The title of this paper clearly has a wide scope. None will dispute the presence of many problems in aircraft development. It was suggested in the first place that this paper might take the form of a description of the workings of a chief designer's mind when a new aeroplane is conceived. Such a paper would deal largely with technical considerations, such as choice of engines, wing loading, plan form and so on.

The diversity of uses to which the modern aeroplane is put gives rise to great differences between types. A jet fighter, for example, will be influenced by considerations far removed from those controlling the design of a long—range civil transport. To attempt to cover in detail such a wide field would result in an encyclopaedia rather than a paper; to tackle it in general terms would beget a vagueness lamentably devoid of interest.

Therefore, it has been thought wiser to attempt an assessment of some of the more fundamental problems confronting aviation.

The problem of flexural vibration of a rotating cantilever blade is non-linear when the vibration takes place in a plane other than perpendicular to the plane of rotation. The non-linearity arises from the effects of Coriolis accelerations due to the rotation of the blade mounted on the periphery of a disc. Lo and Renbarger neglected the non-linear terms in their investigation and arrived at a simple relation for the frequency of vibration in terms of the stationary frequency of the beam, the rotating speed and the orientation of the plane of vibration. Lo in his analysis simplified the non-linear problem by assuming the blade to be rigid everywhere excepting at the root and presented the solution in a phase plane. Isakson and Eisley have considered a similar model to determine the natural frequencies in coupled bending and torsion of twisted rotating cantilever blades.

In the “Peak Amplitude” method of resonance testing the total amplitude of vibration is plotted against frequency or (frequency). The method is unreliable for systems which are heavily damped and for systems with close natural frequencies (for both of which a vector plot such as that suggested by Kennedy and Pancu should be used) but for many systems it provides a rough guide to the resonance properties. One of the principal difficulties attending the analysis of the amplitude plots is to find which part of a given peak is due to the resonant mode and which to the off-resonant vibration. In the note a method is described for analysing a peak assuming that:—

• (i) only a single resonant mode is involved

• (ii) the contribution to the amplitude from the off-resonant vibration is constant in the neighbourhood of the peak.

A theory of the incompressible flow through two- and three-dimensional cascade actuator discs has been developed by several workers over the past ten years, and its accuracy has been confirmed in several experiments. This theory is briefly reviewed, and a parallel theory for subsonic compressible flow through actuator discs is developed. Approximate solutions for several examples are considered, including a compressible shear flow through a two-dimensional cascade, and a compressible flow through an annular cascade of guide vanes.

Although a numerical solution of the turbulent boundary-layer equations has been achieved by Mellor and Gibson for equilibrium layers, there are many occasions on which it is desirable to have closed-form expressions representing the velocity profile. Probably the best known and most widely used representation of both equilibrium and non-equilibrium layers is that of Coles. However, when velocity profiles are examined in detail it becomes apparent that considerable care is necessary in applying Coles's formulation, and it seems to be worthwhile to draw attention to some of the errors and inconsistencies which may arise if care is not exercised. This will be done mainly by the consideration of experimental data. In the work on constant pressure layers, emphasis tends to fall heavily on the author's own data previously reported in ref. 1, because the details of the measurements are readily available; other experimental work is introduced where the required values can be obtained easily from the published papers.

This paper presents the results of an investigation of the flow around rectangular and chamfered cavities at high subsonic and low supersonic speeds. Pressures measured on the faces of the cavities are integrated to find the pressure drag of the cavities. The types of cavity tested range from simple sawcuts to cavities so long that the two ends can be regarded as independent and the results for these are compared with the sum of the drags of isolated forward and rear facing steps.

Although the Reynolds numbers of the tests are similar to those in flight conditions the maximum depth of the cavities tested is only 6 mm so that the pressure resolution on the vertical faces of the cavities is limited. In spite of this it is estimated that the maximum error in the drag of any particular cavity is less than the skin friction drag on a smooth surface equal in area to half the plan area of the cavity.

The research paper addresses the problem of estimating aerodynamic parameters using a Gauss-Newton-based optimisation method. The process of the optimisation method lies on the principle of minimising the residual error between the measured and simulated responses of the system. Usually, the simulated response is obtained by integrating the dynamic equations of the system, which is found to be susceptible to the initial values, and the integration method. With the advent of the feedforward neural network, the data-driven regression methods have been widely used for identification of the system. Among them, a variant of feedforward neural network, extreme learning machine, which has proven the performance in terms of computational cost, generalisation, and so forth, has been addressed to predict the responses in the present study. The real flight data of longitudinal and lateral-directional motion have been considered to estimate their respective aerodynamic parameters. Furthermore, the estimates have been validated with the values of the classical estimation methods, such as the equation-error and filter-error methods. The sample standard deviations of the estimates demonstrate the effectiveness of the proposed method. Lastly, the proof-of-match exercise has been conducted with the other set of flight data to validate the estimated parameters.

