Optimisation of aircraft ground operations to reduce airport emissions can reduce resultant local air quality impacts. Single engine taxiing (SET), where only half of the installed number of engines are used for the majority of the taxi duration, offers the opportunity to reduce fuel consumption, and emissions of NOX, CO and HC. Using 3510 flight data records, this paper develops a model for SET operations and presents a case study of London Heathrow, where we show that SET is regularly implemented during taxi-in. The model predicts fuel consumption and pollutant emissions with greater accuracy than previous studies that used simplistic assumptions. Without SET during taxi-in, fuel consumption and pollutant emissions would increase by up to 50%. Reducing the time before SET is initiated to the 25th percentile of recorded values would reduce fuel consumption and pollutant emissions by 7–14%, respectively, relative to current operations. Future research should investigate the practicalities of reducing the time before SET initialisation so that additional benefits of reduced fuel loadings, which would decrease fuel consumption across the whole flight, can be achieved.

A new rating scale for capturing pilot subjective assessment of simulation fidelity is described in this paper. The scale has been developed through a series of flight and simulation trials using six test pilots from a variety of backgrounds, and is based on the methodology utilised with the Cooper-Harper Handling Qualities Rating scale and the concepts of transfer of training, comparative task performance and task strategy adaptation. The development of the new rating scale has been undertaken using simulations of rotary-wing aircraft on the University of Liverpool’s HELIFLIGHT-R research simulator, in conjunction with the Canadian Flight Research Laboratory’s Bell 412 ASRA in-flight simulator. The utility of the scale applied to locating fidelity boundaries for quantitative metrics is illustrated for an inter-axis coupling criterion. The work described in this paper is preliminary in nature, and research activities are on-going to continue the validation of the fidelity rating scale.

The work presented here forms part of a larger project on Large-Eddy Simulation (LES) of aeroengine aeroacoustic interactions. In this paper, we concentrate on LES of near-field flow over an isolated NACA0012 aerofoil at zero angle-of-attack and a chord based Reynolds number of Rec = 2 × 105. A wall-resolved compressible Numerical Large Eddy Simulation (NLES) approach is employed to resolve streak-like structures in the near-wall flow regions. The calculated unsteady pressure/velocity field will be imported into an analyticallybased scheme for far-field trailing-edge noise prediction later. The boundary-layer mean and root-mean-square (rms) velocity profiles, the surface pressure fluctuation over the aerofoil, and the wake flow development are compared with experimental data and previous computational simulations in our research group. It is found that the results from the wall-resolved compressible NLES are very encouraging as they correlate well with test data. The main features of the wall-resolved compressible NLES, as well as the advantages of such compressible NLES over previous incompressible LES performed in our research group, are also discussed.

Three analyses are presented for predicting the development of an axisymmetric turbulent jet issuing into a co-flowing external air stream. The first analysis is analogous to a method used by Patel to predict the growth of a two-dimensional jet in an external air stream. The method is found to be inadequate when the excess velocity on the axis of the jet becomes small compared with the external stream velocity. The second analysis assumes that the turbulence structure is similar at different streamwise stations but it breaks down when the advection of turbulent energy becomes comparable with the turbulent energy production. In the third approach, a two-parameter model of turbulence developed by Rodi and Spalding, which uses two differential equations for the turbulent energy and the length scale of the turbulence respectively, is found to predict closely the experimental results of Antonia and Bilger for a ratio of jet to external stream velocity of 3.0. The success of this last method emphasises the non-similar character of turbulence.

A simple method is given for calculating, in incompressible flow, the spanwise and chordwise loadings on thin swept wings with symmetrical aerofoil sections and of moderate and large aspect ratios. The present note is based on, and continues, the investigation described in an earlier paper. The design of swept wings with given spanwise and chordwise load distributions, obtained by modifying the plan form of the wing or by adding camber and twist, is briefly considered, as well as the design of swept wings which give minimum induced drag.

