The so-called ejector nozzle has received a large amount of experimental study as a means for achieving efficient supersonic expansion of the main jet without the difficulties of a mechanically variable convergentdivergent nozzle. No satisfactory theory for the calculation of its performance has yet been put forward and it is the object of this paper partly to remedy this defect.

A decade has elapsed since triangular elements under constant strain and stress were independently introduced within the matrix Displacement Method by the Turner group at Boeing and the Aeronautical Structures Department at Imperial College. They proved a considerable success in the practical analysis of complex membrane shapes as occur in modern aircraft wings. However, the original formulation based on a local cartesian system of axes and the classical concept of stress and strain inevitably leads to an unsymmetrical and aesthetically (as well as computationally) unfavourable expression for the stiffness matrix k. Moreover, the method was restricted to isotropic media. The necessity of extending the theory to finite displacements ultimately showed the inadequacy of the past approach and guided the writer to the creation of new concepts, e.g. the natural stiffness of an element. This natural stiffness, which ignores the rigid body motions and is hence of dimensions (3 X 3) for a triangle with three nodal points, relates in a concise manner the nodal force components parallel to the sides and the elongations of the sides themselves. A powerful tool in its derivation and further development was a new definition of stress and strain vectors which excludes shearing contributions and only works with direct components parallel to the sides. The idea has also been extended to tetrahedra and many standard elements which are incorporated as a basic feature in the structural language ASKA (Automatic System for Kinematic Analysis) developed at ISD by the Applied Programming Group under Dr. H. Kamel. These fundamental ideas were emphasised in the author's recent lecture to the Society.

In this paper lower bounds are obtained for the intensity of the sonic bang resulting from a slender aircraft flying straight and level at supersonic speeds. The lower bounds are obtained for an aircraft subject to various constraints on its total lift, volume, area distribution and centre-of-pressure position. The work described proves the results quoted in an earlier paper, corrects the results given in that paper for the general configuration, and extends the results to cover more constraints, mainly the centre-of-pressure position.

On unswept wings, or wings with small amounts of sweep, the favourable pressure gradient round the leading edge, where the flow is rapidly accelerated away from the stagnation line, ensures a certain amount of laminar flow, provided the wing surface is sufficiently smooth. On highly swept wings, however, it has been found that turbulent flow can exist on the attachment line itself and there are therefore no naturally occurring regions of laminar flow. This trouble arises from the turbulence at the root of the wing, which sweeps along the attachment line. If the Reynolds number of this turbulent attachment line boundary layer is greater than some critical value, the whole attachment line boundary layer remains turbulent and the complete wing is contaminated. But if the Reynolds number is below the critical value, the turbulence decays along the leading edge and the boundary layer on the attachment line reverts back to the laminar state. This situation arises when the leading edge radius is small and the wing is only slightly swept. The attachment line boundary layer Reynolds number, Rθ, is given by the following equation:

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.

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.

This paper will address the engineering performance differentiators for an F1 car and highlight the difference aerodynamics can make to that performance. It will also consider some basic aerodynamic challenges and the main tools used for aerodynamic exploration by teams.

The characteristics of the flow field around a circular cylinder in cross-flow placed at various distances from a plane, parallel both to the flow and to the cylinder axis, were analysed using a hot wire anemometer. Experiments were performed in a wind tunnel with Reynolds numbers ranging from 0.85×105 to 3×105. The spectra of the hot wire signals were obtained using a Fast Fourier Transform technique programmed on a PDP 11/40 computer. As regards a smooth cylinder, the main features of the vortex shedding mechanism in the subcritical regime remained unaltered for distances from the plane greater than approximately 0.4 diameters; in particular the Strouhal frequency did not show any significant variation relative to the typical value for an isolated cylinder. As for lower values of the distance from the plane, the regular vortex shedding disappeared and the hot wire spectra showed typical turbulent features. The possibility of obtaining supercritical conditions by roughening the cylinder surface was confirmed together with the importance of the Reynolds number based on the typical roughness size, Rk, in the evaluation of the flow regime around the cylinder. In the case of roughened cylinders, and with values of Rk below-350, the regular vortex shedding disappeared at a distance from the plane smaller than 0.3 diameters. This fact suggests that, at least in part of the supercritical regime, the influence of the plane can be smaller than in the subcritical regime.

Conventional commercial aircraft fuselages use all-aluminium semi-monocoque structures where the skin carries the external loads, the internal fuselage pressurisation and is strengthen using frames and stringers. Environmental and economic issues force aircraft designers to minimise weight and costs to keep air transport competitive and safe. But as metal designs have reached a high degree of perfection, extraordinary weight and cost savings are unlikely in the future. Carbon composite materials combined with lattice structures and the use of topology optimisation have the potential to offer such weight reductions. The EU FP7 project Advanced Lattice Structures for Composite Airframes (ALaSCA) was started to investigate this. This article present some of this research which has now led to the development of a new airframe concept which consists of: a load carrying inner skin; transverse frames; CFRP-metal hybrid stiffeners helically arranged in a grid configuration; insulating foam and an additional aerodynamic outer skin.

