This paper presents performance analysis and design of ducted propellers for lighter-than-air vehicles. High-fidelity computational fluid dynamics simulations were first performed on a detailed model of the propulsor, and the results were in very good agreement with available experimental data. Additional simulations were performed using a simplified geometry, to quantify the effect of the duct and of the blade twist on the propeller performance. It was shown that the duct is particularly effective at low flight speed and that the blades with relatively high twist have better performance over the flight envelope. Design of the optimal twist distribution and of the duct shape was also attempted by coupling the flow solver with a quasi-Newton optimisation method. Flow gradients were computed by solving the discrete adjoint of the Reynolds-averaged Navier-Stokes equations using a fixed-point iteration scheme or a nested Krylov method with deflated restarting. The results show that the ducted propeller propulsive efficiency can be increased by 2%.

The paper discusses the development and application activities within the Aero-Physics and Measurement Technology Laboratory at the University of Manchester on pressure sensitive paint and thermographic phosphor optical imaging systems for gas dynamic applications. It provides a brief review of the basic principles, fundamental theory, properties, chemical characteristics and bonding technologies associated with the two systems. A number of case studies are presented, which exhibit the range of applicability, limitations and potential for further development of the technologies.

As civil aviation expands, environmental aspects and fuel savings are becoming increasingly important. Amongst technologies proposed for more efficient flight, air-to-air refuelling (AAR), ‘hopping’ and flying in close formation (drag reduction), all have significant possibilities. It will be interesting to know also how these technologies may co-exist e.g. AAR and formation flying.

In military use, AAR is virtually indispensable. Its benefits are real and largely proven in hostile and demanding scenarios. We present a case for applying AAR in a civil context to show that substantial reductions in fuel burn for long-range missions are achievable. Overall savings, including the fuel used during the tanker missions, would be of the order of 30-40% fuel and 35-40% financial. These are very significant in terms of the impact on aviation’s contribution to reducing atmospheric pollution.

AAR allows smaller, efficient (greener) aircraft optimised for about 3,000nm range to fulfil long-range route requirements. This implies greater usage of smaller airports, relieving congestion and ATC demands on Hub airports. Problems due to shed vortices and wakes at airports are reduced. Smaller engines will be needed.

Integrated (accepted) AAR could lead to further benefits. Aircraft could take-off ‘light’, with minimum fuel and reserves and a planned AAR a few minutes into the flight. The ‘light’ aircraft would not require over-rating of the engines during take-off and would therefore be less noisy during take-off and climb-out, permitting more acceptable night operations.

The availability of civil AAR will enable opportunities for hitherto borderline technologies to be utilised in future aircraft. Laminar flow will provide fuel savings and increased efficiency in its own right but could be significantly enhanced within a civil AAR environment. Similarly, supersonic transport may become an acceptable economic option.

AAR affords the possibility of a complete widening of the design space and this should appeal to the imagination of current and future designers.

The objective of our paper is to demonstrate the use of spectral/hp element technology in unravelling global flow instability mechanisms. Understanding these mechanisms is central to devising flow control approaches based on theoretically-founded physical principles. Global instability theory is concerned with prediction and control of linear and nonlinear disturbances developing in flows that are inhomogeneous in more than one spatial direction. As such, this theory encompasses the classic analysis of Tollmien which is valid for simple geometries of academic interest (e.g. a flat-plate) and therefore broadens the scope of this well-established but simplified methodology to include realistic problems encountered in aeronautical engineering. Compared with a direct numerical simulation approach, global instability theory can be used to explore efficiently far wider parameters ranges and deliver physical information to be used as handle for effective flow control.

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.

