This paper describes the results of an investigation of the flow over two cylinders in which the diameter of the upstream cylinder is half that of the downstream cylinder. Measurements were made of the surface pressure distribution around the downstream cylinder, and of the base pressures and Strouhal numbers of the two cylinders for separations between 1·32 D and 3·88 D, where D is the diameter of the downstream cylinder, and for angles of stagger between 0° and 45°.

With the cylinders in tandem two different flows were found, and it is suggested that there is a step change between the flows at a critical separation.

The behaviour of the measured quantities with a change of the angle of stagger is found, in general, to depend on whether the separation is less or greater than the critical separation. For each separation an angle of stagger is found at which the pressure distribution has a strong asymmetry over the rear of die cylinder. These pressure distributions are similar to each other, and they indicate the presence of an enhanced flow over the upper half of the cylinder. Also the lift force, for a given separation, then has its maximum value.

It is found that for separations greater than the critical separation the Strouhal numbers of the vortex shedding from the two cylinders are identical over a considerable range in the angle of stagger.

Inlet start/unstart detection is one of the most important issues of hypersonic inlets and is also the foundation of protection controls of scramjets. In ground and flight tests, it is inevitably to introduce the sensor noises to the measurement system. How to overcome or weaken the influence of the sensor noises and the outer disturbances is an important issue to the control system of the engine. To solve this problem, the 2D inner steady flow of hypersonic inlets was numerically simulated in different freestream conditions and backpressures, and two different inlet unstart phenomena were analysed. The membership function for hypersonic inlet start/unstart can be obtained by using probabilistic output support vector machine, and the algorithm of multiple classifiers fusion is introduced. The variations of the classification accuracy with the intensity of the sensor noises and the number of the classifier were discussed respectively. In conclusion, it is useful to introduce the algorithm of support vector machine and multiple classifiers fusion to overcome or weaken the influence of the sensor noises on the classification accuracy of hypersonic inlet start/unstart. The number of the practical fusion classifiers needs a tradeoff between the fusion classification accuracy and the complexity of the classification system.

The separation or breakaway of over-expanded gas from a nozzle wall plays an important role in the estimation of the thrust from a rocket engine when run at altitudes considerably less than the design. A number of reports have been written on the subject but the agreement of test results has not appeared to be good.

In an effort to clarify the position, the Rocket Department of Bristol Siddeley Engines Ltd. has carried out tests on experimental combustion chambers having static pressure tappings in the nozzle divergent section. These tests have shown that, contrary to some previous suggestions, the pressure at which the gas separates is neither constant nor independent of the nozzle length. The paper, which includes results from other workers, divides the field into separation adjacent to the nozzle exit and also at positions where the area ratio is less than 80 per cent of the exit area ratio.

For the most part, consideration is given to conical nozzles having divergence half angles of 15° to 17°, but comment is made on the effects of divergence and other variables. A tentative suggestion as to the mechanism of separation is also made.

Finally, the effect of separation on the sea-level thrust of a nozzle designed for high altitude Is shown.

It is both an honour and a responsibility to deliver a lecture commemorating the names of two scientists who have made unique contributions to fluid dynamics, and whose names perhaps above all others will remain immortal in our subject. The range of topics that would be suitable for such a lecture is very wide, because there is scarcely a branch of fluid dynamics in which the work of Reynolds and Prandtl does not continue to play a significant part. In a lecture of current interest there is, therefore, no problem of relating the subject matter to the names to be commemorated; the influence of Reynolds and Prandtl remains as strong as when they were alive, and many of the problems that engaged their attention are even today only partially solved. The lecturer must, therefore, use other considerations for guidance in the selection of his subject from the limited range on which he is qualified to speak, for few could hope to work over the wide field covered by Reynolds and Prandtl. He must if possible select a, topic in which major contributions have originated from Germany and the United Kingdom, on which research is actively in progress in both countries, and for which progress has relied on the combination of theoretical and experimental work to provide acceptable simplifying assumptions that characterised many of the investigations of Reynolds and Prandtl.

The first basis for the design of any power-driven mechanism is an exact knowledge of the characteristics and capabilities of the power unit available. Without such information no designer is able to start on the initial sketch of the mechanism. The man powered aircraft is no exception and its designers have suffered in the past by insufficient knowledge about man's power. This may seem remarkable because until about two centuries ago almost all the world's work was done by muscle-power, and much of the muscle was human.

A problem of some practical importance concerns the interference between a wing surface in a free stream and a nearby streamwise vortex. This situation can arise for example, either in a helicopter rotor system, or when concentrated vortices, which are shed from a fuselage nose or from forward canard controls, interfere with an aft lifting surface.

