From Kármán’s similarity hypothesis, we derive the equation which describes the mixing length in terms of the turbulent shear stress. For a boundary layer with linear stress distribution, the equation is in reasonable agreement with Bradshaw’s measurements. For a boundary layer with injection, it is shown that injection has an appreciable effect upon the mixing length when (vw/2) /(τ/ρ)1/2 becomes comparable to the Kármán constant. Close similarity is also pointed out between the hypotheses due to Kármán and Townsend. Moreover, the diffusion constant in Townsend’s hypothesis is determined to be 0.25, which is in good agreement with the value 0.2 obtained by Townsend from one experiment.

Early in 1946, when work on guided weapons had only just begun in the United Kingdom, it became clear that there was a future need for a long Range, preferably over land, for the testing and development of guided weapons. Such a Range could not be located in the British Isles, and a survey throughout the British Commonwealth indicated that Australia was most likely to provide a suitable site. Therefore, a party of Ministry of Supply and Ministry of Aircraft Production personnel under the leadership of Lieut. General (now Sir John) Evetts carried out a preliminary survey in 1946, selected a site in South Australia which appeared most suitable and submitted a detailed report to the Ministry of Supply.

Tension bolts are widely used in engines, in machinery of various kinds, and in structures. They have unique advantages in their contribution to ease of assembly and dismantling. They also give a comparatively rigid and compact joint, and from the standpoint of static strength are extremely efficient. Unfortunately they have not got a good reputation from the standpoint of fatigue. Some designers go to considerable trouble to avoid them. Others accept their limitations and ensure safety by care in detail design and by adopting conservative figures for design stresses. Where necessary, a limitation is placed on working life, at the expiration of which the nut and bolt are replaced by new ones.

Although the Hatfield tunnel has been doing some useful testing for just over eighteen months, the experience so far obtained is still limited as each new model brings its own special problems. It is perhaps still too early to give an account of the experience obtained, but because of the interest in this tunnel it was thought that an account, even incomplete, would be useful.

Almost every facet of experience so far obtained is related to some of the special design features of the tunnel and I propose to review some of the problems encountered during the design stages of the tunnel dealing with these design features, before reviewing the actual model testing experience.

Waverider serves as a good candidate for hypersonic vehicles. The typical waverider has sharp leading edge and no control face, which is inappropriate for practical use. This paper puts forward a method modifying the waverider, and the modification impact on the performance of waverider at hypersonic flow conditions is studied. The modification is based on blunted waverider, includes cutting the tip and introducing two control wings. The modification’s effect on aerodynamic performance is obtained and analysed through Computational Fluid Dynamics (CFD) techniques. When blunted with 2cm radius, the waverider retains its good aerodynamic performance and the heat flux at the stagnation point can be managed. Three factors of the introduced wing are argued, the fixed angle, aspect ratio and wing area. Results show that influence on the aerodynamic coefficient is slight and the vehicle retains its high lift-to-drag ratio. The main influences of the modification are the control ability and trim efficiency, which is the motivation of this work and can be adapted when designing a practical waverider.

Three theories of bending are available to the designer for the determination of bending stresses in curved beams. (1) The Euler-Bernoulli hypothesis (often referred to as the simple theory of bending), which assumes a linear distribution of bending stress, having a zero stress value at the mid-plane, i.e. the centroidal and neutral axes are coincident. (2) The Winkler theory, which predicts a depression of the neutral axis away from the centroidal axis and towards the centre of curvature. This theory produces a hyperbolic distribution of bending stress across the section, having a maximum amplitude at the fibre nearest the centre of curvature. (3) The Golovin theory, which is exact.

