Advances in space activity are linked to reductions in launch cost. Air-breathing propulsion-assisted flight systems offer the potential for revolutionary change of the space operations paradigm. Horizontal launch of a space-access system provides mission flexibility, responsiveness, and affordability. One way to reduce launch cost is to increase the Mach number at which a launch vehicle is staged from a carrier aircraft. Without exceeding the engine and airframe design limits, the pre-compressor cooling technology allows an operational aircraft to operate at Mach numbers and altitudes beyond its basic operational limits. This is an essential, near-term technology for reducing launch cost to place small-weight payloads in low Earth orbit. The advantage of this technology is assessed with a modified McDonnell Douglas QF-4C aircraft. Payloads are unachievable or marginal with an unmodified QF-4C. However, payloads weighing around 150 pounds are plausible with this aircraft when incorporating the water injection pre-compressor cooling (WIPCC) technology.

This paper investigates the problem of coning motion stability of spinning missiles equipped with strapdown seekers. During model derivation, it is found that the scaling factor error between the strapdown seeker and the onboard gyro introduces an undesired parasitic loop in the guidance system and, therefore, results in stability issues. Through stability analysis, a sufficient and necessary condition for the stability of spinning missiles with strapdown seekers is proposed analytically. Theoretical and numerical results reveal that the scaling factor error, spinning rate and navigation ratio play important roles in stable regions of the guidance system. Consequently, autopilot gains must be checked carefully to satisfy the stability conditions.

The application of optical imaging techniques to hypersonic facilities is discussed and examples of experimental measurements are provided. Traditional Schlieren and shadowgraph techniques still remain as inexpensive and easy to use flow visualisation techniques. With the advent of faster cameras, these methods are becoming increasingly important for time-resolved high-speed imaging. Interferometry’s quantitative nature is regularly used to obtain density information about hypersonic flows. Recent developments have seen an extension of the types of flows that can be imaged and the measurement of other flow parameters such as ionisation level. Planar laser induced fluorescence has been used to visualise complex flows and to measure such quantities as temperature and velocity. Future directions for optical imaging are discussed.

The SMARTFIRE computational fluid dynamics (CFD) software was used to predict the ‘possible’ behaviour of airflow as well as the spread of fire and smoke within a Swissair configured McDonnell Douglas MD-11 commercial transport aircraft. This work was undertaken by the Fire Safety Engineering Group (FSEG) of the University of Greenwich as part of Transportation Safety Board (TSB) of Canada, Fire & Explosion Group’s investigation into the in-flight fire occurrence onboard Swissair Flight 111 (SR111): TSB Report Number A98H0003. The main aims of the CFD analysis were to develop a better understanding of the possible effects, or lack thereof, of numerous variables relating to the in-flight fire. This assisted investigators in assessing possible fire dynamics for cause and origin determination. In Part 1, the numerical analyses to pre-fire airflow patterns within the cockpit and its vicinity are presented. The pre-fire simulations serve two ends. One is to provide insight into the flow patterns within the cockpit and its vicinity and further supportive numerical evidence for the airflow flight test observations. The other is to provide plausible initial flow conditions for fire simulations. In this paper, some flow patterns at a number of primary locations within the cockpit and its vicinity are highlighted and the predicted flow patterns are compared with the findings from the airflow flight tests. The predicted patterns are found to be in good qualitative agreement with the experimental test findings.

In this study, prediction of aircraft Estimated Time of Arrival (ETA) is proposed using machine learning algorithms. Accurate prediction of ETA is important for management of delay and air traffic flow, runway assignment, gate assignment, collaborative decision making (CDM), coordination of ground personnel and equipment, and optimisation of arrival sequence etc. Machine learning is able to learn from experience and make predictions with weak assumptions or no assumptions at all. In the proposed approach, general flight information, trajectory data and weather data were obtained from different sources in various formats. Raw data were converted to tidy data and inserted into a relational database. To obtain the features for training the machine learning models, the data were explored, cleaned and transformed into convenient features. New features were also derived from the available data. Random forests and deep neural networks were used to train the machine learning models. Both models can predict the ETA with a mean absolute error (MAE) less than 6min after departure, and less than 3min after terminal manoeuvring area (TMA) entrance. Additionally, a web application was developed to dynamically predict the ETA using proposed models.