An approximate non-linear analytical study is described in Ref. 1 in which the three-dimensional loading on a two-dimensional non-swept wing underneath a streamwise line vortex in low speed flow (as shown in Fig. 1) is deduced. Since publication of Ref. 1 a number of linearised theoretical studies have come to light.

This paper describes the current status of knowledge regarding the impact of aviation on the atmosphere. The growth of the aviation industry is likely to continue in the future and at present there are significant concerns that this will adversely affect climate and local air quality in the vicinity of airports. Indeed, it is possible that air quality may become a constraint to increased capacity of some large hub airports. Currently the radiative forcing impact from aircraft emissions, other than CO2, are not accounted for in the International Civil Aviation Organization emission trading scheme proposals. Taking only this approach, emissions trading could, in theory, increase the radiative forcing, rather than decrease it. The study of improved ‘environmentally friendly’ flight is still in its infancy, however any proposed and actual new developments, in terms of technology and operational practice, should include an environmental assessment.

The paper presents an experimental study of vortex systems generated in the region where an incompressible flow with spanwise varying velocity moves about a circular cylinder connected to a flat plate. The role that the various flow and geometric parameters play in the generation of the horseshoe vortex at the front of the cylinder is discussed and the dimensionless similarity numbers which describe the flow are derived.

Trailing edge blowing over a Coanda surface has been utilised as a circulation control mechanism for increasing the lift of an aircraft wing. Typically, high energy air is blown from a narrow spanwise slot over the rounded trailing edge of a wing and the air supply is modulated to effect a degree of lift control on the wing. This configuration produces an aerodynamic force in a uni-directional sense only. An alternative novel flow control actuator is described which utilises a simple variable geometry Coanda surface with upper and lower spanwise blowing slots to achieve fully proportional bi-directional control in the manner of a conventional flap. A prototype device has been wind-tunnel tested and is shown to have substantially linear response characteristics and to be as efficient as an equivalent flap surface. The performance of a flow control actuator suitable for small UAV applications is described.

The long-term goal is to design and manufacture optimal ‘on-demand’ vortex generators, ‘dimples’ that can produce vortices of prescribed strength and duration for the real-time control of aerodynamic flows that are either undergoing transition or are fully turbulent, attached or separating. Electro-active polymers (EAP) are ideal for a dimple control surface, offering high strain rate, fast response, and high electromechanical efficiency. EAP can also be used as the basis of a resistanc – or capacitance – change pressure sensor, development of which has just begun. In terms of manufacture, inkjet printing of EAP also offers a paradigm shift such that a monolithic control surface is a very real possibility. Important features for integration into a control system are robustness and a predictable, repeatable motion. With these objectives in mind, the suitability of EAP-based actuators is assessed both mechanically and aerodynamically. The ultimate goal is to integrate these devices, along with shear-stress and pressure sensors and distributed control, also under development, into a flexible ‘smart skin’ which could be incorporated into an airframe structure. The response of a laminar boundary layer to forcing is investiagted using mechanical dimples.

Real-time flight data from the Automatic Dependent Surveillance–Broadcast (ADS-B) has been integrated, through a data interface, with a flight performance computer program to predict aviation emissions at altitude. The ADS-B, along with data from Mode-S, are then used to ‘fly’ selected long-range aircraft models (Airbus A380-841, A330-343 and A350-900) and one turboprop (ATR72). Over 2,500 flight trajectories have been processed to demonstrate the integration between databases and software systems. Emissions are calculated for altitudes greater than 3,000 feet (609m) and exclude landing and take-off cycles. This proof of concept fills a gap in the aviation emissions inventories, since it uses real-time flights and produces estimates at a very granular level. It can be used to analyse emissions of gases such as carbon dioxide ($\mathrm{CO}_2$), carbon monoxide (CO), nitrogen oxides ($\mathrm{NO}_x$) and water vapour on a specific route (city pair), for a specific aircraft, for an entire fleet, or on a seasonal basis. It is shown how $\mathrm{NO}_x$ and water vapour emissions concentrate around tropospheric altitudes only for long-range flights, and that the cruise range is the biggest discriminator in the absolute value of these and other exhaust emissions.

In the spring of 1904, when echoes of the Brothers Wright first flights began to be known in Europe, the author founded at Koutchino, near Moscow, on his family estate, the Aerodynamic Institute of Koutchino.