Increasingly stringent attitude stabilisation requirements are the current trend in both experimental and commercial satellites as is seen in the current Intelsat, ESRO and UK communication studies. These craft must be lightweight, compact and rugged during the launch phase but after mission capture such requirements no longer apply. The increasingly high power requirements of such craft are met by the use of large flexible solar arrays which are packed away during launch and unfurl when the craft becomes operational. For the three-axis stabilised craft being studied reaction jets are used to achieve the high pointing accuracy required. Such actuators may be hot or cold gas systems or in the future may be electric engines’. The broad spectral content of these actuators will inevitably excite modes of vibration in a wide range of frequencies. The influence of these highly resonant modes on the performance of on-board controllers, needing a relatively high bandwidth of 0-10 Hz, to achieve high pointing accuracy of up to a few seconds of arc may lead to design difficulties.

Numerical simulations were performed to analyse the possibility of damping abrupt incoming free stream disturbances upon a porous aerofoil using an unsteady Reynolds-averaged Navier-Stokes (URANS) model. To mimic the turbulence disturbance levels that are typically encountered in the atmosphere, two flow configurations were considered. In the first configuration, the unsteadiness of the flow was created with vortices shed from a circular cylinder installed ahead of a WTEA-TE1 aerofoil. The continuous von Kármán shedding vortices contained within the cylinder wake were convected downstream and projected upon the aerofoil. In the second configuration, an instantaneous pair of discrete vortices was created by a rotational snapping of a flat plate, installed upstream of the aerofoil. Solid and porous aerofoil configurations, with porosity settings of 11 and 22%, were applied on 50% of the chord of the aerofoil starting from the leading edge. Both steady and unsteady flow simulations were performed to assess the performance of the porosity under steady and unsteady effects. The steady state flow simulations revealed a noticeable reduction in the aerofoil lift coefficient for the porous aerofoil. For unsteady solutions with a continuous or distinct series of vortices interacting with the aerofoil, the porosity showed insignificant damping of the lift coefficient amplitude. The porosity values investigated in the current exercise had indiscernible effect upon the unsteady lift-load alleviations caused by free stream disturbances.

On the basis of well-known literature, an analytical tool named LEAF (linear elastic analysis of fracture) was developed by the authors to predict the damage tolerance (DT) proprieties of aeronautical stiffened panels. The tool is based on the linear elastic fracture mechanics and the displacement compatibility method. By means of LEAF, an extensive parametric analysis of stiffened panels, representative of typical aeronautical constructions, was performed to provide meaningful design guidelines. The effects of riveted, integral and adhesively bonded stringers on the fatigue crack propagation performances of stiffened panels were investigated, as well as the crack retarder contribution using metallic straps (named doublers) bonded in the middle of the stringers bays. The effect of both perfectly bonded and partially debonded doublers was investigated as well. Adhesively bonded stiffeners showed the best DT properties in comparison with riveted and integral ones. A great reduction of the skin crack growth propagation rate can be achieved with the adoption of additional doublers bonded between the stringers.

A method based on the shedding and convecting of discrete vortices has been developed to calculate the transient loading on an aerofoil-spoiler configuration when the spoiler is suddenly deployed to a particular angle. An innovation in the method is the introduction of Kutta conditions at the separation points.

Although the variation of CL with time is highly irregular, qualitative trends can be deduced. For a range of aerofoil angle of incidence and spoiler angle deployment adverse lift effects (ie, when the lift initially increases) have been obtained, similar adverse lift effects have been measured in experiments.

This paper dwells upon investigating the effect of aspect ratio (AR) variation on the aerodynamic performance of unconventional control surfaces called grid fins by virtue of a series of subsonic experiments on a simplified grid fin variant called the cascade fin. Wind tunnel tests were performed for different AR (variable span) grid fins. The same had been investigated for different gap-to-chord ratio (g/c) variants. Results demonstrated a tangible increase in the aerodynamic efficiency as well as stall angle reduction for higher AR. Moreover, higher AR leads to increased pitching moment, which emphasizes elevated hinge moment requirements. The study ensued the presence of higher deviation between the low AR fins, that is $AR<2$ compared to the pertinent deviations between the high AR fins, that is $AR\geq2$ . The effect associated with these variations was termed as span effect in this paper. It was established that, the deviations arising due to this phenomena were lesser for higher g/c and higher AR. The analysis of AR variation for different g/c presented a limiting value of AR reduction for stall performance enhancement. Thus, optimised selection of the g/c and AR values can lead to enhanced aerodynamic efficiency alongside an improved stalling characteristic.