Various types of micro-vortex generators (μVGs) are investigated for control of a supersonic turbulent boundary layer subject to an oblique shock impingement, which causes flow separation. The micro-vortex generators are embedded in the boundary layer to avoid excessive wave drag while still creating strong streamwise vortices to energise the boundary layer. Several different types of µVGs were considered including micro-ramps and micro-vanes. These were investigated computationally in a supersonic boundary layer at Mach 3 using monotone integrated large eddy simulations (MILES). The results showed that vortices generated from μVGs can partially eliminate shock induced flow separation and can continue to entrain high momentum flux for boundary-layer recovery downstream. The micro-ramps resulted in thinner downstream displacement thickness in comparison to the micro-vanes. However, the strength of the streamwise vorticity for the micro-ramps decayed faster due to dissipation especially after the shock interaction. In addition, the close spanwise distance between each vortex for the ramp geometry causes the vortex cores to move upwards from the wall due to induced upwash effects. Micro-vanes, on the other hand, yielded an increased spanwise spacing of the streamwise vortices at the point of formation. This resulted in streamwise vortices staying closer to the floor with less circulation decay, and the reduction in overall flow separation is attributed to these effects. Two hybrid concepts, named ‘thick-vane’ and ‘split-ramp’, were also studied where the former is a vane with side supports and the latter has a uniform spacing along the centreline of the baseline ramp. These geometries behaved similar to the micro-vanes in terms of the streamwise vorticity and the ability to reduce flow separation, but are more physically robust than the thin vanes.

Four buckling problems are considered, namely: —

• (a) A rectangular orthotropic plate, with all edges simply supported, subjected to compression on its ends and a known compression or tension on its sides;

• (b) the same as (a) but with the ends clamped and sides simply supported;

• (c) a rectangular orthotropic plate, with the ends simply supported and sides clamped, subjected to compression on its ends; and

• (d) the same as (c) but with all edges clamped.

In each case it is shown that the variables involved can all be combined into such a non-dimensional form that a single curve serves to give the value of the end compression required to cause buckling. This curve is identical with the curve of buckling stress coefficient against side ratio for a corresponding isotropic plate under uni-axial compression.

A cartesian cut cell method for static body problems was presented in Part A (pp 47-56). Here, we extend the method to unsteady compressible flows involving arbitrarily moving bodies. The moving bodies are allowed to cross a stationary background cartesian mesh. So problems such as mesh distortion and/or restrictions on body motion which may affect other mesh approaches do not occur.

A MUSCL-Hancock finite volume scheme has been modified for moving boundary problems. The upwind fluxes on the interfaces of static cells are updated using an HLLC approximate Riemann solver and an exact Riemann solution for a moving piston is used to update moving solid boundaries (faces). A cell merging technique has been developed to maintain numerical stability in the presence of arbitrarily small cut cells and to retain strict conservation at moving boundaries.

The method has been validated against some well-known two dimensional test problems and applied to practical examples involving an exploding pressure vessel and a store release into a Mach 1.5 stream.

Due to constraints of natural condition, cost and of available time associated with model fabrication and for extensive wind-tunnel tests or flight tests, Computational Fluid Dynamics (CFD) simulation was considered an alternative means of providing air vehicle icing simulation and aeromechanic performance analysis. Full-scale icing experiments and, therefore, certification and cost can be significantly reduced by developing full-numerical simulation methods to evaluate the air vehicle performance for a wide range of icing conditions. This paper summarises helicopter icing simulation methods that include the development of helicopter aerodynamics, calculation methods of helicopter icing, icing protection system performance, icing effects on the helicopter performance, and some challenges in helicopter icing simulation.

For all who are specialists of one kind or another, it is a very healthy exercise to read books written by scientists or engineers of the calibre and breadth of F. W. Lanchester. He was at the same time truly a pioneer of flight, an expert in fluid mechanics, in flight mechanics, in the practical, and (well ahead of his time) in the military applications of flight, in the novel use of materials, and an inventor of devices and mechanisms. However, it was with particular surprise that I found, on studying the literature under the name F. W. Lanchester, a book published in 1935 entitled Relativity and two earlier papers on the subject published in 1921. Lanchester had met Minkowski in 1908 at the house of his friend Carl Runge, the applied mathematician, at Göttingen, and following a discussion with Minkowski had taken such an interest in the field that he was able to make these early English contributions.

Perhaps the most important property that makes such pioneers is the courage to think truly new thoughts. His fascination with the new philosophy in physics may be easily understood in terms of this bold style of thought.

Currently, there is great emphasis, worldwide, on environmental issues. This will have an impact on civil aircraft design, manufacture and operation.

