Evaluation of wave drag reduction by flow control

An analytical expression is proposed to estimate the wave drag of an aerofoil equipped with shock control. The analysis extends the conventional approach for a single normal shock wave, based on the knowledge that all types of successful shock control on transonic aerofoils cause bifurcated λ-shock structures. The influence of surface curvature on the λ-shock structure has been taken into account. The extended method has been found to produce fairly good agreement with the results obtained by CFD methods while requiring negligible computational effort. This new formulation is expected to be beneficial in the industrial design process of transonic aerofoils and wings where a large number of computational simulations have to be performed.

Toward designing with three-dimensional bumps for lift/drag improvement and buffet alleviation

The desire to design more efficient transport aircraft has led to many different attempts to minimize drag. One approach is the use of three-dimensional shock control bumps, which have gained popularity in the research community as simple, efficient and robust devices capable of reducing the wave drag of transonic wings. This paper presents a computational study of the performance of three-dimensional bumps, relating key bump design variables to the overall wing aerodynamic performance. An efficient parameterization scheme allows three-dimensional bumps to be directly compared to two-dimensional designs, indicating that two-dimensional bumps are capable of greater design point aerodynamic performance in the transonic regime. An advantage of three-dimensional bumps lies in the production of streamwise vortices, such that, while two-dimensional bumps are capable of superior performance near the design point, three-dimensional bumps are capable of breakingup regions of separated flow at high Mach numbers, suggesting improvement in terms of buffet margin. A range of bump designs are developed that exhibit a tradeoff between design point aerodynamic efficiency and improvementinbuffet margin, indicating the potential for bespoke designs to be generated for different sections of a wing based on its flow characteristics. Copyright © 2012 by Jeremy Eastwood and Jerome Jarrett.

Normal shock boundary layer control with various vortex generator geometries

Various vortex generators which include ramp, split-ramp and a new hybrid concept "ramped-vane" are investigated under normal shock conditions with a diffuser at Mach number of 1.3. The dimensions of the computational domain were designed using Reynolds Average Navier-Stokes studies to be representative of the flow in an external-compression supersonic inlet. Using this flow geometry, various vortex generator concepts were studied with Implicit Large Eddy Simulation. In general, the ramped-vane provided increased vorticity compared to the other devices and reduced the separation length downstream of the device centerline. In addition, the size, edge gap and streamwise position respect to the shock were studied for the ramped-vane and it was found that a height of about half the boundary thickness and a large trailing edge gap yielded a fully attached flow downstream of the device. This ramped-vane also provided the largest reduction in the turbulent kinetic energy and pressure fluctuations. Additional benefits include negligible drag while the reductions in boundary layer displacement thickness and shape factor were seen compared to other devices. © 2010 by Sang Lee.

Normal shock boundary layer control with various vortex generator geometries

Various vortex generators which include ramp, split-ramp and a new hybrid concept "ramped-vane" are investigated under normal shock conditions with a diffuser at Mach number of 1.3. The dimensions of the computational domain were designed using Reynolds Average Navier-Stokes studies to be representative of the flow in an external-compression supersonic inlet. Using this flow geometry, various vortex generator concepts were studied with Implicit Large Eddy Simulation. In general, the ramped-vane provided increased vorticity compared to the other devices and reduced the separation length downstream of the device centerline. In addition, the size, edge gap and streamwise position respect to the shock were studied for the ramped-vane and it was found that a height of about half the boundary thickness and a large trailing edge gap yielded a fully attached flow downstream of the device. This ramped-vane also provided the largest reduction in the turbulent kinetic energy and pressure fluctuations. Additional benefits include negligible drag while the reductions in boundary layer displacement thickness and shape factor were seen compared to other devices. © 2011 Elsevier Ltd.

Absorption of sound in cellular foams

To further enhance the sound absorption of metal foams via combining the high sound absorption and good heat conductivity of the cellular foam metals, the use and acoustic modeling of these materials are reviewed. The predictions made by three viscous models developed by the authors for the propagation of sound through open-cell metal foams are compared with an experiment both for the metal foams and for the polymer substrates used to manufacture the foam. All models are valid in the limit of low Reynold's number which is valid for the typical cell dimensions found in metal foams provided the amplitude of the waves is below 160 dB. The first model considers the drag experienced by acoustic waves as they propagate passing rigid cylinders parallel to their axes, the second considers the propagation normal to their axes, and the third considers the propagation passing the spherical joints. All three are combined together to give a general model of the acoustic behavior of the foams. In particular, the sound absorption is found to be significant and well predicted by the combined model. In addition, a post-processing technique is described for the experiment used to extract the fundamental wave propagation characteristics of the material.

Vortex generators near shock/ boundary layer interactions

Experiments have been performed in a blowdown supersonic wind tunnel to investigate the effect of arrays of sub-boundary layer vortex generators placed upstream of a normal shock/ boundary layer interaction. The investigation makes use of a recovery shock wave and the naturally grown turbulent boundary layer on the wind tunnel floor. Experiments were performed at Mach numbers of 1.5 and 1.3 and a freestream Reynolds number of 28 × 106. Two types of vortex generators were investigated - wedge-shaped and arrays of counter-rotating vanes. It was found that at Mach 1.5 the vane-type VGs eliminated and the wedge-type VGs greatly reduced the separation bubble under the shock. When placed in the supersonic part of the flow both VGs caused a wave pattern consisting of a shock, re-expansion and shock. The re-expansion and double shocks are undesirable features since they equate to increased total pressure losses and hence increased -wave drag. Furthermore there are indications that the vortex intensity is reduced by the normal shock/ boundary layer interaction. When the shock was located directly over the VGs there was no re-expansion present, but the 'damping' effect of the shock on the vortex persisted. It appears that the vortices produced by the wedge-shaped VGs lift off the surface more rapidly. Similar results were observed at Mach 1.3, where the flow was unseparated.

Three-dimensional SBLI control for transonic airfoils

The application of shock control to transonic airfoils and wings has been demonstrated widely to have the potential to reduce wave drag. Most of the suggested control devices are two-dimensional, that is they are of uniform geometry in spanwise direction. Examples of such techniques include contour bumps and passive control. Recently it has been observed that a spanwise array of discrete three-dimensional controls can have similar benefits but also offer advantages in terms of installation complexity and drag. This paper describes research carried out in Cambridge into various three-dimensional devices, such as slots, grooves and bumps. In all cases the control device is applied to the interaction of a normal shock wave (M=1.3) with a turbulent boundary layer. Theoretical considerations are proposed to determine how such fundamental experiments can provide estimates of control performance on a transonic wing. The potential of each class of three-dimensional device for wave drag reduction on airfoils is discussed and surface bumps in particular are identified as offering potential drag savings for typical transonic wing applications under cruise conditions.