Separated and vortical flows are dominant over various unmanned air vehicles (UAVs). In this article, issues and challenges of vortical flows for future UAVs are reviewed. These include shear layer instabilities, vortex breakdown and wing stall, vortex interactions, nonslender vortices, multiple vortices, and manoeuvring wing vortices. There are also issues relating to vortical flows in certain flow/structure interactions, as well as in aerodynamics/propulsion interactions. Separated and vortical flows are even more dominant at low Reynolds number flows. The main features of vortical flows, unsteady aerodynamics, and propulsion related vortical flow isssues relevant to mini- and micro air vehicles, are discussed.

Direct optimisation techniques using different methods are presented and compared for the solution of two common flows: a two dimensional diffuser and a drag minimisation problem of a fixed area body. The methods studied are a truncated Newton algorithm (gradient method), a simplex approach (direct search method) and a genetic algorithm (stochastic method). The diffuser problem has a known solution supported by experimental data, it has one design performance measure (the pressure coefficient) and two design variables. The fixed area body also has one performance measure (the drag coefficient), but this time there are four design variables; no experimental data is available, this computation is performed to assess the speed/progression of solution.

In all cases the direct search approach (simplex method) required significantly smaller number of evaluations than the generic algorithm method. The simplest approach, the gradient method (Newton) performed equally to the simplex approach for the diffuser problem but it was unable to provide a solution to the four-variable problem of a fixed area body drag minimisation. The level of robustness obtained by the use of generic algorithm is in principle superior to the other methods, but a large price in terms of evaluations has to be paid.

This paper focuses on the design of a cascade within a cold stream thrust reverser during the early, conceptual stage of the product development process. A reliable procedure is developed for the exchange of geometric and load data between a two dimensional aerodynamic model and a three dimensional structural model. Aerodynamic and structural simulations are carried out using realistic operating conditions, for three different design configurations with a view to minimising weight for equivalent or improved aerodynamic and structural performance. For normal operational conditions the simulations show that total reverse thrust is unaffected when the performance of the deformed vanes is compared to the un-deformed case. This shows that for the conditions tested, the minimal deformation of the cascade vanes has no significant affect on aerodynamic efficiency and that there is scope for reducing the weight of the cascade. The pressure distribution through a two dimensional thrust reverser section is determined for two additional cascade vane configurations and it is shown that with a small decrease in total reverse thrust, it is possible to reduce weight and eliminate supersonic flow regimes through the nacelle section. By increasing vane sections in high pressure areas and decreasing sections in low pressure areas the structural performance of the cascade vanes in the weight reduced designs, is improved with significantly reduced levels of vane displacement and stress.

The European Supersonic Commercial Transport’s control surface configuration is based largely on Concorde’s and has been scaled to provide comparable un-augmented stability and manoeuvre performance. Hence, optimising the surface size could provide significant performance benefits in terms of reduced drag, noise, structural and actuator power requirements. Adequate control power will be required to meet current civil aviation regulations whose primary aim is to ensure the aircraft can be flown safely during both normal and emergency operation. Additional design criteria, combined with the optimum longitudinal control laws, are required to ensure desirable handling qualities with minimum pilot workload. Two critical low-speed flight conditions, normal and emergency, together with associated aircraft configurations for control surface sizing have been identified. The rudder must provide sufficient control power to achieve positive heading changes subsequent to a double asymmetric engine failure during normal operation. The fin should be sized to satisfy Dutch roll stability criteria with the un-augmented aircraft in its emergency configuration. The dual functionality of the elevons require that they are sized using both pitch and roll performance and handling quality criteria. The bank angle capture requirement provides the most critical elevon design case, the satisfaction of which also ensures adequate pitch control power. Validation using ‘pilot-in-the-loop’ simulation will be required whilst more explicit control surface size optimisation would require the definition of limiting airspeeds and operating conditions applicable to the European Supersonic Commercial Transport. Additional studies of control power requirements during transonic and supersonic cruise may also be required.

As one aspect of a balance research and development programme, engineers at the Aircraft Research Association (ARA) have addressed the unique problems associated with the design of an internal single piece balance suitable for small military type aircraft models that experience very high loads during a wind tunnel test. The paper details the nature of the problems and the stages undertaken in the design, manufacture and commissioning of an initial prototype balance followed by a re-designed version to meet the stringent requirements. Data from recent tests indicate the quality of the measurements and the suitability of the balance for the testing of fighter type aircraft.