Since the advent of the jet engine and swept wing aircraft, the trends have naturally tended towards greater productivity through increasing speed and payload. The cruise speed of conventional civil aircraft is unlikely to increase beyond current levels. Further increases in productivity are achieved by increasing payloads. This has led towards larger aircraft with the capability for increased ranges. It is shown that designing aircraft for longer ranges increases fuel burn significantly.

A series of aircraft operational parameters have been analysed. Selected data and established trends for current and future aircraft are presented. The data has been interpreted into efficiency terms, relating payload, range, fuel consumed and a measure of unit costs. It is shown that ‘value’ (cost) and noise effective efficiencies decrease dramatically with increasing range.

Environmental and economic considerations, in the future, may well demand greater efficiency in preference to productivity. One solution for long-range services is to use short-range hops. Another is via air-to-air refuelling. This will be addressed, in more detail, in a future paper.

The drag coefficient and the laminar-to-turbulent transition for the aerofoil component of a wing model are optimised using an adaptive upper surface with two actuation points. The effects of the new shaped aerofoils on the global drag coefficient of the wing model are also studied. The aerofoil was optimised with an ‘in-house’ genetic algorithm program coupled with a cubic spline aerofoil shape reconstruction and XFoil 6.96 open-source aerodynamic solver. The wing model analysis was performed with the open-source solver XFLR5 and the 3D Panel Method was used for the aerodynamic calculation. The results of the aerofoil optimisation indicate improvements of both the drag coefficient and transition delay of 2% to 4%. These improvements in the aerofoil characteristics affect the global drag of the wing model, reducing it by up to 2%. The analyses were conducted for a single Reynolds number and speed over a range of angles of attack. The same cases will also be used in the experimental testing of the manufactured morphing wing model.

Supersonic civil aircraft present a unique noise certification challenge. High specific thrust required for supersonic cruise results in high engine exhaust velocity and high levels of jet noise during take-off. Aerodynamics of thin, low-aspect-ratio wings equipped with relatively simple flap systems deepen the challenge. Advanced noise reduction procedures have been proposed for supersonic aircraft. These procedures promise to reduce certification noise levels, but they may require departures from normal reference procedures defined in noise regulations. The subject of this article is a take-off performance and noise assessment of a notional supersonic business jet. Analytical models of an airframe and a supersonic engine derived from a contemporary subsonic turbofan core are developed. These models are used to predict take-off trajectories and certification noise levels. Results indicate advanced take-off procedures are helpful in reducing noise along lateral sidelines.

The mixing promoting capability of right-angled triangular tab with sharp and truncated vertex has been investigated by placing two identical tabs at the exit of a Mach 2 axi-symmetric nozzle. The mixing promoting efficiency of these tabs have been quantified in the presence of adverse and marginally favourable pressure gradients at the nozzle exit. It was found that, at all levels of expansion of the present study though the core length reduction caused by both the tabs are appreciable, but the mixing caused by the truncated tab is superior. The mixing promoting efficiency of the truncated tab is found to increase with increase of nozzle pressure ratio (that is, decrease of adverse pressure gradient). For all the nozzle pressure ratios of the present study, the core length reduction caused by the truncated vertex tab is more than that of sharp vertex tab. As high as 84% reduction in core length is achieved with truncated vertex right-angled triangular tabs at moderately overexpanded level, corresponding to expansion level pe/pa = 0·90. The corresponding core length reduction for right-angled triangular tabs with sharp vertex and rectangular tabs are 65% and 31%, respectively. The present results clearly show that the mixing promoting capability of the triangular tab is best than that of rectangular tabs at identical blockage and flow conditions.

The fundamental frequencies of flexural vibration are determined for uniform isotropic rectangular plates that have free edges and pinpoint supports at the four corners. For the particular case of a square plate the lowest five frequencies and mode shapes have been determined. The second frequency is most unusual, since an infinite number of mode shapes exist with identical frequencies. Finite difference expressions, which simplify the treatment of the free boundaries for definite values of Poisson's ratio, are used in conjunction with extrapolation procedures to obtain the approximate solutions.