A review of recent developments in helicopter noise reduction is presented. The paper includes a discussion of the present state of understanding of helicopter noise radiation mechanisms and a review of actual noise reductions achieved by the manufacturers over the past decade. An analysis of progress achieved in meeting certification limits is given, together with an evaluation of the impact on design. Future trends in helicopter noise reduction based on current noise research programmes are discussed and suggestions advanced for possible additional areas of study.

A three-dimensional moving mesh method for solving the Euler equations describing the compressible flow about a wing undergoing arbitrary motions and deformations is described. A finite-volume formulation is chosen where the volumes distort as the wing moves or deforms. By using transfinite interpolation, a technique for generating the required sequence of grids has been developed. Furthermore, as the speeds of the grid at the vertices of the finite volumes are required by the flow solver, transfinite interpolation is also used to obtain these by interpolation of the boundary speeds. A two-dimensional version of the method has also been developed and results for both two- and three-dimensional transonic flows are presented and compared with experimental data where available.

The CAA has recognised the need to develop civil standards for unmanned aerial vehicles (UAVs). Accordingly the CAA Directorate of Airspace Policy has published the document ‘CAP 722 –Unmanned Air Vehicle Operation in UK Airspace – Guidance’, (which can be accessed via the CAA website — http://www.caa.co.uk/docs/33/CAP722.pdf). The document itemises the top-level military and civil regulations which impact upon the operation of UAVs in the UK. One of the items listed within the document is ‘Civil Certification’ of UAVs. This particular aspect is the responsibility of the Design and Production Standards Division, (D&PSD), of the CAA Safety Regulation Group. The purpose of this paper is to present and explain the current position of D&PSD regarding the appropriate design standards to be applied for the civil certification of the airworthiness of UAVs.

This paper presents the CAA position that UAVs should be granted permission to fly by qualifying for certificates of airworthiness, by demonstration of compliance with defined airworthiness standards comparable to, and derived from, those applied to manned aircraft. The paper also presents a method for determining, to a first approximation, the level of airworthiness standards which should be applied to any particular design of UAV by reference to the existing codes of airworthiness requirements for manned aircraft.

Several workers have found evidence for the presence of longitudinal vortices in wind tunnel wall boundary layers near the centre line of each wall, although the only actual publication appears to be the reproduction of Bearman's work in Ref. 3. The phenomenon usually observed is shown in Fig. 1: the boundary layer is considerably thickened near the centre line. (This symmetrical and localised perturbation should be distinguished from the effect of non-uniformities in the wind tunnel screens which apparently also leads to longitudinal vortices, randomly spaced across the width of the working section). The hypothetical vortex pattern is also shown in Fig. 1: it arises because lateral pressure gradients in the wind tunnel contraction naturally deflect the slow-moving fluid in the boundary layer towards the centre line more strongly than the main ‘inviscid’ part of the flow, implying the generation of longitudinal vorticity: this is secondary flow of Prandtl's first kind.

At a sufficiently high Reynolds number small separation bubbles, as for example at the leading-edge of an aerofoil, are characterised by laminar flow separation and turbulent reattachment. The transition to turbulence in the separated-flow region through the introduction of a disturbance at the boundary, localised in time and space, is examined. It is found that transition to turbulence is promoted by rapidly growing convective instabilities, and not an absolute instability at the point of application of the disturbance.

Lean techniques are having a major impact on aerospace manufacturing. However, the cost and value of aerospace (and many other) products is determined primarily in product development. Migrating lean to engineering processes is ongoing in the industry, and a subject of study at the MIT Lean Aerospace Initiative. This paper summarises findings to date, with references to both research literature and successful implementation examples. To implement lean engineering, a three-part approach is needed: Creating the right products, with effective lifecycle and enterprise integration, using efficient engineering processes.

