A new combustor concept referred as the trapped vortex combustor (TVC) employs a vortex that is trapped inside a cavity to stabilise the flame. The cavity is formed between two axisymmetric disks mounted in tandem. TVC offers many advantages when compared to conventional swirl stabilisers. In the present work, numerical investigation of cold flow (non-reacting) through trapped vortex combustor is performed. The numerical simulation involves passive flow through TVC to obtain an optimum cavity size to trap stable vortices inside the cavity and to observe the important characteristics of TVC. One of the main objectives is to evaluate various two equation turbulence models for the aerodynamic predictions of TVC. Commercial CFD software Fluent is used for the present study. In addition to many models available, Non-linear k-ω and modified k-ω models are incorporated through user defined functions. Results obtained include streamlines, residence time and entrainments for all models. The reattachment length obtained by non-linear k-ω model closely matches with that obtained by DNS in the case of forebody-spindle alone. Non-linear k-ω model alone captures the corner vortices while all the other models failed to capture. From the entrainment characteristics study, it is inferred that the primary air needs to be injected for accommodating the decrease in oxidizer inside the cavity to obtain better performance from the TVC.

A review of morphing concepts with a strong focus on morphing skins is presented. Morphing technology on aircraft has found increased interest over the last decade because it is likely to enhance performance and efficiency over a wider range of flight conditions. For example, a radical change in configuration, i.e. wing geometry in flight may improve overall flight performance when cruise and dash are important considerations. Although many morphing aircraft concepts have been elaborated only a few deal with the problems relating to a smooth and continuous cover that simultaneously deforms and carries loads. It is found that anisotropic and variable stiffness structures offer potential for shape change and small area increase on aircraft wings. Concepts herein focus on those structures where primary loads are transmitted in the spanwise direction and a morphing function is achieved via chordwise flexibility. To meet desirable shape changes, stiffnesses can either be tailored or actively controlled to guarantee flexibility in the chordwise (or spanwise) direction with tailored actuation forces. Hence, corrugated structures, segmented structures, reinforced elastomers or flexible matrix composite tubes embedded in a low modulus membrane are all possible structures for morphing skins. For large wing area changes a particularly attractive solution could adopt deployable structures as no internal stresses are generated when their surface area is increased.

Blade-element models are the most common models for the analysis of propeller aerodynamics, performance calculations and propeller design. In spite of their simplicity these models are very efficient and accurate. Blade-element models use the local induced velocities as an input thus they should be combined with another model in order to calculate these induced velocities. Various models are used for the calculation of the induced velocity, where the most popular ones include: momentum, simplified-momentum, lifting-line (prescribed and free wake), and vortex (McCormick and Theodorsen) models. The paper describes the various models, compares their results and discusses the advantages and disadvantages of each one. The results indicate that the Bladeelement/simplified-momentum model offers very good accuracy together with high efficiency. For propeller performance calculations during steady axial flight, where most of the cross-sections do not experience stall, detailed and complicated models for calculating the induced velocities do not show advantages over the simple bladeelement/simplified-momentum model,

In this paper, the results of a series of experiments funded by DARPA to determine the feasibility of using small actuators to provide directional control for a supersonic projectile are presented. Controlling the flight of the projectile was accomplished by taking advantage of complex shock-boundary-layer interactions produced by mechanical devices. Experimental tests were conducted at GTRI to screen several control concepts and actuator locations. Further experiments were conducted on a scale projectile in a supersonic stream to investigate the rise time of the forces. Several different mechanical actuators were tested which served to provide guidance for future actuator designs. CFD results were also used to predict the results in flight as well as gain insights into the fluid mechanics involved. Flight tests of a Mach 4 round proved the viability of the guidance actuator.

The current understanding of periodic transonic flow is reviewed briefly. The effects of boundary-layer transition, non-adiabatic wall conditions and modifications to the aerofoil surface geometry at the shock interactions on periodic transonic flow are discussed. Through the methods presented, it is proposed that the frequency of periodic motion can be predicted with reasonable accuracy, but there are limitations on the prediction of buffet boundaries associated with periodic transonic flows. Several methods have been proposed by which the periodic motion may be virtually eliminated, most relevantly by altering the position of transition fix, contouring the aerofoils surface or adding a porous surface and a cavity in the region of shock interaction. In addition, it has been shown that heat transfer can have a significant effect on buffet.

