—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.

Results obtained from previous studies of the three-shock confluence are applied to a simple wedge intake with the objective of obtaining a qualitative description of the flow for given initial conditions (M0, δ0). A simple mapping procedure is used, so that when M0, δ0 varies slowly and continuously the resulting effect on the shock geometry can be discussed in a systematic fashion. A number of special curves are computed to help classify various regions of initial conditions.

Stratford's analysis of the laminar boundary layer near separation uses two physical ideas. In the outer part of the boundary layer, where viscous effects are small, the development is given by the condition that the total head is constant along streamlines, apart from a second-order correction for viscosity. Near the wall, however, viscous forces must balance the pressure forces, and the profile adjusts itself accordingly. Quantitatively these ideas yield a simple formula for predicting separation, which has been found to be particularly accurate.

In this paper it is indicated how the same approach may be used to yield the full distribution of skin friction along the wall. Further, the effects of suction may be incorporated into the method. Physically, suction affects the outer part of the boundary layer in that the streamlines are drawn towards the wall when suction is applied. At the wall, the balance between viscous and pressure forces is influenced by the momentum of the fluid which is sucked away. When these effects are accounted for quantitatively, the resulting formula for the skin friction is still very simple.

Several examples are considered, and comparison is made with exact theory and with approximate results by other methods. It is indicated that the method has a useful range of validity.

A general-purpose low turbulence wind tunnel was constructed using the design method of Bradshaw. Sound absorbent material was used in every four corners to decrease sound intensity produced by a fan. The longitudinal component of turbulence intensity at the centre of the closed working section is less than 0.02% of the mean velocity in the speed range between 18 m/s arid 53 m/s. The mean velocity variations across the working section are within ±0.1% of the mean velocity.

Performance measurements have been done at representative tunnel cross sections to clarify the behaviour of flow in the tunnel. This work differs from previous studies in the sense that emphasis is placed not only on velocity distributions, but also on turbulence intensity distributions at several cross sections of the tunnel.

The critical Reynolds number for a flat plate at zero incidence, Rec = 3.5 x 106, measured in a stream of the very low turbulence intensity of 0.016%, is larger than that reported by Schubauer and Skramstad.

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.

This study examines the performance of two axisymmetric nozzles which were designed to produce uniform, parallel flow with nominal Mach numbers of 4 and 8. A free-piston-driven shock tube was used to supply the nozzle with high-temperature, high-pressure test gas. The inviscid design procedure treated the nozzle expansion in two stages. Close to the nozzle throat, the nozzle wall was specified as conical and the gas flow was treated as a quasi-one-dimensional chemically-reacting flow. At the end of the conical expansion, the gas was assumed to be calorically perfect and a contoured wall was designed (using Method-of-Characteristics) to convert the source flow into a uniform and parallel flow at the end of the nozzle. Performance was assessed by measuring Pitot pressures across the exit plane of the nozzles and, over the range of operating conditions examined, the nozzles produced satisfactory test flows. However, there were flow disturbances in the Mach 8 nozzle flow that persisted for significant times after flow initiation.

This article presents the research status and development trends of Vertical Take-off and Landing hybrid Unmanned Aerial Vehicles. In this research, a special emphasis is laid on the design philosophies, analysis techniques, dynamic modeling, and control laws of hybrid VTOL UAVs. It studies and compares various design configurations of hybrid VTOL UAVs, based on key design features such as aerodynamic performance, flight stability, structural strength, propulsive power, avionics systems, flight controls, autonomy, ease of fabrication and flight transition mechanisms. The benefits and shortcomings of each design configuration are expressed in detail. A selection problem is formulated in a fuzzy environment and the Multi-Attribute Decision-Making technique is employed. Ongoing research projects in the field are discussed and a novel design of tail sitter hybrid VTOL UAV is presented by the authors. This work serves as a useful guide for the prospective explorers of this challenging field of research.

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.

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 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.

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.

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.

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.

International air travel becomes more affordable over the years with the introduction of low-cost carriers (LCCs) in Asia Pacific. As a result, this region has seen an exponential growth in its air traffic movements. Asia Pacific has been recording faster air traffic growths, compared to regions such as Europe and North America. This is largely due to emerging markets and developing economies in Asia Pacific. This is especially seen in the demand and supply for LCCs in Southeast Asia. Changi Airport, being one of the major air traffic hubs in Asia, is facing overloaded traffic in the future due to this exponential growth. Although a new runway and new terminals are proposed and in the process of being built, the coming years could still bring a huge challenge to the airport, if its existing capacity is not maximised to cope with the increasing traffic demand. Hence, we studied the impacts of increasing LCC flights on the airport runway capacity. Different runway operational scenarios were simulated to investigate their advantages and disadvantages. It is found that the segregated parallel arrival-departure runway operation is beneficial. By simulating various operational scenarios with the projected increasing flight movements, it is found that a dedicated runway to medium-sized arrival aircraft may be more beneficial, as this would possibly enable it to have a higher runway capacity.

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.

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.

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.

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.

Product development, especially in aerospace, has become more and more interconnected with its operational environment. In a constant changing world, the operational environment will be subjected to changes during the life cycle of the product. The operational environment will be affected by not only technical and non-technical perturbations, but also economical, managerial and regulatory decisions, thus requiring a more global product development approach. One way to try tackling such complex and intertwined problem advocates studying the envisioned product or system in the context of system of systems (SoS) engineering. SoSs are all around us, probably in any field of engineering, ranging from integrated transport systems, public infrastructure systems to modern homes equipped with sensors and smart appliances; from cities filling with autonomous vehicle to defence systems.

Since also aerospace systems are certainly affected, this work will present a holistic approach to aerospace product development that tries spanning from needs to technology assessment. The proposed approach will be presented and analysed and key enablers and future research directions will be highlighted from an interdisciplinary point of view. Consideration of the surrounding world will require to look beyond classical engineering disciplines.