This paper describes the current status of knowledge regarding the impact of aviation on the atmosphere. The growth of the aviation industry is likely to continue in the future and at present there are significant concerns that this will adversely affect climate and local air quality in the vicinity of airports. Indeed, it is possible that air quality may become a constraint to increased capacity of some large hub airports. Currently the radiative forcing impact from aircraft emissions, other than CO2, are not accounted for in the International Civil Aviation Organization emission trading scheme proposals. Taking only this approach, emissions trading could, in theory, increase the radiative forcing, rather than decrease it. The study of improved ‘environmentally friendly’ flight is still in its infancy, however any proposed and actual new developments, in terms of technology and operational practice, should include an environmental assessment.

The aim of this work is to investigate how the wake vortex hazard generated by an aircraft can be reduced by modifying the wing lift distribution. The wake associated with a given lift distribution is simulated with a two-dimensional, in viscid ‘vortex method’. A numerical optimisation technique is used to guide the modifications so that a hazard measure, based on the rolling moment induced on a following aircraft encountering the wake, is minimised. The results show that a significant (~20–25%) reduction in hazard compared with conventional high-lift configurations can be achieved, via lift distributions that have a local minimum between the regions contributing to the flap and tip vortices. The counter-signed vorticity associated with this minimum appears to prevent the tip and flap vortices from merging, resulting in a safer wake. Furthermore, the hazard reduction is robust to variations in other features of the lift distribution and should thus be achievable in practice.

Competitive pressures and economic constraints are driving aircraft manufacturers towards an ever-increasing exploitation of CFD for design, optimisation and prediction of off-design conditions. Such exploitation is favoured by rapid advances in meshing technology, numerical algorithms, visualisation tools and computer hardware. In contrast, the predictive capabilities of mathematical models of turbulence are limited — indeed, are often poor in regions of complex strain — and improve only slowly. The intuitive nature of turbulence modelling, its strong reliance on calibration and validation and the extreme sensitivity of model performance to seemingly minor variations in modelling details and flow conditions all conspire to make turbulence modelling an especially challenging component of CFD, but one that is crucially important for the correct prediction of complex flows.

Unsteady spiked body flows were simulated by a second order time-accurate CFD method. Laminar, axisymmetric flow was considered at Mach 2.21 and Mach 6 freestreams and Reynolds’ numbers based on the blunt body diameter of 0.12 million and 0.13 million, respectively. A spiked forward facing cylinder with spike lengths between LID = 1.00 and LID = 2.40 was used as the model geometry. Following the numerical method’s verification, the individual flow modes of oscillation and pulsation were examined. The frequency of the events was found in good agreement with the experiment, while the pressure amplitudes were overpredicted in the Mach 6 cases. Analysis of the numerical results showed that the oscillation flow mode was driven by a viscous mechanism, whereas the pulsation by an inviscid one. The hysteresis phenomenon in the transition between the two flow modes was predicted qualitatively.

This paper examines the influence of exit separation, exit availability and seating configuration on aircraft evacuation efficiency and evacuation time. The purpose of this analysis is to explore how these parameters influence the 60-foot exit separation requirement found in aircraft certification rules. The analysis makes use of the airEXODUS evacuation model and is based on a typical wide-body aircraft cabin section involving two pairs of Type-A exits located at either end of the section with a maximum permissible loading of 220 passengers located between the exits. The analysis reveals that there is a complex relationship between exit separation and evacuation efficiency. A main finding of this work is that for the cabin section examined, with a maximum passenger load of 220 and under certification conditions, exit separations up to 170ft will result in approximately constant total evacuation times and average personal evacuation times. This practical exit separation threshold is decreased to 114ft if another combination of exits is selected. While other factors must also be considered when determining maximum allowable exit separations, these results suggest it is not possible to mandate a maximum exit separation without taking into consideration exit type, exit availability and aircraft configuration.

The trailing vortex wake shed by large transport aircraft not only has detrimental effects on their aerodynamic efficiency through the production of induced drag but can cause a considerable safety hazard to smaller following aircraft. This hazard is a major factor in determining the minimum allowable spacing between aircraft during airport operations with separations of up to five nautical miles being required between heavy aircraft and smaller aircraft following them on the approach for landing. It has been estimated that the minimum safe operating distance based on considerations other than vortex wake hazard is approximately two nautical miles’. The potential economic saving from the reduction in operating times at busy airports as well as the improvement in aerodynamic efficiency has led to the investigation of many schemes for the alleviation of these trailing vortices.

