This series of lectures recalls the achievements of that outstanding pioneer of rotary wing flight Juan de la Cierva. At around the age of 23, he began to develop the concept of improved safety in flying machines, and included in his studies the concept of the helicopter. Having carried out many tests of various possible configurations, he discarded the concept of true helicopter flight as not being achievable with the then current state-of-the-art.

However, he became convinced that his basic objectives of short takeoff and landing, and safe low speed flight could be achieved by replacing the wing of a conventional aircraft with an autorotating rotor.

An investigation was carried out to study the parameters affecting single-jet ground-effect hover lift loss and to identify the cause(s) behind the large discrepancies in lift loss between the experiments of Wyatt of RAE Bedford and Corsiglia et alof Nasa Ames. The cause of the discrepancies was traced to a single geometrical parameter: baffle plate edge geometry, which significantly affected the flow separation underneath the plate. In contrast, the out-of-ground lift loss is insensitive to the edge geometry, but can be significantly affected by room size restriction. The investigations have highlighted the importance of accurately representing the edge geometry in ground effect testing of Vstol aircraft, and the importance of the test environment when conducting out-of-ground hover experiments.

An experiment is described which was undertaken specifically to provide CFD validation data for the case of transonic flow over nozzle afterbodies. The tests were undertaken with the AGARD standard 10° and 15° axisymmetric boat-tail geometries. Onset Mach numbers in the range 0·80-0·99 and subsonic and under-expanded jet plumes were employed in the tests. Test conditions were selected which provided a range of afterbody flow features from largely attached to shock-induced separated flows. A uniquely detailed set of surface pressure and flowfield data are presented. The flow data were acquired with a two-component laser Doppler anemometer (LDA) and define the mean and fluctuating flow components at about 500 spatial locations for each of these complex transonic flowfields. Additional information was recorded which fully defines the required computational boundary conditions.

Also presented is a detailed study of the necessary attributes of windtunnel CFD validation data. It is demonstrated that relatively high blockage experiments using cost-effective windtunnels can be used to generate CFD validation data if proper account is taken of the model/tunnel interference.

Aircraft yaw control at high angles of attack by tangential forebody blowing has been investigated experimentally. Tests were performed in the University of Bath 21 m x 1.5 m low speed wind tunnel using a 6% scale generic combat aircraft model fitted with blowing slots in the nose cone. Six component strain gauge balance force and moment data were measured for angles of attack up to 90° for various slot geometries and locations. The effect of slot azimuthal location is demonstrated and a slot stall phenomenon described. A geometry dependent forebody/wing flowfield coupling has been identified which can lead to unexpected yawing and rolling moments. The primary source of yawing moment is shown to be the enhanced area of attached flow on the blown side of the forebody rather than direct vortex influence.

The optimum slot extent and location depend on the angle of attack range over which control is required. For regions where steady vortex asymmetry is present, slots near the apex of the forebody produce severe control reversals at low blowing rates. These reversals can be minimised by placing the slots away from the apex. For control in regions where the flow is dominated by periodic vortex shedding, long slots offer efficient control to 90° angle of attack. The most suitable compromise for wide range control would appear to be a short slot placed away from the apex of the forebody.