Recent interest in aircraft vortex wakes has highlighted the need to identify appropriate algorithms for their simulation. The problem is not well-suited to grid-based techniques, due to the large domain and the need to avoid introducing artificial viscosity. A promising alternative is offered by Lagrangian formulations. This work aims to assess the usefulness of one such approach - the 2D vortex method -via comparison with experimental results. One in-house and two external (from DLR and ONERA) data sets are used. The results show generally good agreement in vortex locations and size, with the calculation performing best when initialised downstream of the strongly-three-dimensional trailing edge region. The neglect of viscosity can lead to some asymmetry in the vortex cores, but, even here, excellent agreement can be obtained when the field is integrated to give circulation (as a function of radius). The 2D vortex method is thus an accurate and efficient tool for wake vortex prediction.

The coupling of independent structural dynamic and inviscid aerodynamic models, in the time domain, is considered. The accuracy and CPU requirements of the two common approaches, namely ‘weak’ and ‘strong’ coupling procedures, are investigated. It is found that the strong coupling scheme is more accurate than the weak coupling approach, and only for large real time-steps is the strong coupling scheme more expensive. The computational method developed is used to perform transonic aeroelastic and aeroservoelastic calculations in the time domain, and used to compute stability (flutter) boundaries of 2D wing sections. A control law is implemented within the aeroelastic solver to investigate active means of flutter suppression via control surface motion. Comparisons of open and closed loop calculations show that the control law can successfully suppress the flutter and results in a significant increase in the allowable speed index in the transonic regime. The effect of structural nonlinearity, in the form of hinge axis backlash is also investigated. The effect is found to be destabilising, but the control law is shown to still alleviate the destabilising effect.

High performance fighter aircraft have problems with handling qualities, at high angles of incidence. Pilots are limiting combat effectiveness by cautiously avoiding hazardous regions of the flight envelope. This paper presents and extends existing work on the F-16 aircraft's handling characteristics during the deep stall condition. Inflight experience shows that there are two types of control problems when flying at high angles-of-attack, the ‘pitch departures’ and the ‘deep stall trim’. This paper investigates the critical deep stall condition, assesses the effectiveness of the proposed method of recovery and suggests an augmented and reliable method of returning to normal flight.

International political and socio-economic developments have led the armed forces of many countries to become more aware of how their increasingly-stringent financial budgets are spent. One major expenditure for military authorities is upon aero-engines, because in-service deterioration in any mechanical device, such as an aircraft's gas-turbine engine, is inevitable. Each deterioration has an adverse effect on the performance and shortens the reliable operational life of the engine, thereby resulting in higher life-cycle costs. For a military aircraft's mission-profiles, the consequences of an aero-engine's deterioration upon the aircraft's operational-effectiveness as well as its fuel consumption and life have been predicted in this project using validated computer-simulations. These help in making wiser management-decisions, so leading to the achievement of improved engine utilisation, lower overall life-cycle costs and optimal mission effectiveness for squadrons of aircraft.

In this paper a new moving-boundary algorithm is presented for the solution of two-dimensional flows on unstructured grids. Application of this algorithm is investigated by simulating the internal ballistics of solid rocket motors (SRM). Solution domain is discretised to a number of triangular elements using control-volume based finite-element (CVFE) method. Euler equations are integrated over the control volumes, and Roe's flux-difference scheme is used to evaluate flux vectors at the control-volume surfaces. To eliminate numerical errors due to grid movement, geometric conservation laws (GCL) are carefully implemented. The present algorithm recognises locations along the moving boundaries where the stretched elements result in a distorted grid. To avoid this, new elements are inserted through a local re-triangulation of elements in these regions. Internal ballistics of three SRMs with cylindrical grains burning from inside, are numerically simulated. Using a heat transfer model, erosive burning is included in calculation of propellant burning rate. The aftend of third SRM is tapered to demonstrate capability of present algorithm in solving problems with moving boundaries approaching each other. Very good agreement is observed between the time variations of head-end chamber pressure, obtained from the present work and the other methods.