It is shown that, for infinite sheared wings with symmetrical aerofoil sections, the value of the suction peak near the nose in incompressible flow depends primarily on the value of the lift coefficient and the nose radius and not on the angle of sweep.

When investigating the suitability of a basic wing section for an aircraft design with comparatively large aspect ratio, it is useful to know up to which lift coefficient one can expect the flow to be attached over the wing surface, and separation to occur at the trailing edge only. For aerofoils of small or moderate thickness-chord ratio, flow separation tends to occur first near the leading edge.

The last occasion on which a lecture was given to the Society on the subject of digital computing was in 1954, when Dr. Goodwin, of the Mathematics Division, N.P.L., spoke on “ The uses of the ACE Pilot Model Computer.” At that time there were very few electronic digital computers in regular operation in this country and only one was primarily engaged on aeronautical work.

An investigation into the turbulent flow characteristics in a smooth-walled annulus has been continued. Velocity and its correlation has already been dealt with in the author's note in the November 1958 JOURNAL (p. 830). In the work presented here, the eddy diffusivity in gas flow is discussed, its distribution predicted, and experimental results for the eddy difiusivity of heat in an insulated annulus with a heated inner wall are included in support.

The distributions of velocity, shear, temperature and heat flux for an annular flow with heat transfer at the inner boundary are shown diagrammatically in Fig. 1. The features of the turbulent velocity profile and methods of correlating velocity distribution are outlined in Ref. 1. The shearing stress T is given by the equation

Data were obtained of the pressure loss characteristics of metal strip matrices constructed by winding together alternate layers of crimped and plain metal strip. The hydraulic diameters of the passages so formed were in the range 0.023 in. to 0.060 in. with passage lengths from 1 in. to 6 in. and metal thickness from 0.002 in. to 0.006 in.

The units were mounted in a 6 in. diameter duct in which air velocities of up to 200 ft./sec. were obtained. The range of passage Reynolds number was from 300 to 6,000. The majority of the experimental results were for atmospheric pressure, but some tests were carried out at sub-atmospheric pressures to test a correlation equation derived from analysis of the results.

In a recent note, Fessler has suggested that some of the practical problems of strut testing can be overcome by using a “tensile strut-testing machine”. As Fessler points out, slender struts are frequently tested with pinned ends, either in the form of knife-edges or ball-ends.

The calculation of damping from the results of resonance and free oscillation experiments has hitherto been based on elementary theory in which the springs or other elastic supports are represented by lumped stiffness parameters. This “ light spring theory” is generalised here to include the elastic stiffness terms in distributed parameter form; i.e. the mass of the springs or other supports is taken account of by including wave equations for the propagation of stress waves in the springs and so on. A record of the free motion of a model in a wind tunnel, shown in Fig. 1., indicates the inertial effects of the spring supports. To minimise this “surging” of the supports one may design them to have high natural frequencies compared with the system frequencies. However, this gives no indication of the error entailed in using light spring theory, nor does it cover experiments which must be performed at high frequencies.

During the Second World War, a great many rockets were made and used operationally by both sides. By the end of the conflict, a pattern had begun to emerge. For applications where burning times were short and performance unimportant when compared with the need for simplicity, solid propellant rockets were used exclusively. Where control of thrust and the ability to start and stop repeatedly were involved, as in aircraft propulsion, or where performance was vital, as in the V2 ballistic missile, liquid propellants were applied. Since the war, this general pattern has persisted. There has been a progressive improvement in the performance of both types. At the same time, the thrust and total impulse of rockets have risen, until engines producing over one million pounds of thrust are envisaged.

The number of available exact solutions to the St. Venant flexure problem is limited, but there are some simple solutions which appear to have been overlooked. The sections covered are a wedge, a double wedge, a biconvex section, and an aerofoil-type section. All the solutions given are restricted to certain values of Poisson's ratio, depending on the section.

