This paper reviews and highlights recent developments of certain aspects of flow control concerned with reducing the drag of, and delaying flow separation on, wings and bodies over which the flow is turbulent. The study is restricted to devices that extend beyond the viscous sub-layer but are on a smaller scale than geometric features of the aircraft (e.g. wing chord). The review is mainly concerned with developments within the UK, although significant developments in other countries are discussed. The review discusses types of flow that need to be controlled, basic features of flow control devices and applications. It concludes with recommendations for future research.

A review of recent developments in flow control

Experiments have been performed in a blowdown supersonic wind tunnel to investigate the effect of subboundary layer vortex generators placed upstream of a normal shock/turbulent boundary layer interaction at a Mach number of 1.5 and a freestream Reynolds number of 28 x 10 6 . The Reynolds number based on the inflow boundary layer displacement thickness was 26,000. Two types of subboundary layer vortex generators were investigated: wedge-shaped and counter-rotating vanes. It was found that the vane-type subboundary layer vortex generators eliminated and the wedge-type subboundary layer vortex generators greatly reduced the shock-induced separation. When placed in the supersonic part of the flow, both types of subboundary layer vortex generators caused a wave pattern consisting of a shock, reexpansion, and shock. The reexpansion and double shocks are undesirable features because they equate to increased total pressure losses. Furthermore there are indications that the vortex intensity is reduced by the normal shock/boundary layer interaction. Overall, the vane-type subboundary layer vortex generators were the more effective devices as they eliminated the shock-induced separation and had the least detrimental effect on the shock structure.

Effect of Microvortex Generators On Seperated Normal Shock/ Boundary Layer Interactions

Three-dimensional bumps have been developed and investigated on transonic wings, aiming to fulfill two major objectives of shock-wave/boundary-layer interaction control, that is, drag reduction and buffet delay. An experimental investigation has been conducted for a rounded bump in channel flow at the University of Cambridge and a computational study has been performed for a spanwise series of rounded bumps mounted on a transonic aerofoil at the University of Stuttgart. In both cases wave drag reduction and mild control effects on the boundary layer have been observed. Control effectiveness has been assessed for various bump configurations. A double configuration of narrow rounded bumps has been found to perform best, considerably reducing wave drag by means of a well-established X-shock structure with little viscous penalty and thus achieving a maximum overall drag reduction of about 30%, especially when significant wave drag is present. Counter-rotating streamwise vortex pairs have been produced by some configurations as a result of local flow separation. On the whole a large potential of three-dimensional control with discrete rounded bumps has been demonstrated both experimentally and numerically.

Shock-Wave/Boundary-Layer Interaction Control Using Three-Dimensional Bumps for Transonic Wings

Pseudo-shock waves and their interactions in high-speed intakes

This work presents an immersed-boundary technique for compressible, turbulent flows and applies the technique to simulate the effects of micro vortex generators in controlling oblique-shock/turbulent boundary-layer interactions. The Reynolds-averaged Navier-Stokes equations, closed using the Menter k-ω turbulence model, are solved in conjunction with the immersed-boundary technique. The approach is validated by comparing solutions obtained using the immersed-boundary technique with solutions obtained on a body-fitted mesh and with experimental laser Doppler anemometry data collected at Cambridge University for Mach 2.5 flow over single micro vortex generators. Simulations of an impinging oblique-shock boundary-layer interaction at Mach 2.5 with and without micro vortex-generator flow control are also performed, considering the development of the flow in the entire wind tunnel. Comparisons are made with experimental laser Doppler anemometry data and surface-pressure measurements from Cambridge University and an analysis of the flow structure is performed. The results show that three dimensional effects initiated by the interaction of the oblique shock with the sidewall boundary layers significantly influence the flow patterns in the actual experiment. The general features of the interactions with and without the micro vortex-generator array are predicted to good accord by the Reynolds-averaged Navier-Stokes/ immersed-boundary model.

