Each year greenhouse gas emissions remain high the climate mitigation and adaptation challenges grow. The economic downturn was already in train in 2008, yet concentrations of CO2 increased unabated. Without concerted effort across all sectors there will be little chance of avoiding ‘dangerous climate change’ and the aviation sector has a clear role to play. Current and forthcoming technologies, operational practices and behavioural change offer widespread opportunities for other sectors to mitigate their CO2 emissions in absolute terms, but as they do so, aviation will become an increasingly important player. By comparing a range of global cross-sector emission scenarios with existing aviation projections, this paper illustrates the importance of understanding the future context with regard to other sectors when assessing the aviation industry’s potential impact. Given growth projections for aviation and the relatively slow pace of technological change, aviation’s proportion of 2050 global CO2 emissions is low only in those global cross-sector emission scenarios where there is a high probability of ‘dangerous climate change’. For a ‘reasonable’ (>50%) chance of avoiding ‘dangerous climate change’, the most technologically radical scenarios for aviation make up 15% of global CO2 in 2050 and conventional scenarios exceed the carbon budget entirely. Only by recognising that aviation’s currently projected emissions are incompatible with avoiding ‘dangerous climate change’ can the industry fully grasp the challenge of accelerating innovation and managing demand to deliver a more sustainable route to 2050 and beyond.

Implicit large eddy simulations (ILES) of two different Royal Navy ships have been conducted as part of the UK Ship/Air Interface Frame-work project using a recently developed very high order accuracy numerical method. Time-accurate CFD data for fourteen flow angles was produced to incorporate into flight simulators for definition of safe helicopter operating limits (SHOLs). This paper discusses the flow phenomenology for the different wind directions and where possible reports on the validation of the ILES results for mean and fluctuating velocity components and spectra against experimental data.

The main objective of the project is to develop and validate the use of a fuel cell based power system for the propulsion of more/all electric aircraft. The Rapid 200-FC two-seater electric-motor-driven aeroplane which is powered by fuel cells is at present being completed and will be validated during a flight test in Autumn 2009. Several configurations have been evaluated in order to install the new energy and propulsion system on board while maintaining the centre of gravity within allowable limits. The fuel cell system and the electric motor are being integrated on board. The FC stack will be able to deliver a maximum continuous power of 20kW. A battery pack has to guarantee another 20kW of maximum continuous power for a limited time period (15 minutes), during take-off, climbing and, in the case of emergency, during landing.

The main goal of the project is to validate the overall high performance of an all electric aircraft system which is capable of remaining aloft for one hour. A parametric analysis has also been carried out to evaluate which key technologies influence the performance of future aircraft to the greatest extent.

This contribution is based on the Lanchester Lecture of the Royal Aeronautical Society held in Manchester, UK, in October 2009. The lecture reflected on further possible perspectives for aeronautical research, in particular with respect to aerodynamics. The air transport system of today has matured to very high performance and safety standards. However, due to the ongoing global developments of economy and ecology, new challenges have to be faced. Thus, the relevance of all engineering aeronautical disciplines contributing to the final aircraft configuration is now closely scrutinised, and especially aerodynamics is challenged whether any investment in further research is still justified. In this contribution, it is argued that, when carefully exploited, aerodynamics still offers a substantial technological perspective, well beyond the concepts for the next generation of aircraft as currently being pursued. As an example, the aerodynamics of a forward swept-wing are discussed and it is outlined that, when combining the transonic characteristics of a forward swept-wing with the advantages of natural laminar-flow, considerable improvements in efficiency may be gained. However, this potential may only be leveraged through a strategic, long-term orientation of fundamental and applied research.

State of the art test capabilities are described that can enable optimised electronic warfare (EW) system development programmes for military air platforms. EW systems are key enablers of air platform survivability and the above is important as defence ministries and industry world-wide wrestle with affordability, technical and industrial challenges. The military end user needs improved capability quickly, with high availability; defence ministries want this at lowest possible cost and with minimum risk. Satisfying these requirements, while remaining in business against a background of fierce international competition and world financial crisis, is indeed challenging. Improvements are described that can enable improved technical performance and lower cost, time and risk programmes than hitherto, with emphasis on RF EW systems and the recent multi-£M upgrade to BAE Systems’ EW Test Facility, with its aircraft-sized anechoic chamber. This upgrade assures a further decade of world-class UK capability for testing manned and unmanned air platforms.

