An analysis of the crossover between local and massive separation on airfoils

Massive separation on airfoils operating at high Reynolds number is an important problem to the aerodynamicist, since its onset generally determines the limiting performance of an airfoil, and it can lead to serious problems related to aircraft control as well as turbomachinery operation. The phenomenon of crossover between local separation and massive separation on realistic airfoil geometries induced by airfoil thickness is investigated for low speed (incompressible) flow. The problem is studied both for the asymptotic limit of infinite Reynolds number using triple-deck theory, and for finite Reynolds number using interacting boundary-layer theory. Numerical results are presented which follow the evolution of the flow as it develops from a mildly separated state to one dominated by the massively separated flow structure as the thickness of the airfoil geometry is systematically increased. The effect of turbulence upon the evolution of the flow is considered, and the impact is significant, with the principal effect being the suppression of the onset of separation. Finally, the effect of surface suction and injection for boundary-layer control is considered. The approach which was developed provides a valuable tool for the analysis of boundary-layer separation up to and beyond stall. Another important conclusion is that interacting boundary-layer theory provides an efficient tool for the analysis of the effect of turbulence and boundary-layer control upon separated vicsous flow

Turbine vane gas film cooling with injection in the leading edge region from a single row of spanwise angled holes

An experimental study of gas film cooling was conducted on a 3X size model turbine vane. Injection in the leading edge region was from a single row of holes angled in a spanwise direction. Measurements of the local heat flux downstream from the row of coolant holes, both with and without film coolant flow, were used to determine the film cooling performance presented in terms of the Stanton number ratio. Results for a range of coolant blowing ratio, M = 0 to 2.0, indicate a reduction in heat flux of up to 15 to 30 percent at a point 10 to 11 hole diameters downstream from injection. An optimum coolant blowing ratio corresponds to a coolant-to-freestream velocity ratio in the range of 0.5. The shallow injection angle resulted in superior cooling performance for injection closest to stagnation, while the effect of injection angle was insignificant for injection further from stagnation

Effects of a synthetic jet acting on a separated flow over a hump

The effects of an oscillatory zero-net-mass-flux jet (i.e. synthetic jet) acting on a separated flow over a hump are investigated in terms of two actuation parameters – actuator position and forcing frequency. By considering the vorticity flux balance and introducing a centroid of vorticity production over the hump surface, lift and drag acting on the hump can be expressed as a function of this centroid and the rate of vorticity production. To study the parametric dependence of lift and drag, direct numerical simulation (DNS) is performed by solving compressible, unsteady, laminar flows over a half-cylindrical hump in two dimensions. The DNS results show that periodic actuation significantly reduces the rate of vorticity production at the wall and shifts the centroid upstream so that the drag is reduced and the lift is increased, respectively. When the actuation parameters are varied, it is found that the lift is governed by the horizontal coordinate of the vorticity-production centroid, while the drag is determined by both the vertical coordinate of the centroid and the rate of vorticity production over the hump. This paper explains by using ideal flow models that the vorticity-production centroid is controlled by two factors: one is the actuator position at which clockwise vorticity is generated, and the other is the point where the separation vortex is pinched off, thereby the clockwise vorticity being absorbed

Secondary flow reduction techniques in linear turbine cascades

This thesis investigates a novel secondary flow reduction method. The inlet boundary layer to a linear turbine cascade is skewed by injection of air through an upstream slot to oppose regular generated negative stream wise vorticity. Other methods from the pertinent literature are reviewed on a broad basis. Detailed measurements of the flowfield in the Durham Linear Cascade facility have shown that substantial reductions in secondary flows and losses are possible. If the kinetic energy required for the blowing is taken into account by means of an availability analysis, no net gain in loss is achieved. Tests are performed at two different angles, of which the higher is typical for film cooling applications, and at a wide range of injection ratios. Calculation of the mixed-out losses show the tangential rather than spanwise momentum of the injected air is more effective in countering the generation of secondary flows. Computations using a state-of-the-art Navier-Stokes solver indicated shortcomings in modelling a flow governed by complex vortex dynamics. Improvements in the turbulence model and injection geometry could remedy this. The evaluation of turbulent and laminar production rates obtained without injection helps to explain total pressure loss generation mechanisms. The comparison of calculated and experimental eddy viscosities reveals the inadequacy of the Boussinesq assumption for high turning flows. The results obtained in this work are relevant to endwall film cooling applications. The tangential injection of air in front of the leading edge provides coolant in an optimum manner whilst possibly reducing secondary losses to a large extent. Disc cooling air, present in a real engine to prevent the ingestion of hot air from the mainstream, could be used to supply the injection

Feasibility of generating an artificial burst in a turbulent boundary layer, phase 2

Various drag accounts for about half of the total drag on commercial aircraft at subsonic cruise conditions. Two avenues are available to achieve drag reduction: either laminar flow control or turbulence manipulation. The present research deals with the latter approach. The primary objective of Phase 2 research was to investigate experimentally the feasibility of substantially reducing the skin-friction drag in a turbulent boundary layer. The method combines the beneficial effects of suction and a longitudinally ribbed surface. At a sufficiently large spanwise separation, the streamwise grooves act as a nucleation site causing a focusing of low-speed streaks over the peaks. Suction is then applied intermittently through longitudinal slots located at selected locations along those peaks to obliterate the low-speed regions and to prevent bursting. Phase 2 research was divided into two tasks. In the first, selective suction from a single streamwise slot was used to eliminate either a single burst-like event or a periodic train of artificially generated bursts in laminar and turbulent boundary layers that develop on a flat plate towed in a water channel. The results indicate that equivalent values of the suction coefficient as low as 0.0006 were sufficient to eliminate the artificially generated bursts in a laminar boundary layer

The near wall effect of synthetic jets in a boundary layer

Copyright @ 2007 Elsevier Inc. All rights reserved.An experimental investigation to analyse the qualitative near wall effect of synthetic jets in a laminar boundary layer has been undertaken for the purpose of identifying the types of vortical structures likely to have delayed separation on a 2D circular cylinder model described in this paper. In the first instance, dye visualisation of the synthetic jet was facilitated in conjunction with a stereoscopic imaging system to provide a unique quasi three-dimensional identification of the vortical structures. Secondly, the impact of synthetic jet structures along the wall was analysed using a thermochromic liquid crystal-based convective heat transfer sensing system in which, liquid crystals change colour in response to the thermal footprints of a passing flow structure. Of the different vortical structures produced as a result of varying actuator operating and freestream conditions, the footprints of hairpin vortices and stretched vortex rings revealed a marked similarity with the oil flow pattern of a vortex pair interacting with the separation line on the cylinder hence suggesting that either of these structures was responsible in delaying separation. Conditions were established for the formation of the different synthetic jet structures in non-dimensional parameter space