The aerodynamic performance of two dimensional (2D) aerofoils and finite aircraft wings is investigated when the leading edge is modified with a standard ice shape. Three ice-shapes, G1, G3 and R7 are selected for this investigation. For the aerofoils it was observed that the presence of strong vortex flows near the leading edge fundamentally changes the flow in that region forcing earlier transition and premature separation. Normal suction peaks designed to provide high aerodynamic lift and prolonged attached flow are replaced with a system of vortices which produce unwanted pressure spikes which influence the development of the boundary layer leading to earlier stall. The present study also includes the three dimensional (3D) effects due to iced leading edges on aircraft surfaces. As different from such studies elsewhere, which only considered the simple sweep back effects on constant cross section wings, the present work includes more realistic wings equipped with twist, taper and non-equal sweep back at leading and trailing edge of the wing. In absence of any experimental data on such real life type configurations, the computational fluid dynamic (CFD) results were first validated against measurements on non-swept wings. It was found that ice formations cause noticeable changes of the flow in the leading edge regions of the wing and promote strong 3D effects, which pervade across the entire span wise direction of the model.

This paper discusses the evolution of airpower and the ever-increasing component of stealth technology in dictating warfare, based on the need felt in the literature to integrate and place them in the right perspective. Hence, the role of stealth aircraft and the dominance of all aspects of stealth technology, especially in recent conflicts, are reviewed. The integration of the fundamental aspects of stealth technology, through its classification, types of aircraft signatures: especially radar, infrared, and visual signatures, their sources of origin, modelling techniques, and methods of signature reduction are discussed. Due to the increasing importance of infrared signatures relative to radar signatures, infrared signatures are also closely examined and analysed. Hence, the lock-on and surveillance ranges of infrared detectors are compared. Future projections in stealth technology, especially based on the role of anti-stealth technologies, are also elaborated.

The second part of this survey closed a little after the termination of the First World War. By then the utility of the aeroplane as an instrument of warfare had been amply demonstrated. Moreover, something of its future potential in civil use had been indicated by such British achievements as the first transatlantic flight by Alcock and Brown in 1919, albeit in a Vickers Vimy biplane of otherwise limited performance. It was also at this time that the new aerodynamic ideas of Kutta, Zhukovskii and Prandtl reached Britain. All such thinking pointed inevitably to the greater speed, efficiency and economy to be expected of the streamlined monoplane.

The effects of aviation on the environment are becoming increasingly important. The world's airlines already have thousands of aircraft serving the passenger and freight markets and the market demand is still increasing. Unless steps are taken to address the problems created by aircraft noise and emissions, the future progress of the air transport industry will be inhibited.

The seriousness of the topic and the factual nature of the Technology Sub Group report are the reasons why this special issue of The Aeronautical Journal is being devoted to “Greener by Design”. Although the report has been carefully read by a number of people outside the Sub Group, it has not been refereed in the normal way (though a number of minor numerical changes have been made to the tables and text of the original report, partly to correct errors and partly for internal consistency). The reports from the other two Sub Groups, covering operations and market-based options, together with the Summary for Policymakers of the Technology Sub Group, will be issued shortly in a single volume but will not necessarily be published in The Aeronautical Journal.

Control surface effectiveness is an important parameter for any aeroplane. For a hypersonic aircraft, though the power required to operate the flaps is determined by low speed flying conditions, it is imperative to know the effect of flaps at hypersonic speeds. Hence, studies have been done on this topic by aerodynamicists for over 40 years. In spite of this, only a limited data is available in the literature on this subject.

This paper discusses the experimental study of the effect of sweep on the aerodynamic characteristics of thin slab delta wings with flaps at hypersonic speeds. For the purpose of this investigation, a novel special thin six-component balance, which has a thickness of 4mm and can be housed inside wings with 8mm thickness, has been designed. The wings had a sweep of 76°, 70° and 65°, t/c of 0.053 and flaps with 12% of wing area and 12% of wing chord. Testing were done at Mach 8.2, Re number of 2.13 x 106 (based on chord), from α = –12° to 12° and flap angle of 20°, 30° and 40°. Separation lengths, measured from Schlieren pictures, clearly show that there is ‘no appreciable’ effect of sweep on them. Also, using a simple local flow field calculation, the separation has been identified to be transitional in nature. These features of separation reflect in the force data. Because of the small separation length, the flaps (inspite of their small size) were very effective in generating additional CN, CM and Cl, which increased with increase in flap angle. In general, the CN, CM and XCP were unaffected by sweep for symmetric flap deflection at positive incidences and asymmetric flap case. For symmetric flap case at negative incidences, only CN was not influenced by the sweep but CM decreased and XCP moved upstream as the sweep is decreased. The wing with lower sweep produces higher CA and lower (L/D)max for both symmetric and asymmetric flaps. The rolling moment and adverse yaw increased with decrease in sweep for asymmetric flap deflection. Newtonian theory is shown to be incapable of predicting the effect of sweep on Cl, Cn and on the incremental values of CN, CM and CA.