This paper concerns aircraft system identification and, in particular, the process of aerodynamic model structure determination. Its application to experimental data from unmanned aerial vehicles (UAVs) is also described. The procedure can be particularly useful for determining an aerodynamic model for aircraft with unconventional airframe configurations, which some unmanned aircraft tend to have. Two model structure determination techniques are outlined. The first is the well-established stepwise regression method, while the second is an adaptation of an existing frequency response approach which instead utilises maximum likelihood estimation. Example applications of the methods are presented for two data sources. The first is a set of UAV flight test data and the second is data recorded from dynamic wind tunnel tests on a UAV configuration. For both examples, the model structures determined using stepwise regression and maximum likelihood analysis matched one another, suggesting that the maximum likelihood approach and the chosen thresholds for its statistical metrics were reliable for the data being analysed.

The paper uses the conventional nine-noded Lagrangian element for studying transient linear response analysis of composite stiffened shells. An improved version of the stiffener modelling has been used in which the stiffener can be placed anywhere inside the element. For the first time, concentric or eccentric stiffeners have been used for composite stiffened shells for solving the transient dynamic response of these structures. These are not available in existing commercial packages. Different types of time dependent loading such as short duration air-blast loading, suddenly applied uniformly distributed step loading and sinusoidally harmonic loading have been considered in this paper. The results of stiffened composite cylindrical shells and doubly curved shells with different boundary conditions and various laminate orientations have been presented for eccentric stiffeners. Parametric studies considering different variables have also been carried out.

Recent developments in the field of Mini-UAVs lead to successful designs in both hovering rotorcraft and fixed wing aircraft. However, a polyvalent MAV capable of stable hovering and fast forward flight is still expected. A promising candidate for such versatile missions consists of a tilt-body tail-sitter configuration. That concept is studied in this paper both from the flight mechanics and control points of view. Developments are based on an existing prototype called Vertigo. It consists of a tail sitter fixed-wing mini-UAV equipped with a contra-rotating pair of propellers in tractor configuration.

A wind-tunnel campaign was carried out to extract experimental results from the Vertigo aerodynamic characteristics. A 6-component sting balance was fitted in the powered model enabling excursion in angles of attack and sideslip angles up to 90°. Thus, a detailed understanding of the transition mechanism could be obtained. An analytical model including propwash effects was derived from experimental results.

The analytical model was used to compute stability modes for specific flight conditions. This allowed an appropriate design of the autopilot capable of stabilisation and control over the whole flight envelope. A gain sequencing technique was chosen to ensure stability while minimising control loop execution time. A MATLAB-based flight simulator including an analytical model for the propeller slipstream has been developed in order to test the validity of airborne control loops.

Simulation results are presented in the paper including hover flight, forward flight and transitions. Flight tests lead to successful inbound and outbound transitions of the Vertigo.

In many engineering applications (e.g. helicopters, turbines, compressors), lifting surfaces experience unsteady motion or are perturbed by unsteady incoming flows. Horizontal axis wind turbine rotors experience large time dependent variations in angle-of-attack as a result of control input angles, blade flapping, structural response and wake inflow. In addition, the blade sections encounter substantial periodic variations in local velocity and sweep angle. Thus, the unsteady aerodynamic behaviour of the blade sections must be properly understood to enable accurate predictions of the air loads and aero elastic response of the rotor system.

A growing interest in constellations of small satellites has recently emerged due to the increasing capability of these platforms and their reduced time and cost of development. However, in the absence of dedicated launch services for these systems, alternative methods for the deployment of these constellations must be considered which can take advantage of the availability of secondary-payload launch opportunities. Furthermore, a means of exploring the effects and tradeoffs in corresponding system architectures is required. This paper presents a methodology to integrate the deployment of constellations of small satellites into the wider design process for these systems. Using a method of design-space exploration, enhanced understanding of the tradespace is supported , whilst identification of system designs for development is enabled by the application of an optimisation process. To demonstrate the method, a simplified analysis framework and a multiobjective genetic algorithm are implemented for three mission case-studies with differing application. The first two cases, modelled on existing constellations, indicate the benefits of design-space exploration, and possible savings which could be made in cost, system mass, or deployment time. The third case, based on a proposed Earth observation nanosatellite constellation, focuses on deployment following launch using a secondary-payload opportunity and demonstrates the breadth of feasible solutions which may not be considered if only point-designs are generated by a priori analysis. These results indicate that the presented method can support the development of future constellations of small satellites by improving the knowledge of different deployment strategies available during the early design phases and through enhanced exploration and identification of promising design alternatives.