Many aspects of parachute behaviour are dominated by the permeability of the canopy cloth. Techniques employing kinematic principles to predict parachute opening behaviour, for example, must first be able to predict the air flow through the canopy material. This paper is concerned with the development of such a predictive capability.

The flow of air through cloth of gauze has been studied by many workers; since the 1850’s, in connection with filters; and since the early twenties, in connection with parachutes. Workers in each of these disciplines seemed unaware of the work being done in the other.

In the case of parachute cloth permeability we find the foundations being laid down by such eminent workers as Glauert in 1932 and Taylor in 1944. Yet most subsequent writers seem not to be aware of this, and since World War II, there have been many papers “rediscovering” basic principles; often with errors which could have been corrected by reference to the earlier authorities.

In this paper we attempt to bring together all the work which has been done in this field, and to rationalize the data by simple mathematical modelling. Individual elements of the model have been proposed before, of course.

We then examine the effect of tension on the geometric porosity of fabric, and hence the pressure loss Ap. A theoretical analysis shows that tension is likely to be a major factor. For cloths which have low porosity when unloaded our simple mathematical model shows that, for a given Ap, a stress equal to 50% of ultimate can increase the air volume flow Q by an order of magnitude. In general, the increase in permeability is greatest at the lower values of Ap, indicating that the viscous term is more influenced by tension than the dynamic term. This theoretical result is confirmed by some rather limited experiments carried out by Lashbrook and Marby. But their data also shows that a cloth made from relatively stiff glass fibers can experience a reduction in permeability when loaded, due to a “closing up” of the weave. The reasons for this are discussed.

The 48th Wilbur Wright Memorial Lecture, “Mathematics and Aeronautics” was given by Mr. M. J. Lighthill, F.R.S., before a large and distinguished audience at the Institution of Mechanical Engineers, 1 Birdcage Walk, on 19th May 1960. Dr. E. S. Moult, C.B.E., B.Sc, F.R.Ae.S., M.I.Mech.E., President of the Society, presided.

As has become the custom, before the lecture was delivered the President presented the awards made by the Council for 1960 for outstanding contributions to aeronautics. In the unavoidable absence of the Secretary he was assisted by Miss E. C. Pike, M.A., A.F.R.Ae.S., Secretary of the Medals and Awards Committee and Editor of The Aeronautical Quarterly. The list of the awards presented on this occasion was published in the June 1960 Journal—(p. XXXII).

The President, introducing the lecturer, reminded the audience that the Wilbur Wright Lecture commemorated the first manned, powered controlled flight by Wilbur Wright at Kitty Hawk, North Carolina on 17th December 1903. The first Wilbur Wright Memorial Lecture was held in 1913 and they had been held each year ever since, through two World Wars; in itself a wonderful record. The custom had been, and would continue to be, to have alternately a lecturer from the United Kingdom and a lecturer from the United States; last year the lecture had been by Mr. McCarthy of Chance Vought and next year again it would be by an American.

The Wilbur Wright Lectures had been a most distinguished series by distinguished lecturers, and he was sure that the lecture that night would be in keeping. His pleasant duty was to introduce Professor Lighthill—or as they had to call him Mr. Lighthill now that he was Director of the Royal Aircraft Establishment, which was an even greater distinction. Mr. Lighthill was the youngest Director that the R.A.E. had ever had—that alone was a tribute to Mr. Lighthill's ability and his high standing in the profession. Before joining the Royal Aircraft Establishment Mr. Lighthill was Professor of Applied Mathematics at Manchester University, and for a time, after leaving Cambridge, had worked at the National Physical Laboratory where he applied the mathematics of Cambridge to the problems of aerodynamics. His lecture was on “Mathematics and Aeronautics” he would ask Mr. Lighthill to deliver the 48th Wilbur Wright Memorial Lecture.

