This article focuses on the aerodynamic design of a morphing aerofoil at cruise conditions using computational fluid dynamics (CFD). The morphing aerofoil has been analysed at a Mach number of 0.8 and Reynolds number of $3 \times 10^{6}$ , which represents the transonic cruise speed of a commercial aircraft. In this research, the NACA0012 aerofoil has been identified as the baseline aerofoil where the analysis has been performed under steady conditions at a range of angles of attack between $0^{^{\kern1pt\circ}}$ and $3.86^{^{\kern1pt\circ}}$ . The performance of the baseline case has been compared to the morphing aerofoil for different morphing deflections ( $w_{te}/c = [0.005 - 0.1]$ ) and start of the morphing locations ( $x_{s}/c = [0.65 - 0.80]$ ). Further, the location of the shock wave on the upper surface has also been investigated due to concerns about the structural integrity of the morphing part of the aerofoil. Based upon this investigation, a most favourable morphed geometry has been presented that offers both, a significant increase in the lift-to-drag ratio against its un-morphed counterpart and has a shock location upstream of the start of the morphing part.

This paper aims to present a vertical take-off and landing unmanned aerial vehicle (VTOL UAV) configuration and numerically simulate its flight transition from hover to cruise and from cruise to hover. It can tilt the canard and wing along with two attached propellers. Additionally, two fixed front propellers are pointing upwards. Multi-body equations of motion are derived for this concept of aircraft, which are used to compute the flight transition trajectory from hover to cruise configuration. Furthermore, a transition control algorithm based on gain scheduling is described, which stabilises the aircraft while it accelerates from hover to cruise, gradually tilting the wing along with its propellers, sequentially switching between equilibrium states, as the stability cost functions thresholds are reached. The transition control algorithm of the conceptual aircraft model is numerically simulated.

This paper has proposed a new robust hybrid nonlinear guidance law, which accounts for a missile’s terminal line-of-sight (LOS) angle constraint, in order to intercept a non-cooperative maneuvering target. The proposed hybrid nonlinear guidance strategy consists of two phases; in the first phase, a guidance law named PIGL is derived from prescribed performance control and the inertial delay control method. In PIGL, a revised prescribed performance function is put forward, and a prescribed performance controller with unknown uncertainties is then derived. The controller smoothly drives both the LOS angle and its rate to a predesigned small region under unknown uncertainties that are induced by target’s maneuvers within a fixed time. Then, a guidance law named SIGL is activated, which is derived from sliding mode control and inertial delay control. By driving the desired sliding mode variable to zero within a finite time, the SIGL guidance law is able to achieve high terminal interception accuracy. The robustness of both of the proposed sub-guidance laws has been proved explicitly in this paper. The hybrid guidance law has the advantage of a tunable convergence rate of the LOS angle and the rate of the LOS angle at the beginning period, by which an excessive large initial maneuver can be avoided. Meanwhile, the hybrid guidance law also has the advantage of lower sensitivity to errors in the estimation of the time-to-go.

An integrated parametric model involving the design of propulsion system, airframe and flight mission is presented. Based hereon, the carbon dioxide (CO2) emission characteristics of advanced direct-drive turbofan and open rotor powered aircraft are analysed against pertinent aircraft and propulsion system design parameters. In addition, initial concept-specific trend statements on nitrogen oxides (NOx) as well as propulsor noise emission characteristics are derived. The obtained results contribute to a better understanding of more appropriate aircraft design attributes for advanced system architectures.

The ground delay program (GDP) is a commonly used tool in air traffic management. Developing a departure flight delay prediction model based on GDP can aid airlines and control authorities in better flight planning and adjusting air traffic control strategies. A model that combines the improved sparrow search algorithm (ISSA) and Multilayer Perceptron (MLP) has been proposed to minimise prediction errors. The ISSA uses tent chaotic mapping, dynamic adaptive weights, and Levy flight strategy to enhance the algorithm’s accuracy for the sparrow search algorithm (SSA). The MLP model’s hyperparameters are optimised using the ISSA to improve the model’s prediction accuracy and generalisation performance. Experiments were performed using actual GDP-generated departure flight delay data and compared with other machine learning techniques and optimisation algorithms. The results of the experiments show that the mean absolute error (MAE) and root mean square error (RMSE) of the ISSA-MLP model are 16.8 and 24.2, respectively. These values are 5.61%, 6.3% and 1.8% higher in MAE and 4.4%, 5.1% and 2.5% higher in RMSE compared to SSA, particle swarm optimisation (PSO) and grey wolf optimisation (GWO). The ISSA-MLP model has been verified to have good predictive and practical value.