This paper presents a high-order sliding mode approach to design variable structure controllers for nonlinear multivariable cross-coupling systems whereby a change in either control input affects the control loops of all the subsystems dynamics. A hybrid-sliding mode control (hybrid-SMC) scheme is constructed combining a new equivalent control algorithm with a high-order discontinuous control algorithm to benefit from the advantages of both control strategies. The hybrid-SMC scheme uses weighting coefficients to weight and combine controllers. The equivalent control algorithm uses a relative degree concept through dynamic constraints imposed on the sliding variables to overcome the limitations of the conventional approaches and to provide an optimum tracking performance. A scale-laboratory helicopter model is used to sum up the main features and demonstrate the effectiveness of the developed control scheme. The proposed hybrid-SMC strategy is compared to existing sliding mode-based control approaches in terms of tracking performance, stability and control efforts. The obtained results demonstrate the validity and efficiency of the proposed hybrid-SMC scheme.

When accepting the invitation to present the Seventh Louis Blèriot Memorial lecture, the author was deeply conscious not only of the honour which was bestowed upon him by this invitation but, in an equal measure, of the responsibility which attached to it.

For the first time one of the eminent memorial lectures is devoted to the rotating wing; the helicopter has been raised to the level of discussion of the great aeronautical papers.

The natural impulse of the gentler sex when faced with a great occasion is invariably: “how shall one dress?” To the author, a somewhat similar problem became manifest: how can the helicopter, this new-comer, best be “ dressed?” How should it be presented, so as to make not only a good impression but also a full and, above all, a true impression?

This paper reviews work on the flexural centre of elastic cantilever beams and contains a number of hitherto unpublished results, including a formula giving the position of the flexural centre in terms of the Prandtl torsional stress function.

The appearance of the note by Jacobs has prompted the preparation of a review of this subject which is all the more desirable as several of the investigations made shortly before the 1939-45 War were never published and others, although published, seem to be in danger of being overlooked. The present paper contains a number of hitherto unpublished results. The question of nomenclature is worthy of mention since general agreement is lacking. Some of the names used are flexural centre, centre of flexure, elastic centrum and centre of shear but a complete search of the literature would probably lead to the discovery of yet other names. The terms “centre of shear” and “elastic centrum” seem vague and fail to indicate that the point referred to has any special relation to flexure. We shall here use the name “flexural centre” exclusively.

There is considerable lack of information regarding shear stresses in tapered beams. In this brief report these stresses are examined by a simple extension of the conventional engineering method for uniform beams, and where possible the results are compared with those of classical theory.

Complete agreement is obtained with the few existing classical solutions. The use of the engineering method is illustrated by its application to various types of beam, and curves are drawn illustrating the general nature of the influence of taper on shear stresses.

A wing of finite span in flight generates trailing vorticity which “rolls up” into two “discrete” vortices. During the rolling up process the trailing vorticity is displaced vertically downward (assuming that the aircraft is in horizontal flight) under the influence of the wing circulation and the self-induced downwash velocities. In this note a method is presented for the prediction of the spatial distribution of the trailing vorticity during the rolling up process, thus giving the overall downwash field behind the wing.

A review is made of recent research in the United States relating to runway slipperiness and to slush-drag or fluid-drag effects on aircraft ground performance. The formation of fluid spray and its drag effects on aircraft are traced from spray origin at the intersection of the tyre and runway to impingement on the aircraft over the complete aircraft ground-velocity range. Six manifestations of tyre hydroplaning are presented and discussed: fluid-drag peak, tyre spin-down, loss of aircraft directional stability, loss of aircraft braking traction, suppression of bow wave, and progressive detachment of the tyre footprint from the fluid-covered runway surface as ground speed is increased. A convenient formula for estimating tyre-hydroplaning speed is indicated to be

where Vp=hydroplaning speed, knots

p=tyre-inflation pressure, lb/sq in.

Comparison with available experimental data is shown. The effects of fluid viscosity and density on runway slipperiness are discussed. The effects of some runway-surface textures, tyre-tread patterns, and aircraft landing-gear wheel arrangements on slipperiness are described. The paper also includes indications of areas of interest that need further study.

Before entering upon the subject of my lecture I beg to express my feelings of gratitude for the honour and appreciation of my efforts, implied in the invitation to speak on metal aeroplane construction before so eminent and learned a body as the Royal Aeronautical Society.

In consideration of the general political situation I deem myself entitled to attribute to this invitation the deeper meaning of a token of amiable disposition given apart and beyond my person, from nation to nation, and to see in it an effort of renewing the ties of a genuine humanity, which is well aware of the unavoidableness of conflict and fight, but does not know hatred, and desires to extinguish the sad traces of a devastating war by hoisting the flag of peaceful competition.

The method of Beecham and Titchener is extended to systems with n degrees of freedom and is shown to be a combination of the averaging principle and the method of variation of parameters. In this extended form, the method provides a powerful solution algorithm for non-linear problems such as those which arise in aircraft structural dynamics and aeroelasticity. The method is exemplified in application to a two-degree-of-freedom damped non-linear oscillator and to a binary (flexure-aileron) non-linear flutter system. The method is finally extended to non-linear differential equations in first-order form such as those which arise commonly in flight mechanics.