Numerical simulations of effects of micro vortex generators using immersed-boundary methods

The effect of streamwise slots on the interaction of a normal shock wave with a turbulent boundary layer has been investigated experimentally at a Mach number of 1.29. The surface-pressure distribution for the controlled interaction was found to feature a distinct plateau. This was caused by a change in shock structure from a typical unseparated normal shock-wave/boundary-layer interaction to a large bifurcated lambda-type shock pattern, which led to a reduction of total pressure losses. A strong spanwise variation of boundary-layer properties was observed downstream of the slots, whereas the modified shock structure was relatively two-dimensional. Surface flow visualization confirmed that the slots introduced a region of recirculation into the boundary layer, similar to passive control with uniform surface ventilation. Surface flow visualization revealed the presence of a pair of counter-rotating vortices, confirmed by crossflow velocity measurements. Because of the reduction of total pressure losses, streamwise slots can reduce aircraft wave drag at transonic cruise while incurring only small viscous penalties. A similar control device can also be of use in supersonic intakes where total pressure losses limit engine performance

Shock-wave/boundary-layer interaction control using streamwise slots in transonic flows

The effect of streamwise slots on the interaction of a normal shock wave / turbulent boundary layer has been investigated experimentally at a Mach number of 1.3. The surface pressure distribution for the controlled interaction was found to be significantly smeared, featuring a distinct plateau. This was due to a change in shock structure from a typical unseparated normal shock wave boundary layer interaction to a large bifurcated Lambda type shock pattern. Boundary layer velocity measurements downstream of the slots revealed a strong spanwise variation of boundary layer properties whereas the modified shock structure was relatively twodimensional. Oil flow visualisation indicated that in the presence of slots the boundary layer surface flow was highly three dimensional and confirmed that the effect of slots was mainly due to suction and blowing similar to that for passive control with uniform surface ventilation. Three hole probe measurements confirmed that the boundary layer was three dimensional and that the slots introduced vortical motion into the flowfield. Results indicate that when applied to an aerofoil, the control device has the potential to reduce wave drag while incurring only small viscous penalties. The introduction of streamwise vorticity may also be beneficial to delay trailing edge separation and the device is thought to be capable of postponing buffet onset. © 2001 by A N Smith.

Control of normal shock wave/turbulent boundary layer interactions using streamwise grooves

The effect of streamwise slots and grooves on a normal shock wave-turbulent boundary-layer interaction has been investigated experimentally at a Mach number of 1.3. The surface pressure distribution for the controlled interaction in the presence of slots featured a distinct plateau. This was due to a change in shock structure from a typical unseparated normal shock wave-boundary-layer interaction to a large bifurcated lambda type shock pattern. Velocity measurements downstream of the slots revealed a strong spanwise variation of boundary-layer properties, whereas the modified shock structure was found to be relatively two-dimensional. Cross flow measurements indicate that slots introduce streamwise vortices into the flow. When applied to an aerofoil, streamwise slots have the potential to reduce wave drag while incurring only small viscous penalties. In the presence of grooves the interaction was initially found to be significantly different. A bifurcated shock structure was observed but the trailing leg appeared stronger and featured a second lambda foot. Oil flow visualisation also revealed differences in the interactions, with the region of suction and blowing being limited to a smaller extent of the grooved control surface. The amount of crossflow present was reduced compared to the slotted control surface. By varying the internal geometry of the grooves it was found that the interaction could be modified to be similar to that in the presence of slots indicating that a more practical control device can be designed.

Normal shock wave-turbulent boundary-layer interactions in the presence of streamwise slots and grooves

An experimental investigation has been undertaken to investigate the potential of streamwise slots as a shock / boundary layer interaction control device. Results indicate that slots delay the onset and reduce the extent of shock induced separation. Due to the reduction in separation, flow breakdown is delayed until higher Mach numbers, resulting in increased lift and decreased drag. As well as the delay in fully separated flow, the shock wave position was found to be practically constant on the aerofoil fitted with control, suggesting that slots may also delay buffet onset.

A. N. Smith

Papers

Shock-wave/boundary-layer interaction control using streamwise slots in transonic flows

Control of normal shock wave/turbulent boundary layer interactions using streamwise grooves

Normal shock wave-turbulent boundary-layer interactions in the presence of streamwise slots and grooves

Experimental Investigation of Transonic Aerofoil Shock/ Boundary Layer Interaction Control Using Streamwise Slots