An Artificial Satellite set to start on a circular orbit round the Earth, will have to be given a known velocity, at right angles to the radius vector, which depends upon its distance from the Earth's centre.

It will, from the start, begin to spiral downwards towards the Earth, owing to the resistance of the atmosphere, but if it begins its path at a sufficiently high altitude, the decrease of height after one revolution will be small.

Should the initial “ tangential ” velocity be greater than the value calculated for a circular orbit, the satellite will, according to the text books, describe an ellipse instead of a circle; and, if at a sufficient distance for the air resistance to be at first ignored, will return after one revolution to the distance at which it started, this point being the perigee of the ellipse. It will never return to a point outside the starting point.

There are two types of shock loading to which aircraft components are subjected, namely: medium shock loads which can heavily damage equipment; and shocks of high impact, such as those imposed during crash landing, causing the equipment to be torn loose from its mounting, with danger of injury or even death to aircraft personnel.

An impact shock machine has been installed recently in a laboratory for studying this important problem. The machine is capable of providing calibrated shock tests up to a maximum peak acceleration of llg, and of simulating the shock loads normally experienced under everyday operating conditions. Apparatus being tested is mounted on a frame bolted rigidly to the table of the impact testing machine, which has been designed to take specimens up to a maximum weight of 400 lb., with maximum overall dimension equivalent to a cube of 30 inches side. A specimen under test is dropped from a maximum height of 5 ft., falling freely.

This note describes an investigation into the secant and tangent moduli for flat panels in shear. Consideration is given to a comparatively recent analysis giving the secant modulus and comparison is made with he Society's Data Sheets.

In the early stages of preliminary design, it is essential to have a method for the comparison of aircraft on an operating cost basis. The standard S.B.A.C. method is too lengthy and, in any case, demands a greater amount of data than is usually available in the early design stages.

The formulae presented in this note were evolved by the author some years ago and the general method was quoted in a paper entitled “The helicopter—has it a future?” which was read to the Belfast Branch in January 1955. (See Journal, November 1955.)

The turbulent boundary layer on an aerofoil is represented by simple distributions of velocity, so as to obtain approximate general formulae for the displacement and momentum thicknesses in terms of two ratios of local velocity to that at the edge of the boundary layer.

Optimum positions for the local velocity are calculated to be at 0·15 and 0·62 of the boundary layer thickness from the aerofoil surface. The accuracy of the formulae is closely examined. A rapid experimental and numerical procedure is given for estimating the displacement and momentum thicknesses within ± 2 per cent. A minimum of seven velocity measurements has been found satisfactory in practice.

A Set of normalised orthogonal polynomial functions suitable for problems concerning doubly built-in uniform beams was given in R. & M. 228l. The nth function of the set was expressed linearly in terms of three Legendre polynomials as follows:—

where

The method referred to in Dr. Coleman's notes was developed with the collaboration of my colleague Mr. J. B. Edwards of Handley Page Research Department. The purpose was to obtain a rational estimate of suction quantities and suction distribution, linked up with measurements of boundary layer profiles and suction quantities on wind tunnel models, and also to assess the effect of a certain degree of roughness of the order to be expected on actual wings. Existing theoretical methods ignore roughness which, however, is a most important parameter not only in wind tunnel tests, but also in flight at higher values of the unit Reynolds number; surface roughness obviously limits the intensity of suction which can be applied at a spanwise suction strip.

It was intuitively assumed that the removal of fluid by suction was equivalent to cutting off the lower portion of the boundary layer profile at the upstream edge of the suction strip and that a rapid re-adjustment of the boundary layer profile within a short distance took place.

Several Attempts have been made lately to reduce the drag due to lift of wings by means of camber and twist. Slender wings in particular have been considered and a typical theory is that by Tsien on the supersonic conical wing of minimum drag, where the lift forces on the wing are calculated from conventional aerofoil theory so that the overall lift of a flat wing turns out to be proportional to the angle of incidence. For convenience, this is called a “ linear theory.” It is usual to consider two cases: (1) Wings where the drag is calculated by the same theory so that the full thrust component and the suction at the leading edge are included. (2) Wings where the drag is not calculated by the same theory as the lift but where the thrust at the leading edge is omitted.

A study has been made of the effect of inlet conditions on the performance of conical diffusers with 4:1 area ratio and 5 and 10 degrees total angle of expansion. The conditions at entry were varied by using different approach lengths of diffuser inlet diameter, and by means of projecting annular screens of woven wire cloth. With this new technique it was possible to vary the velocity distribution substantially within moderate settling lengths, and to produce velocity profiles with inflections. The suitability of the annular screen method of boundary layer generation for diffuser investigations was confirmed from the comparison with the approach length results.

Energy and pressure coefficients, as well as diffuser energy efficiency and the conversion efficiency, were found to depend on the diffuser angle β and the momentum thickness ratio at inlet θ/D0. The latter emerged as one of the chief parameters controlling diffuser performance. Variation in the inlet shape parameter H of the order of 20 per cent did not significantly affect the pressure recovery or the losses in the diffuser. For moderately thick boundary layers, θ/D0 X β<0·1, the diffuser angle could be eliminated as a parameter by plotting the pressure coefficient against θ/D0 X β. The Reynolds number based on diffuser inlet diameter was, for essentially incompressible flow, 2·5 X 105 in all tests.