The effects of flap deflection, configuration incidence and leading edge bluntness on boundary layer separation and transition have been studied experimentally. A quasi two-dimensional flat plate equipped with a full span trailing edge control flap has been employed for these tests.

The studies have been carried out in a hypersonic gun tunnel at M∞ = 8·2 and Re∞/cm = 9·0 × 104. The flap deflection angles studied were in the range 0° ≤ β ≤ 30°. The incidence range was from zero to α = 10° (positive α is nose down). Leading edge bluntness effects were simulated by the introduction of a hemi-cylindrical leading edge.

The flow structure was studied using high speed Schlieren photography as well as surface pressure and heat transfer measurements. Liquid crystals were employed to study the threedimensionality of the flow structure for selected configurations. Analytical theories have been developed to estimate the flap pressure and heat transfer levels for the sharp and blunt configurations. These are compared against experimental measurements.

Experiments are performed on the effects of a turbulent boundary layer on surface pressure and heat transfer, using a cylindrical model fitted with a variety of flares at Mach 5. The flow over compression corners of up to 20° (below the expected incipient separation value) is examined. Liquid crystal thermography indicates an unexpected peak in heat flux close to the corner. It is suggested that this is due to a small scale separation occurring earlier, i.e. at smaller corner angles, than previously observed. An estimate of the size of such a separation bubble is made and it is suggested that this is a predominantly laminar phenomenon.

This paper describes a simple method to accurately simulate the flow field around a wing in surface effect ship (Wises) model that involves the use of a contraction and a slot to suck the boundary layer from a flat plate. Theoretical aspects of the design are discussed with relation to the required level of suction and the slot width.

The effectiveness is ascertained from measurements of the boundary layer height on the plate through measurement of the mean velocity profile and a reduction in height of 50% was generally observed when suction was used.

Comparison of the forces on a low aspect-ratio wing, obtained from the proposed and the image method, verified the proposed method for wing heights above 2·5% of the mean chord. At lower heights, separation of the boundary layer on the ground plate is thought to have altered the pressure distribution on the wing.

The paper uses the conventional nine-noded Lagrangian element for studying transient linear response analysis of composite stiffened shells. An improved version of the stiffener modelling has been used in which the stiffener can be placed anywhere inside the element. For the first time, concentric or eccentric stiffeners have been used for composite stiffened shells for solving the transient dynamic response of these structures. These are not available in existing commercial packages. Different types of time dependent loading such as short duration air-blast loading, suddenly applied uniformly distributed step loading and sinusoidally harmonic loading have been considered in this paper. The results of stiffened composite cylindrical shells and doubly curved shells with different boundary conditions and various laminate orientations have been presented for eccentric stiffeners. Parametric studies considering different variables have also been carried out.

The work reported here extends the work of Nikuradse on the categorisation of surface roughness. Nikuradse’s work considered pipe flow, whilst in this paper boundary layer flow is investigated. Important and fundamental differences between these two types of flow are shown. An experiment upon turbulent boundary layers over sandpaper-type roughnesses enabled a fluid-mechanic categorisation of these surfaces to be obtained for this flow. This is compared with some shape and size categorisations of the roughnesses from detailed measurements.

The purpose of this paper is to describe an approach which performs data fusion on the output of multiple, spatially separate, sensors engaged in the real time tracking of obstacles in a helicopter's environment. The generated information can be used either as a flight director aid or as feedback required by an automatic collision avoidance system. Obstacle track estimation has been commonly carried out using the Kalman filter (KF) for linear estimation, or the extended Kalman filter (EKF) for use on nonlinear problems. However, certain assumptions made in the derivation of the EKF algorithms render it sub-optimal for aerial obstacle track estimation. Additionally, the EKF has problems with initialisation and divergence (stability) for many non-linear processes.

Research at the University of Southampton has highlighted a link between fuzzy networks and associative memory neural networks. This link is important as it allows new learning rules to be developed for training fuzzy rules, and learning convergence to be proved. This paper explores methods of fusion of estimates using neuro-fuzzy models, and addresses some of the weakness of the Kalman filter approximation introduced by the assumptions made in its derivation.

A theoretical study was undertaken to determine the effect of taper ratio on the trailing vortex structure of a high aspect ratio unswept wing. The wing was modelled using lifting line theory, and the vortex was assumed to have a Rankine distribution of circulation. The results show that reducing the taper ratio of the wing, decreases the maximum tangential velocity of the vortex as well as the vortex span while the core's diameter increases. Data is also presented showing the sensitivity of induced drag to wake vortex parameters.