An aeroelastic formulation of 2D lifting surfaces in an incompressible flow field via the use of aerodynamic indicial functions is presented. The approach carried out in time and frequency domains yields the proper aerodynamic loads necessary to the study of the subcritical aeroelastic response and flutter of lifting surfaces, respectively. The expressions of the unsteady lift and aerodynamic twist moment in the frequency domain are given in terms of the Theodorsen’s function, while in the time domain, these are obtained directly with the help of the Wagner’s function. Closed form expressions of unsteady aerodynamic derivatives, helpful in an unified approach of response and flutter analyses, and results related to aeroelastic response to gust and blast loads are supplied. In this context, a novel representation of the aeroelastic response in the phase - space was supplied and pertinent conclusions on the implications of some basic parameters have been outlined.

The displacement effects on a set of five rectangular mouthed pitot tubes of a constant thickness ratio, 0-6, were investigated in fully developed turbulent flow in a plain tube. The results show that the displacement due to shear only, δ1, is given by δ1/b=0.19, where b is the outside thickness of the pitot tube. The additional wall effect is found to have a dependence on Reynolds number or, equivalently, wall shear within the range investigated of 27 500 < Rev < 58 300. The wall effect results may be presented as a single curve, however, if consideration is given to the distance up to which the wall effect is present.

The author mentioned in his Main Lecture(1) the success achieved in the analysis of three-dimensional media, for small and large displacements, as well as anisotropic and non-elastic behaviour, by the introduction of tetrahedron elements of constant strain and stress(2), see also technical note 1 of this series(3). A cardinal point of the theory is the specification of natural strains, stresses and stiffness. At the same time attention was drawn to certain difficulties arising in the interpretation of the stresses at the nodal or other points, which are more severe than for constant strain triangles, the corresponding elements in the two-dimensional case. It was suggested in the lecture that a considerable improvement might be achieved by the specification of a linearly varying strain or stress state within the tetrahedron. The solution of this problem, limited to small displacements, is summarised in this fifth technical note and its application is to be demonstrated on an example in the printed lecture.

—Since flow separation occurs readily from a highly swept leading edge, but gives rise in general to a steady flow, it is proposed that the rational approach to the aerodynamic design of slender wings is to attempt to control, rather than to suppress, these separations. This leads to the suggestion that the leading edges should be sharp, and that the wing should be shaped so as to make them attachment lines at one attitude (the design attitude) at which classical wing theory can be applied. It is argued, further, that if the leading edge separations are to develop regularly with change of attitude of the wing, separation must occur only from the trailing edge at the design attitude; and the velocity field favourable to boundary layer development without separation forward of the trailing edge is discussed. Subject to the restrictions thus imposed on the design, low drag is sought at the design attitude. This leads to the consideration of a particular class of doubly curved mean surfaces satisfying the leading edge condition, onto which thickness distributions are superposed so as to provide favourable velocity fields together with low drag. A number of examples are considered, using slender thin wing theory for flexibility, to illustrate the manner in which plan form and thickness distributions affect the pressure distribution, and to indicate the relatively high lift/drag ratios which seem feasible. Some consideration is given to the limitations of the theory used and to the further developments which seem desirable.

This paper is intended as a general introduction to the requirements for future passenger aircraft design. The needs of the 21st century are addressed to meet the important requirements of the customer airlines as well as those of the general public. In particular, the impact on two traditional major requirements are reviewed, the Design Mission and the operating costs. The effect of aircraft on the environment and the increases in the cost of fuel will have a substantial effect on the way future aircraft are optimised. These demands are summarised before moving on to the basic equations affecting how the aircraft design must respond. Very similar targets driving research work have been set in both Europe and the United States, and some of the new technologies that we can expect to be incorporated are outlined. Finally, a glimpse is given of the possible future aircraft configurations we may see in the skies in response to the new demands.

The effect of depth taper on the flexural vibration characteristics of a beam of uniform width is investigated in this paper. The frequency parameters and mode shapes for the first five modes of vibration of tapered beams are presented for a wide range of depth taper. The Euler-Bernoulli equation of a beam is reduced to an eigenvalue problem and its eigenvalues and eigenvectors are obtained by using a digital computer. The theoretical results are compared with those of other authors and with the experimental results.

A neuro-fuzzy controller method for smart material actuator (SMA) hysteresis modelling is presented, conceived for a morphing wing application. The controller correlates each set of forces and electrical currents that are applied to the smart material actuators with the actuator elongation. The actuator is experimentally tested for four forces, using a variable electrical current. The final controller is obtained through the Matlab/Simulink integration of three independent neuro-fuzzy controllers, designed for the increase and decrease of electrical current, and for null electrical current in the cooling phase of the actuator. This final controller gives a very small error with respect to the experimental values.