This paper describes the application of low-order models to the prediction of the steady performance of axial compressors at sub-idle conditions. An Euler body-force method employing sub-idle performance correlations is developed and presented alongside a mean-line approach employing the same set of correlations. The low-order tools are used to generate the characteristic lines of the compressor in the locked-rotor and zero-torque windmilling conditions. The results are compared against steady-state operating points from three-dimensional (3D) Reynolds-averaged Navier–Stokes (RANS) computational fluid dynamics (CFD) simulations. The accuracy of the low-order tools in reproducing the results from high-fidelity CFD is analysed, and the trade-off with the computational cost of each method is discussed. The low-order tools presented are shown to offer a fast alternative to traditional CFD which can be used to predict the performance in sub-idle conditions of a new compressor design during early development stages.

In an analysis by the matrix force method of a structure containing thin tubes of open cross-section, it is commonly assumed that the torsional stiffness of such members may be taken into account by considering the St. Venant shear stresses only. This effectively introduces the assumption that the cross-section of an open tube is able to warp freely, whereas in practice the warping may be constrained by the surrounding structure.

In this note a method is presented which allows such warping constraint effects to be taken into account in a finite element analysis using the Force Method. Because of the assumption of linear variation of direct stress along the element the full flexibility matrix does not have to be used, which may be an advantage in certain cases. The “exact”, but more complex, flexibility matrix has been derived and will be presented in a later note.

It is well-known that, at large distances from the flight path, the waveform of the sonic bang from an aircraft, without lift at any rate, has the characteristic N-wave shape, as illustrated in Fig. 1. When we depict a waveform such as this we are stating what would occur at large distances in the absence of viscosity and other diffusive effects, such as turbulence. We know, of course, that in reality, because of these effects, the bow and stern shocks will be “softened”, i.e. the pressure jumps associated with them will not be instantaneous, but will have a finite, but small, “rise time”. However, this is not the matter under discussion. The essential feature of the N-wave illustrated in Fig. 1 is that the pressure jumps associated with the bow and stern shocks are equal, and that the positive and negative phases are of equal duration. The arguments for the waveform to take up this shape at large distances are put in a classical paper by Whitham.

Some observers believe that there will be such growth in the travelling public and such constraint on airport and airspace usage that, by early in the 2000s. very large passenger carrying aircraft will be required. By ‘very large’ is meant anything from 1,000 to 2,000 seaters. Whilst the problems which come to mind when musing on this possibility are mostly those of ‘passenger flow’ in the widest sense and rather less related to the vehicle itself, there are some issues related to the vehicle which suggest themselves. One is the choice of fuselage cross section.

Because cabins have to be pressurised (10 psi is a typical figure for cruise altitudes up to say 40,000 ft), it has been the norm to utilise circular cross sections (or double-bubble type) since the pressure forces are contained by hoop loads only, without frame bending, which affords a light and fatigue-insensitive structure (Fig. 1).

A recent paper(1) gives details of an investigation on the low endurance fatigue behaviour of an aluminium-zinc-magnesium alloy (D.T.D. 683). It becomes apparent in subsequent discussion that information on another aluminium alloy to B.S. L65C would satisfy a wider practical interest than the former. Although the equipment used for the earlier work was engaged on another project, it was decided that the latter should be halted for a period to enable a brief programme on low endurance fatigue to be conducted on B.S. L65C material which had previously been kindly provided by the Aluminium Development Association.