In some situations, closed loop control by the pilot can result in the combined pilot-aircraft system becoming marginally stable or even unstable. This can happen whether the pilot is controlling attitude or flight path. In this paper, an investigation into helicopter stability under flight-path constraint below the minimum-power speed is reported. The work provides a theoretical basis for flight path handling qualities criteria particularly for flight on the, so-called, back-side of the power curve. The research uses the theory of weakly coupled systems by partitioning the helicopter longitudinal dynamics to investigate three interacting subsystems – classically the surge mode, the phugoid mode and the heave mode. Under certain conditions, strong control of flight path or vertical speed is shown to drive the aircraft-pilot system unstable and a conflict is shown to exist between feedback gain values to guarantee stability of both the surge and the flight path motions. This conflict constitutes a potential source of adverse rotorcraft-pilot couplings. The problems are exacerbated in cases when the use of collective control is restricted. The phenomenon is explored in both ground based simulation and flight test to provide a verification of the theory.

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 aerodynamic design of the Miles M.52 experimental supersonic aircraft is reviewed relative to the state of knowledge at its time of 1943 – 1946. Drawing on widely-ranging material, much not previously published, this study enlarges upon, and in places amends, previous accounts of the project.

Based on advice collated from the fields of aerodynamics and ballistics, Miles conceived an original and forward-looking design, backed by an extensive test programme. Novel solutions to new requirements in the areas of structures and systems provided a robust airframe, showing fair prospects of being stable and controllable throughout the specified flight envelope. An equally innovative power plant was devised for it by Whittle’s company, Power Jets Ltd.

Contradictory predictions of its performance were made by Miles and the RAE, through differences in their estimations of fuselage drag. A new evaluation suggests that the available information would have shown the aircraft to be capable of providing vital aerodynamic data for the transonic and early supersonic regimes at a time when no other sources were available, though further engine development would be needed for it to reach its full potential.

The Royal Netherlands Air Force (RNLAF) conducts frequent operations of Chinook helicopters with external slung loads. The normal RNLAF practice is for large external loads to be underslung by means of a two-strop suspension system backed up by a third point of suspension, the so-called ‘redundant HUSLE (Helicopter Underslung Load Equipment)’ comprising a redundant set of slings which come into action if one of the normal strops fails. The redundant HUSLE is relatively expensive in terms of time and operating costs. The present work has investigated the behaviour of a Chinook helicopter with an underslung load following failure of a front or rear strop, to determine whether the three-point suspension system can be safely replaced by a two-point suspension, eliminating the redundant HUSLE. The paper will demonstrate that although in general the redundant HUSLE results in a less violent helicopter reaction to a cable failure, it does not necessarily guarantee safety. It is concluded that flying with loads up to around two tonnes could be done safely with a two-point suspension system. A novel feature of this analysis of helicopter operations with slung loads is the integrated approach used for simulating pilot-in-the-loop load failures using a full non-linear model.

The layout design of a satellite module is a complex mechanical layout problem. Its main difficulties lie in combinatorial explosion of computational complexity, engineering complexity, and applicability in engineering practice. Inspired by the human-computer cooperation ideas, a human–computer co-operative co-evolutionary method for optimising layout design of a satellite module is developed. This method constructs the diversity reference set by using the diversity intelligence solutions (DIs) that are created by using the combinatorial operators of differential evolution (DE) and the blend crossover operator (BLX-a). During the co-evolution process of the presented method, the AIs, the DIs and the algorithm solutions are expressed by unified encoding strings and incorporated together to create new co-operative solutions. An instance of a satellite module layout design is presented to demonstrate the feasibility and effectiveness of the proposed method. Compared with the co-evolutionary approach and the all-at-once optimisation approaches, computational results show that the proposed method not only can produce better solutions, but also can better balance the conflicting objectives on the trade-off issues.

In this numerical study, the role of hyper-mixers on supersonic mixing is investigated for six different strut configurations. To this end, 3D, compressible, turbulent, non-reacting flow calculations with air as the secondary injectant have been carried out. A qualitative comparison of the predictions with experimental results is made through Schlieren and Mie scattering images. A quantitative evaluation of the predictions is made by comparison with experimentally measured exit stagnation pressure, wall static pressure and the degree of unmixedness. Based on these results, three strut configurations have been selected for carrying out simulations with hydrogen as the injectant. Results from the hydrogen simulations are compared with the predictions using air and also across the strut configurations. The results clearly demonstrate that castellated strut configurations are very effective in enhancing mixing in such high speed flows.