Large-eddy simulations (LES) have been employed to investigate the far-field four-vortex wake vortex evolution over 10min behind an aircraft formation. In formation flight scenarios, the wake vortex behaviour was found to be much more complex, chaotic and also diverse than in the classical single aircraft case, depending very sensitively on the formation geometry, i.e. the lateral and vertical offset of the two involved aircraft. Even though the case-by-case variability of the wake vortex behaviour across the various formation flight scenarios is large, the final plume dimensions after vortex dissolution are in general substantially different from those of single aircraft scenarios. The plumes are around 170 to 250m deep and 400m to 680m broad, whereas a single A350/B777 aircraft would produce a 480m deep and 330m broad plume. Formation flight plumes are thus not as deep, yet they are broader, as the vortices do not only propagate vertically but also in span-wise direction. Two different LES models have been employed independently and show consistent results suggesting the robustness of the findings. Notably, $CO_{2}$ emissions are only one contribution to the aviation climate impact among several others like contrails and emission of water vapour and nitrogen oxides, which would be all affected by the implementation of formation flight. Thus, we also highlight the differences in ice microphysical and geometrical properties of young formation flight contrails relative to the classical single aircraft case.

In this paper the computed flow and performance characteristics at low angle-of-attack (AOA) of an integrated diverterless supersonic inlet (DSI) are presented. The subsonic characteristics are evaluated at M∞ = 0·8 while the supersonic characteristics are evaluated at M∞= 1·7, which is near the design Mach number for the intake. In addition to the external flow features, the internal intake duct flow behaviour is also evaluated. The results of this study indicate effective boundary layer diversion due to the ‘bump’ compression surface in both subsonic and supersonic regimes. At M∞ = 1·7, the shockwave structure (oblique/normal shockwave) on the ‘bump’ compression surface and intake inlet is satisfactory at design (critical) mass flow rate. The intake duct flow behaviour at subsonic and supersonic conditions is generally consistent with ‘Y’ shaped intake duct of the present configuration. The secondary flow structure inside the duct has been effectively captured by present computations. The computed intake total pressure recovery at M∞ = 1·7 exhibits higher-than-conventional behaviour at low mass flow ratios, which is attributed to unique inlet design. Overall computed subsonic and supersonic total pressure recovery characteristics are satisfactory under the evaluated conditions and are also in agreement with wind tunnel test data.

An analysis of results from reports on droplet evaporation indicated a marked difference in behaviour depending upon the suspension technique used. Few results were obtained with freely falling droplets, the condition which is of interest in aerial spraying operations. Test results obtained at Cranfield by dropping droplets down a climatic tower showed two unexpected results.

The diameter of water droplets reduced linearly with time until they reached a size of about 150 micrometers and then the rate of reduction in diameter increased by about 27%. It is shown that this change occurs when the fall Reynolds number of the droplet becomes less than about 4.

Above this Reynolds number the airflow is separated from the base of the droplet and the trapped air in the toroidal vortex becomes fully saturated and no evaporation occurs over the base region. Below a Reynolds number of 4 the flow is attached everywhere and evaporation occurs from the whole surface area.

When small amounts of molasses are added to a spray formulation the droplets evaporated initially in a way similar to water droplets. As a result of falling, a toroidal vortex motion is induced within the droplet. This causes the same small fraction of the volume of the droplet to cycle close to the surface.

Hence, this fraction of the volume looses all its water, becomes far more viscous as a result and the consequence is the toroidal motion almost stops. In effect the molasses forms a skin over the droplet surface inhibiting further evaporation of the remaining volume of water.

Flowfield measurements of the blade tip vortices from a rotating helicopter rotor model were performed by three component laser-Doppler velocimetry (3D-LDV) and conventional (two component) particle image velocimetry (PIV). In general, the results are in good correspondence, but also illustrate the different properties of both techniques: LDV offers the capability of three-component measurements, whereas PIV captures the unsteadyness of the flowfield.

