In this paper, we describe the structures that produce a spike-type route to rotating stall and explain the physical mechanism for their formation. The descriptions and explanations are based on numerical simulations, complemented and corroborated by experiments. It is found that spikes are caused by a separation at the leading edge due to high incidence. The separation gives rise to shedding of vorticity from the leading edge and the consequent formation of vortices that span between the suction surface and the casing. As seen in the rotor frame of reference, near the casing the vortex convects toward the pressure surface of the adjacent blade. The approach of the vortex to the adjacent blade triggers a separation on that blade so the structure propagates. The above sequence of events constitutes a spike. The computed structure of the spike is shown to be consistent with rotor leading edge pressure measurements from the casing of several compressors: the centre of the vortex is responsible for a pressure drop and the partially blocked passages associated with leading edge separations produce a pressure rise. The simulations show leading edge separation and shed vortices over a range of tip clearances including zero. The implication, in accord with recent experimental findings, is that they are not part of the tip clearance vortex. Although the computations always show high incidence to be the cause of the spike, the conditions that give rise to this incidence (e.g., blockage from a corner separation or the tip leakage jet from the adjacent blade) do depend on the details of the compressor.

In this paper we describe the structures that produce a spike-type route to rotating stall and explain the physical mechanism for their formation. The descriptions and explanations are based on numerical simulations, complemented and corroborated by experiments. It is found that spikes are caused by a loss of pressure rise capability in the rotor tip region, due to flow separation resulting from high incidence. The separation gives rise to shedding of vorticity from the leading edge and the consequent formation of vortices that span between the suction surface and the casing. As seen in the rotor frame of reference, near the casing the vortex convects toward the pressure surface of the adjacent blade. The approach of the vortex to the adjacent blade triggers a separation on that blade so the structure propagates. The above sequence of events constitutes a spike. The simulations show shed vortices over a range of tip clearances including zero. The implication is that they are not part of the tip clearance vortex, in accord with recent experimental findings. Evidence is presented for the existence of these vortex structures immediately prior to spike-type stall and, more strongly, for their causal connection with spike-type stall inception. Data from several compressors are shown to reproduce the pressure and velocity signature of the spike-type stall inception seen in the simulations.

At low mass flow rates, axial compressors suffer from flow instabilities leading to stall and surge. The inception process of these instabilities has been widely researched in the past---primarily with the aim of predicting or averting stall onset. In recent times, attention has shifted to conditions well before stall and has focused on the level of irregularity in the blade passing signature in the rotor tip region. In general, the irregularity increases in intensity as the flow rate through the compressor is reduced. Attempts have been made to develop stall warning/avoidance procedures based on the level of flow irregularity, but little effort has been made to characterize the irregularity itself, or to understand its underlying cause. Work on this project has revealed for the first time that the increase in irregularity in the blade passing signature is highly dependent on both tip-clearance size and eccentricity. In a compressor with small, uniform, tip-clearance, the increase in blade passing irregularity that accompanies a reduction in flow rate will be modest. If the tip-clearance is enlarged, however, there will be a sharp rise in irregularity at all circumferential locations. In a compressor with eccentric tip-clearance, the increase in irregularity will only occur in the part of the annulus where the tip-clearance is largest, regardless of the average clearance level. In this paper, some attention is also given to the question of whether the irregularity observed in the prestall flow field is due to random turbulence or to some form of coherent flow structure. Detailed flow measurements reveal that the latter is the case. From these findings, it is clear that a stall warning system based on blade passing signature irregularity would be difficult to implement in an aero-engine where tip-clearance size and eccentricity change during each flight cycle and over the life of the compressor.