For the purpose of the design and certification of inflight icing protection systems for transport and general aviation aircraft, the eventual re-definition/expansion of the icing environment of FAR 25/JAR 25, Appendix C is under consideration. Such a re-definition will be aided by gathering as much inflight icing event data as reasonably possible, from widely-different geographic locations. The results of a 12-month pilot programme of icing event data gathering are presented. Using non-instrumented turboprop aircraft flying upon mid-altitude routine air transport operations, the programme has gathered observational data from across the British Isles and central France. By observing a number of metrics, notably windscreen lower-corner ice impingement limits, against an opposing corner vortex-flow, supported by wing leading edge impingement limits, the observed icing events have been classified as ‘small’, ‘medium’ or ‘large’ droplet. Using the guidance of droplet trajectory modelling, MVD values for the three droplet size bins have been conjectured to be 15, 40 and 80mm. Hence, the ‘large’ droplet category would be in exceedance of FAR/JAR 25, Appendix C.

Data sets of 117 winter-season and 55 summer-season icing events have been statistically analysed. As defined above, the data sets include 11 winter and five summer large droplet icing encounters. Icing events included ‘sandpaper’ icing from short-duration ‘large’ droplets, and a singular ridge formation icing event in ‘large’ droplet. The frequency of ‘large’ droplet icing events amounted to 1 in 20 flight hours in winter and 1 in 35 flight hours in summer. These figures reflect ‘large’ droplet icing encounter probabilities perhaps substantially greater than previously considered. The ‘large’ droplet events were quite localised, mean scale-size being about 6nm.

The emergence of electronic computers has brought about a revolution in the analysis of aircraft structures in the past decade which is now generally accepted throughout the Industry. But the major impact of computers on design has, in my opinion, not yet been realised and I hope, in this short paper, to show how powerful a tool they can become in the hands of imaginative design engineers. The direct benefits to the designer lie in the ability of the computers to perform large amounts of routine arithmetic in a short space of time. This capability enables us to obtain an insight into structural behaviour and to consider the influence of the aircraft environment and design alternatives from the earliest stages of design. By the development of general purpose computer programmes for handling routine calculations the structural designer can be freed from much of the laborious calculation associated with complex structures and he is able to devote more time to creative work. An even more important indirect benefit is the basic re-thinking of design processes which is necessary in order to reduce them to routine arithmetic. We begin to realise that many aspects of efficient design such as continuity, cost and weight can be introduced as quantitative parameters instead of relying on intuitive compromise.

Features of aircraft which influence flow separations, and hence the onset of buffeting, should be always of interest to an aerodynamicist. The present compilation of such features is based on notes made during a visit to the USAF Aircraft Museum in Dayton, Ohio. Although two low-speed aircraft with high aspect ratio wings are considered, the emphasis is primarily on transonic and supersonic military aircraft, with wings of lower aspect ratio.

Some recommendations are made to stimulate research into improved methods to control flow separation, particularly for future transonic and supersonic aircraft.

The analysis of risks per phases of flight is fundamental for safe helicopter operations, of which night-time offshore oil- and gas-related missions form an important part. The safe execution of such missions also depends on pilots’ recent flying practice and a stable visual approach segment prior to landing. However, the poor quality of the safety data currently available prevents accurate analysis of risk on a per-phase-of-flight basis, establishment of a meaningful flying recency requirement and identification of any preferable visual approach design. To redress these problems, this paper develops a bespoke taxonomy of phases of offshore helicopter flights and uses it as the basis for a questionnaire survey on the phase-specific risk levels experienced by pilots in the night-time, perceived optimal flying recency requirement and preferred visual approach design. With the responses obtained from pilots located in seven countries, extensive statistical hypothesis testing shows that the phases involving visual scan techniques at high speed regimes are problematic, especially the visual segment of instrument approaches. Moreover, the between-night-flights time gaps required for assured flying recency were found considerably shorter than currently standardised across the industry. Finally, no preferred visual approach technique was identified. A number of important implications have been highlighted and should form the basis for future safety interventions.