The free-wing tilt-body aircraft refers to a vehicle configuration in which the wing, fuselage, and empennage are in a longitudinally articulated connection. This allows the main wing to freely rotate relative to the body, while the empennage, which is in the form of a long twin boom connected to the rear of the body, changes its incidence angle relative to the body in response to external commands. The principal advantages claimed for the configuration are short takeoff and landing capability, and reduced gust sensitivity. The aerodynamics of the free-wing tilt-body configuration has been previously studied, but analysis of its flight mechanics is limited. In this paper we present derivations of the flight dynamic equations of motion using multi-body dynamic modelling techniques, and combine the resulting equations of motion with experimental aerodynamic data to achieve a nonlinear mathematical model for flight simulation of a generic free-wing tilt-body vehicle. The mathematical model is suitable for the study of detailed dynamic characteristics as well as for model based control law synthesis. Key flight performance, and stability and control characteristics of a generic configuration are obtained from the mathematical model.

Aviation emissions have an impact on the global climate, and this is consequently an active area of research worldwide. By adapting replicable and transparent systematic review methods from the field of evidence-based medicine, we aim to synthesise available data on the effects of aviation emissions on climate. From these data, we aim to calculate lower and upper bounds for estimates of the effect of aviation on climate in an objective manner.

For the systematic review an appropriate protocol was developed and applied by two independent reviewers, to identify research that met the inclusion criteria. These included all aviation types, original research studies, climate models with aviation as a specific component, with outcomes for emissions, radiative forcing, global warming potential and/or surface temperature changes. These studies were prioritised and data extracted using a standard process. The 35 studies reviewed here reported radiative forcing, global warming potential and/or temperature changes as outcomes, allowing direct comparisons to be made.

Tabulated results and a narrative commentary were provided for overall effects on climate, and the individual effects of carbon dioxide, water, contrails, cirrus clouds, ozone, nitrogen oxides, methane, soot and sulphur oxides. Lower and upper bounds for these effects, and their relative contributions compared to overall radiative forcing and surface temperature changes, have been described.

This review shows that the most recent estimates for the contribution of aviation to global climate are highly dependent on the level of scientific understanding and modelling, and predicted scenarios for social and economic growth. Estimates for the future contribution of aviation to global radiative forcing in 2015 range from 5·31% to 8·04%. For 2050 the estimates have a wider spread, from 2·12% to 17·33%, the latter being for the most extreme technology and growth scenario. These global estimates should be considered within the context of uncertainties in accounting for the direct and indirect effects of different contributions. Variations between lower and upper bounds for estimates of radiative forcing are relatively low for carbon dioxide, around 131% to 800% for cirrus clouds effects, and 1,044% for soot. Advances in climate research, particularly in the area of contrail and cloud effects, has led to some revision of the 1999 IPCC estimates(1), and demonstrates that the research community is actively working to further understand the underlying science.

The approaches assumptions, limitations and future work were discussed in detail. We have demonstrated how the systematic review methodology can be applied to climate science, in a replicable and transparent manner.

Solar powered aircraft are becoming more and more interesting for future long endurance missions at high altitudes, because they could provide Earth monitoring, telecommunications, etc. without any atmospheric pollution and, hopefully in the near future, with competitive costs compared with satellites. The research activities carried out till now have been mainly focused on flying wings or conventional aircraft configurations, with a great emphasis on the technological aspects. The present paper aims to define a mathematical model for solar powered aircraft preliminary design, valid independently of the aerodynamic configuration. A preliminary analysis is carried out in order to simulate Helios and the results are compared with those available from the flights of this aircraft. The proposed mathematical model is used also to compare four different aircraft configurations, namely: a flying wing, a conventional aircraft, a twin boom aircraft and a biplane aircraft. The results obtained are discussed in the paper and an optimum aircraft is analysed.

Handling qualities are expressed as requirements at the interface of the pilot and the machine. In this way, the key functionality questions facing the design engineer are seen from the perspective of the interaction of the human pilot with the aircraft system and the environment in which it operates. In this paper, the author takes a ‘virtual engineering’ approach to handling qualities, emphasising the importance of conducting ‘requirements capture’ and preliminary design as an iterative process. When stretched capabilities are required, this approach minimises the risk to finding appropriate technology solutions, through developing explicit relationships between capability and design parameters, thus facilitating fully informed trade studies and predictions. Case studies from the development of a civil tiltrotor aircraft are presented that show how the difficult challenges facing the designer first need to be structured in terms of HQ predictions and assignments. This then provides the basis on which handling qualities improvements can be constructed within the multidisciplinary context of rotorcraft engineering.