Air is the fluid meant by the term “aerodynamic” and artificial flight was the most important of the applications included in the programme of the Institute; foreseeing, however, the importance of comparative researches in air and water, the author chose intentionally a location particularly suitable for hydraulic experiments. A small river, dammed so as to form a pond, afforded the possibility of regulating the speed of the current at will, and since 1905 parallel researches were carried on in the wind tunnel and in the stream of water.

The idea of such an institute was very favourably received in the Russian scientific circles. When the author asked N. E. Joukovsky, V . V . Kouznetzoff and S. S. Nejdanovsky to participate in the organisation of this institution, they willingly gave their assent.

The noise problem associated with an aircraft flying at supersonic speeds is shown to depend primarily on the shock wave pattern formed by the aircraft. The noise intensity received by a ground observer from a supersonic aircraft flying at high as well as low altitudes, is shown to be high, although it is of a transient nature.

A study of the shock wave patterns around an aircraft in accelerated and retarded flight is shown to lead to an explanation of the one or more booms, of short duration, heard by ground observers after an aircraft has dived at supersonic speeds.

The shock wave patterns associated with an aircraft flying in accelerated or retarded flight at transonic speeds are shown in certain cases to be very different from the corresponding patterns observed in steady flight. The significance of these results, with reference to problems of flight at supersonic speeds is briefly discussed.

The Galerkin method for the approximate solution of elasticity problems (see e.g. Ref. 1) is usually presented as an alternative to the Rayleigh-Ritz method. The main distinction between the two methods is stated to be that the former begins with an equation of equilibrium, whereas the latter begins with a total potential energy expression. The two methods are not always equivalent, and the conditions for equivalence are given in Duncan's lucid presentation of the method (Refs. 2 and 3, which unfortunately have been out of print for some time). These conditions, however, are generally omitted from discussions of the Galerkin method, and it is the purpose of this note to re-emphasise them in a slightly different form.

Measurements of mean velocity and Reynolds shear stress have been made in a turbulent boundary layer downstream of a small step change in surface roughness. Upstream of the step the surface is smooth, while downstream it consists of a d-type rough wall made up by a series of two-dimensional elements of square cross section placed transversely across the flow and spaced one element width apart in the direction of the flow. The calculated mean velocity and Reynolds shear stress profiles obtained using the method of Bradshaw, Ferriss and Atwell are in good agreement with the measurements throughout the relaxation region of the layer. Well downstream the calculation method adequately reproduces the self-preserving features of a d-type rough wall.

Fitting a spike on a blunt body provides a drag reduction at supersonic and hypersonic speeds. In this study, the laminar flow over a spiked, conical body terminated by a spherical cap, inspired by the Apollo re-entry capsule design, was investigated using a hypersonic wind tunnel. Schlieren pictures revealed the absence of flow unsteadiness for the range of spike lengths tested, and force measurements showed a maximum reduction of 77% of the unspiked body drag.

A simple theoretical model based on the pressure drag generated by a solid cone showed good agreement with the experimental data. The measured shock stand-off distance agreed well with predictions.

A method of design is given for wind tunnel contractions for two-dimensional flow and for flow with axial symmetry. The two-dimensional designs are based on a boundary chosen in the hodograph plane for which the flow is found by the method of images. The three-dimensional method uses the velocity potential and the stream function of the two-dimensional flow as independent variables and the equation for the three-dimensional stream function is solved approximately. The accuracy of the approximate method is checked by comparison with a solution obtained by Southwell's relaxation method.

In both the two and the three-dimensional designs the curved wall is of finite length with parallel sections upstream and downstream. The effects of the parallel parts of the channel on the rise of pressure near the wall at the start of the contraction and on the velocity distribution across the working section can therefore be estimated.

The impact of an axisymmetric jet onto a ground board is a flow of some interest in VTOL aircraft aerodynamics. This paper presents a simple argument for correlating data obtained in the impact region from different nozzle configurations and for different nozzle heights above ground. The argument is applied, in particular, to the static pressure on the ground board and to the peak dynamic head in the flow over the ground board.

A method is developed enabling rapid determination of the optimum cross-sectional dimensions of compression surfaces having unflanged integral stiffeners, and consideration is given to the effects of practical limitations on the design. The theoretical efficiency of the optimum integral design is found to be only 85 per cent, of that of optimum Z-stringer design.

Large deflection analysis of a cantilever beam under a tip concentrated rotational load governed by a second order non-linear differential equation is solved using a fourth-order Runge-Kutta integration scheme. Initially the two point boundary value problem is converted to an initial value problem by estimating one of the two required initial conditions in an iterative process, so as to satisfy the other boundary condition. The details of load deflection curves are presented.