This paper describes and compares various analyses leading to the development of approximate models for the linear stability modes of the non-rigid airship. The progress of the analyses was variously dependent on assumptions made from the detailed scrutiny of linear numerical models for three airships. For each airship studied, the linear models were obtained from non-linear simulation models by linearising about a number of chosen trim speeds representative of a typical speed envelope. The decoupled linear models comprised the longitudinal and lateral-directional state equations of the neutrally buoyant airship, for speeds from the hover to 30m/sec. Since the fidelity of the earliest airship models was not known, the principal purpose of this paper is to re-visit the original analysis using a later airship model of known excellent fidelity.

The longitudinal modes of the airship comprise the surge mode, the heave-pitch subsidence mode and the oscillatory pitch-incidence mode. The lateral-directional modes of the airship comprise the sideslip subsidence mode, the yaw subsidence mode and the oscillatory roll pendulum mode. Approximate models for these modes are derived and expressed in terms of concise aerodynamic stability derivatives. The mode characteristics are discussed, and the approximate models are compared with the actual airship modes over the typical airspeed envelope.

Measurements were made of the pressures in flattened, blunt-nosed pitot tubes, in an air stream at Reynolds numbers from 13-1,000 (based on the external height of the tube). Three tubes were used, having cross sections with ratios of width to height (W/H) of 3, 7 and 11.

The photoelastic “stress freezing” technique has been employed to evaluate the elastic stress concentration factors associated with the fillet blend radius in a number of shouldered shafts. The full range of practical sizes of blend radius and depth of shoulder has been examined. Comprehensive results for the shaft subjected to torsion, pure bending and axial load are given in this (Part I) and in two subsequent papers (Parts II and III).

The accuracy of the graphs of stress concentration factors is better than six per cent. Comparison has been made with the existing theoretical and experimental results for each mode of loading. The results of an investigation into die limiting value of the stress concentration factor (for a particular shoulder radius) as the depth of the shoulder is increased to infinity are included in Part III.

Dual-mode scramjet is one of the candidates for hypersonic flight propulsion system which will be used in wide range of flight Mach numbers from 4 to 12 or higher, wherein dual-mode scramjet should be well designed to be suitable for subsonic/supersonic combustion operation according to the flight conditions. Therefore this system is required to operate in a finite number of operational modes that necessitate robust, stable, and smooth transitions between them by which selective operability of supersonic/subsonic combustion modes and efficient combustor operation in these modes may be realised. A key issue in making mode transition efficient and stable is mode transition control. The major problem in mode transition control is the handling of the various flow and combustion coupling effects of dual-mode scramjet whose physical states are spatially coupled and whose governing equations are partial differential equations. Involving these distributed parameter issues, our basic idea is using the shape control theory to study the control problems of mode transition for dual-mode scramjet with the aim of achieving the desirable design properties and increasing control reliabilities. This specific approach is motivated by the promise of novel techniques in control theory developed in recent years. Concrete control arithmetic of this approach, such as shape control model, sensitivity analysis and gradient-based optimisation procedure, are given in this paper. Simulation results for an axisymmetric, wall-injection dual-mode scramjet show the feasibility and validity of the method.

To improve the crashworthiness of civil aircraft, the design concept of energy absorption structure for civil aircraft is investigated. Two typical different design principles could be identified. The first category includes Helicopter and Light fixed-wing Aircraft (HLA), and Transport, Mid-size and Commuter type Aircraft (TMCA) are classified into the second group. Frame, strut and bottom structure are the three kinds of energy absorption structure for TMCA. The strut layout of conventional civil aircraft is studied and some energy absorption devices are adopted. High efficiency energy absorption structures such as the foam and sine-wave beam are employed as the bottom structure for both of HLA and LMCA. The finite element method is used to analyse and design energy absorption structure in aircraft crashworthiness problem. Results show that the crashworthiness of civil aircraft could be largely improved by using proper strut layout and excellent energy absorption device. The stiffness combination of frame and strut should be considered to get better global aircraft deformation. Supporting platform and failure model are the two core problems of bottom energy absorption structure design. Foam and sine-wave beam under the lifted frame could improve the crashworthiness of civil aircraft.