In conclusion, it can be said that a small flap is generally adequate for hypersonic aeroplanes provided they operate at altitudes where transitional and turbulent separation can be expected to occur. This would make the flaps effective and thus enable ample control authority.

A computational fluid dynamics (CFD) model of three different axisymmetric convergent nozzle designs was created. Two of the nozzles had trailing-edge modifications (TEMs), which consisted of four regularly-spaced tooth-gap castellation pairs. One of the castellated nozzles had a divergent gap profile (regular castellated) and the other had a convergent gap profile (convergent-chamfered castellated). The remaining plain nozzle had no castellations.

Calculated CFD results were examined at a nozzle pressure ratio (NPR) of 6·5 in the near field of the regular castellated and convergent-chamfered castellated nozzles. Jet fluid was seen to be ejected radially from the jet core through the gaps between the castellations at the nozzle exit plane. Contra-rotating streamwise vortices were seen to develop at the interface between the gap and castellation. These were of the sense that acts to pump jet fluid out in the plane of the gap and entrain ambient fluid in the plane of the tooth. Peak increases of approximately 45% in non-dimensional mass flow rate over the plain nozzle were seen.

The ejections from the four-tooth regular-castellated nozzles extended far enough radially that they became detached from the main jet core. The contra-rotating vortices moved closer together with downstream distance reinforcing the pumping action between them. The jet ejections from the convergent-chamfered castellated nozzles did not extend far enough radially to detach from the main core (at the NPR tested) nor did the contra-rotating vortices move significantly closer together. Peak increases of approximately 45% and 25% (for the regular-castellated and convergent-chamfered castellated nozzles respectively) in non-dimensional mass flow rate over the plain nozzle were seen.

The strength of the contra-rotating vortices had dropped by two orders of magnitude at five nozzle diameters downstream from the exit plane. This suggested that the extent of fluid ejection may be more important than the initial strength of the vortices (although stronger vortices may help to eject fluid further out radially). Thus, configurations of nozzle geometry which cause the resulting streamwise vortices to interact favourably in ejecting fluid, would be desirable.

The interaction between slender body vortices and a single vertical fin located down the axis of the body is investigated numerically for angle of attack of 30 deg and Reynolds number of 6000. The present research includes a parametric study on the effect of fin location and fin azimuthal position. A numerical method based on the pseudo-compressibility is used for solving the three-dimensional incompressible Navier-Stokes equations using the Lower-Upper Symmetric Gauss-Seidel implicit scheme. The numerical results show that the vortices remain very coherent and attached to the body until they reach the fin section, where they become less coherent and begin to separate from the body. Also, the results show that the fin location does not affect the upstream development of the vortices, but it does affect the location at which the vortices separate from the body. The effect of azimuthal fin positions was also investigated. As the azimuthal angle of the fin increased, the size of the vortex on the port side decreased, but the starboard side vortex grew in size and moved across the leeward ray to the port side. The computed results are found to agree well with the available experimental data. The effect of the grid size on the numerical solutions was also investigated.

This paper presents an algorithm to solve the three-dimensional unsteady Euler equations on unstructured tetrahedral meshes using a dynamic mesh algorithm. The driving algorithm is an upwind biased implicit second order accurate cell-centered finite volume scheme. The spatial discretisation technique involves a naturally dissipative flux-split approach that accounts for the local wave propagation characteristics of the flow and captures shock waves sharply. A continuously differentiable flux limiter has been employed to eliminate the spurious oscillations near shock waves, generally arising in calculations involving upwind biased schemes. The temporal discretisation is also second order accurate and uses a Newton linearisation for unsteady calculations. To calculate time dependent flows a dynamic mesh algorithm has been implemented in which the mesh is moved to conform to the instantaneous position of the body by modelling each edge of each cell by a spring. The paper presents a description of the solver and the grid movement algorithm along with results and comparison that assess their capabilities.