At low mass flow rates axial compressors suffer from flow instabilities leading to stall and surge. The inception process of these instabilities has been widely researched in the past — primarily with the aim of predicting or averting stall onset. In recent times, attention has shifted to conditions well before stall and has focussed on the level of irregularity in the blade passing signature in the rotor tip region. In general, this irregularity increases in intensity as the flow rate through the compressor is reduced. Attempts have been made to develop stall warning/avoidance procedures based on the level of the flow irregularity, but little effort has been made to characterise the irregularity, or to understand its underlying causes. Work on this project has revealed for the first time that the increase in irregularity in the blade passing signature is highly dependent on both tip-clearance and eccentricity. In a compressor with small, uniform, tip-clearance, the increase in blade passing irregularity which accompanies a reduction in flow rate will be modest. If the tip-clearance is enlarged, however, there will be a sharp rise in irregularity at all circumferential locations. In a compressor with eccentric tip-clearance, the increase in irregularity will only occur in the part of the annulus where the tip-clearance is largest, regardless of the average clearance level. In this paper, some attention is also given to the question of whether this irregularity observed in the pre-stall flow field is due to random turbulence, or to some form of coherent flow structure. Detailed flow measurements reveal that the latter is the case. From these findings, it is clear that a stall warning system based on blade passing signature irregularity will not be viable in an aero-engine where tip-clearance size and eccentricity change during each flight cycle and over the life of the compressor.

Two-dimensional (2D) strip-theory modelling of unsteady gust-aerofoil interaction is standard practice in many industrial applications, but the limits of applicability of 2D unsteady flow modelling on 3D wing and rotor geometries are not well understood. This paper investigates the effects of 3D geometry features, such as finite span, taper, sweep and rotation, on the unsteady lift response to gusts, and the flow-physical differences between 2D and 3D geometries in unsteady flow. A frequency-domain inviscid vortex lattice model is validated and used for the 3D analysis. The results are compared to unsteady transfer functions from 2D linear analytic theory (e.g. Theodorsen and Sears functions). The study agrees with previous research findings that 3D effects are most significant at low reduced frequencies and low aspect ratios, as well as near the wing tips. The driving cause of 3D response is shown to be the wake vorticity: both streamwise and spanwise components of unsteady wake vorticity must be modelled. The study concludes by investigating whether unsteady response of more complex 3D wing and rotor geometries can be represented by the response of a rectangular wing. The results indicate that this is possible for tapered wings and rotating blades, but not for swept wings. Nomenclature Influence matrix Semi-chord (m) c Chord (m) Cl Local lift coefficient CL Total lift coefficient Force (N) Reduced frequency Lift (N) Number of chordwise lattice panels Surface normal vector Number of blades Radius

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In order to make the extraction of tidal current energy economically viable, the power production per turbine must be optimised in each tidal array. Furthermore, the impact of power extraction on the marine flow environment must be understood. These two aims mean that designers must be able to model different configurations of a tidal array in order to create the most efficient, least invasive arrangement. In this paper, an analytical model is developed for array design in idealised rectangular tidal channels with idealised turbines. The model includes the effects of (i) local blockage, (ii) surface deformation and (iii) added drag due to the installation of the array. While these effects have been accounted for individually in past work, the model presented here is the first to include all three such that the interaction between different effects can be understood. Results are presented for optimal local blockage and turbine resistance as functions of inherent channel drag coefficient, channel length and Froude number at various global blockage values. It will be shown that it is necessary to model the effects of local blockage and added drag simultaneously in order to obtain the design parameters of a tidal array (global blockage, local blockage and turbine resistance), which will maximise the power extraction per turbine. Neglecting either effect will lead to an array design with lower power extraction than the optimum, the addition of unnecessary extra turbines and higher lost power from the array.

The two main obstacles to the commercial deployment of tidal turbines are cost and reliability. Behind both of these lies the ability of a turbine to withstand the hostile marine environment for long periods without failure or costly maintenance. Unsteadiness in the incoming flow seen by a tidal turbine causes large fluctuations in thrust and torque, which generate fatigue loads on the blade and drivetrain respectively and thus significantly reduce turbine lifespan. This paper will present the development of flow control devices to keep thrust and torque constant despite changes in flow speed. In particular, it will be shown that torque can be controlled using a small trip near the leading edge of the blade, and that thrust can be controlled via a flap at the trailing edge. Data from a small-scale turbine will be used to show the effectiveness of both devices, and a reduction in unsteady loading of up to 75% will be demonstrated. This could translate into an order of magnitude increase in fatigue life.