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 x 10 5 . 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.
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.
Wind tunnel measurements were done for the separated and reattaching flow formed over a backward-facing step. Turbulent energy and turbulent normal stress balances were estimated from the measured data of mean velocities, Reynolds stresses, and turbulent triple products. The main objective of the experiments is to analyze the turbulent structure inside the reattaching shear layer, including the reverse flow, in detail. The results indicated different turbulent structures in three classified regions of the backward-facing step flow, first the dead air region, second the near wall region just upstream and downstream of the reattachment, and third the separated shear layer region. The second region is the one in which the transverse diffusion by turbulence in the turbulent energy balance equation plays an important role in comparison with other regions. In the reverse flow region just upstream of the reattachment included in the second region, the production term by shear stress, the transverse diffusion term by turbulence, and the advection term similarly help to balance the dissipation term.
Low-speed wind-tunnel tests have been made on a number of vortex flap and vortex plate configurations at the Cranfield Institute of Technology. The objectives of the experiment are to assess the benefits of these devices on the lift/drag ratio improvement of delta wings. The force and surface pressure measurements were made on a 1.15-m span, 60-deg delta wing model. Results indicate that the vortex flap deflection angle, which causes the flow to attach on the flap surface without any large separation, shows a much higher lift/drag ratio than the flap deflection angle which forms a leading-edge separation vortex over the flap surface. The performance of a vortex plate protruding from the leading edge of the datum delta wing is comparable to that of the vortex flap. However, when the vortex plate is used with the vortex flap deflected, it showed no benefit in these tests. Nomenclatureusing Cr and measured about xlCr = 0.4 Cp = pressure coefficient Cr -wing centerline chord, m g ~ vortex plate leading-edge position measured from leading edge of the wing in the chordwise direction, m LID = lift/drag ratio Re Cr -Reynolds number based on wing centerline chord U x = freestream velocity, m/s x -chordwise coordinate measured from the apex of the delta wing, m y = spanwise coordinate orthogonal to x, measured from the wing centerline, m a -wing incidence, deg a g = geometrical wing incidence (i.e., without tunnel corrections), deg 8 f -vortex flap deflection angle measured normal to the hinge line, deg
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