Significant progress has been made to date in modelling, computationally, the formation and development of the dust cloud that forms in the air surrounding the rotorcraft under brownout conditions. Modern computational methods are able to replicate not only the development of the dust cloud in appropriate operational scenarios, but also the sensitivity of the shape and density of the dust cloud to the detailed design of the rotorcraft. Results so far suggest that attempts to ameliorate brownout by aerodynamic means, for instance by modifying the rotor properties, will be frustrated to some extent by the inherent instability of the flow field that is produced by the helicopter. Nonetheless, very recent advances in understanding the fundamental mechanisms that lead to the formation of the dust cloud may allow substantial progress to be made once certain elements of the basic physics of the problem are more fully understood and better quantified.

A helicopter model has been identified and validated for flight conditions defined by altitudes, speeds, loadings and centre of gravity positions. To identify the helicopter models, 2-3-1-1 multistep control inputs were performed by the pilot to excite all helicopter modes. Then, each estimated signal has to remain in tolerance margins defined by the Federal Aviation Administration. Three methods were used to observe the system outputs from its states: a fuzzy logic method, a linear method optimised with a neural network algorithm and a classical method. Because of random effects when gathering data, classical method did not give good enough results. The fuzzy logic method was not robust enough so that output plots showed peaks that could be felt by the pilot. Then, because the model could be implemented in a simulator for the pilot training, the pilot feedback is very useful in order to compare the reality with the results of the mathematical model. When the outputs are obtained from the measured state variables, the linear method gave the best results.

The aerodynamic design optimisation of a Micro Air Vehicle (MAV) wing is performed to obtain the optimal anti-symmetric wing twist distribution for the roll control of the MAV’s wing instead of using conventional ailerons. This twist distribution should produce minimum induced drag and achieve a better roll response. The implementation of several anti-symmetric load distributions such as the half lemniscates and the Horten distributions is studied leading to an initial solution for the optimal distribution that could achieve better roll requirements. Multhopp’s method based on Prandtl’s classical lifting line theory is used for the determination of the spanwise load distribution required during the optimisation process. The optimisation process is based on the modified feasible directions gradient based optimisation algorithm implemented in the optimisation system, VisualDOC, given by Dr. Garret Vanderplaats. The proposed optimisation process is applied to the ‘BARQ’developed MAV which has successful flight in July 2009.

Generic wrap-around shape optimisation technology is presented, and is applied to a helicopter rotor in hover, using compressible CFD as the aerodynamic model. An efficient domain element shape parameterisation method is used as the surface control and deformation method, and is linked to a radial basis function global interpolation, to provide direct transfer of domain element movements into deformations of the design surface and the CFD volume mesh. This method is independent of mesh type and size, and optimisation independence from the flow solver is achieved by obtaining sensitivity information for an advanced parallel gradient-based algorithm by finite-difference. The method is applied here to a two-bladed hovering rotor, with transonic tip speed, comparing the effects of global twist parameters and more local planform parameters. Significant torque reductions are achieved in both cases, and it is shown that using only three global twist parameters are extremely effective. Using 63 local and global parameters, large geometric changes are demonstrated, with both induced power and wave drag reduced.

In this note, the experimental results of wind-tunnel measurements of lift coefficient and lift-curve slope for aerofoil NACA 0012 obtained from the VTI-Institute Zarkovo, are presented. The results were obtained from tests and integrations of surface static-pressure data over a model of the NACA 0012 aerofoil section. The data were obtained for a free-stream Mach number range of 0·25-0·8 and a chord Reynolds number range of 2-25MRe. The essential results of these measurements along with the results from other authors are presented and evaluated. The principal factors which influence the accuracy of two-dimensional wind-tunnel test results are analysed. The influences of Reynolds number, Mach number and wall interference with reference to solid and flow blockage as well as the influence of side-wall boundary-layer control are analysed. Interesting results brought to light the Reynolds number effects of the test model versus Reynolds number effects of the facility in subsonic and transonic flow as well as the effects of the side-wall boundary-layer control and wall interference.

Adding a small aerodisk at the tip of the spike was proved to improve the spike’s performance in reducing the drag. If designed properly, the aerodisk spike can offer better performance in drag reduction in comparison with the pointed sharp-nosed spike. However, the impact of the aerodisk on the aeroheating reduction capabilities of the spike was not investigated in detail before. In the present work, a numerical investigation of the aerothermodynamic characteristics of spiked blunt bodies at Mach 6 flight conditions is conducted. It was found that, adding an aerodisk degrades the performance of the spike in reducing the aeroheating level. Nonetheless, the latter remains below that of the unspiked body.

The design of wings with morphing capabilities is known to give aerodynamic benefits. These aero-dynamic benefits come from both the use of hinge-less surfaces and the greater adaptability to flight conditions. This paper describes the structural design of a twisting wing to be used for an unmanned air vehicle (UAV) and presents finite element analysis and experiment results. This is part of a research project carried out at the University of Southampton in which one of the goals is to compare different novel wing designs and technologies to determine which one of them gives the best performance. The twisting capability provides roll control without hinged surfaces hence providing aerodynamic improvement. The wing is manufactured using polystyrene foam and is cut out of block of this material using a hot wire machine. In order to link this foam structure to a main spar, ABS plastic inserts were manufactured using a 3D printer. The mechanisms used to actuate the wing are also made from this material. A full scale UAV wing has been manufactured and tested in order to compare with FEA results.

A procedure for developing efficient aeroelastic reduced order models (ROMs) for aerospace structures containing geometric nonlinearities is described. The structural modelling is based upon a combined modal/FE approach that describes the non-linear stiffening effects from results of non-linear static analyses for a range of prescribed inputs. Once the structural ROM has been defined, it is coupled to the rational fraction approximation of the doublet lattice aerodynamic model corresponding to the wing planform. The aeroelastic model can then be used to predict the dynamic aeroelastic behaviour of the defined structure. The methodology is demonstrated on the aeroelastic model of a flexible high aspect ratio wing with the static deflections, LCO behaviour and gust response being predicted.