The behaviour of streams of solid particles when ejected into a fluid crossflow suggests that there are considerable forces caused by aerodynamic interaction between closely spaced particles. The overall effect is a dispersion of the solids normal to the fluid flow direction and a bunching up along the axis. This effect is apparent in aerial work where particulate solids are released from aircraft and also in dilute two phase flow especially at the onset of choking or saltation. The experimental work reported here contrives to measure these forces for spheres at the appropriate Reynolds numbers. The results are considered to be sufficiently accurate and comprehensive to give some understanding of the phenomena occurring and to permit an explanation, at least qualitatively, of the behaviour mentioned.

An analysis is presented for the post-buckling behaviour of simply-supported, laminated cross-ply rectangular plates subjected to biaxial compression. The solution to von Kármán-type large deflection equations of the plate is expressed as a double sine series for the transverse deflection and a double series of clamped-clamped beam functions for the force function. All the boundary conditions, including those involving the normal bending moment at the edges, are satisfied exactly. The series solution is found to converge rapidly. Using only the first few terms in the series, numerical results for square graphite-epoxy unsymmetric laminates under uniaxial compression is presented graphically.

This paper presents a method with strong convergence characteristics for the determination of eigenvalues and eigenvectors of continuous systems. The limitation on the number of undetermined constants in the displacement functions introduced by the conditions at the ends of a segment is removed by the introduction of points of freedom within the segment.

This improves the convergence of eigenvalues and eigenvectors very rapidly with the number of segments, especially in torsional vibration problems where the convergence with the usual Matrix Displacement method is very poor. The continuous medium is successively approximated by the use of sub-systems with finite numbers of degrees of freedom. The principles upon which the method is based and the convergence of the results are discussed and illustrated by a series of examples.

This paper reviews some of the research that has been carried out at the University of Liverpool where the Flight Science and Technology Research Group has developed its Heliflight-R full-motion research simulator to create a simulation environment for the launch and recovery of maritime helicopters to ships. HELIFLIGHT-R has been used to conduct flight trials to produce simulated Ship-Helicopter Operating Limits (SHOLs). This virtual engineering approach has led to a much greater understanding of how the dynamic interface between the ship and the helicopter contributes to the pilot's workload and the aircraft's handling qualities and will inform the conduct of future real-world SHOL trials. The paper also describes how modelling and simulation has been applied to the design of a ship's superstructure to improve the aerodynamic flow field in which the helicopter has to operate. The superstructure aerodynamics also affects the placement of the ship's anemometers and the dispersion of the ship's hot exhaust gases, both of which affect the operational envelope of the helicopter, and both of which can be investigated through simulation.

A number of modern “aerodynamic” bicycle wheels were tested at full scale speeds in a windtunnel. The effects of wheel geometry, rotational speed and cross winds on drag force, side force and yawing moment were investigated.

Compared with a conventional spoked wheel, all the “aerodynamic” wheels were found to give similar improvements in drag of approximately 25%. In cross wind conditions, aerodynamic characteristics were strongly dependent on wheel geometry, but essentially independent of wheel rotational speed.

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.

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.

The decay of a trailing vortex containing a turbulent core is examined according to the argument that the turbulence is partly sustained by, and interacts with, the irrotational flow through viscous diffusion of eddies across the core boundary. The region separating the core from the outer irrotational flow, in which the interaction between the turbulence and the mean flow occurs, expands with time like (time)1/2, characteristic of viscous diffusion.

On the basis of a simple representation of the radial distribution of mean circumferential velocity, the decay of the vortex is calculated and is found to behave in a way similar to what Squire predicted, on the assumption of a constant eddy viscosity, in 1954. The essential difference from Squire’s solution found here is that the apparent viscosity is not simply proportional to the initial circulation about the vortex as he proposed, but depends on the Reynolds number defined with respect to the circulation.

Wind tunnel and flight observations of vortex decay can be more closely reconciled than Squire’s theory allowed; those differences that exist are attributed to the vastly different ages of the vortices in the laboratory and flight experiments. The principal conclusion of practical interest is that the maximum circumferential velocity in a trailing vortex generated by an aeroplane in flight is

where Γ is the circulation about large circuits enclosing the vortex, v the kinematic viscosity of the air and t the (sufficiently great) age of the vortex.

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 effect of base slant on the base pressure distribution, drag coefficient and vortex shedding characteristics of a model consisting of an axisymmetric main body with an ellipsoidal nose have been investigated for three fineness ratios; 3, 6 and 9. A sudden change in the drag coefficient and separated flow pattern is observed at a critical slant angle (for constant incidence) or at a critical angle of incidence (for a constant base slant angle). The tests confirm that the value of the maximum drag coefficient is extremely sensitive to angle of incidence. Measurements of the frequency of vortex shedding are presented and the structure of the wake is investigated using smoke visualization and hot-wire correlation measurements. The wake is found to be far less stable than that from a two-dimensional bluff body and the vortex structures are sometimes in-phase and sometimes out of phase across the wake. The effect of free-stream turbulence on this family of body shapes is observed to be different to that on three-dimensional blunt-faced bluff bodies. Free-stream turbulence is found to have a minimal effect on base pressure for slant angles giving a recirculating type near wake flow. When longitudinal vortices are present the addition of free-stream turbulence slightly reduces the magnitude of the peak suctions recorded on the base but has little effect on base drag.

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.