight parallel vortex lines in the wake for subcritical values of the cross-flow Reynolds number component and for cross-flow components of the free stream Mach number up to at least 1·4. In ref. 1 numerous such schlieren photographs were analysed using the impulse flow analogy (refs. 2 and 3 for example). According to the analogy the progressive development of the wake along the body when viewed in cross-flow planes is similar to the growth with time of the flow behind a two-dimensional cylinder started impulsively from rest with the same cross-flow conditions.