The aerodynamic performance of two dimensional (2D) aerofoils and finite aircraft wings is investigated when the leading edge is modified with a standard ice shape. Three ice-shapes, G1, G3 and R7 are selected for this investigation. For the aerofoils it was observed that the presence of strong vortex flows near the leading edge fundamentally changes the flow in that region forcing earlier transition and premature separation. Normal suction peaks designed to provide high aerodynamic lift and prolonged attached flow are replaced with a system of vortices which produce unwanted pressure spikes which influence the development of the boundary layer leading to earlier stall. The present study also includes the three dimensional (3D) effects due to iced leading edges on aircraft surfaces. As different from such studies elsewhere, which only considered the simple sweep back effects on constant cross section wings, the present work includes more realistic wings equipped with twist, taper and non-equal sweep back at leading and trailing edge of the wing. In absence of any experimental data on such real life type configurations, the computational fluid dynamic (CFD) results were first validated against measurements on non-swept wings. It was found that ice formations cause noticeable changes of the flow in the leading edge regions of the wing and promote strong 3D effects, which pervade across the entire span wise direction of the model.

Preliminary investigations of the effect of tracer particle characteristics on flow visualisation in a steady delta wing leading-edge vortex are presented, using both 3D simulations of particle trajectories and a simplified analysis of the radial equilibrium conditions. It is shown that many of the flow structures apparently seen in wind and water tunnel visualisation studies of delta wing vortices may in fact be artefacts of the tracer particle characteristics. In particular the ‘smoke tube’ or ‘hole’ effect often seen in smoke flow visualisation in wind tunnels does not (contrary to popular usage) indicate the size of the viscous subcore, and is not a reliable indicator of the vortex structure. The implications for both visualisation studies and optical flow measurement techniques are discussed.

This paper describes the application of the theory of pulse responses to a linearised Euler scheme on a moving grid. The impulse responses of such a linear system are memory functions or temporal representations of the manner in which, and the time over which a perturbation remains active in the response of the system. The exact response of the linear system to an arbitrary input can then be predicted via convolution. The linearised Euler scheme is derived from a full nonlinear Euler method based on Jameson’s cell-centred scheme which has been modified to make it time-accurate and has the necessary terms to account for grid motion. The discrete form of the full Euler equations are linearised about the corresponding nonlinear steady mean flow assuming the unsteadiness in the flow and grid are small. The resulting set of differential equations for the flow perturbations are solved at each time step using a dual-time scheme. These equations can be easily solved to obtain pulse responses, which capture the entire frequency content of the linear system. These pulse responses then provide a simple and efficient method for calculating the system response to many different inputs via convolution. A further use of the pulse responses is their potential to allow a reduced-order linear model to be constructed, which preserves some of the accuracyof the original system. Preliminary work carried out in this area is presented. Results of calculations for heave, pitch and ramp test cases are presented.

The effect of streamwise slots and grooves on a normal shock wave-turbulent boundary-layer interaction has been investigated experimentally at a Mach number of 1.3. The surface pressure distribution for the controlled interaction in the presence of slots featured a distinct plateau. This was due to a change in shock structure from a typical unseparated normal shock wave-boundary-layer interaction to a large bifurcated lambda type shock pattern. Velocity measurements downstream of the slots revealed a strong spanwise variation of boundary-layer properties, whereas the modified shock structure was found to be relatively two-dimensional. Cross flow measurements indicate that slots introduce streamwise vortices into the flow. When applied to an aerofoil, streamwise slots have the potential to reduce wave drag while incurring only small viscous penalties. In the presence of grooves the interaction was initially found to be significantly different. A bifurcated shock structure was observed but the trailing leg appeared stronger and featured a second lambda foot. Oil flow visualisation also revealed differences in the interactions, with the region of suction and blowing being limited to a smaller extent of the grooved control surface. The amount of crossflow present was reduced compared to the slotted control surface. By varying the internal geometry of the grooves it was found that the interaction could be modified to be similar to that in the presence of slots indicating that a more practical control device can be designed.

Noise reduction in the inlet of stationary gasturbines or aeroengines is usually achieved by passive absorbers (acoustic liners) the efficiency of which is limited due to the confined length. Higher sound level reductions can be obtained — in particular for the tonal components -by using additionally active noise control techniques (ANC): The rotorspeed-coherent tonal components of the primary sound field are superimposed with an anti-phase sound field which is produced by secondary acoustic sources.

A detailed numerical study has been performed to investigate the origin and mechanism of the formation of windward shocks that have been observed on inclined slender bodies at supersonic speeds inside the bow shock wave around the body. It is shown that the feature is associated with the virtual double cone-like deflection of the supersonic stream by the primary vortices and, as such, can be named the ‘vortex shock’.