The aerodynamic effectiveness of a cylinder-flare body at zero incidence under laminar and turbulent boundary-layer conditions has been studied experimentally. Two nose geometries, namely a 10° half-angle sharp cone and a hemisphere, were used. The study has been carried out in a hypersonic gun tunnel at a Mach number of 8·2 and a Reynolds number of 158,100, based on the cylinder diameter, at flare angles 0°, 10°, 20°, 30° and 45°. The surface flow was studied using oil-dot and liquid crystal techniques. Some information regarding the shock layer was obtained from schlieren pictures. The effects of entropy layer and boundary-layer state on flare effectiveness were deduced from pressure measurements over the cylinder and the flare.

A survey is presented of results obtained from experiments in the RAE 5m low-speed pressurised wind tunnel. The tunnel is capable of operation over a range of pressures (from one to three atmospheres) so that the effects of Mach and Reynolds number may be separated. This decoupling of scale and compressibility effects has made possible reliable extrapolation of test results to full-scale conditions (where this is necessary) as well as giving greater insight into the underlying flow mechanisms. The large size of the tunnel, combined with pressurisation to three atmospheres, has enabled some tests, on small combat/trainer aircraft and on bomb-like stores, to be carried out at full scale.

Three facets of the work of the 5m tunnel are described — work aimed at predicting or improving the high-lift performance of specific aircraft, work of a more fundamental nature concerned particularly with the optimisation of high-lift devices on generalised research models, and work on the drag of stores carried externally on combat aircraft.

The effectiveness of any braking system is ultimately limited by the point at which the wheels lock and slide over the surface on which the stop is being made. When surfaces are clean and dry the adhesion with a rubber tyre is high enough to meet most requirements of braking without risk of skidding. In wet, or in icy conditions, however, adhesion may be reduced to such an extent that, even with efficient brakes, stopping distances are greatly increased and dangerous conditions can arise. Under wet conditions there are special difficulties. Different surfaces may have very different skid-resisting properties and, in addition, when surfaces are wet, the adhesion with a tyre depends on speed, generally diminishing as speed is raised.

Inverse dynamics is investigated as an efficient means of generating large angle satellite attitude slew manoeuvres. The required slew manoeuvres are specified using high order polynomials which result in a smooth control torque profile. Such smooth torque profiles are of importance in avoiding excitation of elastic modes of the satellite structure. In addition, the polynomial functions may be used to define a slew trajectory between arbitrary initial and final states allowing track-to-track manoeuvres. To compensate for uncertainties in the satellite dynamics, the inverse control is extended to include feedback linearisation about the nominal reference trajectory. It is shown that for rest-to-rest slew manoeuvres the resulting composite control is robust to uncertainties in the satellite inertia matrix and to actuator degradation. In addition, it is shown that complex paths may be generated through the use of high order polynomial functions which contain all the information required to enforce user defined constraints and boundary conditions in a compact, computationally efficient form.

A symmetric Mach 12·76 hypersonic flow in a 90° corner, formed by 30° swept back intersecting 8° wedges, was investigated in detail through numerical simulation using locally conical Navier-Stokes equations. Three different numerical schemes for spatial discretisation, MacCormack central differencing, van Leer’s flux vector splitting and Osher’s flux difference splitting, were studied to compare their capabilities to capture both strong shock waves and thin shear layers. Comparison with experimental data was made to validate the simulation. The numerical simulation provided further insight into the flowfield and a pair of counter-rotating vortices were discovered near the junction of the corner.

There is very little information on the stress concentration which is caused by a notch on one side of a plate in tension and the authors are not aware of any published data based on experiment. The theoretical solutions due to Neuber are for notches of hyperbolic or elliptic form and only give an approximate result for U-notches of varying depth.

The ability of microramps to control shock - boundary layer interaction at the vicinity of an axisymmetric compression corner was investigated computationally in a Mach 4 flow. A cylinder/flare model with a flare angle of 25° was chosen for this study. Height (h) of the microramp device was 22% of the undisturbed boundary layer thickness (δ) obtained at the compression corner location. A single array of these microramps with an inter-device spacing of 7.5h was placed at three different streamwise locations viz. 5δ, 10δ and 15δ (22.7h, 45.41h and 68.12h in terms of the device height) upstream of the corner and the variations in the flowfield characteristics were observed. These devices modified the separation bubble structure noticeably by producing alternate upwash and downwash regions in the boundary layer. Variations in the separation bubble’s length and height were observed along the spanwise (circumferential) direction due to these devices.

