A theoretical model had been previously developed to predict the behaviour of a semi-rigid helicopter rotor blade during rotor engagement and disengagement under high wind conditions. This has now been extended to cater for an articulated rotor by modelling the droop and anti-flap stop mechanisms by high rate linear springs.. Much increased blade deflections are predicted, as anticipated, along with the possibility of blade tip/tail boom contact at a high enough wind speed. The importance of the higher modes in the description of the phenomenon is demonstrated. A comparison of the theoretical model with an experiment using a full size Puma aircraft is presented.

The prime need to know Mach number in flight is likely to be for the efficient operation and control of air-breathing propulsion units. The errors incurred by the use of the analytical expressions, which have been developed for supersonic flight, are quantified and shown to be increasingly serious for speeds in excess of M = 5.5 approximately. Systems for calculating this parameter in flight are reviewed and compared. It is shown that three out of the four systems identified are capable, potentially, of registering Mach number to a good precision. In each case, to achieve optimum precision requires a good knowledge of ambient air temperature. Such knowledge is likely to be available from meteorological satellites.

This paper considers the dynamic stall vortex of importance in helicopter rotor aerodynamics and discusses previous measurements of its convection speed. It emerges that an anomaly exists between the available data sets, i.e. that some workers find that the convection speed is dependent upon the aerofoil motion, while others find that this is not the case. Measurements of the convection speed from data gathered at Glasgow University for a variety of aerofoil shapes and motion types are then presented, which support the conclusion that the dynamic stall vortex convection speed is independent of aerofoil type and motion type to a first order.

A simple flow model has been developed to predict the size of the separation bubble which forms at the sharp leading edge of a flat plate aerofoil at incidence a which are less than about 8°, and the corresponding lift and drag coefficients. This flow model is based on irrotational flow and source singularities which are determined indirectly from experiment. The bubble length is shown to be proportional to α2with a slope which depends on the rate of growth of the separated shear layer. A thin double-wedge aerofoil has been tested in a wind tunnel. It is concluded that these and previous measurements agree fairly well with the theory for chord Reynolds numbers greater than about 0.5 × 106.

A new integral equation, and the corresponding boundary-element method, is presented for the aerodynamic analysis of compressible, subsonic, potential flows around bodies in arbitrary motion, with applications to helicopter rotors in hover. The formulation is closely related to that introduced by Morino for the aerodynamic analysis of aircraft: in that paper, the surface of the aircraft is assumed to be in arbitrary motion with respect to a frame of reference that moves in uniform translation. Such a formulation is very useful for the analysis of aircraft undergoing small oscillations with respect to a reference configuration. However, for the analysis of more complex motions (e.g. helicopter rotors in forward flight), it is more convenient to use a frame of reference that moves in arbitrary motion (in particular, for the case of a body moving in rigid-body motion, a frame of reference rigidly connected with the body).

The paper reports the first computational explorations of the flow in the vicinity of riblets with a two-equation turbulence model. For simplicity fully developed flow in a plane channel is considered though the results should be applicable to boundary-layer flow, as the riblets are confined within the (virtually) constantstress near-wall sublayer. The study of idealised L-shaped (or blade) riblets showed larger levels of drag reduction (up to 22%) than have hitherto been reported. When the usual triangular-profile riblets were considered, however, the maximum drag reduction was reduced to about 10%, in line with experiment.

Although the maximum drag reduction is well predicted, there are serious differences from experiment for the values of h+ at which the optimum performance is achieved. Moreover, the computations suggest that at low Re for h+ < 10, drag increases result. These anomalies highlight two distinct and counteracting weaknesses in the model of turbulence: there is an insufficient sensitivity of the equation to three-dimensional straining in the jow-Reynolds-number region while use of an isotropic viscosity model enforces zero secondary circulation in the vicinity of the riblets.

The development of the European spaceplane Hermes is based on new technologies under investigation. These technologies remain valid to the European aerospace database whatever the political future of the programme. Results achieved have been continuously implemented in the design process, which allowed the consolidation of the baseline configuration. Major technologies within the programme comprise re-entry Computational Fluid Dynamics and hypersonic Experimental Fluid Dynamics, the hot structure, thermal protection and material development, and development of new subsystem technologies for manned spaceflight.

A summary of the technology results achieved during the pre-design phase is made and implementation into the consolidated pre-design configuration is shown.