We revisit our earlier study of targeted energy transfer (TET) mechanisms for aeroelastic instability suppression by employing time-domain nonlinear system identification based on the equivalence between analytical and empirical slow flows. Performing multiscale partitions of the dynamics directly on measured (or simulated) time series without any presumptions regarding the type and strength of the system nonlinearity, we derive nonlinear interaction models (NIMs) as sets of intrinsic modal oscillators (IMOs). The eigenfre-quencies of IMOs are characterised by the ‘fast’ dynamics of the problem and their forcing terms represent slowly-varying nonlinear modal interactions across the different time scales of the dynamics. We demonstrate that NIMs not only provide information on modal energy exchanges under nonlinear resonant interactions, but also directly dictate robustness behaviour of TET mechanisms for suppressing aeroelastic instabilities. Finally, we discuss the usefulness of NIMs in constructing frequency-energy plots that reveal global features of the dynamics to distinguish between different TET mechanisms and to study robustness of aeroelastic instability suppression.

When defining a new product like an aircraft, space access vehicle or space mission, the Advanced Projects Group evaluates the available design space and compares it with the design space required to accomplish the specified mission. As with any product development process, the general life-cycle characteristics are established first during the conceptual design (CD) phase, clearly before a design proposal can be released to the follow-on design phases such as preliminary design (PD), detail design (DD), flight test (FT), and finally operation and disposal. As a rule of thumb, it can be assumed that around 80% of the flight vehicle configuration and mission tandem are determined during the CD phase alone, which is the key phase where the initial brainstorming has to take place. Clearly, it is the responsibility of the CD team to simulate the entire life-cycle of the project from ‘cradle to grave’ where the focus is on correctness rather accuracy in order to identify the design space and offer an overall proof of design convergence. Currently, the important primary aerospace vehicle and mission design decisions at CD level are still made using extremely simple analysis and heuristics. A reason for this scenario is the difficulty in synthesising the range of individual design disciplines for both, classical and novel aerospace vehicle conceptual designs, in more than an ad hoc fashion. Although the CD segment is seen as the most important step in the product development phase due to its pre-defining function, it is the least well understood part of the entire product evolution process due to its level of abstraction. This paper presents the roadmap towards the next generation of aerospace life-cycle synthesis systems, a software and management process capable to immediately calculate cost and time implications while simultaneously linking design, manufacturing, testing, and operation. A historical review of how design has been accomplished until today is presented. The design approaches are categorised and the characteristics of today’s state-of-the-art design synthesis systems are discussed. A specification for the new class of intelligent generic design synthesis systems is presented capable of satisfying the demands imposed by the new breed of high-performance aircraft, space access vehicles, space missions, and others. Finally, the development status of the next generation aerospace vehicle design synthesis (AVDS-PrADO) simulation-based acquisition environment is presented.

This paper presents the results of experimental and computational investigations on the effect of slot cavities on a supersonic projectile. Experimental work was carried out to show the effects of the slots on the drag at Mach numbers M = 1·36, 1·65, 1·83. The computational analysis was done at M = 1·36. A single configuration of the slot pattern was used with two different slot widths (0·5mm and 2·0mm). Flow features were investigated in the slotted area and at the base. The analysis presented includes the pressure distribution, the supersonic cavity flow and the effect of the slots on the overall aerodynamic drag. Unlike the case of slots at transonic speeds, the suction and blowing mechanism is not found at supersonic Mach numbers. Streamwise cavity slots cause a small base drag reduction. The reduction in total drag is modest when the width is 0·5mm. However, the experiments showed that with the wider slots (2mm) the drag actually increases at Mach numbers from 1·36 to 1·83.

This paper reviews recent results on the mechanics and aerodynamics of perching in a large bird of prey, the Steppe Eagle Aquila nipalensis. Data collected using onboard and high-speed video cameras are used to examine gross morphing of the wing planform by the flight muscles, and smaller-scale morphing of the wing profile by aeroelastic deflection of the feathers, Carruthers et al. High-resolution still images are used to reconstruct the shape of the wing using multi-station photogrammetry, and the performance of the measured wing profile is analysed using a panel code, Carruthers et al. In bringing these lines of research together, we examine the role of aeroelastic feather deflection, and show that the key to perching in birds lies not in high-lift aerodynamics, but in the way in which the wings and tail morph to allow the bird to transition quickly from a steady glide into a deep stall.