A three-dimensional integral formulation for the solution of the full potential equation and the associated numerical algorithm, the field-integral method, are presented. The method is a direct extension of a two-dimensional method and in particular retains the simple grid generation requirements noted in that method. Results are presented for the flow over body shapes and a complex winglet configuration, and are compared with existing transonic methods and experiments with good agreement. The further work necessary to provide a fast, robust method for use in design is outlined.

A theoretical and experimental investigation has been made into the effect of swirl on a supersonic jet. In certain circumstances a core of reversed flow can occur along the axis of symmetry and the jet mixes much more quickly with the ambient air than does the corresponding non-rotating jet. A characteristic calculation has been performed for the first cell in the case where the axial flow is not reversed. This indicates a shortening of the cell length and the formation of shocks nearer the nozzle.

This paper assesses the financial performance of 35 European airports for the decade 1990 to 2000, comparing those subject to partial or full privatisation with those still in public ownership. In contrast to earlier research, the outcomes of partial factor productivity (PFP), financial ratio (FRA) and data envelopment analysis (DEA) are evaluated, in order to investigate differences attributable to the degree of privatisation. Changes in performance after a change in ownership structure are reviewed. The analysis of sample data reveals economically meaningful and statistically significant differences between publicly owned and privatised airports. The major differences lie in operating efficiency, capital productivity and capital structure. Although partially and fully privatised airports operate more efficiently, this does not translate into significantly higher returns on shareholders’ funds. Due to their at least indirectly government-backed credit standing, publicly owned airport companies can assume more debt relative to their respective equity. This results in considerably higher gearing and financial leverage, which compensates for the comparatively low return rate on assets.

This paper presents the first results from research into active control of structural load alleviation (SLA) for tiltrotor aircraft carried out in the European ‘critical technology’ RHILP project. The importance of and the need for SLA in tiltrotors are discussed, drawing on previous US experience reported in the open literature. The paper addresses the modelling aspects in some detail; hence forming the foundation for both the FLIGHTLAB simulated XV-15 and EUROTILT configurations. The primary focus of attention is the suppression of in-plane rotor yoke loads for pitch manoeuvres in airplane mode; without suppression these loads would result in a very high level of fatigue damage. Multi-variable control law design methods are used to develop controller schemes and load suppression of 80-90% is demonstrated using rotor cyclic control, albeit at a 20-30% performance penalty. However, rotor flapping transients tend to increase by the action of the SLA system. A dual-objective control design approach demonstrates the effectiveness of suppressing both loads and flapping simultaneously.

Wind tunnel tests have been made on the effect of bottom wall and sidewall fences of various sizes and combinations in reducing the swirl of an S-duct at high angle of attack. It is shown that large reductions can be made: in some cases the swirl is eliminated or even reversed. Fences of modest size can be chosen to combine a large swirl reduction with an improvement of the total pressure distortion at high incidence and only a small effect on mean pressure recovery at low incidence. Some further observations are made on the nature of the general swirl phenomenon.

Automated manufacturing of thermoplastic composites has found increased interest in aerospace applications over the past three decades because of its great potential in low-cost, high rate, repeatable production of high performance composite structures. Experimental validation is a key element in the development of structures made using this emerging technology. In this work, a $750\times640\times240$ mm variable-stiffness unitised integrated-stiffener out-of-autoclave thermoplastic composite wingbox is tested for a combined shear-bending-torsion induced buckling load. The wingbox is manufactured by in-situ consolidation using a laser-assisted automated tape placement technique. It is made and tested as a demonstrator section located at 85% of the wing semi-span of a B-737/A320 sized aircraft. A bespoke in-house test rig and two aluminium dummy wingboxes are also designed and manufactured for testing the wingbox assembly which spans more than 3m. Prior to testing, the wingbox assembly and the test rig were analysed using a high fidelity finite element method to minimise the failure risk due to the applied load case. The experimental test results of the wingbox are also compared with the predictions made by a numerical study performed by nonlinear finite element analysis showing less than 5% difference in load-displacement behaviour and buckling load and full agreement in predicting the buckling mode shape.