The recent period has been filled with exceptionally interesting developments and advances, resulting in high-performance conventional and non-conventional manned and unmanned aircraft. Although those vehicles seem to comply well with specific mission performance requirements, one is still confronted with an apparent weakness to reliably stabilise and control throughout the flight envelope. Since the provision of satisfactory stability and control characteristics invariably compromises flight performance, it becomes essential to identify and integrate performance-optimal stability and control design solutions early during the flight vehicle definition phase. In particular, the conceptual design of integrated control effectors for advanced aircraft is far from being trivial. Never before have we been presented with such tremendous wealth of specialised data and information suitable for detail design of controls. In contrast, never before has it been necessary to approach any one of the primary design disciplines still as entirely ad hoc and inconsistent as in the case of designing controls during the conceptual design phase. This need initiated the development of a configuration independent (generic) stability and control methodology capable of sizing primary control effectors of fixed wing subsonic to hypersonic designs of conventional and unconventional, symmetric and asymmetric configuration layouts. This paper summarises the methodology concept and demonstrates its versatility and validity by analyzing selected stability and control characteristics of the Northrop YB-49 flying wing.

Based on two-phase flow theory, an Eulerian method to simulate rime ice accretions on an aerofoil has been developed. The SIMPLE (semi-implicit method for pressure linked equations) algorithm on a collocated grid is employed to solve the governing equations for the airflow. In order to simulate droplets impinging on an aerofoil, a permeable wall is proposed to solve the governing equations for supercooled droplets. The collection efficiency and impingement limits are obtained from the droplets’ flowfield. The process of ice accretion is simulated using the assumption that ice accumulates layer-by-layer and the ice shape is predicted with the assumption that ice grows in the direction normal to the aerofoil surface. The rime ice accretions on a NACA0012 aerofoil at 0° and 4° angles-of-attack have been investigated and there is agreement between the simulated results and previously published experimental data. The change of the pressure coefficient along the iced aerofoil is also analysed.

Extensive experiments were conducted to study the effect of various parameters on the surface pressure distribution and transition point of an aerofoil section used in a wind turbine blade. In this paper details of the variation of transition point on the aforementioned aerofoil are presented. The aerofoil spanned the wind-tunnel test section and was oscillated sinusoidally in pitch about the quarter chord. The imposed variables of the experiments were free stream velocity, amplitude of motion, mean angle-of-attack, and oscillation frequency.

The spatial-temporal progressions of the leading-edge transition point and the state of the unsteady boundary-layer were measured using eight closely-spaced, hot-film sensors (HFS). The measurements show that:

(i) Reduced frequency has a pronounced effect on the variations of the transition point.

(ii) There exists a hysteresis loop in the dynamic transition location and its shape varies with the reduced frequency and mean angle-of-attack.

This paper celebrates the centenary of Queen Mary College’s involvement in aeronautics, a celebration with a unique distinction since it was this College’s immediate forebear which was the first British higher education institution to begin teaching and research in this subject. Thus the emphasis is on the early years from 1907 until the 1950s, a period ripe for recording before it recedes beyond living memory, but also the period during which the degree course in aeronautical engineering became firmly established and its parent Department acquired its reputation for research. Section 2.0 gives a brief history of the College’s origins in the East London College. Subsequent sections deal with the foundation of the aeronautical laboratory there, from which the aeronautical department grew, and the activities of the two men who led these developments, A.P. Thurston and N.A.V. Tonnstein who changed his name to Piercy.

Space transportation remains in the pioneering stages. What might this century bring if we had a ‘railroad to space’ that embodied the characteristics of the transcontinental undertaking? The X-33 and Venture Star projects were one attempt to achieve the characteristics of that transcontinental railroad. There are others, here and in other countries, but perhaps we need to begin with a smaller first step, a small, commercial reusable rocket with ballistic ascent to space altitude with a hypersonic glider return? Our challenge in space today is to develop vehicles that are in continuous use, maintained and operated on a fixed schedule despite weather or environmental hazards, which move payloads not only into space but back again. The X PRIZE was a $10 million prize awarded to Scaled Composites as the first privately financed spaceship that launched the equivalent of three persons to an altitude of at least 100 kilometers on two consecutive flights within two weeks. What about an analogous vehicle that flies two or three times a week, every week for a number of years? A major difference is that this challenge is to be accomplished without government support or government developed vehicles. The aerospace vehicle design (AVD) Laboratory team at the University of Texas at Arlington is developing a generic space access vehicle (SAV) design synthesis environment with focus on the conceptual design phase. The AVD Lab has applied elements of this toolbox to the study of a tourist aerospace vehicle under a grant from Rocketplane Limited, Inc. The development of a low-cost tourist vehicle based on the adaptation of a Learjet 25/35/45 series aircraft is the focus of this paper.

