Since 1993, research at Loughborough University has included examining the commercial air safety record of Australia in an attempt to rise to the challenge posed by Professor James Reason: “Should we not be studying what makes organisations relatively safe rather than focusing upon their moments of unsafety? Would it not be a good idea to try and find out what makes them good and whether or not these ingredients could be bottled and handed on?”.

The objective is not to prove Australia to be the safest place to fly, nor its airlines the safest in the world. It is simply to examine what sort of defences are in place that may be strengthened or emulated.

Screen separations as little as 0.02 chamber diameters do not adversely affect suppression of mean-flow non-uniformity provided the overall pressure drop coefficient, K, does not exceed about 3.5. At larger K a velocity overshoot develops near the wall boundary layers, which decreases only slowly with wider separation. A simple approximate extension can be made to the theory of Taylor and Batchelor for any number of closely-spaced screens. This shows good agreement with measurements and with low-drag screens predicts perfect attenuation when the overall pressure-drop coefficient is 3.2 or 7.6. Some inferences regarding the effect of close screens on turbulence suppression are also drawn.

A computational study has been undertaken of 3D vortical separation from the curved surface of a prolate spheroid at high angle of attack (10° and 30°). Attention focuses on the predictive capabilities of a new variant of non-linear, low Re eddy-viscosity model and full second-moment closure, the latter coupled to a low Re k-ε Boussinesq-viscosity model which is applied to the semi-viscous near-wall region. The study demonstrates that both anisotropy-resolving formulations return very similar predictive performance which is in several respects superior to that achieved with the k-ε model based on the linear stress-strain relations. At the higher incidence angle, transition is free, and this is the source of considerable uncertainty in respect of the sensitivity of the predicted leeward flow to the location of transition. While none of the models is fundamentally capable of capturing natural transition, the ability of the non-linear model to suppress turbulence generation by irrotational straining at the windward impingement region, thereby preventing (bypass) transition by diffusion of turbulence from the freestream to the laminar boundary layer, is exploited to demonstrate that the sensitivity of the separated leeward flow to the location of transition is weak.

The flight test phase of the Nasa-Army UH-60A Airloads Program was completed in February 1994. The objective of the project was to measure the airloads in flight on an UH-60A Black Hawk using 221 blade pressure transducers installed in nine radial arrays. In addition, 21 transducers were installed along the blade leading edge to better identify blade-vortex interaction phenomena. This paper briefly describes the program's background and history, instrumentation, and data acquisition. Airloads data are presented for two test conditions with different levels of complexity: steady, level flight and a diving turn. These flight data provide useful information of relatively simple and complex nonlinear flow for the validation of theoretical analyses. Evidence of shock waves, reverse flow, dynamic stall and blade vortex interactions are presented.

The use of large scale windtunnel models offers considerable advantages to an experimenter

• 1. the possibility of obtaining high Reynolds number, certainly significant and sometimes equal to that in flight, for moderate stagnation pressures

• 2. an accurate representation of geometric elements, i.e. good form similitude (it is necessary to emphasise the importance of being very thorough in this area)

• 3. large internal volumes which enable numerous transducers to be implanted, ambitious experimental devices and means of remote control of model motorisation

• 4. test conditions remain those corresponding to industrial manufacturing methods and so avoid problems related to extreme miniaturisation.

Additionally, there exist some test conditions for which scale reduction is limited or even impossible; for example, because of physical phenomena, limitation in industrial methods of miniaturisation, or quite simply because the subject is itself the flying object. These conditions are applicable to all tests whether industrial or research. They are especially true for industrial tests where these concern important decisions. They are also true for research tests whenever they need to be sufficiently close to flight reality. Wind-tunnels are always a means to reduce the costs and risks of aeronautical project developments. The risk taken in extrapolating from windtunnel tests to flight are lower when the test is more realistic and this gives a clear advantage to tests on large sized models.

Results of a comparison of dynamic stall onset are presented, as assessed from low Mach number windtunnel data contained in the University of Glasgow's database and that from the well established Beddoes’ model. The model, which was originally developed to reconstruct higher Mach number dynamic stall characteristics, exhibited a lower stall onset incidence than that assessed from the windtunnel data. The differences and speculations on the physical reasoning underlying the two assessments are discussed. An addition to the Beddoes model is proposed which yields an improved reconstruction of the Glasgow windtunnel data.

This paper derives innovative techniques for use in the trimming and stability analysis of advanced rotorcraft simulations. It begins by exploiting the symmetry of the rotor to produce an efficient definition of periodic trim which is applicable to rotorcraft simulations. This definition is then expanded to produce a trimming algorithm which is capable of concurrently ascertaining the initial conditions and control inputs necessary to trim latest generation simulation models to a specified periodic trim state. The algorithm is based on a periodic shooting approach with Newton-Raphson iteration and exploits the symmetry of the rotor to minimise computational workload. The definition of periodic trim is then further developed to produce a technique by which the stability characteristics of rotorcraft can be ascertained from advanced simulation models. This technique is based on a Floquet approach and again exploits the symmetry of the rotor to reduce computational burden. The paper concludes by presenting results obtained when the stability characteristics of a tiltrotor simulation model are investigated.

A computer program for use in the conceptual stage of aircraft design has been developed. The program obtains minimum mass designs for high aspect ratio, composite wings, subject to constraints on flutter speed, divergence speed and material stress. The wing is modelled as a series of composite beam elements and both flutter speed and divergence speed are calculated using a normal mode approach. Modal analysis is carried out by applying the Wittrick-Williams algorithm to the dynamic stiffness method, whereas unsteady aerodynamic loads are calculated from strip theory, although an option which uses lifting-surface theory is also presented. A previously published example is given to validate the analysis. Single level optimisation is carried out using a sequential quadratic programming strategy combined with the modified methods of feasible directions optimizer, for which flutter sensitivities are obtained by an efficient determinant interpolation technique. Design variables include topological variables such as spar and engine positions as well as layer thicknesses, which are modelled using quadratic functions. The wing of a regional turboprop aircraft is optimized to illustrate the use of the program. The problem was modelled using 10 elements and had 43 design variables, 162 constraints and required just over 20 minutes of CPU time on a workstation. This, coupled with the fact that a full three-dimensional FE model of the same wing would require over 1000 elements, illustrates the suitability of the dynamic stiffness method to the conceptual design stage.