The current flight management system (FMS), CMA-9000, from CMC Electronics-Esterline, only optimises the vertical flight profile in terms of the speed of the aircraft. This article defines a methodology that optimises the speeds and altitudes for the vertical profile, obtaining a trajectory that reduces the global flight cost.

The performance database (PDB) provided by CMC Electronics-Esterline is presently used on aircraft such as the Lockheed L-1011, the Airbus A310 and the Sukhoi Superjet 100 Russian regional jet. The PDB is used as the reference to design different trajectory optimisation algorithms to obtain the altitude where the aircraft fuel efficiency is the best. These algorithms are compared with the part-task trainer (PTT), simulator that represents the FMS CMA-9000, supplied by CMC Electronics-Esterline as well.

To validate the results, the FlightSIM® software is used, which considers a complete aircraft aerodynamic model for its simulations, giving accurate results and very close to reality.

Shockwaves are present in a flow as soon as the Mach number becomes supersonic. Being viscous phenomena, Shockwaves are a source of drag which can be predominant when the Mach number is significantly higher than one. In supersonic air intakes, the production of entropy by shocks is felt as a loss in efficiency. At high Mach numbers, Shockwaves produce a considerable temperature rise leading to severe heating problems, complicated by real gas effects. The intersection - or interference - of two shocks gives rise to complex wave patterns containing slip-lines and associated shear layers whose impingement on a nearby surface can cause detrimental pressure and heat transfer loads. The impact of a Shockwave on a boundary layer is the origin of strong viscous interactions which remain a limiting factor in the design of transonic wings, supersonic air intakes, propulsive nozzles and compressor cascades. More effort is needed to improve prediction of these interactions and to devise new techniques to control such phenomena.

A computational study is presented, which examines the performance of variants of second-moment closure and non-linear eddy-viscosity models when used to predict attached and separated flows over a high lift aerofoil for a range of incidence angles. The capabilities of both model types, especially in respect of resolving the onset of suction-side separation at high incidence, are contrasted with those of a low-Re k-ε model based on the linear Boussinesq stress-strain relationship. The second-moment model contains a conventional linear approximation of the pressure straining process; a cubic (realisable) variant has been investigated in an earlier study and found to offer no advantages. The quadratic eddy-viscosity model features coefficients which are sensitised to the strain and vorticity invariants. While both models, in the form originally proposed, are superior to the linear eddy-viscosity variant, neither performs well in respect of resolving separation, unless modified so as to return the requisite low level of shear stress in the boundary layer approaching separation. Once separation is resolved with sufficient realism, the near wake aft of the trailing edge is also well represented. All models return poor representations of the far wake which is characterised by low levels of turbulence production to dissipation ratio.

A method for calculating the growth of a turbulent wall jet in streaming flow has been developed. The flow is assumed to be two-dimensional, incompressible and over a plane, smooth wall. Downstream variations of pressure are permitted and separation in an adverse pressure gradient may be predicted. The method incorporates procedures for matching the flow to that at the blowing slot, although it is postulated that the upstream boundary layer there is thin enough that the wall jet develops without an unmixed wake (i.e. there is not a minimum in the mean-velocity profile).

The method incorporates four integral momentum equations taken from the wall to various points in the flow. The calculation of the outer shearing stress, although empirical, is based on the large-eddy equilibrium hypothesis and therefore has some foundation. The remaining empiricism in the method is based on measurements in self-preserving wall jets.

The method has been used to predict the jet-momentum coefficient required to suppress separation over a trailing-edge flap attached to a thin aerofoil. Plausible curves have been obtained Using assumed values of upstream boundary layer at the slot. Of some practical interest is the indication that large savings in power are possible if the upstream boundary layer is removed. This indicates that blowing combined with upstream suction, or multiple-slot blowing, may give useful savings in the application of blowing to prevent separation.

Evidence is presented to show that the universal law of the wall has a wider range of validity than the assumption l = ky, with k a universal constant. If an effective value of k is defined for the wall region its value is shown to vary between wide limits, and keff can be correlated with other parameters describing the flow in the wall region.