The components of shear stress parallel and perpendicular to the direction of loading are given by:

The Farnboro Indicator gives a large permanent pressure-time diagram with a linear time base and pressure scale. In its latest versions accurate diagrams can be taken over a wide pressure range. The major disadvantage is that it can only be used for cyclic repetitive processes such as those occurring in engines or specially designed machines. During the course of an investigation of pressure wave phenomena in an exhaust pipe of a high-pressure-charged two-stroke cycle engine model the Farnboro indicator was used to record shock waves travelling along the pipe and to measure the total head pressure downstream of a nozzle placed at the end of the pipe and hence to calculate the transient velocity. In this note some of the pressure records taken in the investigation are presented.

In his excellent review of the “state of the art” of effusion cooling (p. 73, February Journal, 1959), Dr. P. Grootenhuis clearly distinguishes the two mechanisms which contribute to the cooling effectiveness; namely, efficient abstraction of heat from the porous wall material by internal transfer to the effusing coolant, and reduction of the rate of external transfer of heat from the stream to the wall as a result of boundary layer injection. An observation that deserves emphasis, however, is that the first of these mechanisms is the predominant one. This fact is illustrated by the good agreement between the theory assuming laminar flow conditions (i.e. involving integration of the Navier-Stokes equations and the use of a one-parameter velocity-profile expression proposed by Schlich-ting for the laminar boundary layer with injection) and the various experiments, all of which involved turbulent stream conditions. This observed agreement indicates that the surface temperature is primarily controlled by the internal heat transfer within the porous wall and is relatively insensitive to the external heat-transfer conditions.

Professor Lighthill's paper has discussed the dissociative behaviour of air at the high temperatures produced when it is compressed in a shock or heated by friction in a boundary layer under hypersonic flight conditions. Just as molecules may dissociate into atoms, so may these molecules and atoms split further into charged ions and electrons. The equilibrium state of this ionisation process is amenable to the same thermodynamical or statistical-mechanical treatment as in the case of dissociation, the Law of Mass Action in this context being referred to as the Saha equation. Consider the simplest case where a monatomic gas is singly ionised, a fraction α of the atoms having lost one electron, so becoming singly charged positive ions.

In connection with the note by Henry Barrow in the November 1958 JOURNAL (page 830) it may be of interest to call attention to a little known reference which describes an extensive series of experiments. A 190 mm. pipe with 10, 35 and 100 mm. inserts was studied up to a Reynolds number of 106. Besides the fully developed profiles some interesting measurements showing the development of the flow in the inlet are included.

The general neglect of these results which constitute an appreciable fraction of the available data for this type of flow is surprising.

If a plane oblique shock wave, inclined to the free stream at the angle ε, is produced in two-dimensional supersonic flow of Mach number M by (for example) a wedge which deflects the flow through an angle δ, the equation connecting these quantities may be written

In this form, δ is given explicitly when M, ε are fixed. Similarly, we may obtain M explicitly when ε, δ are fixed; equation (1) may be written (see, for example, Liepmann and Puckett, Equation 4.27)

The intention of this paper is to point out certain aerodynamic problems that may be of future interest, by trying to argue in general terms about their possible context. Perhaps the most distinctive feature of any project analysis of a vehicle designed to carry a man into an orbit in space and return him safely to the ground, is the need to limit the accelerations and retardations of his passage, and the temperature of his environment. There are of course innumerable other considerations which have to do with his comfort and safety, but it is his inability to withstand excessive heat or high g which distinguishes man from more rugged but less versatile instruments.

The first of these conferences was held in 1953, the second in 1954 and the third in November 1958 at Milwaukee. This third conference was sponsored by the Agricultural Research Service of the United States Department of Agriculture and, to promote worthwhile discussions, attendance was by invitation only.

Twenty papers were presented on four main themes, economics, progress as regards agricultural applications, progress as regards forest applications and, finally, a review of present and proposed research.