The aviation community is currently faced with various approaches for the determination of Target Levels of Safety (TLS). The targets are usually derived for a specific airspace region, for a specific type of operation or a specific phase of flight. Therefore, current practices support the determination of TLS for specific aviation components in isolation. This paper argues that the setting of a new safety target for aviation has to be driven by an integrated system approach. Relevant past research on TLS is reviewed and augmented with the results from operational reports from two countries. A possible safety target for the year 2020 is suggested and scoped down to estimate the safety budget for air traffic control (ATC) equipment. The paper concludes with a discussion of the results and recommends useful practices to achieve the proposed integrated safety approach.

The following paper forms a continuation of a recent work by Prandtl (I) “On the Calculation of Boundary Layers.” It deals with the problem of developing in detail the method, proposed in that paper, for continuing a given velocity profile in a laminar boundary layer when the pressure distribution is given; the method is tried out in practice by numerical evaluation of an example.

The theme of flutter has not often been the subject of lectures before this Society, but the earliest of these was given as long ago as 1929 by my friend and former colleague Dr. R. A. Frazer. At that time a secure foundation for flutter theory had been laid and the principal methods for flutter prevention were already known. However, both theory and practice have developed much since 1929 and flutter has become of greater relative importance.

The title of this paper is not strictly correct since it is not really concerned with “efficiency” in the accepted sense of the word. It has in the past been considered that the sole duty of an engine air intake is to deliver air to the engine at a high efficiency, that is at a mean total pressure as nearly as possible equal to the aircraft pitot pressure. More recently other considerations have arisen, such as the necessity of avoiding asymmetric flow in twin duct air intakes. These aspects of air intake design have frequently been discussed, for example in Dr. Seddon’s lecture to the Society in December 1951, but the particular subject of this paper has been rather neglected—the velocity distribution of the air delivered to the engine.

This report describes a series of steady and time dependent pressure measurements on a half-model of a rectangular wing of aspect ratio 4. The wing was mounted on a halfbody attached to a sidewall of the RAE 8 ft x 8 ft windtunnel and the tests were made at Mach numbers of 0·60, 0·70 and 0·80 with fixed transition and a Reynolds number of about 2·4 x 106 at the highest Mach number. The wing was oscillated about its mid-chord axis at four frequencies giving frequency parameters up to 0·27 at M = 0·80. The static angle of incidence was varied from -10° to + 10°; here interest is centred on angles of incidence for which the flow on the upper surface of the wing is separated.

Selected mean and oscillatory measurements are offered as a challenge to computational fluid dynamicists for this simple threedimensional configuration. The measurements are of particular interest because the transonic flows include both incipient and well developed separations.

In various studies involving subsonic transition cavity flows, terms like ‘transitional-open’ and ‘transitional-closed’ cavity flow regimes are used in the literature. However, the exact fluid dynamic criteria to distinguish them have not been firmly established. The present work attempts to clarify this. Experiments to measure static pressure and fluctuating pressure distribution in a rectangular cavity with length-to-depth (L/D) ratio of 3 to 20 were performed. Based on pressure measurements on the front, bottom and rear faces of the cavity, additional criteria are established to clearly distinguish the range of critical L/D (or (L/D)cri) where these different transitional cavity flow regimes occur. The present work ascertains that the (L/D)cri for open and transitional-open cavity flow is 6·0-6·5, for transitional-open and transitional-closed cavity flow is 9-10 and for transitional-closed and closed cavity flow is 14-15. Flow visualisation qualitatively supports the flow pattern in the different cavity flow regimes.

Total air transport operating costs are divided into three general classifications : (1) Significant operating costs which are defined as those resulting from the flight of an aeroplane or those affected by variation of aeroplane size, power, initial cost, and aerodynamic efficiency. Equations for the development of significant costs are presented. (2) Passenger service costs which are defined as those which vary with the number of passengers aboard an aeroplane are enumerated. (3) Overhead costs in air transportation are also dwelled upon and it is concluded that these costs change their proportion to significant costs when the physical and performance characteristics of the aeroplane being considered change.

The marked effect on significant cost of varying aeroplane utilisation is brought out. Utilisation is also shown to be chiefly a function of trip length, with block speed, seasonal schedule variation and other factors being secondary functions.

The effect on significant cost of varying block speed either by increasing cruising power or improving aerodynamic efficiency is investigated for one type of transport aeroplane. Increased block speed obtained by increasing cruising power is found to increase significant cost very rapidly. Increased block speed obtained by improving aerodynamic efficiency is found to decrease significant cost appreciably and relatively large increases in aeroplane first cost are found to be justified if aerodynamic efficiency is improved. The adverse effect on significant cost of increasing trip length is investigated. The effect of changing specific fuel consumption is discussed.

It is pointed out that economic factors apart from those associated with fundamental aeroplane design and performance, such as interior arrangement, safety records, competition and general economic conditions, play the major role in the economic outcome of air transportation.