This paper adds data to help establish fidelity criteria to accompany the simulator motion system diagnostic test specified by the International Civil Aviation Organisation. Twelve airline transport pilots flew three tasks in the NASA Vertical Motion Simulator under four different motion conditions. The experiment used three different hexapod motion configurations, each with a different trade-off between motion filter gain and break frequency, and one large motion configuration that utilised as much of the simulator's motion space as possible. The motion condition significantly affected (1) pilot motion fidelity ratings, and sink rate and lateral deviation at touchdown for the approach and landing task, (2) pilot motion fidelity ratings, roll deviations, maximum pitch rate, and number of stick shaker activations in the stall task, and (3) heading deviation after an engine failure in the take-off task. Significant differences in pilot-vehicle performance were used to define initial objective motion cueing criteria boundaries. These initial fidelity boundaries show promise but need refinement.

The performance of a helicopter rotor has always been dogged by the dissymmetry of air velocity over the blades caused by its translation in a direction substantially parallel to the plane of the rotor disc. Furthermore, in order to make the rotor even usable, this dissymmetry of lift has caused the introduction of major mechanical complications to the rotor hub. In forward flight the rotor is aerodynamically split into two halves either side of the flight direction. The advancing side, where the rotor rotation is in the same sense as the forward velocity, will see a greater dynamic head than the retreating side, where the rotor rotation and forward speed are in opposition. A rigid rotor would consequently suffer a major rolling moment and if a viable proposition is to be achieved the rolling moment must be avoided.

Flight simulators are integral to the design/development, testing/qualification, training and research communities and their utilisation is ever expanding. The use of flight simulation to provide a safe environment for pilot training, and in research and development, must be underpinned by quantification of simulator fidelity. While regulatory simulator standards exist for flight training simulators and new standards are in development, previous research has shown that current standards do not provide a fully quantitative approach for assessing simulation fidelity, especially in a research environment. This paper reports on progress made in a research project at the University of Liverpool (Lifting Standards), in which new predicted and perceptual measures of simulator fidelity have been developed. The new metrics have been derived from handling qualities engineering practice. Results from flight tests on the National Research Council (Canada) Bell 412 ASRA research aircraft and piloted simulation trials using the HELIFLIGHT-R simulator at Liverpool are presented to show the efficacy of adopting a handling qualities approach for fidelity assessment. Analysis of the new metrics has shown an appropriate degree of sensitivity to differences between flight and simulation.

It is well known that the performance at high subsonic and transonic speeds of a swept-back wing-body combination in which the wing is untwisted and has the same section at all stations along the span and in which the body is not specially shaped to allow for the presence of the wing, falls far short of what would be predicted for the corresponding infinite sheared wing. For example, with a sweep of 45° and a thickness/chord ratio of 6 per cent it has been found experimentally that a rapid shock-induced increase in drag occurs above a Mach number of about 0·95 and a peak value of CD is obtained at Mach numbers slightly in excess of 1·0, whereas it can be estimated that for the corresponding infinite sheared wing, sonic speed in a direction perpendicular to the isobars (the lines joining points where the pressure is equal) would not be obtained until a Mach number of 1·18 was reached. The poorer performance of the finite swept-back wing results principally from the fact that the pressure distributions for sections near the root and tip are distorted in shape from what would be obtained on an infinite sheared wing and, as a result, the isobars tend to lose some or all of their sweep. With a moderate aspect ratio such as 3, such effects extend over most of the span at high subsonic speeds.

The prediction and characterisation of the limit cycle oscillation (LCO) behaviour of non-linear aeroelastic systems has become of great interest recently. However, much of this work has concentrated on determining the existence of LCOs. This paper concentrates on LCO stability. By considering the energy present in different limit cycles, and also using the harmonic balance method, it is shown how the stability of limit cycles can be determined. The analysis is then extended to show that limit cycles can be controlled, or even suppressed, by the use of suitable excitation signals. A basic control scheme is developed to achieve this, and is demonstrated on a simple simulated non-linear aeroelastic system.

Although a great deal has been said and written in the past eight years or so about engine noise, previous discussion has been very largely concerned with the physical aspects of jet noise and with the possibility of suppressing it by means of special propelling nozzles and ground mufflers. There is however another, and in general a better, way of dealing with jet noise, namely by using engines of low specific thrust, i.e. of low exhaust velocity and one purpose of the paper is to examine the possibilities of such engines. This leads to a consideration of blade noise from fans and turbines for as jet noise is reduced this tends to become the dominant factor. But first of all it is necessary to re-examine the nature of the noise problem.