Shock-wave/turbulent boundary-layer interactions (SWTBLIs) with separation are known to be inherently unsteady but their physical mechanisms are still not totally understood. An experimental investigation has been performed in a supersonic wind tunnel at a freestream flow Mach number of 2·42. The interaction between a shock wave created by a shock generator (α = 3°, α = 9°, α = 13° and α = 15° deflection angles) and a turbulent boundary layer with thickness δ = 5mm has been studied. High-speed Schlieren visualisations have been obtained and used to measure shock wave unsteadiness by means of digital image processing. In the interactions with separation, the reflected shock’s unsteadiness has been in the order of 102Hz. High-speed wall pressure measurements have also been obtained with fast-response micro-transducers along the interactions. Most of the energy of the incoming turbulent boundary layer is broadband and at high frequencies (>104Hz). An addition of low-frequency (<104Hz) fluctuation energy is found at separation. Along the interaction region, the shock impingement results in an amplification of fluctuation energy due to the increase in pressure. Under the main recirculation region core there is only an increase in high frequency energy (>104Hz). Amplification of lower frequency fluctuation energy (>103Hz) is also observed close to the separation and reattachment regions.

A forward facing spike attached to a hemispherical body significantly changes its flow field and influences aerodynamic drag and wall heat flux in a high speed flow. The dynamic pressure in the recirculation area is highly reduced and this leads to the decrease in the aerodynamic drag and heat load on the surface. Consequently, the geometry, that is, the length and shape of the spike, has to be simulated in order to obtain a large conical recirculation region in front of the blunt body to get beneficial drag reduction. It is, therefore, a potential candidate for aerodynamic drag reduction for a future high speed vehicle. Axisymmetric compressible laminar Navier-Stokes equations are solved using a finite volume discretisation in conjunction with a multistage Runge-Kutta time stepping scheme. The effect of the spike length and shape, and the spike nose configuration on the reduction of drag is numerically evaluated at Mach 6 at a zero angle-of-attack. The computed density contours agree well with the schlieren images. Additional modification to the tip of the spike to get different types of flow field such as the formation of a shock wave, separation area and reattachment point are examined. The spike geometries include the conical spike, the flat-disk spike and the hemispherical disk spike of different length to diameter ratios attached to the blunt body.

For aerodynamic design optimisation as well as for multidisciplinary design optimisation studies, it is very desirable to limit the number of the geometric design variables. In Ref. 1, a ‘fundamental’ parametric aerofoil geometry representation method was presented. The method included the introduction of a geometric ‘class function/shape function’ transformation technique, CST, such that round nose/sharp aft end geometries as well as other classes of geometries could be represented exactly by analytic well behaved and simple mathematical functions having easily observed physical features. The CST method was shown to describe an essentially limitless design space composed entirely of analytically smooth geometries. In Ref. 2, the CST methodology was extended to more general three dimensional applications such as wing, body, ducts and nacelles. It was shown that any general 3D geometry can be represented by a distribution of fundamental shapes, and that the ‘shape function/class function’ methodology can be used to describe the fundamental shapes as well as the distributions of the fundamental shapes. A number of applications of the ‘CST’ method to nacelles, ducts, wings and bodies were presented to illustrate the versatility of this new methodology. In this paper, the CST method is extended to include geometric warping such as variable camber, simple flap, aeroelastic and flutter deflections. The use of the CST method for geometric morphing of one geometric shape into another is also shown. The use of CST analytic wings in design optimisation will also be discussed.

A NACA 641-412 aerofoil with circle and star damage and also three repair configurations has been numerically investigated. Two different methods of mesh generation were employed: multi structured mesh for the star damaged aerofoil and unstructured mesh for the other aerofoils. The results show that the damage will cause a reduction in lift coefficient of the aerofoil and also a different stall angle relative to that of the undamaged aerofoil. Each kind of repair improves the aerodynamic characteristics of the aerofoil considerably. The flow Field inside the damage hole and the cavity caused by the repair sheets was also investigated. Finally, the numerical solution was qualitatively and quantitatively validated using the available experimental results.