Preliminary investigations of the effect of tracer particle characteristics on flow visualisation in a steady delta wing leading-edge vortex are presented, using both 3D simulations of particle trajectories and a simplified analysis of the radial equilibrium conditions. It is shown that many of the flow structures apparently seen in wind and water tunnel visualisation studies of delta wing vortices may in fact be artefacts of the tracer particle characteristics. In particular the ‘smoke tube’ or ‘hole’ effect often seen in smoke flow visualisation in wind tunnels does not (contrary to popular usage) indicate the size of the viscous subcore, and is not a reliable indicator of the vortex structure. The implications for both visualisation studies and optical flow measurement techniques are discussed.

The use of rule-based fuzzy logic system for estimating the stability and control derivatives for the longitudinal aircraft motion is proposed. The capabilities of the fuzzy logic system in estimating both the short-period and the phugoid mode of motions are explored. The flight data used in the estimation process were generated using the three nonlinear longitudinal equation of motion for a small remotely piloted vehicle with all the aerodynamic coefficients obtained from the wind-tunnel tests. The preferred method of perturbation of the aircraft elevator for data collection is also highlighted. The stability and control derivatives are estimated as the change in the aerodynamic force or moment due to small variation in one of the motion or control variables about its nominal value when the rest of the variables are held constant at their respective nominal values. The changes in the aerodynamic force and moment are predicted using the fuzzy logic system. The results show that the fuzzy logic system has a good potential as alternative tools for parameter estimation from flight data. The proposed method does not require any guesses of the initial values of the flight parameters.

Autonomous shipboard launching of UAVs in a hostile atmospheric environment would greatly extend the usefulness of these aircraft. Currently a majority of UAVs are remotely launched under limited conditions. This paper presents control and guidance strategies that will allow autonomous launching under a much greater set of atmospheric hazards. Two modern control techniques are compared for use in this case. The main problem when dealing with low speed UAVs is found to be high steady winds causing stall. Several strategies are presented that can allow stall to be avoided and recovered.

Over the years there have been aircraft accidents and incidents where crews have been found in error. Many times the errors were found with hindsight to have been compounded by poor checklists and inadequate design. Of particular concern to the author is the dilemma of ‘smoke and fire of unknown origin’. This is a serious emergency with little time for complex troubleshooting. It is a human factors classic ‘fight or flee’ scenario. Research among the major UK airlines, assisted by the UK Flight Safety Committee, has revealed that there is an inconsistency in direction as to when to recommend divert and land as of paramount importance. Thus a gap is created and those believed to be responsible identified. The author has experienced several instances and takes the opportunity to describe one of them. Lessons are included on work being done in modifying aircrew training to cater for the removal of the flight engineer. Design recommendations are made and warnings given about the future. Should recommendations in the wake of accidents and incidents be ‘risk managed’ by the regulators? Ultimately, the crews are forced to rely on their checklist handling and limited technical knowledge that sometimes means that the gap is too wide and an incident occurs. The paper is a collaborative effort prepared by the author on behalf of the RAeS Flight Operations Group (FOG) but the author alone is responsible for the version published.

The development of micro-air vehicles poses a considerable challenge to the materials scientist. Vehicles that may occupy less than 10cm in any direction require many structural components with dimensions well below 1mm when conventional engineering and assembly techniques are employed. As with conventional unmanned air vehicles (UAVs), and full-scale aircraft, materials that display a huge breadth of characteristics are required to build the airframe and to engineer sub-components, while meeting stringent mass and volume budgets. In particular, high modulus, high toughness, easy formability and jointing are desirable qualities. For larger structures, conventional foamed materials can meet some of these properties. However, the pore sizes in these conventional materials are too large to allow manufacture of micro-scale components, airframe sections and skins. This paper describes the manufacture and properties of some micro-foamed polymers, and polymer composites, metals and oxide ceramics that are being developed at DERA specifically for micro-vehicle applications.

Closed-form equations for the evaluation of forces on a body in compressible, inviscid and viscous flows are presented. The expressions which are used to calculate lift and drag require only knowledge of the velocity field (and its derivatives) on an arbitrarily chosen contour enclosing the body. Various forms of the force equations are presented, both for two-dimensional and three-dimensional flows. The work is an extension of earlier research by Noca) developed for incompressible flows. The equations are implemented on a number of two-dimensional aerofoil cases and the results show excellent agreement with surface and wake integration approaches.