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.

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.

The past century has witnessed the rise and maturity of the flying machine, starting with the Wright brothers flyer to today’s modern passenger aircrafts and warfighters. At the start of this century, yet another achievement in flying vehicle technology was seen with the launch of the Boeing 787 aircraft, which has a significant portion by weight of polymer matrix fibre composites. This paper, therefore, addresses the effects of the manufacturing process of fibre reinforced polymer matrix composites on mechanical performance. Computations are carried out using the Finite Element (FE) method at the microscale where Representative Volume Elements (RVEs) are analysed with Periodic Boundary Conditions (PBCs). Straight fibre pre-preg-based composites and textile composites are considered. The commercial code ABAQUS is used as the solver for the FE equations, supplemented by user-written subroutines. The transition from a continuum to damage/failure is effected by using the Bažant-Oh crack band model, which preserves mesh objectivity. Results are presented for RVEs that are first subjected to curing and subsequently to mechanical loading. The effect of the fibre packing randomness on the microstructure is examined by considering multifibre RVEs where fibre volume fraction is held constant but with random packing of fibres. Plain weave textile composites are also cured first and then subjected to mechanical loads. The possibility of failure is accommodated throughout the analysis – failure can take place during the curing process even prior to the application of in-service mechanical loads. The analysis shows the differences in both the cured RVE strength and stiffness, when cure-induced damage has and has not been taken into account.

In connection with gliding flight enough measurements of “ lift ” and “ drag ” have been made to enable us to calculate the conditions for success of an aeroplane fitted with wings of standard sections; but no attempt has been made, as far as the present writer is aware, to ascertain what would happen if a flying machine were fitted with wings of standard section and these were flapped in a rhythmical manner. Would it support and propel itself? Several authors, including M. F. Fitzgerald and Colonel J. D. Fullerton in the previous issue of this journal, have discussed various portions of this problem; but instead of appealing to wind-tunnel determinations the latter has used such expressions as pSUV or pSUV2 for the pressure on a wing of area S, where p is the density of the air and V, U are the component velocities along and at right angles to its surface.

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.

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.

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.

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.

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.

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.

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.

The many methods by which Pontryagin's Maximum Principle is applied in optimal control problems can be divided into two groups, termed direct and indirect. The indirect methods use the conditions required for mathematical optimality as the starting point and attempt to satisfy the boundary conditions. The direct methods on the other hand satisfy the equations of motion and boundary conditions and then attempt to improve the performance index. It has been reported that the direct methods, while normally converging from very bad initial assumptions, have the inherent disadvantage of very slow convergence in the neighbourhood of the optimal solution and, so, complicated second order methods have recently been devised.

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.

In this paper, a morphing carbon fibre composite aerofoil concept with an active trailing edge is proposed. This aerofoil features of camber morphing with multiple degrees of freedom. The shape morphing is enabled by an innovative structure driven by an electrical actuation system that uses linear ultrasonic motors (LUSM) with compliant runners, enabling full control of multiple degrees of freedom. The compliant runners also serve as structural components that carry the aerodynamic load and maintain a smooth skin curvature. The morphing structure with compliant truss is shown to exhibit a satisfactory flexibility and loading capacity in both numerical simulations and static loading tests. This design is capable of providing a pitching moment control independent of lift and higher L/D ratios within a wider angle-of-attack range. Such multiple morphing configurations could expand the flight envelope of future unmanned aerial vehicles. A small prototype is built to illustrate the concept, but as no off-the-shelf LUSMs can be integrated into this benchtop model, two servos are employed as actuators, providing two controlled degrees of freedom.

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.

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.

This paper presents an open framework, through which, conventional general purpose aerial munitions can be converted into smart munitions. The retrofit consists of a smart adaptation kit (SAK) having a dedicated Guidance and Control Module (GCM). The adaptation kit along with the GCM ensures that the SAK glide optimally towards the designated target. To reduce cost, the number of control surfaces of the SAK has been kept to a bare minimum, which resulted in an under actuated system. The methodology proposed utilises the theory of gain-scheduled control and leads to an efficient procedure for the design of the controllers, which accurately track reference trajectories defined in an inertial reference frame. The paper illustrates the application of this procedure to the design of stabilisation and tracking controller for the SAK. The design phase is summarised, and the performance of the resulting controllers is assessed in simulation using dynamic model of the vehicle. Simulation results show that apart from improved circular error probable (CEP) of hitting the target, munition ballistic range has also significantly increased with the proposed modification.

A method of deriving a dynamic stiffness matrix for any non-uniform beam is presented. In particular, the case of a linearly tapered cantilever beam is considered, and excellent results are found with the use of only a few elements.

Corrugated structures offer a potential solution for morphing wing skin applications due to their anisotropic behaviour that allows chordwise camber and length changes. Aerofoils with corrugated skins in the aft 1/3 of the chordwise section have been studied experimentally and computationally using various corrugation shapes and forms (sinusoidal, trapezoidal and triangular) at different Reynolds numbers. The study showed that the aerodynamic performance is highly dependent on corrugation amplitude, wavelength, gradient (combination of amplitude and wavelength) and Reynolds number. Evidence is given highlighting that penalties for having a non-smooth surface in the aft 1/3 of the chordwise section of an aerofoil can be eliminated for the lift curve slope and minimised for the zero lift drag coefficient.