Particle image velocimetry (PIV) measurements of the flow around a wing section are employed as a basis for non-intrusive aerodynamic mean loads characterisation, providing sectional lift, drag and pitching moment. The technique relies upon the application of control-volume approaches in combination with the deduction of the pressure from the PIV experimental data through application of the momentum equation. The treatment can also be applied when the flow is unsteady; in that case time-mean loads are obtained from velocity statistics, through the use of Reynolds-averaged formulation of the governing equations. The procedure was applied in the experimental investigation of a NACA 642A015 aerofoil, in which the PIV approach is validated against standard pressure-based methods (surface pressure distribution and wake rake). The chord Reynolds number considered in the investigation ranges between 1 – 7 × 105. In addition, the consistency and potential performance of the method was assessed by means of synthetic velocity field data obtained from a numerical flow simulation.

As commissioned to demonstrate the feasibility of jet propulsion, the E28/39 needed to exceed the performance of contemporary fighters. But Carter, the chief designer, took the opportunity to look further ahead, and devised an aircraft in which the onset of compressibility effects was taken into account from the beginning of the design. Successful operation over a wide speed range required a shrewd synthesis of previous experience and practice with uncertain material emerging from the research domain. The resulting aircraft showed no significant aerodynamic vices, requiring only minor modifications from its first flight to its participation in diving trials, that took it into hitherto unexplored regions of high subsonic speed. It proved to be fully worthy of its pivotal role at the beginning of a new era in aeronautics.

The aerodynamic features of Carter’s design are reviewed in relation to the limited state of knowledge at the time. Drawing upon fragmented material, much not previously published, this study enlarges upon, and in places amends, previous accounts of this notable machine.

The first generation of supersonic commercial transportation has seen three serious attempts to arrive at an economically and environmentally viable aircraft. The US B2707-200/300 design was cancelled early before a prototype could emerge; the Russian Tu-144 design succeeded to become the first supersonic transport but spanned only a few years of restricted airline service; the Anglo-French Concorde endured more than 27 glamorous airline service years until the last of its species was retired on 30 August 2003. This first generation was followed by a second generation of supersonic commercial transport projects in the time period between 1986 until about 1999, designs which did not proceed towards the production hardware stage. This study critically examines the anatomy of two generations of supersonic commercial transport design failures and successes in order to arrive at lessons learned free of ‘wishful thinking’. The design conditions leading to the identification of the product ‘solution space’ for an economically and environmentally acceptable supersonic commercial transport are discussed. Having assembled an understanding of the product metrics valid for supersonic commercial transports, the paper then provides an outlook for the first generation of supersonic corporate and cargo jet projects. This first generation of supersonic business jet (SSBJ) and supersonic cargo jet (SSCJ) projects spans a period of nearly two decades of development, starting from 1988 until today. The present study identifies that the product development metrics of this class of aircraft is radically different compared to the metrics valid for supersonic commercial transports. The challenges in VIP transportation and dedicated freight transportation at supersonic speeds are portrayed leading to two principal trains of thought targeting the development of the first supersonic business jet and/or supersonic cargo jet hardware: the development based on a new airframe, and alternatively the development based on an existing airframe.

This paper presents the application of a relatively new technique of fluidic thrust-vectoring (FTV), named Co-flow, for a small gas-turbines. The performance is obtained via experiment and computational fluid dynamics (CFD). The effects of a few selected parameters including the engine throttle setting, the secondary air mass-flow rate and the secondary slot height upon thrust-vectoring performance are provided. Thrust vectoring performance is characterised by the ability of the system to deflect the engine thrust with respect to the delivered secondary air mass-flow rate. The experimental study was conducted under static conditions in an outdoor environment at Cranfield University workshop that was especially designed for this purpose. As part of this investigation, the system was modelled by CFD techniques, using Pointwise’s Gridgen software and the three-dimensional flow solver, Fluent. Also, Cranfield’s gas-turbine performance code (TurboMatch) was utilised to estimate boundary conditions for the CFD analysis with respect to the integrated nozzle. The presented technique is easy-to-use approach and offers better result for thrust-vectoring problems than previously published works. Experimental results do show the overall viability of the blowing slot mechanism as a means of vectoring the engine thrust, with the current configuration. Computational predictions are shown to be consistent with the experimental observations and make the CFD model a reliable tool for predicting Co-flow fluidic thrust-vectoring performance of similar systems.