The problem of aircraft formation dynamics and control is investigated from the viewpoint of formation architecture. Three different formation structures, leader-wingman, virtual leader and behavioural approaches are introduced. A comparative study is made using a unified approach through a suitable control law. The formation systems are analyzed on a quantitative basis and objective results are made available for the designer. The trade-off between system performance and complexity is indicated. A complete nonlinear simulation involving a flight-path change and a heading change manoeuvre is discussed. Results show the superiority of a behavioural approach to maintain close formations of vehicles.

This paper presents identification, control synthesis and simulation results for an YF-22 aircraft model designed, built, and instrumented at West Virginia University. The ultimate goal of the project is the experimental demonstration of formation flight for a set of 3 of the above models. In the planned flight configuration, a pilot on the ground maintains controls of the leader aircraft while a wingman aircraft is required to maintain a pre-defined position and orientation with respect to the leader. The identification of both a linear model and a nonlinear model of the aircraft from flight data is discussed first. Then, the design of the control scheme is presented and discussed with an emphasis on the amount of information, relative to the leader aircraft, needed by the wingman to maintain formation. Using the developed nonlinear model, the control laws for a maneuvered flight of the formation are then simulated with Simulink® and displayed with the Virtual Reality Toolbox®. Simulation studies have been performed to evaluate the effects of specific parameters and the system robustness to atmospheric turbulence. The conclusions from this analysis have allowed the formulation of specific guidelines for the design of the electronic payload for formation flight.

Experimental investigations were carried out to study the acoustic radiation from a rectangular wall mounted cavity in a confined supersonic flow. The free-stream Mach number was maintained at 1·5 and the cavity length-to-depth ratio was varied from 0·43 to 5·0. Acoustic measurements made on the top wall show jumps in the dominant frequency as the cavity behaviour changes from shallow-to-square-to-deep cavity. Numerical simulations of this unsteady two-dimensional flow using the commercially available software FLUENT have also been carried out. Unsteady pressure data at the same location in the flow field as the pressure transducers in the experiments was collected. FFT analysis of the unsteady pressure data was performed to obtain the dominant acoustic frequencies. The values for these dominant frequencies predicted by the numerical calculations agree well with experimental data. The numerical study also predicts the frequency jump observed in experiments.

The work described here concentrates on under-expanded, axisym-metric turbulent jets issuing into quiescent conditions. Under-expanded turbulent jets are applicable to most aircraft propulsion applications that use convergent nozzles. Experimental studies used laser doppler velocimetry (LDV) and pitot probe measurements along the jet centreline. These measurements were made for two nozzle pressure ratios (2·5 and 4·0) and at various streamwise positions up to 10 nozzle diameters downstream of the nozzle exit plane. A computational fluid dynamics (CFD) model was developed using the Fluent code and utilised the RNG K-ε two-equation turbulence model. A mesh resolution of approximately one hundredth of nozzle exit diameter was found to be sufficient to establish a mesh independent solution.

Comparison of the jet centreline axial velocity (LDV data) and pressure ratio (pitot probe data) showed good agreement with the CFD model. The correct number of shock cells had been predicted and the shock strength agreed well between the experiments and numerical model. The CFD model was, however, found to over-predict the shock cell length resulting in a longer supersonic core. There was some evidence, based on analysis of the LDV measurements that indicates the presence of swirl and jet unsteadiness, which could contribute to a shortening of the shock cell length. These effects were not modelled in the CFD. Correlation between the LDV and pitot probe measurements was generally good, however, there was some evidence that probe interference may have caused the premature decay of the jet. Overall, this work has indicated the good agreement between a CFD simulation using the RNG k-ε turbulence model and experimental data when applied to the prediction of the flowfield generated by under-expanded turbulent jets. The suitability of the LDV technique to jet flows with velocities up to 500ms-1 has also been demonstrated.

Laminar separation bubbles formed on NACA 0012 aerofoil near the onset of a stall were investigated to clarify the behaviour of the laminar separation bubble. Measurements were done at a chord Reynolds number of 1·3 × 105. Mean velocity measurements indicate that the long bubble of about 35% chord length is formed at α = 11·5° after the short bubble burst occurred. However, the instantaneous flow visualisation picture indicates that the flow is strongly oscillating at this angle of attack. The phase averaging technique has been applied to analyse this oscillating behaviour. The results indicate that the flow is oscillating between a small separation-reattachment bubble formed near the leading-edge at about a 10% chord length and a large separated region extending over the aerofoil surface. It is suggested that this small separation-reattachment bubble has a similar flow structure to that of the short bubble formed at a lower angle of attack.