This contribution presents a novel simulation for a fixed-wing aircraft powered by gas turbine engines and advanced propellers (turboprops). The work is part of a large framework for the simulation of aircraft flight through a multi-disciplinary approach. Novel numerical methods are presented for flight mechanics, turboprop engine simulation (in direct and inverse mode), and propeller dynamics. We present in detail the integration of the propeller with the airframe, aircraft and tonal noise model. At the basic level, we address a shortfall in multi-disciplinary integration in turboprop-powered aircraft, including economical operations and environmental emissions (exhausts and noise). The models introduced are based on first principles, supplied with semi-empirical correlations, if required. Validation strategies are presented for component-level analysis and system integration. Results are presented for aerodynamics, specific air range, optimal cruise conditions, payload-range performance, and propeller noise. Selected results are shown for the ATR 72-500, powered by PW127M turboprop engines and F568-1 propellers.

I am very appreciative of the honour that has been accorded me in being invited to deliver the Second Lanchester Memorial Lecture: appreciative, but all too conscious of my own inadequacy. This is especially true, since in the first Lecture we heard a tribute to a very great man from another such: and the prospect of following von Karman to this lectern is one to deter perhaps all but the most foolhardy. Since von Karman’s address discussed very fully the contributions of Lanchester to aeronautics, little is left for subsequent lecturers on this subject, and I have therefore chosen to speak of a problem which has been my special interest and concern throughout my professional life, namely, aeroelasticity—a field in which, as in so many others, Lanchester was a pioneer.

Lean Blow-Out (LBO) limits are critically important in the operation of aero engines. Previously, Lefebvre's LBO empirical correlation has been extended to the flame volume concept by the authors. Flame volume takes into account the effects of geometric configuration, spatial interaction of mixing jets, turbulence, heat transfer and combustion processes inside the gas turbine combustion chamber. For these reasons, LBO predictions based on flame volume are more accurate. Although LBO prediction accuracy has improved, it poses a challenge associated with Vf estimation in real gas turbine combustors. This work extends the approach of flame volume prediction based on fuel iterative approximation with cold flow simulations to reactive flow simulations. Flame volume for 11 combustor configurations were simulated and validated against experimental data. To make prediction methodology robust, as required in preliminary design stage, reactive flow simulations were carried out with the combination of presumed Probability Density Function (PDF) and discrete phase model (DPM) in Fluent 15.0 The criterion for flame identification was defined. Two important parameters—critical injection diameter (Dp,crit) and critical temperature (Tcrit)—were identified and their influence on reactive flow simulation was studied for Vf estimation. Results exhibit ±15% error in Vf estimation with experimental data.

The objective of this work is to make an overview of opportunities and issues related to the aeronautical application of solid oxide fuel cell hybrid systems. The great interest on fuel cells comes from their capacity of producing electric energy with high efficiency at low pollutant production. The application of these systems as full-time auxiliary power units is an interesting alternative in a future scenario, which is supposed to include a More Electric Aircraft and more restrictive environmental standards. A review of the technological aspects of this application is presented. The physical models found in literature were investigated and the results were compared and discussed.

A knowledge based expert system for the design and analysis of composite laminated struts is described. This system is part of a Composite Design EXpert (CODEX) system that not only performs analysis and design of composite laminated plates and struts, but also incorporates the modelling of uncertainties of information, definition and decision when assessing competing designs. The analysis part of the system is based on a minimum weight optimisation that satisfies the failure criteria of local buckling, overall buckling and maximum strain or maximum stress. Available in the system for analysis and assessment are seven types of strut shapes, six composite materials and aluminium alloy. Examples of strut design are carried out and assessed using this system.

Wing tip sails are small lifting surfaces fitted to the wings in such a way that they use the local air flows about the wing tips, induced by the generation of lift on the wing, to produce a thrust. They act in a manner similar to that of the sail of a yacht when it is sailing close to the wind, hence their name.

Previous work, described in references 1, 2 and 3, has investigated the effects of adding sails to existing wings, extending their span and wetted area. It has long been realised that the correct use of sails is to replace part of the outer wing area with them, in other words to include them in the initial design of the wing. This note describes wind tunnel tests on a simple untapered, unswept wing to examine the benefits from such a design approach and subsequent flight tests on a Cessna Centurion to confirm them.

In recent years a number of authors have analysed the bending of clamped skew plates under uniform pressure. Many of these solutions are very approximate. Favre and Dorman adopted the Ritz method and their expressions for displacement are restrictive. Based on the generalised method postulated by Lardy, Mirsky has presented a solution which involves considerable algebraic and arithmetic work. Komatsu's paper contains a solution using conformal mapping. Quinlan has given a powerful method to solve skew plate problems involving different boundary conditions and various types of loadings, but no numerical work has been done.