The title of this article may surprise those who look upon composites as a recent innovation. The first major step was taken in 1940 when Dr Norman de Bruyne's Aero Research Ltd of Duxford, now part of the Ciba-Geigy Group, built a Spitfire fuselage in Aerolite, a material developed by them comprising flax thread skeined and impregnated with a phenolic resin and formed in a heated press. The project arose as a safeguard against a possible war-time shortage of aluminium alloys. Tests at RAE proved entirely satisfactory whilst the weight was the same as the metal version, see Figs 1-3.

In 1947 Turner Brothers developed a chrysotile asbestos fibre reinforced phenolic resin called Durestos. The wings of the experimental Fairey Delta E10/47 were built from it and were the first primary composite structures to be flown.

Since those days immense strides have been made in the use of advanced composites, most modern civil and military aircraft taking advantage of their unique properties.

A small, but increasingly active, lighter-than-air industry has developed in the United Kingdom over the last 20 years. One of the five companies now involved has built more conventional helium filled airships than the rest of the world's manufacturers put together. Two others have supplied over 100 hot air airships — mostly to orders received from outside the UK.

Current LTA designs and activity in the UK are described.

A comparison between computed and experimental pressure distributions on straight and forward-swept wings placed in interference with a fore canard surface at M = 0·3 and Re ≈ 2·8x106 is presented. It is shown that a numerical code, based on a non-linear vortex lattice method and expressly developed for the analysis of interfering lifting surfaces, is capable, in spite of its simplicity, of very accurate predictions in all configurations which do not correspond to sufficiently high angles of attack and to a close interference between the fore wake and the wing surface. Furthermore, even in the latter cases the predictions are acceptable, and the code is shown to be extremely robust as regards the variation of all its free parameters.

The next challenge for future engines is to increase the bypass ratio in order to improve the propulsive efficiency. These very high bypass ratio engines improve the fuel consumption but create new technical challenges. These challenges include the fact that larger engines require bigger nacelles, hence the external drag and weight increase; furthermore the installation of these engines on civil aircraft is more challenging. Therefore important work has been carried out to optimise the design of these new nacelles and their installation on the airframe.

The present work is based on a new engine, with a very high bypass ratio. For this engine Snecma has designed an advanced nacelle with a low diameter-low length fan cowl and a core cowl mounted thrust reverser. A very short secondary flow nozzle is required to improve accessibility to the core engine and its accessories. The fan cowl length is dramatically reduced.

Different solutions were investigated to design the inlet and the fan cowl. Based on 2D potential inverse codes and 3D Euler codes, CFD calculations were made to optimise the aerodynamic lines for cruise conditions while keeping good off-design performance. A highly challenging exhaust nozzle was designed to maintain good efficiency with very tough geometric constraints. For this Snecma used 2D Navier-Stokes CFD codes. New reverse systems were studied for this type of nacelle, for example core cowl mounted thrust reversers. The optimisation of these highly 3D geometries was performed using 2D and 3D Navier-Stokes calculations.

An experimental and computational investigation of hypervelocity, hypersonic combustion of hydrogen with air is described which was intended to confirm the presence of combustion at these conditions and then to gauge the pressure rise associated with the heat release of combustion. Experiments were conducted in a square cross section duct and wall pressure was measured. It was found that combustion did occur and that the maximum pressure increase was approximately 72% over the intake pressure.

A semi-inverse viscous/inviscid coupling technique for the calculation of compressible flows is presented. Euler equations in integral form are solved for the inviscid part and integral compressible boundary-layer equations in direct and inverse form are solved for the viscous one. Interaction is realised using the equivalent sources approach.

The technique has been applied to the calculation of the aerodynamic characteristics of aerofoils both in cruise and take-off or landing configurations. To treat the complex geometries involved in the latter case, a multi-block approach, based on structured ‘H’ grid topology, has been developed and validated for different configurations.

Some of the experimental features of the dynamic response of a cantilever subjected to two streams of air at right angles to each other are reported. The cantilever is mounted across a mainstream flow while subjected to another jet injected parallel to the test cylinder.

The dimensionless vibration displacement amplitude was measured in the transverse direction (lift force direction) as well as in the streamwise direction (drag force direction) while varying the secondary jet flow velocity. It was found that in the vicinity of the cylinder a secondary jet can be used to reduce the dimensionless transverse vibration displacement amplitude in the lock-in region and increase it outside the lock-in region. The effect of jet injection on the streamwise direction vibration is to increase it for all reduced velocities considered.

