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No description available.

The complex flowfield which is associated with a rotor wake gives rise to the multitude of aerodynamic interactions that may occur during rotorcraft operation. These interactions may give rise to undesirable noise and lead to an unacceptable performance degradation, and as such the investigation of the fundamental mechanics of such interactions, that which occurs between the tail rotor and the trailing tip vortices shed from the main rotor assembly, is the focus of the current investigation. As the purpose of the tail rotor is to provide balance for the torque of the main rotor, these types of interaction will adversely impact on the overall rotorcraft performance. The basis of the present thesis has been an experimental investigation of the orthogonal BVI, in which the axis of the interacting vortex (in the plane of the vortex core axial flow) is nominally orthogonal to the interacting blade chordline, representing the tail rotor interaction. The tests have been conducted using a specifically designed facility at the University of Glasgow, with the flow interrogated using a Particle Image Velocimetry (PVI) technique. The PVI method allows global flowfield information to be obtained pertaining to the nature of the interaction. The methodology was benchmarked against synthetic flowfields, and with the accuracy of the flowfield measurements improved dramatically with the implementation of the Forward/Reverse Tile Test (FRTT), which improved the accuracy in the flowfields to 3% in two-dimensional interrigation, and 5% in three-dimensional. The interrogation of the flowfield around the representative tail rotor blade demonstrated that the characteristics imparted vortex due to the BVI event could be attributed to the manner in which the axial flow component of the vortex was affected by the interaction. The results for the isolated flow conditions agreed well with those from previous measurements of the vortical structure, and the post interaction structure clearly indicated distinct differences determined by the direction of the axial flow relative to the blade chordline. Initial testing indicated that the thickness ratio had a marked effect on the progression of the OBVI, and for a suitably high thickness ratio, there was little evidence to suggest that the vortex core axial flow is 'cut' by the interacting body in the manner observed for the lower thickness ratios. For lower thickness ratios, as the vortex core is blocked by the interacting blade surface, the retardation of the axial component on the blade lower surface leads to rapid redistribution of the fluid into the surrounding flow, and the corresponding enlargement and distortion to the vortex tangential velocity components promoted by the radial outflow. On the upper blade side, regions of negative axial flow velocity indicate the presence of some fluid passing down through the core towards the surface of the blade, which are accompanied by a split divergence pattern around the vortex core. The effects immediately behind the trailing edge continue to be of interest due to the manner in which the vortex might be regenerated after the interaction and before any subsequent interactions with following blades. A relative lack of distortion within the out-of-plane component indicates that a rapid regeneration of the axial flow component may occur once the vortex has passed over the trailing edge. The use of passive control techniques in reduction of the effects associated with the orthoganal BVI have also been addressed, considering the effect of a counter-rotating vortex pair on the progression of the interaction. Although the inclusion a notch in the leading edge and outboard sweep on the rotor blade producing the representative trailing tip vortex did produce a well defined inboard vortex structure, there is evidence to suggest that this structure is ingested into the outboard tip vortex, as there is no significant modification to the progression of the OBVI.

The influence of trailing-edge separation on the dynamic stall characteristics of a typical rotor section is at present unclear. Although previous research has given a fundamental understanding of the unsteady stall process, the variety of aerofoils tested has made it difficult to isolate the effect of trailing-edge separation. Further investigation into this field may be carried out by testing two similar aerofoils which differ only in their trailing-edge separation characteristics. The early part of the work concentrated on the development of a numerical method whereby the theoretical pressure gradient over the trailing-edge upper surface of a given aerofoil may be modified to either enhance or reduce such separation. Since previous work at the University of Glasgow had included a detailed unsteady aerodynamic study of a NACA 23012 aerofoil, this was the appropriate profile for modification. The above technique was applied to this aerofoil with the objective of modifying the geometry in such a manner that would retain the leading-edge pressure distribution whilst forcing an earlier and more gradual trailing-edge separation growth. The subsequently designed aerofoil, designated the NACA 23012(A), was shown to display an enhancement of the trailing-edge separation characteristics via both boundary-layer calculations and oil-flow visualisation tests. On comparison with unsteady data previously collected for the NACA 23012, several systematic methods of estimating the effects of trailing-edge separation on the dynamic stall process are presented. During oscillatory tests the NACA 23012(A) displayed a more stable damping characteristic which was attributed to the enhanced trailing-edge separation producing an earlier pitching-moment break. Based on the analysis of pressure-time histories obtained during ramp tests, it was deduced that a consequence of significant trailing-edge separation was to delay the initiation of the dynamic stall vortex. Detailed analysis of hot-film data led to the conclusion that aerofoils which display a tendency to stall in steady conditions, via separation growth from the trailing-edge, will experience vortex initiation by the breakdown of a thin layer of reversed flow travelling upstream beneath a stable shear layer which remains in close proximity to the aerofoil's surface contour.

Focusing on helicopter aerodynamics, it is known that the aerodynamic performance of the retreating side of a rotor disk is mainly dictated by the stall characteristics of the blade. Stall under dynamic conditions (Dynamic Stall) is the dominant phenomenon encountered on heavily loaded fast-flying rotors, resulting in an extra lift and excessive pitching moments. Dynamic stall (DS) can be idealised as the pitching motion of a finite wing and this is the focus of the present work which includes three main stages. At first, comparisons between available experimental data with CFD simulations were performed for 3D DS cases. This work is the first detailed CFD study of 3D Dynamic Stall and has produced results indicating that DS can be predicted and analysed using CFD. The CFD results were validated against all known experimental investigations. In addition, a comprehensive set of CFD results was generated and used to enhance our understanding of 3D DS. Straight, tapered and swept-tip wings of various aspect ratios were used at a range of Reynolds and Mach numbers and flow conditions. For all cases where experimental data were available effort was put to obtain the original data and process these in exactly the same ways as the CFD results. Special care was put to represent exactly the motion of the lifting surfaces, its geometry and the boundary conditions of the problem. Secondly, the evolution of the Ω-shaped DS vortex observed in experimental works as well as its interaction with the tip vortices were investigated. Both pitching and pitching/rotating blade conditions were considered. Finally, the potential of training a Neural network as a model for DS was assessed in an attempt to reduce the required CPU time for modelling 3D DS. Neural networks have a proven track record in applications involving pattern recognition but so far have seen little application in unsteady aerodynamics. In this work, two different NN models were developed and assessed in a variety of conditions involving DS. Both experimental and CFD data were used during these investigations. The dependence of the quality of the predictions of the NN on the choice of the training data was then assessed and thoughts towards the correct strategy behind this choice were laid out.

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