Lift alleviation by a mini-spoiler on aerofoils, unswept and swept wings encountering an isolated counter-clockwise vortical gust was investigated by means of force and velocity measurements. The flow separation region behind the spoiler remains little affected during the gust encounter. The maximum lift reduction is found for the static stall angle of attack. The change in the maximum lift during the gust encounter is approximately equal to that in steady freestream. The comparison with plunging aerofoils reveals that, for the same maximum gust and plunge velocity, the effectiveness of the mini-spoiler is much better in travelling gusts. This reveals the importance of the streamwise length scale of the incident gust. For the unswept wing, there is some three-dimensionality of the flow separation induced by the mini-spoiler near the wing tip. The magnitude of the lift reduction can be estimated using the aerofoil data and by making an aspect ratio correction for the reduced effective angle of attack. For the swept wing, the mini-spoiler can disrupt the formation of a leading-edge vortex induced by the incident vortex on the clean wing and can still reduce the maximum lift.

In a paper published in 1956 the senior author analysed the problem of the creep bending and buckling of a reinforced circular cylindrical shell subjected to the action of a bending moment constant in time and along the axis of the shell. With the aid of the principle of virtual displacements he found that the problem is governed by two simultaneous non-linear first-order differential equations. The equations were integrated step by step with the aid of a desk calculator and satisfactory agreement was found between theory and experiment when the results were compared with experimental data obtained in the laboratory of the Polytechnic Institute of Brooklyn. The experimental investigation and its results were described in ref. 1 as well as in a research memorandum of the National Advisory Committee for Aeronautics and the theory was presented in detail at the Symposium on Structural Problems of High-Speed Flight held at Cranfield, England, in August 1956 and published in the Journal of the Royal Aeronautical Society in 1957.

Three of the primary considerations involved in the design of an aeroplane are safety, weight and serviceaability. The structure weight being a major item in the design which can influence its efficiency as a flying machine, the considerations of safety and serviceability have always to be studied with respect to their effect on the weight of the structure. The history of the development of aircraft structural design indicates clearly the tremendous amout of work which has been devoted to the development of higher strength materials and more accurate stressing and testing methods, always with the aim of ensuring adequate safety and reliability consistent with efficiency of design, construction, and operation.

A previous paper was concerned with the uniform finite bending of a flat rectangular elastic plate about an axis parallel to one of its pairs of opposite edges. Expressions were obtained for the shape assumed by sections of the plate containing the axis of curvature, and the applied bending moment. In the present paper the work is extended to deal with a plate having shallow regular corrugations of arbitrary shape running along it at right angles to the axis of curvature. Such a plate may be of interest as it exhibits the type of instability characteristic of curved thin-walled members. Its solution can be derived quickly from the previous work.

A method is presented for determining the potential flow around bodies of revolution, at incidence, in uniform, incompressible, flow. This method utilises polynomial distributions of both sources and doublets, in the manner of Fuhrmann, and gives considerable advantage over existing techniques. In particular, when compared with the method of Nielsen, in a typical case, the computer storage requirement is reduced by a factor of eleven. Compared with panel methods the saving is very much greater.

Unlike the method of Landweber, both body surface and flow field conditions are obtainable and it would appear that the new method is ideally suited to the ‘store trajectory’ application in place of the existing discrete source method.

A unified analytical method is presented for evaluating the generic buckling behaviours of simply-supported thick orthotropic cylindrical shells subjected to both axial compression and lateral pressure. The thick shell theory and Flugge type equations are employed. The equilibrium equations of shells become particularly simple in the chosen space by considering an appropriate transformation and utilizing some previous material constant definitions due to Brunelle. Hence, comprehensive solutions are found to this problem so that the curves presented in the text are generic rather than specific. The analysis is for the restricted case of a single homogeneous layer. The effects of various parameters on buckling loads are studied. The results indicate that the buckling coefficient increases with increasing generalised rigidity ratio D* but decreases with increasing generalised Poisson's ratio ε. The results of a reduced set of equations agree with those obtained by previous investigators.

In 1909 Monsieur Louis Blériot forged a link between our two countries, one might say an aeronautical link. He did so in an aeroplane which by today's standards took only a short take-off run and a short landing.

Today there are signs of new links being forged between the two countries and among them is co-operation in the development of vertical take-off and landing aircraft. There can, therefore, be no more suitable subject for a lecture in commemoration of Louis Blériot than one on “The Role of Jet Lift.”

—Measurements of pressure distribution and skin friction have been made on an 11% thick aerofoil in the presence of a variable base bleed flow. These tests show that the effects of bleed decay rapidly upstream of the base and are measurable only over the rear 40% of the chord. The integrated profile drag values also confirm that wake traverse measurements are subject to error in a strong vortex street.