The present work investigates the structure associated with flow past two tandem circular cylinders with a diameter ratio of 0·33. The smaller cylinder is always upstream. The two parameters varied in the present work are the spacing between the cylinders denoted by l/d2where l is the centre to centre spacing and d2 is the diameter of the downstream cylinder, and the Reynolds number Re based on d2 Just like the case of flow past two equal size cylinders, a critical spacing was found to exist. For spacing less than the critical value, shear layers that separate from the upstream cylinder reattach onto the downstream cylinder (reattached flow regime) whereas for spacing larger than the critical value both cylinders shed vortices (co-shedding flow regime). For diameter ratio dl/d2 of 0·33 and at Re ≈ 6·27 × 104, the critical spacing is at l/d2 ≈ 1·8 to 2·2. In 1·8 ≤ l/d2 ≤ 2·2, bi-stable flow situation where the flow structure changes intermittently between the reattached and the co-shedding types was observed and the probability that the flow is co-shedding within the bi-stable regime is found to follow a normal distribution.It was also found that the reattached flow regime can be sub-divided into two sub-regimes. In 0·66 ≤ l/d2 ≤ 1·2 (sub-regime 1), the vortex formation length, lf, of vortices shed by the downstream cylinder appears to be shorter, resulting in larger drag force on the downstream cylinder and larger r.m.s. pressure around it. In 1·4 ≤ l/d2 ≤ 1·18 (subregime 2), lf appears to be larger which results in a smaller drag force and smaller r.m.s. pressure distribution when compared with sub-regime 1. Only the downstream cylinder has a Strouhal number S2 associated with it in the reattached flow regime. The general trend is that of a slight reduction in S2 with increasing l/d2 which is consistent with the relation of larger lf at larger l/d2 in the reattached flow regime. In the coshedding regime, the Strouhal numbers of both cylinders are similar.

The Reynolds number was varied in the range 3·15 × 104 ≤ Re ≤ 8·81 × 104. It was observed that both the mean drag and the r.m.s. lift and drag forces reduce in magnitude when Re was increased. From the spectra of the lift force, reduction in C'L was found to be due to a reduction in the regularity of the vortex shedding process. From the pressure distribution of the downstream cylinder at constant l/d2, it was found that the reduction in mean drag with increasing Re is caused by an increase in base pressure which in turn is likely to be the consequence of an increase in vortex formation length. The upper and lower limits of the range in the lldi at which bistable flow was detected were found to decrease with increasing Re. At l/d2 = 2·0, the flow was of the reattached type at Re = 3·15 × 104 and 4·75 × 104 but became bi-stable at Re ≥ 6·27 × 104.

A general analytical solution for the buckling of laminated composite plates and shell panels for all combinations of free, simply supported and clamped edge conditions is presented. The solution is formulated in terms of stress function and transverse deflection. All the boundary conditions for the stress function are exactly satisfied. For the transverse deflection the prescribed boundary conditions are satisfied by including them in the expression for the energy integral. Several types of boundary conditions in which at least a pair of opposite edges of the panel are free and subjected to axial compression are considered. The Galerkin method, with beam vibration functions used as shape functions, is utilised to determine the buckling loads. The results obtained are used to examine the effect of bending-extensional coupling, aspect ratio and the curvature parameter on the buckling load.

An economical numerical method has been developed to calculate unsteady aerodynamic characteristics of aircraft configurations, undergoing small amplitude simple harmonic motion in various laterally symmetric modes of oscillation, in inviscid incompressible flow, by taking advantage of the computational efficiency of internal distribution of singularities. The only approximation introduced in the method is the linearisation of the problem with respect to the amplitude of oscillation. A number of comparisons have been made to demonstrate the effects of different aircraft components on wing or tail pressure, and lift distributions.

The dynamic stability of blunt cones at angles of attack and high supersonic Mach numbers has been investigated along lines similar to the pointed cone solution derived by the present author. The steady pressure distribution on a blunt cone is complemented with a first and second order azimuth contribution resulting from an angle of. attack displacement. The unsteady perturbation is then superimposed on the steady solution by accounting for the streamline deflection in an oscillating flow. The closed form expressions obtained for the dynamic stability of blunt cones reduce to their pointed cone counterparts once the bluntness parameter is equated to zero. The comparison between the theoretical and measured results is quite encouraging.